Results of the Fourth Saturn Launch Vehicle Test Flight SA-4

8/6/2019 Results of the Fourth Saturn Launch Vehicle Test Flight SA-4

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- -MPll-5AT-63-6May 10. 1963

.

I E S I L T S O F l iE F O I I I I SARli IL 1 I I C H _mT E S TRaJ

SA-4...".......... Ai~JII".


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GEORGE C. ~IAnSIt.\LL SPACE FLIGHT CENTEH

!\IPH-S.\ T - G 3 - 0V

HESl'LTS OF THE Fot'HTIf SATtlHN I LAllNCIf VEIfICLE TEST FLIGHT ".4J--)By Saturn Flight Evaluation Working Group

ABSTHACT1113~

This report presents the results of the Earty En-gtneermg Evaluatton ofthe SaturnSA-4 test flight. Theperformance ofeach major vehicle system is discussedwith special emphasis on malfunctions and deviations.

The SA-4 flight test was a complete success, withall missions ofthe flightbeing accompl.shed, Xomajormalfunction or deviation whichwould be considered aserious system failure or design deficieney occurred.SA-4 was the last test of the Block I series of SaturnI vehicles, ~ "f* '(/!!fP-Il-

Any questions or comments pertaining to the in-formation contained inthis report are invited andshouldbe directed to:

Dtrector; GeorgeC. Marshall Space Flight CenterHuntsville, AlabamaAttention: Chairman, Flight Evaluation Working

Group, M-AERO-F (Phone 876-2701)


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GEORGE C. MARSHALL SPACE FLIGHT CENTER

MPR-SA T-63-6

RESULTS OF THE FOURTH SATURN I LAUNCH VEHICLZ TEST FLIGHT

SATURN FLIGHT EVALUATIONWORKING GROUP


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ACKNOWLEDGEMENT

Acknowledgement is made to the vr.rtous divisionsand elements of MSFC and Launch O!.erations Centerfor their contributions to thit. report. Without the jointefforts and asststance of these elements, this inte-grated report wouldnot have been possible. The SaturnFlight Evaluation WorkingGroup is especially indebtedto the following MSFC branches (or their major con-tributions:

Aeroballistics DivisionAerodynamics BranchAerophysics and Astrophysics BranchDynamic Analysis BranchFlight Evaluation Branch

Astrionics DivisionElectrical Systems Integration BranchFlight Dynamics BranchGuidance and Control Systems BranchGyro and Stabilizer BranchInstrumentation Development Branch

Computation DivisionData Reduction BranchDigital Computer Branch

Launch Vehicle Operations DtvtstonElectrical Engineering, Measuringand Track-ing OfficeMechanical, Structural and Propulsion Office

Propulston and Vehicle Engineering DivisionPropulsion and Mechanics BranchStructures BranchVehicle Engineering Branch


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O RtG IN Al PA GE~ A(~!( A N O W HIT E PHOTOORAr-~

~ .t-.~,. , - .


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SECTION L

SECTDN II.SECTION m,

SECTION IV.

SECTIONV.

TABLE CF CONTENTS

FLIGHT TEST SUMMARY.............1.1 Flight Test Results ......................1.2 Test Objectives ...................INTRODUCTION ....................LAUNCHOPERATIONS...............3. 1 Summary ..3.2 Prelaunch Milestones .....................3.3 Prelaunch Atmosphertc Surface Conditions ..............3. 4 Countdown... . . . . . . . . . . . . . . . . . . . . . . .3. 5 Propellant Loading . . . . . . . . . . . . . . . .3.6 Holddown...... . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3.7 Ground SUPPOl"tEquipment ........................3.8 Launch Facility l\{easurenlc,lt.s ...............TRAJECTORY. . . . . . . . . . . . . . . . . . . . . . 4.1 Summa ry ........................4.2 Tracking Analysis .....................4.3 Actual and Nominal Trajectory ...................

4. 3. 1 Powered Flight ...................4.3. 2 Thrust Decay ....................

4. 4 Retro Rockets. . . . . . . . . . . . . . . . . . . . . HPROPULSION . . . . . . . . . . . .5. 1 Summary ..5.2 Individual Engine Performance .................5.3 Vehicle Propulsion System Performance .............5. 4 Pressu rtzatton Systenl s . . . . . . . . . . . . . . .

5. 4. 1 Fuel Tank Pressurization . . . . . . . . . . . . .5.4.2 LOX Tank Pressurization ................5. 4. 3 Control Pressure System . . . . . . . ..5.4.4 Air Bearing Supply............

5.5 Vehicle Propellant tTtilization .............5. 6 Hydraulic SyStelll . . . . . . . . . . . . . . . . _ . 5.7 Retro Rocket Pe rformanee ................

SECTION VL MASSCHARACTERISTICS.......6.1 Vehic le \Veights . . . . . . . . . . . . . . . . . 6.2 Vehicle Center of Gravity and Moments of Inertia ......

SECTION VII. CONTROL .........7. 1 Summal")' ..7. 2 Control Analysis ..............

7.2. 1 Pitch Plal'K! ..7. 2. 2 Yaw Plane .. .. . . .. .. . . . . .. .. .. .. . . .. . . .. .. .. .. .. .. . . .. . . .. .. . . .. . . .. . . .. .. .. .. .. .. ..7. 2. 3 Roll Plane .... .. .. .. .. .. .. .. .. .. .. .. .. .. . . .. .. . . .. .. . .. . . .. . . .. .. .. .. .. .. . . . . . . 7.2.4 Altitude and Con~rol MLer Cutoff .........

7. 3 Functional Analysis ...........7. 3. 1 Control Sensors .........7. 3. 2 Control Computer : . . . . . . . 7.3.3 Actlla.tors......... . . . . . . . . . . . . . . . . . . 7.3.4 ST-124P stabilized Platform Attitude Comparison ..

7. 4 Propellant Sloshing . . . . . . . . . . . . . . . . . . . . . . . . . . .

6777

12121214151515161616171719191921212121232526272729293030

iii ........ PAGE ILANK NOT f ILMED


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TARLE OF ('ONTEXTS ('OXT'D,

Pa!l,~'

SEC nON VID. GUIDANCE ...................s, 1 Summary .ti.::! Dcs(:rlption of the Guidance System ..

1'.2.1 ST-90 Guidance System ....1'.2.2 ST-12-tP Guidance System .....

~.:1 Platfo rm and Acceler-ometer- Alignments1'1.4 Operational Analysts .............

S. 4.1 Guidance Intefhgence Errors .!i.4. 2 Aceeleromete r Outputs (ST-90, .1'1.4.3 .\c~derometer Outputs (S~-12-tP' .

s.;) Functional Analysrs ...........

- , . ," -r.)-'I.," -',,).)-:n

~. 5. 1~..5.2

Guidance Sensors. . . . . . . . . . . . .Velodty Encoder-s and Guidance SignalProcessor Repeater's ....Stabilized Platforms .

:1;,

b.5.3SECTION IX. VEHICLE ELECTRI(,A L SYSTEM ........

9. 1 Summary .............9.2 Flight Results .....................

SECTION X. STRUCTlJRES ..... .,-."10. 1 Summary .... .. . . .. . . .. .. .. .. . . .. .. . . .. .. . . .. . . .. .. . . .. . . .. . . .. .. .. .. . . .. . . .. .. .. . . .. . . .. .. .. .. . . .. .. .. . . ;1,10. 2 Bending )'oment and Normal Load Factors. . . . . . . . . . . . . . . . . . . . . . . . . :1';

10. :!. 1 Instrumentation ., .10.2. 2 ~Iolnenl Loads, . . . . . . ..........................

10.3 Longitudinal Loads ....................................10.4 Retro Rocket Alignment ...............................10.5 Bending Oscillations ................................10.6 Vibrations ......................................

10.6. 1 Summary of Vibration Data ..........................10.6. 2 Instr-umentation ......10.6.:1 Dtscusaton of Vibration Measurements ...............

.,-. . .

10." Vehicle A


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TABLE OF CCNTE1\TS(C01\,'m

57SECTION XII. INSTRUMENTATION . 5712.1 Summary . 57~. . .. .. " .. .. .. '" "" . . . . . " . . . .

12.2 Measuring Analysis . 5712. 3 Tele111et~ . 5712.3. t S~sCems .

12. :L2 Airborne Tape Re('order .......................... , 5'"12.3.3 Experimental Links 6 (PCMI and 9 (UHF) , ....... 5'"5912.3." Performance Analysis .............................RF System Analysis ............................... , 5912.4.1 Telemetry ...................................... , ~~12.4.2 Azusa ...............12.4.3 UOOP ........................ 6112.4.4 Ce-Band Radar .................................. , 616212.4.5 Altinleter ....................................... 62Engineering Sequential Camera Coverage .......................

12.4

12.5SECTION >"'11l SUMMARY OF MALFUNCnoNS AND DEVIATIO~'5 .................... , 64

65SECTION XIV. SPECIAL MlSSf()NS .................................... 6514.1 Horizon Sensor ..................................... 6514.2 Ca.-city Microphone (Micro-Meteorite Detector . 6514.3 Other Special Missions ......................................Appendix A - Atmospheric Summary {or SA-4 ....................... 66

v


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FigureI-I3-1

LIST OF ILLUSTRATIONSTitle Page

Saturn 8-1 Stage Polarity Chart 0 0 0 0 0 SA-4 Holds 0 0 0 0 0 0 0 0 0 0

26

3-2 RP-I Speci!ic Weight versus Temperature. 0 0 64-14-24-34-44-55-1

5-85-95-105-115-126-17-17- 27-37- 47-57-67-77-87- 97-107-117-12'4-137-147-15

Altimete r Deviation. . . . . . . . . . . . . . . . . 8Trajectory . 0 0 10Earth- Fixed Veloc tty . 0 0 0 0 0 10Mach and ~'Dar.lic Pressure. . . . . . . . . . . . toLongitudinal Acceleration .. 0 0 10Individual Engine Deviation from Predicted Thrust . . . . . . . . . . . . 12Indi"idual Engine Deviation from Predicted Specific Impulse ... 12Chamber Thrust Buildup. . . . . . . . . 13Outboard Engine Thrust Decay .... 0 14Vehicle Longitudinal Thrust and Specific Impulse . 0 0 0 14Vehicle Mixture natio and Total Flow Rate .. 0 0 0 0 0 0 15Gas Pressure in Fuel Tank 0 0 0 0 15Current Block n Pressuriz~ Sequence 0 16Revised mock n Pressurization SequebCe . . . 16Prelaunch I"(}X Tank Pressure ... 0 0 0 0 16Gas Pressure in l.(}X Tanks .. 0 0 o . 16Typical Retro Rocket Cbanlber Thrust. 0 0 17Vehicle Weight, Longitudinal Center of Gravity and Mass Moments of Inertia . .. o . 19Pitch Attitude, Angular Velocity and Average Actuator Position .... 0 21Tilt Program and l'itcb Veiocity Vector Angle ...... 0 22Pitch Plane Wind Component and Free-Stream Angle-of-Attack .... 22Comparison of Pitch Parameters lII'itbDesign Criteria. . . . . . . 23Yaw Attitude, Angular \~elocity and Average Actuator Position. . . .. . ...... 23Yaw Plane Wind Compaoent and Free-Stream Angle-of-Attack. . . . . . . . . 24Comparison of Yaw Plane Simulation Results lII'i!;hTelemettred ~3;)-95sec) 24Comparison of Yalil' Parameters with Design Criteria .............. 25Roll Attitude and'Aver:tg!e Actuator POsitions .. . . . . . . . 25Pitd. and Vaw Control AcceJerometerdtcr OECO ...... 26Angular Velocity Roll. . . . . . ... 0 0 27Angles-of-Attack During Reentry 0 0 0 27Telemetered andCalculak>d Normal Aecelentions o..ring Reeatry. . . . . ~Pitcl1 and Va-'Control .t\cceieratioDfs. . , 0 21Representative ..\ctu:dor Leads.. . . .0 0 0 29

vi


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Figure7-167-17

LIST OF ILLlTSTRATk>NS (Cont'd)T~ ~~

Non-Control Actuator Loads. SA-.f. . . . . . . . . . . . . . 29AttitudE' Difierences Between ST-90 and ST-12.fP . . . . . . . . . . . . . . . 30

7-18 Sloshing Amplitudes after 95 Seconds . . . . . . . . . . . . . . . 317-19 SA-4 Response to Sloshing ..................... " 318-1 ST-90 Platform Alignment Relative to Vehicle Axes. . . . . . . . 328-2 Guidance Comparisons. ST-90 rTelemetered-Caleulatedj , . . . . . . . . . . . . . . 338-3 Guidance Oomparfsons , ST-124P Guidance System . . . . . . . . . . . .. 348-.f Cross-Range Velocity r Telemetered) . . . . . . . . . . . . . 359-1 Power Supply and Distributor Networks. . . . . . . . . . . . 3610-1 Bending Moment and Normal Load Factor. . . . . . . . . . . . .. 37!0-2 Bending Moment at Station 979, Angle-of-Attack and Gimbal Angle Comparison. . . . 3810-3 Pitch Moment Prior to Loss of Signal. . . . . . . .. 3810-4 Yaw Moment Prior to Loss of Signal ...... , . 3810-5 SA-4 Maximum Dynamic Response. . . . . . . . . . . . . . . 3910-6 Retro Pocket Assembly ...... ' ....... " 3910-7 Nose Acceleration Envelope and Frequency Trend, Pitch. . . . . . . . . . . 4010-8 SA-4 Bending Modes - First Mode, Pitch. . . . . . . . . . . . . 4110-9 SA-4 Bending Modes - First Mode, Pitch. . . . . . . . . . . .fl10-1010-1110-1210-13100H10-15

Spectral Frequency Distribution and Composite GaMS Level on Spider Beam . . 42SA-4 Vib' ation Envelopes. . . . . . . . . . . . . . . . . . . . . . 42Spectral Frequency Distribution and Composite GallS Level for the Engine Area. . . . 43Spectral Frequency Distribution and Composite GRMS Level on ST-124P Roll Axis Gimbal.. 44SA-4 Vibration Envelopes. . . . . . . . . . . 44Overall Time-History of the Saturn Acoustical Measurement at Station 889(insidel on Fin IV . . . . . . . . 4;;

11-1 Base Pressure Minus Ambient Pressure. . . . . . . . . . . . . . 4611-2 Ratio of Base Pressure to Ambient Pressure. . . . . . . . . . . . A611-3 Differential Pressure - Engine Compartment. . . . . . . . . . . . . . . . 471-4 Compartment Pressure Locations. . . . . . . . . . . . . . . . . . . 4711-5 Differential Pressure, t;per Compartment Minus Shroud. . . . . . . . . . . 471-6 Differential Pressure. Engine Compartment Minus Shroud. . . . . . . . . . . . . . . 471-71-81-91-101-111-12

SA-4 Base npe ratures . . . . . . . . . . . . . 48SA- 4 Heating Rates . . . . . . . . . . . . . . 49SA-4 Hea..ting IQ.tes ,. _ _ 49SA-4 Base Temperatures . . . . . . . . . . . . . 50Typical FOA-ard Side Heat Shield Temperature History. . . . . . . . . . . ;,0SA-4 Propellant Tank Temperatures. . . . . . . . . . . . . . . 51


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Figure11-1311-1411-1'511-1611-1711-1811-1911-2011-2111-2211-23U-2412-1

LIST OF ILLlTSTRATIOl'l.'S rConr'd)Title

Pressure Ratio, SurfaC to Ambient ..........SA-4 Interstage Pressure and Sound Measurements ....................Surface Preesure Coefficient ........................Surface Pressure Coefficient .......................Power Spectral Density .....................Acoustic Level (L64-20) ...................Acoustic Level (L65-20) ......................S-IVD and Interstage Thermal Measurements ..................

Page5151525 : ; :535~5454

Temperature History, S-IVD Stage Skin. . . . . 55Temperature History, S-IVD Protuberance A rea ................S-IVD and Interstage Skin Temperatures ................Temperature of Thermal Sensors. . . . . . . .............Typical Signal Strength. . . . . . . . . . . . . . . . . . . . . . . . . .

55

5659

12-2 UHF and Azusa Signal Strength . . . . . . . . . . . . . . . . . . 6012-3 L1>OPSignaIStrength .................... 6112-4 Radar Signal Stre.ngth. . . . . . . . . . . . . . . . . . . . . . . . . 61A-I Low Altitude SA-4 Launcb Time Wind Components. . . . . . . . . . . 66A-2 Higb Altttude SA-4 Launch Time Wind Components . . . . . . . . . . . . . 66A-3 Low Altitude Wind Shear Components (1000 01). 67A-4 Higb Altitude Wind Shear Components 0000 011 67


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Table1-13-14-1

:

LIST OF TABLESTitle Page

T'Imes of Events ..... ,. . 369

Propellant Loading Values .................Cutoff Conditions. . . . . . . . . . . . . . .......... ,..........

4-n Significant Events . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114-111 V eloc ity lJ:a.in . . . . . . ,. .. ,. . . . . ,. . . . . . . . .. I f5-1 Engine Ignition and Cutoff Information. . . . . . . . . . . . . . .. 135-11 Engine Cutoff Impulse. . . . . . . . . . . . . . . . . . . . . . 145-m Retrv Rocket Parameters. . . . . . . 186-1 SA-4 Vehicle \\'eight. . . . . . . . . . . . . . . . . . . . . 206-U Mass Characteristics Comparison. . . . . . . . . . . . . . . . . . 21)7-1 Maximllrn Pitch Plane Control Parameters. . . . . . . . . . .. 217-117-m7-fiJ7-V~-I10-111-1l1-U1 2 - 112-1114-1

Maximun. Yaw Plane Control ParametersMaximum Roll Plane Control Parameters 25

26263339

Roll Mment.. . . . . . . Thrust Vector Angularity During Cutoff Decay ..............Gu.ida.nce Compa.risons . ,. . . . . . . Retro Rocket Alignment Tolerances .............Comparing SA-3, SA-4, S-IVD Skin Temperatures ...........Comparing SA-3, SA-4, S-IVD Healing Rates .........Mea&urement Malfunctions .....................

5555575865

SA-4 Telemetry LinksSA-4 Special Missions


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AbbreviationAEDCAFCAGeAMARDCC rvCpEBWFMGBIG6CGG[ECOMFVKlSTRAM11LVOECO

qRFlOISs-I5-IVDS-VDSPGGSSTIlTPTCOOPUHFUPWTVHF

ABBREVIA nONS AND SYMBOLS

DefinitionArnold Engineering and Development Centerautorr.atic freq..,ncy controlautomatic gain controlamplitude modulationAir Research and Development Centerthrust coefficientpressure coefficientexploding bridge wirefrequency modulationGrand Bahama IslandGuidance and Controlgasg~ratorinboard engine cutoffmain fuel valveMissile Trajectory Measurementmain WX valveou.tboard engine cutoffchamber pressarepWse ampliblde modulationpaise code modIIlatiOllpower spectral densitypropeUul ulliliUtiondylamic pressureradio freqaeacyroot mean 8qIaftfirst (booster) stage of 1heSatum Ivehiclesecorad (dllmmy) 8Iqe of the Saturn I~clethird (dammy) sa.ge of theSdun Ivehiclesolid pnpellalll gas generalnrsiJl&lesi .....telelDelier

Ultra 1 I i & I t FI\JII eDCYDoppleraItn. IIiP fnq.eDCY


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CONVERSIONFACTORSFOR PREFERRED MSFCMEASURINGUNITS

Quantity Multiply ! ! Y To ObtainAcceleration f t / s 2

-1 m/s2. 3.0480Oxl02. Area in2 6.4516xl0""t m23. Density Ib-s2 5.25539xl01 kg-rnr- m

slug/ftll 5.25539xl01 kg-rm4. Energy BTU 2. 51996xl0-t kcal5. Force Ib 4. 53592xl0-1 kg6. Length in 2. 5400Oxl0-2 m~ Mass

Ib-s21.48816 kg-s

2 = TMUI. ft m8. Mass Flow Rate Ib-s 1.48816 kg-sft m

Ib/in2 -2 kglcm29. Pressure 7.03067xl010. Temperature DF_32D -I DC. 55556xl0U. Velocity ft/s 3.0480Oxl~-1 m/s12. Volume gal (U.S.) -3 mll3.78543xl0

W 2.83168xl0-2 m313. Volume Flow Wis 2.83168xl0' mli/s

gal/s -li mli/s. 78543xl0


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GEORGEC. MARSHALLSPACE FLIGHTCENTER

MPR-SAT-63-6

RESULTSOF THE FOURTHSATURNI LAUNCHVEHICLETEST FLIGHTBy Saturn Flight Evaluation Working Group

SECTIONL FUGHT TEST S'JMMARY

FLIGHTTEST RESULTS

Saturn launchvelticle SA-4was launched at 1511:55urs EST on Man:h 28, 1963. The flight test was amplete success, as were the first three Saturn flightts. The flight test did not reveal any malfunctionsdeviations whichcould be considered a serious sys-failure or design deficiency. SA-4 completed the

ght testing of the Saturn I, Block I vehicles, with allissions assigned tothis blockofvehif;:les having been

SA-4was launched approximately eight weeks afterrival of theS-Istage at cape canaveral. The sched-ed lO-hour countdown began at 0330 EST on March, 1963. Three holds were encountered during theuntdown. The first hold was called at T minus 100inutes to evaluate an out-of-tolerance attitude indi-tion in the ST-90 Gyro Platfonn. The second holds called at Tminus 65minutes due to problems withe theodolite and the telemetry ground calibrationstem. The thirdand final holdwas called at Tminusminutes because the "open" indication of the LOXbbling valve was not received. All automatic pro-llant loading sequencing processes were within ex-cted tolerances. Launch preparations, execution ofcountdown, and the launch itself8.l*-indemonstratede compatibility of the ground support equipment ande space vehicle. The launch complex and supportuipment suffered only the mir>ordamagIe normallypected for a launch of this nature.The actual flight path of SA-4 was close to nom-

al. Slightly higher acceleration and an early cutoffthe actual altitude to be higher and the actual

nge to be less than nominal throughout the nigllt.e loss of telemetry (vehicle breakup. occurred at.05 sec at an altitude of 26. 95kilometers.Theperformance of thevehicle propulsion system

s wellwithin satisfactory limits throughout the night

test. Performance of individual engines, hydraulicsystems, and propellant tank pressurization systemsdid not deviate significantly from predicted values.The vehicle longitudinal thrust and specifiC impulsewere approximately 1.4 and 0.7 percent higher tharpredicted, respectively. No adverse effects werenoted from the programmed cutoff of Engine 5 at 100sec of flight.

The accelerometer control, flown for the firsttime on SA-4, functioned properly. An 8-engine tiltprogram, biased for a 50 percent March zonal wind,was used for the first time on this vehicle. This re-sulted in a decrease, compared to earlier nights, ofcontrol angles experienced in the pitch plane duringthe period of maximum dynamic pressure.

Operation of the hydraulic actuators and the con-trol computer were satisfactory.

The Saturn SA-4 vehicle was flownwithout activepath guidance. However, passenger hardware for bothST-90and ST-124P (Prototype) guidance systeml'" wason board to establish operational capabilities of theguidance equipment in the Saturn night environment.The telemeterer' data as well as trajectory compari-sons confirmed satisfactory petformance of both plat-forms.

The last aligDmentcheck before liftoff showed theST-90 platform biased 0.023 leg about the X-axis.This misaligDment is substantiated by an error of ap-proximately 1.0 mls in cross-range velocity at out-!Joard engine cutoff.

In general. the vibration levels for SA-4 weresimilar to those recorded on previous Saturn nights.Onlyone deviation from expected levels was recordedduring flight. This was a measurement on Engine 1turbopump which indicated a level two times that forany other comparable night measurement.


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uCC IIE II 0.> t"'I ~~+ ~~. '0 N >.. . . . . t"'IC II I c o .. . . I>. . . . 0. c . . . . . . . >.to . N0. ... c o .:s . . . c I. . .

C I I C I I'" . . . . . r >.e C l . . . .C C C C~o .. ....~ . . . .. . . . . . . e . . . . . 0.0.>'+ ~~ + - c : *

..if'+

. . . .....r-""--+-~~~~~""----- ....

H~ir~ ~{

.--


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Strain gauge data indicated a very smooth flight,was evidenced by the r elatfvcly low bending mo-nts encountered. The SA-4 bending oscillationsowed no si!!'l1ificant difference from the response of

First-mode bending compared well with the-D test and calculated predictions.

The base region =nvtronment during the SA-4ht was similar to that observed on the previousturn flights except fOI temperatures on the base,

indicated absolute levels conststently 100 to0C cooler. Analysis of the temperature data indi-es that water was present on the forward side ofheat shield,Heating rates in the general area of Engme 5 de-

ased with engine shutdown as expected,Static pressure measurements taken on the sur-

e of the S-I and 8-IVD stages agreed well with ex-ted values. Fluctuating pressures measured nearI-beam fairing showed a maximum power spectralsity (PSD, of 4-05.4 (kg/mz) z/cps " 10-7 at ap-

200 cps for the time interval between 7074 seconds.A total of 610 flight measurements were flown on

Of tbese, 3 were completely unusable, 7 weretially usable, and 9 were questionable. The signalngth of all RF systems was verv close to the ex-ted values. 'TEST OBJECTIVESThe objectives of the Saturn SA-4 flight test were:

st Objedivc - Boostel'Prove the propulston system. structural design.control system of the high thrust booster.

ond Objective - Ground Sueport EquipmentProve the operational concept of the associated

launch facilities (or Saturn class vehicles,ch include propellant systems, automatic checkoutipment, special instrumentation, launch pedestalholddown arms, and other necessary handling andching equipment. Achieved.

rd Objective - Vehicle in Flight(a~ Aeroballistics

Confirm valueso( aerodynamic cbaracteris-s, correlating predicted stability and performancethat encountered in flight. Achieved.

f b) Propuls ionProve that the booster stage is capable of

providing the prope r thrust to propel the Block I ve-hicle through the desired trajectory at the requiredvelocity. Determine the in-flight performance of alleight engines. the cont rotl ing movements of the fouroutboard gimballed engines, engines' cutoff, propel-lant uttl izat ion, and othe r desired propulsion data.Achieved,Ic) Structural and Mechanical

Verify the structural integrity of the Block 1airframe by correlating theoretical calculations andspecification requtrements with conditions encounteredduring flight. Specifically, deter mine the inflightstress, vibration levels, and associated frequencycontent at various locatrons throughout the vehiclestructure, so that the dynamic increments to the shearand bending moments can be calculated and componentvibration environment can be determined. Measurethe O\'\. .-all structural response to define critical dy-namic occurrences. E\ ~hate the presence of any ex-cessive strainor body bending effects, andaccumulatedata which can be used to determine the mode shape ofthe bending curve during flight. Achieved.

(d) Guidance and ControlDetermine the capability of the G&C system

(a modified ST-90stabilized platform) to perform therequired control, guidance, and operational sequencefor the Block 1 flight tests. Specifically, prove thatthe system will establish an accurate space-fixed co-ordinate reference for determ ining vehicle attitude andproviding an accurate coordinate velocity signal.Achieved.Fourth Objective - Engine Out

Demonstrate the capability of the 8-1 booster toperform a lmtted mission in the e\'ent of an enginefailure after liftoff. Achie\'ed.

TABLE i-t, TIMES or EVENTSr----~_~_~~~~- ___i- -- ---~,~~-__,_--- -~.~,,-~_..,~ .~ .- - ,---'-~~--'..... nh--- -~

-~":'..!~-~-_


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SECTION11. INTRODUCTION

Saturn launchvehicle 8A-4was launchedat 1511:55ESTon March 2", 1963, from Saturn LaunchComplex34, Atlantic Missile R:'nge, Cape Canaveral, Florida.SA-4 was the fourth vehicle to be flight tested in theSaturn I R" D program, The major objective of thistest was to evaluate the designs of the propulsion sys-tem, control system, and structure of the 590, 000k~(1. 3 million Ib] thrust booster.

This report presents the results of the Early En-gineering Evaluation of the 8A-4 test flight. The per-formance of each major vehicle system is discussed

with special emphasis on malfunctions and deviations.

This report is publishedbythe Saturn Flight Evalu-ation Working Group, whose members are represent-attvos from all Marshall Space Flight Center Divisions.Therefore, the report represents the official MSFCposition at this time. This report will not be followedby a stmllarly integrated report unless continuedanalysis and/or new evidence should prove the conclu-sions jJlesented here partly or wholly wrong, Finalevaluation reports may. however. be published by theMSFC Divisions covering some of the major systemsand/or special subjects as required.


19/87

SECTIONW. LAU~CHOPU:ATlONS

3.1 SUMMARYSA-4was launched on an aztmuth of 100deg East

of North from Complex 34;Geodetic Latitude 28.52153deg North and Longitude 80.56136 deg West.

The scheduled 1O-hourcountdownbegan 0330hoursEST Thursday, March 28. Three holds were calledtotaling 1 hour and 42 minutes. The first hold lasted20 minutes due to an out-of-tolerance indication in theST-90 gyro platform (remedied bya bias adjustment);the second, lasting 40 minutes, was called because ofproblems with the ST-90 theodulite and the telemetryground calibration system: the third and last hold re-sulted from the lack of an open indication in the LOXbubblingvalvecircuitry. Anelectrical bypass was ef-fected in the blockhouse permitting resumption of thecount after the 42 minute hold.

AUautomatic operations, sequenced from Tminus364 sec through liftoff, performed normally includingthe manually-operated LOXbubbling interlock. Launchpreparations, execution of the countdown, and launchwere as expected and successfully demonstrated thecompatibility between the ground support equipmentand the flight configuration.3.2 PRELAUNCHMILESTONES

DateFebruary 2, 1963 SA-4 (all stages) arrived at

Cape Canaveral on Saturnbarge "Promise"

February 4, 1963 S-I booster erected on launchpedestal at Pad 34.

February 5. 1963 Dummy stages S-IV, S-V. andpayload assembled to the S-Ibooster

February 20. 1963 Fuel loading simulation andfuel malfunction sequencetests completed.

February 25, 1963 IA)X loading simulation testand RF subsystem test com-pleted.

Mareb 2, 1963 Networks - Powerplalll overalltest number 1 completed.

March 4. 1963 Networks - Powerplantoveralltest number 2 completed.

Marc~ 15, 1963 Guidance and Control overalltest number 3 completed.

March 20, 1963 Guidance andControl plugdropoverall test number 4 com-pleted.

March 27, 1963 Simulated flight test com-pleted.

March 28, 1963 Launch.

3.3 PRELAUNCH ATMOSPHERIC SURFACE CON-DITIONSWeather conditic .s around Cape canaveral at the

time of launch were good. The visibility was 16km(10 miles). Cloudcoverage along the flight path con-sisted of I-tenth cumulus, base 760 m (2500 feet) andI-tenth cirrus, base 1830m (6000 feet). BarometriCpressure was 763mm of mercury (1017.6 mbs) , rel-ative humidity 70 percent, and temperature 24.4C.Surface winds were from 45 deg at ;)m/s.

3.4 COUNTDOWN. The launch countdown began at T minus 600 min-

utes at 0330 hours EST on March 28, 1963and wascontinuous except for the followtng holds:Holds

1. The first holdwas called at T minus 100min-utes of the countdown (1150 hours EST) to evaluate anout-of-tolerance attitude indication inthe ST-90 stabi-lized platform. This condition was remedied by biasadjustment. The indicated bias was O. 12 deg in yaw.Acorrection of O. 04


20/87

:!. The thi rd and final countdown hold was calledat T m ir.us 1!1 minutes (I-t! 1 hours ESTl for -t2 min-lilt'S because till' "open" ind icut ion of till' LOX bubbl ing"31,'c was not rece-ived, even though the valve wasobviously olll -rat ing properly, An e-lect rtcal bypasswas effected in the blockhouse. a llow in~ the count toIX' resumed at H;j:l hour-s EST and continued untillaunch, Fi~Jl'{' :1-1 shows till' ucoumulutod SA-4 holdtimc ve-rsus countdown time.

Ar cumu let t~~'Hold TimP (mtn)Ico

100 L OX B uh t1 1 iflllV.. Ive

In(~1cat 10n

60The-odoll teand 'nil Gr-oundCal rbr-et ion

20

oL-----~J'~------~~~~~----~------~T600 r-i -o r-so

Countdovn Ti_ (min)T-40 T-O

FIGURE 3-1. SA-4 HOLlEAutomatic Countdown

The automatic countdown sequence was initiated bythe firing command at Tminus 364 sec prior to ignitioncommand (T minus 01. All automatic ope rat ions , se-quenced from this time through liftoff, performed nor-mally including the manually-operated LOX bubblinginterlock.3. j PROPE LL.\XT LOADIXG

The Saturn propellant loading system was designedto tank propellants to a given total weight at a ratio toprovide simultaneous propellant depletion at outboardengine cutoff. A final fuel density check is performedprior to the flight. Based on the final fuel densitycheck, fuel is drained and LOXis tanked to the propor-tions necessary to provide the total propellant load de-sired. The system was designed to load propellants towithin O. 25 percent total weight accuracy.

6

The prope llam Ioading s_"stem indicates that a tot;11of2!'i,(i-t-t kg (h;l4, H" Ih) of propellants wen' I ldld.Liquid 1(_'v(_,1robe data, in conjuru-t ion with stmulutcdflowrate dat.a, incuc-atc that a tot.rl of 2"(;, !J!J7 kg (i:l2,7:!1 lb] of propcl lants wen' IO:Hk'd. Loss of centertank liquid leve Iprobe signals and the fi r-st probe suma Ifrom an outboard tank tend to dec-rea se the va liditv of. .the level probe data: however , this weight agrees towithin approximately o.:! pcrr-cnt 01 the loading svste mweight abovo, thus ve I'ifyi ngthe I!.llllI'("{'nt propel lantloadi nl{ systc-m accu racy.

Fue-l wa s loaded tho day hpfor(_' launch to 0\'1..'1' 100pe rcent of the desil't'd fill va lue , LOX \\ as loaded theday of launch to under 1no perc-ent of the dcst red fiII\';1 IUt'. Ba sed on the fina Ifuel den si tv c-hec-k, a -0. OU"WIkg!t'm2 (-0.11;1 psi) IucI co rrcc-t iou and a +0.00;):!7kglem2 (+0.07:; psO LOX cor-rcct ion we re dialed intothe re spccttvc tanking computer-s. The LOX c-or-rec-tionautomatically initiated till' LOX fill seque-nc-e to tank tothe desired propellant ratio, The adjust leve-l dra insequence \\ as then manua llv init iatr-d to drum the tanksof cxr-e ss Iuo l,

Table :;-1 lists the readings of the final tankingcomputer, Iue l density c-omputer , backup manometer,also computer readinu deviat ions lrom backup ma-nometer.

T.\BLE :;-1. pnOPELL\XT LO.'\DIXG'.\LlJE~

Back-upCornpute rDeviat ionfrom

Computet'(kg! cm2,

Manometer :\131101llt'tel'( kghm", ____;(,-,~~Q-"-_

LO::-:! P 1. 1527 1. 1542 n.l:l0Fuel.lP O.~t-lii O. ~,,!I!) O.l:ljFuel Density.lP 0.7962 O.79(i-t O.02jNote: Computer readings not avarlable: value listed is

from propellant loading tables.

Fuel density versus fuel temperature ( Fib'1I1'e3-2)was determined from fuel chemical analysis data.

.Of,

Ijetermme-o tl'l. Cht:!nl('al AruL~:;;15o BlOC'kbOUll l" Temp. l ll i. . ao _ " 'Y ' f " & g { ' . .. uft'd PnmJ.JInlet Te-mp. (f\f'r F"h.mt~ ~ :~ of Bloc""ou... T


21/87

Blockhouse temperature data, density probe data,average of blockhouse temperature and density probedata, and average measured pump inlet temperatureover flight are included in Figure 3-2. The differenceinblockhouse and density probe data temperature was1.81 C which corresponds to a density difference of0.96 kg/m33.6 HOLDOOWN

Engine start and transition were smooth with allengines receiving ignition from a LOXlead in the gasgenerator ignition system. All blockhouse redlinevalues were within limits, and no holds were encoun-tered because of redline discrepancies,

3. 7 GROUJl.TDUPPORT EQUIPMENTThe general condition of the ground support

equipment was satisfactory. Only mtnor damagenormally sustained for a launch of this nature was ex-per+enced, The flame deflectors sustained no appre-ciable damage. The LOXfueling mast was damaged,and there was some scorching of paint at the lower endof the umbilical tower.

~. 8 LAUNCHFACIUTY MEASUREMENTSThe static pressureandvibration measurements,

located on the umbrlical tower during SA-3 and SA-4night, indicate that the gas dynatdic pressure actingon the tower may impose a severe loading condition.Whenthe engines ignite, the air surrounding the toweris sucked into the rocket engine exhaust stream, re-sulting in a flow pattern across the tower. Under-pressure recorded on SA-3 and substantiated by theSA-4 flight indicates that the magnitude of the loaddistribution imposed on the tower is significant enoughto warrant further investigation.

The vibration measuremects, located on theegress beam, also indicate the presence of & . h e flowfield across the tower. There is a vibration buildupat ignition, with a steady state level being maintainedthrough holddown release. Afurther buildup occurswhen the base of the vehicle reaches the 54.3 m (178ft) level of thE' tower. The Vibration level showed nochange at liftoff, indicating very little vibration beingfedthrough the support pedestal to the base ofthe tower.The vibration. for the most part, is being caused bythe acoustic and gas dynamic environment.

!---~- ..- .... -- ~


22/87

SECTIONIV. TRAJECTORY

4.1 SUMMARYThe actual fligtu path of SA--twas close to nom-

inal. Due to higher acceleration and an early cutoff,the actual altitude was higher and the range was slightlyless than nominal throughout the flight. The loss oftelemetry occurred at 398.05 sec at an altitude of26. 95 kilometers.

At OECO the actual altitude was 1.55 km higher,the range 0.21 km shorter, and the velocity 12.7 m/sgreater than nominal.

The first engineering tests of the MISTRAMrack-ing system and the radar altimeter for the Saturn ve-hicle were successful.4.2 TRACKINGANALYSIS

Tracking data was available from first motionuntil loss of telemetry at 398.05 seconds. The re-duced tracking data (especially Azusa and llDOP) ,after40sec of flight. bad smaller random errors thanon anyprevious flight. Thedifferences betweenAzusa,MISTRAM.andUDOPtracking systems (corrected fortheir respective antenna locations), during the poweredportion of flight. were 5 meters or less.

The postflight trajectory was established fromfixed camera data (0 to 15 sec), UDOP (15 to 95 sec),telemetered guidance (95 to 132 sec), Azusa (132 to150 sec), and a computed trajectory (150 to 397. 8 sec)using Azusa data at 150 sec as initial conditions.

SA-4 was the first engineering test (\f theMlSTRAM tracking system on the Saturn vehicle.Tracking data from the Valkaria site was reduced inthe following intervals:

35. 75 to 125.75 seconds14-4.35 to 172.70 seconds173.70 to 182.20 seconds219.30 to 313. 30 seconds

The tracking data obtained was comparable inc;aality to that obtained from Azusa and tJDOP. Thedeviations between MlSTRAMand the other two sys-tems were 5meters or less during the powered flight.

An engilleering test of the radar altimeter wasflown on SA-4. The altimeter was designed for usewith the Block n vehicles, where the vehicle is es-sentially flying parallel to the earth's surface. Nodata was obtained from the altimeter until about 105

sec range time, when the vehicle had reached a tiltangle of ..bout 42.6 deg from the vertical. This datawas not obtained from the main lobe of the antenna,but from oneof the s ide lobes; consequently, the qualityo f. the data was not the same as can be expected fromBlock II flights.

The output rate of the altimeter was 36 samplesper second, with half-second time markers given.These time markers are not r.eferenced to range time,but the necessary correlation can be determined afterthe flight from telemetry data. Valid output from thealtimeter was obtained from about 105to 125seo rangetime.

The general trend of the radar altimeter outputshows that the output has: t) bursts of noise that oc-cur about every second, 2) an altitude of about 30moccurring at OECO, and 3) a random error of about 30 m, which is within the altimeter specifications.The cause of the bursts of noise and the altitude shiftat .OECOhave not yet been completely determined.

The output of the alttmeter is compared with theactual trajectory in Figure 4-1.

FIGURE4-1. ALTIMETER DEVIATIONThe cable associated with the altimeter and the

distance between the center of gravity and the aitimeterantenna contribute about +15to +20m bias in the dif-ferences. The oscillations about the mean are causedby the noise in the altimeter output. The deviations,altimeter minus actual trajectory, vary from +40 to-20 meters.4. 3 ACTUALANDNOMINAL TRAJECTORY4.3. 1 POWEREDFLIGHT

Actual and nominal altitude, range. and crossrange (Ze) are sho.... in Figure 4-2. The actual al-titude was higher than nominal, but the range wasshorter than nominal. The cross range (Ze) was 0."km left of nominal at IECO {see Table 4-11. About

8


23/87

O.l26km of this deviation was due to the duference inalignment of the platform and vehicle (see Chapter8.0), and approximately O. 208km was caused by thelateral winds. The remaining difference (0.066 kmtis due to other small effects, The nominal trajectoryis presented in Reference 1.

The actual and nominal earth-fixed velocity isshown inFigure 4-3. Machnumber and dynamicpres-sure are shown in Figure 4-4. These two parameterswere calculated using measured meteorological datato an altitude of 31.5 kilometers. Between31.5 and47.0 km altitude, the measured data was adjusted tothe 1959ARDC~'lmosphere, above which the 1959ARDCwas used. The actual peak dynamic pressurewas slightly less (0.002 kg/cm!) than nominal.

A comparison of actual and nominal parametersat the three cutoff signals is shown inTable 4-1. Out-boardengine cutoff signal was initiated by WX deple-tion, thus allowing the cutoff signal to occur slightlyafter thrust decay had begun. At OECO, altitude was1.55 kin higher, range was 0.21 km shorter, and theearth-fixed velocttyw 3S 12.7m/shigher than nominal.

The longitudinal acceleratron was slightly higher

from first motion until IECO (see Figure 4-5). Theactual andnominal longitudinalacceleration during thecutoff period is shown in the lower portion of Figure4-5. A comparison of actual and nominal parametersat significant event times is given in Table 4-11.4. 3. 2 THRUSTDECAY

Thevelocity gain from the thrust deca.. of thevarious engines after cutoff signal is shown in Table4 -m.

Thevelocity gain fromoutboard engine thrust de-cay has no significance for cutoff impulse from theseengines, since thecutoff signal occurred slightly afterthrust decay had already begun.4. 4 RETROROCKETSThe retro rockets ignited and burned for about

2.14sec andimparted a veloctty loss of about 7.8 m/sto the entire "A-4 vehicle, since no separation wasscheduled. Thelongitudinal acceleration during retrorocket operation is shown in the lower portion ofFig-ure 4-6: Themaximumdeceleration during this peri-odwas 3.8m/s!.

TABLE4-1 CUTOFFCONDmONS

~c.o. rrco O[COParameter

Actual Sominal 1 Acr-Nom Ac-tual Nominal Act-~om Actual :Sominal Acr-jcomRange Time (sec] 100. sz lOU . 5~ * . . . . ._lI~ 11:). " ' ~ l 11-1.-I" -(I. ,r; 121.12 121. 67 -0 ?).)Rang


24/87

. . . . . Altlt.dto(Ia) ( .. . )' ' i f ) '4i- f -''''' " i : O f ) . + -II

"

~.,p

""1 -~~;(l{ "" - + - ~ -Iivo M i100 40 ~-~ -.

'" 2 .o o 4 . '0 ten

c.._~ ....... Z . Cia), .

---_---.-- +----~----

10FIGURE 4-5. LONGI'romNA L ACCELERATION


25/87

~\ l III1 Fll ~l ,t.tronI(--I, 1\1:'h

~-~-----\ Ma..\'11l1Ll!ll Dvnamn Pl'l'~:"'>Ui'(',I

Ma.'.,tllUIll La rth-f ixe-d\'l'loelt~ (P\I\l,t_'!'I,d Phn sc-

- - - - - - - - - - 1

TAB LE 4-ll SIGNTFlC A~" EVEKTS

Actual :-';ommal- - - - - + O{I\1. HI H. -illHangt' Ttou (~t_., , -. ;)0. l:J :12. 21 -2.oti

AltltUt'k' (~mi -0.24+ Ran:'"


26/87

SECTIONV. PROPULSION

5.1 SUMMARYThe pe rformanee of the vehicle propulsion sys-

tem was wellwithin satisfactory limits throughout theflight test 0 1 " Saturn SA-4. Performance of individualengmes , hydraulic systems, andpropeltant tank pres-surtzatton systems did r . deviate sign.ficant.ly fromthe predicted values. The vehicle longitudinal thrustwas 1.3 percent and specific impulse 0.7 percent" :"It", .. ihan corresponding predicted values.

All missions, Ineludtng primary, secondary, andspecial, weve accomplished. Results of special mis-sions o( particular sigll~ficance to the vehicle propul-sion system are described below:

1. The shutdown of Engine 5 at 100sec pre-sented noproblems tothe propulsion system andprovedthe feasibility of the "engine out" concept.

2. The thrust OK cutoff of outboard enginesdue to LOXdepletion again showed a significant in-crease in propellant utilization wth no problems inengine shutdown and vehicle control.

3. The retro rockets again ignited and opera-ted satisfactorily at the ead of the S-I stage poweredRight.5.2 INDIVIDtTALENGINEPERFORMANCE

The performance of the individual engines on theSA-4 flightwas satisfactory, The maximum deviationin engine thrust between that calculated (rom flightdata and predicted values was approximately 3. 3 per-cent, occurring on engine position 2. The deviations(or the other engines varied (rom -0. 3 to +2.7percentas compared to ~he predicted thrust (see Fig. 5-1).

FIGURE5-1. Thil>IVIDUALNGINEDEVIATIONFROMPREDICTEDTHRUST

The engine-to-engine deviation from the actual meanthrust was (rom -1.3 to -1.9 percent.

Themaximumdeviation inengine specific impulsebetween that calculated from flight data and the pre-dictedvalues was approximately +1.7 percent, oecu"ring on engine position 1. The deviations for the utherengines varied (rom-i.61.0+0. 9 percent as comparedto the predicted impulse (see Fig. 5-2). The engtnc-to-er 6ine deviation from the actual mean specittc im-pulse was from -1.3 to +0.7 percent.

. - - ;. . .-_.... . . . . .-I

--~~~L-~~~~~~~~~-~ . . . . . .FIGURE5-2. INDMDUAL ENGINEDEVIATION

FROMPREDICTEDSPECIFIC IMPUlBEIt should be noted that a considerable amount o(

devtationbetween predicted and actual thrust and spe-cific impulse is due to di((erenct;s in the thrust coef-ficient (C(v) used in the predicted data and that usedin flight evaluation. The predicted data incorporatedthe C(v's obtained (rom engine log data; however, (orthe following reasons, the nominal C(v was used forflight evaluation: (1) burning time was shorter thanpredicted and cutoff velocity was higher than predic-ted: (2) the prediction of engine parameters, such asengine combustion chamber pressure and propellantpump speeds, agreed with flight data; and (3) the av-erage individual engine C(v (engine log data) was 1.0percent below nominal.

Engine main propellant valve opening and closingtimes (Table 5-1) were acceptable (or all engines.Two small discrepancies, neither o( which is con-sidered significant, occurred during the ignitionphase.Engine position 7 experienced an ignition delay in theSolid Propellant Gas Generator o( about 45 ms whichresulted in an altered starting sequence (see Fig.5-3). The chamber pressure measurement on engineposition 6 indicated a decrease in chamber pressureduring transition. Sibci! this decay is not supported

12. f : . . .


27/87

by gas generator injection JJl~~bsure.the only relaudcondition capable oj' reproducing this tranf'lient ,chamber prCs8UT'{' measurcmera failure, is not 8US-.peca;ed. 1 '1' tX't'f'Ot'manOC'indicated a vaHI'f meas-urement on this etCine.

Ay"kt_. vah'(' was incot',JOratt'd on ttwSA-4 ('ngin{'gas gClleJ':.tt.o)' syst('m tor the first t.imo. This va lvereplaces the Independent LOX and I"U(" cracking checkvalves used on pr('\'iOIl8(ng-ineeonHgUl',ltion8. Thei~pcndent action of the check valves Pl'ts(>nt.ed tlwJJOssibility of exces stve LOX lead t ime in th(' tIS g('n-erator(.QG) and tHlbsequent tur'bineef'08ion a8SOC iat


28/87

. ,"was derived early in the Saturn program and is cur-rently being re-evaluated based ondata fromthe BlockIflights. The equivalent velocity gains during thrustdecay arc gtven in Section 4.3. 2.

TABLE 5-1I. ENGINE CUTOFF IMPULSE

4

that1M' Cutoff c-..rhOft vitta ... l.. 1 . ._1..{ ~ - . . c , {Ib-.ee, !kg- sec , I (lb- eee ,26,399 ~R.:'OO S~f" lI(!t( 3 . .n.40S 51 ,bOO20.1]


29/87

--S ...J.b4PI_----Pt.41~ . .. . . .._,-, . . .

FIGURE 5-6. VEHICLE MIXTURE RATIO ANDTOTAL FlDW RATE

There were two approaches us~ to evaluate thevehicle propulsion system performance. The firstmethod comaares propulsion system inflight meas-urements with corresponding predicted information:The vehicle thrust was calculated from measuredcombustion chamber pressures utilizing the nominalthrust coefficients (Cfy). The vehicle total propel-lant flow is defined as the total propellant expendedby the vehicle to include engine flow, lube fuel flows,and vented GOX. The engine flows are determinedby using gain relations of propellant pump speed andpump inlet conditions and comparing with thepropel-lants loaded and the residuals after thrust decay. Theseflows are considered more accurate than the flowsdetermined from flow meters. The vebicle specificimpulse was determined from tlte Vehicle thrust andtotal propellant flow described above. The second ap-proach is through the flight simulation method, wbicbis a computer program with a differential correctionprocedare used toobtain adjustments to the propulS10Uparameter iDputs, wbicbwillproduce a trajectory thatm1tches the actual trajectory.

The average percent deviation from predicted,along with the estimated accuracy limitati\JIdJ of eachparameter from both approaches. is shown below:

ThrustTotal Flow RateSpecific Impulse

Flight Propulsion+1.3(1.0) ,.+0.7(1.0) "+0. 7(d. 0) .,.

Flight SimulationH +0.67.,.*+0.99"'*

* Altbough the flight simulation was not completelysatisfactory, the accuracieti are less than %1~rcent.

Tbe deviations shown above are computed bysub-traettng predicted from actual and dividing by predic-ted. Tbe largest deviation between the two approachesis approximately O. 4 percent , whichwould be expectedof two independent methods.5.4 PRESSURIZATIONSYSTEMS5.4.1 FUEL TANKPRESSURIZATION

The fuel tank pressurization operated satis-factorily during flight. Gaseous nitrogen, suppliedby 48 high-pressure spheres, showed a pressure of204 kg/cml (2900 psi gauge) at liftoff and decayed asexpected to approximately 112kg/cm! (1600psi gauge)at OECO. Fuel tank gas pressure during poweredflight is shown in Figure 5-7. The n.easured pres-sure varied between 1.1 kg/cm! (15 psi gauge) and1.18 kg/cm! (16.8 psi gauge)

." . .. . ~ (kf.!Cl'llll.UIW')u

0.'

:0t - ~ - - + ~ + l - -I_'.~,dDl-t'J)!---J>",.:l1ct~d, ,L- _L ~1 ~ __ ~~ __ ~ __ ~0

120

I .10 100

FIGURE 5-7. GAS PRESSURE IN FUEL TANK5.4.2 LOX TANK PRESSURIZATION

LOX tank pressu.re was required during flightto provide structural rigidity to the tanks and to pro-vide adequate pressure at the engine pump inlet. Thepressurization system was designed tomaintain a con-stant pressure of 4. 2 kg/em! (60 psi). Prelaunchpressurization was acbievedwith helium from a groundsource andmaintained during flight with GOX from theengine beat exchangers.

Prelaunch instructions required that a pressuri-zation time of 80%3 sec be met during pressurizationchecks. This was accomplished by adjustment of thehelium pressure regulator.

LOX tank pressurization was initiated at -129. 33sec, and adequate pressure was attained at -38.60seconds. This pressurizing time of 90.73 sec waswltbin the required 75to 95sec. but exceeded the 80% 3sec that was established during prelaunch checkout.

This slow pressurization presented no problemsin the SA-4 cCMllltdown;however, it did create an a-wareness of a potential problem area for Block n

-_._------_-- --- 15- f-~---'"",


30/87

vehicles. The existing automatic sequence for LOXpressurization 011 Block n vehicles is shown in Fig-ure 5-8.

LOX HuH'i inV(.'nt l. Rl-licolValy(- Clo"eJ

SI"n31 LOX T;ln~: Pr-c s su r ize tt on

FIGURE5-8. CURRENTBLOCKnPRESSURIZATIONSBQUENCELOX tank pressurtzatton is dependent upon the

preceding signals, and if the LOX tank pressurizedsignal is not received within 150see of initiating LOXbubbling, the sequence halts. To minimize this pos-sibility and the problems it would introduce to the 8-1and S-IV stages at this part of the countdown, the se-quence was altered in the following manner:

The LOX bubbling was shortened to 50 sec, theLOXpressurization will no longer depend upona re-ceipt of the "valve closed" signal, and the pressuriz-ing line will be ortf'iced for a 75 sec pressurizationtime. This provides for a 25 sec safety margin, andgives the sequence shown in Figure 5-9.

LOXlIubb Uag LOXPre.surization

o 50 125 150FIGURE5-9. REVISEDBLOCKn PRESSURIZATION

SEQUENCEThe prelaunch LOX tank pressure is shown in

Figure 5-10. The SA-2 prelaunch pressure is includ-ed (or comparison because the LOXtanks' ullage vol-ume was approximately the same for both vehicles.

~ft < . . . )11 0

40

~---_1~~----_1~OO----_~IIO----~~--~----~---400.... ,... (-


31/87

of Engine 5. The high percentage of propellant utili-zation, as on the SA-3 flight, resulted from the out-board engines being allowed to deplete the LOXtankabefore being cutoff bya "thrust OK" pressure switch.Center LOXtank depletion (gas breakthrough), which8hould have occurred near IECO, occurred approxi-mately 1.0 sec after IECO, due to a 0.8 kg/cml (1.2psi) lower-than-predicted differential pressure be-tween center and outer LOXtanks.

Data from several sources may be used to deter-mine propellant utilization: flowmeter data, propel-lant utilization system data, and discrete level probedata inconjunction with simulated flowrate data. Basedonresults of this andprevious flights, flowmeter databas not been sufficiently accurate and was thereforenot considered in determining propellant utilization.A propellant utilization system was test-flown to de-termine system performance and reliability, and wasnot a control feature of the Saturn first stage. Sincere8ults from the propellant utilization system meas-8uremenls were invalid, tbis method of determiningpropellant utilization was not used in establishing pro-pellant weights and residuals. The most reliablemethod of determining propellant utilization was dis-crete level probe data used in conjunction with sim-ulated nowrate data. The performance of the overallvehicle PU was established using this method.

An evaluation of vehicle propellant utilization in-dicates that 2115 l-'t (4662 Ib) of LOX and 3478 kg(7668Ib) of fuel' !lined on board the vehicle at theend of outboan'eol.De thrust decay. This compareswell with the predicted residuals of lB18 kg (4007 Ib)LOXand 2666 kg (5877 lb) fuel. Of the 3478 kg (7668Ib) of fuel left on board, approximately 900 kg (2000Ib) were loaded as extra fuel, part of which is consid-ered bias to insure burning of aU usable LOX andthereby iAsuring LOXdepletion. Tbe following tablecompares the residuals of SA-4 with thOf;eof SA-3.

LOXFuel

SA-32,145 kg3,892 kg

SA-42,115 kg3,478 kg

The table shows that good repeatability of resid-uals can be expected wben a LOXdepletion cutoff isemployed. These residuals represent a substantialincrease in overall vehicle performance wben com-pared with tbe residuals from the timer cutoff (6 secbetween IECO and OECO) used on SA-I .H I SA-2.For example, if the 6 sec timer cutoff had been usedon SA-4, the residuals at end of thl'USt decay wouldhave been 3358 kg (7403 Ib) LOXand 4046 kg (8921Ib) fuel; bowever, the fuel load would not have beenbiased for LOXdepletion.5.6 HYDRAULICSYSTEM

The operauon of the four independent hydraulicsystems WIiS satisfactory with no deviations irom ex-pected performance. T~e cutoL "f inboard engine po-sition 5 during the flight presented no excessive bur-den on the hydraulic systems. The average actuatordemand flows were less than the pump capacitythroughout the flight. This is evident from relativelyconstant source pressures and la ck of level changesassociated with excess flow demand.

5.7 RETRO. 'OCKET PERFORMANCEFour solid propellant retro rockets were flown on

the SA-4vehicle; SA-4flight was the second flight testof the retro rockets. The retro rockets were the onlyactive part of the 8-l/8-1V stage separation systemflight tested onSA-4. The retro rockets were mounted90 deg apart 01'1 the spider beam at the top of the 8-1stage. Retro rocket thrust vectors were directedthrough the 8-1stage eenter of percussion. The rocketmotors were directed downward and canted 12 degfrom the vehicle centerline. Retro rocket locationsare shown in Figure 5-12. Retro rocket firing com-mand (125.46 sec range time) was given, as sched-uled, 12-see after inboard engine cutoff.

. JiI~..j1 -.~~~~~~~~~~,~:_t'.-t_"

FIGURE 5-12. TYPICAL RETRO ROCKETCHAMBERTHRUSTA typical retro rocket thrust curve is shown in

Figure 5-12. Telemetered retro rocket chamberpressure data indicated satisfactory retro rocket per-formance levels for the four rockets (see Table 5-llI) Shorter-than-predicted burning time and higher-than-predicted average thrust indicate a higher-than-pre-dieted propdlant grain temperature (15.6 C) for aUthe retro rockets. Measured retro rocket burningtimes varied from 2.04 to 2. 12 sec, but were withinexpected limits of 1.94to 2.36seconds. Short burningtime and highaveragetbrustof retro rocket 3 indicatea higher propellant grain temperature than on 1 2, or4; it is probable that radiant beating from the suncaused retro rocket 3 propellant grain temperature tobe higher, because retro rocket 3was more exposedto the sua prior to launch. AU retro roeket thrustvalues after ipution and prior to decay were withinthe expected limits of 15,400 to 20,200 kilograms.

17


32/87

".14.1621.IIW ETABLE s-m RETRO ROCKET PARAMETERS

ActualParameter Retro Rocket Number

1 2 3 4 Total PredictedBurning Time (sec) 2.12 2.10 2.04 2.10 -- 2.15Total Impulse (kg-sec) 34.300 34.600 34.300 33.700 133.900 33.800

(lb-sec) 75.600 76.300 75.600 74.300 301.800 74.500Average 'lbrust (kg) 16,170 16.480 16.810 16.050 65.510 15.720

(lb) 35.660 36.330 37.060 35.380 144.430 34.650Average Pressure (kg!eml) 96.8 98.6 100.6 96.1 -- - -

(psi) 1.377 1.403 1.431 1.367 - - --FiriDg Command (sec 125.46 125.46 125.46 125.46 -- - - -

range time)Time of Pressure Buildup 125.43 125.44 125.44 125.44 -- --(sec range time)

Definition of Terms1. Burning Time - Time interval between the intersection points on the zero thrust line described by a line

tangent 10 the rise of thrust at the point of inflection extended to intersect the zero thrust line and by aline tangent to the decaying tIarust curve at the point of inflection extended to intersect the zero thrust line.

2. TotalJmpulse - Area under thrust versus time curve.3. Average 'lbrust - Total impulse divided by burning time.4. Average Pressure - Area under pressure versus time curve divided by the burning time.

This was commutated measurement and could be in error :i: 83 milliseconds.Predicted nlues were based OIl propellant grain temperature of 15.6C and an altitude or 76.2 kilometers.

G G t I I.U :I A l18


33/87

SECTIO:\ VI. MASS CIJ..\RACTEIUSTICS

6.1 VEIfICLE WEIGHTS 6.2 VEHICLE CENTEH OF GHAVITY A:-';U :'1101\1-ENTS OF INEHTIA

The tota l vehicle weight was approxunatcly4"", !"I5!)kg (956,715 lb) at ignition command, Ap-proximately 21\1,mIl; kg 1619,711 101 of propcl lantwe re consumed during the S-llx)\\eJ'('d phase of f:ightIsee Figure 6-1). Table 6-1 indicates weights atvarious night events;

Longitudinal and radid contc of gravity and pitchand roll moments of ine rt.ia are given in Table H-II.These parameters are also plotted \'l'I'SUS range timein Figure 6-1.

Vehc+leW'i$t (kg)6 x 10'

Longrtudianl Conte-r 01 (;I'an1y(calibers f rom gimhal stauon l

,J,O

IIL- ~L- ~L_ ~ ~ ~ __ , ~ ~II l:!O0 GO 1011

~, ":3 . "

~, "

Holl Inertia (kg-m-s


34/87

TABLE 6-L SA-4 VEHICLE WEIGHTSICIIlTlOll " .-ID OUI'IOAID am or!'111ft aRWID marr~ l1:I. CUl 'OPF _I. curorr l1:I. CUl 'OFT ftIUft D_T. . . . . "ctua1 .. . . ".tnal ~ _41 .. 1 ...- 41" . . 1 r~.. Act . . ) . . . . . . W1I&l_T I l l E oee -3.43 -3.38 0.40 .40 I 9 .. 11 II~ 1:_. 9 IZ.ll 124.90 1242.0

IIElClll'S (ka)On Vehicle 144.'8(, 14S,loo 144,986 145,100 144,986 145,100 144,9. 14~.IOO 144,986 14~, 100 144,986 145,100L O i : )97,986 197,667 192, SO~ 1~2.142 28,69S 26,C)42 8,005 8,520 Z,108 2,341 1,818 2,115_1 M,S9' 89,978 81,480 87,851 n,389 15,OS] 6,113 6,852 3,188 3,841 = ' , 6 6 6 3,478Caa tft LOXeont..1nt'r 336 336 340 377 1,243 1,147 1,147 1,236 1,390 1,290 l.]~. 1,2'2G112 398 398 398 398 398 398 398 398 J98 398 3 )98II1dr.ullc Otl 27 27 27 27 27 27 27 27 27 27 27 27TOTAL 43:],132 433,506 42St 736 425,M!! I QO, 7 38 188,667 rso, 8i'r~ 16' 133 152,0197 152,4J97 151,281 15:',410IlEICIII'S(lb)Dry V.-hiel. 319,639 lllJ,890 319,63' 319,890 319,63' JI',890 la,6l9 319,890 119,61' ))C~."O 319,639 319,89C- 436,485 435,781 424,400 423,602 61,262 59,397 17,649 18,783 4,647 5,162 4,007 4,662_ 197,5]2 lQa,367 192,860 1",679 33,928 n,l86 13,477 15,106 7,028 ',468 5,877 7,668ca. 1n LOI OMta_r 740 740 7"9 832 2,740 2,529 2,"0 2.724 3,06~ 2,845 3,068 ',848:'Jnuuc on 87C 8n 877 877 877 877 877 877 877 877 877 87760 60 60 60 60 60 60 60 60 64l 60 60-. 955,333 955,115 938,585 938,"'0 420,5006 015,939 35J,,672 3507,440 335,316 337,102 333,528 n6,OO~IIJI'ES: 1. Act-uel dry wlgl, t include. 87,022 kg (191.850 1b) .... t .t" ball t.

2. Predicted dry V!'lIht tnclude. 86,795 q (191. 'SO Ib) _tet" beUael.3. GOI vented accounted fOT.4. QlZ vented fn fuel conteiRe".S. tee at-c_x.tlon, .ppr('l"i.mud~!.Sl..b kg (1.000 lh) .t liftoff, not included.6. tsnttlon w@lht doe. not illclu~ Jacket pftflll.7. Pre41ctH propellant __lght. baaed oe ftM.ldbty of 102.0 q:/.3 (500.07 Ib/ft3).

Actual ,r:opellant _~ht. baHd OIl ~l deaett,. of BOlt.6Iq;/.' (50.23 Ib/ft').8. ....1 con--.d {adud 0.227 qf.(0.50 Ib/ eec) h. fuel flow per ... '_.9. Ga. ia t.ca cont.1Mn doe. not iActude 282 k (622 1.) he ll . ..

"ftdlcted. CharactertaUc. are t.... ntpOned tft 11-5V!.!S-18-63; tMftfoft. the pftCflctedU of ewat. differ fhlll the....... 1 tnJectory tn Cha,te-r 4.0.

TABLE 6-1L MASS CHARACTERISTICS COM-PARISONS

. . . . . .

124.1JO

Do >

0'

...,.TDIE

IADlALC.C.

PtTCllIiIIIIW!'OF IJIDTlA

0.'D!~ V 1 < . ' I O n~O.'

e.- Ie. 1 1 - eM . ., I" " I

0' l!..CJ):!' I~ !

~tct.rd lin o...ractertdtctl .... tt.- ~ hII-I'IIfI-.-JI-'J; u .n f. _ . ~ . . . . lI:tNu-.f ..... tffet f_ de--. .. 1 ttaJect:o..,. ill~r 4.8.WftS! 1..... ictM fty_taM. bel.... _"5" (l'l.1501", . .. .. . "11- '.

2. Ad.. l ~ _~t ~l"" .'.022 .. OU,15O 1. ' -... "1l8et.

-7;"-


35/87

SECTIO~ VU. CONTROL

Aut tc . . . . . - - IICO GICO , I ;1\ I I _,A. . . . . . . . _ A I- - ~ ~V" r- II I110 I I10 20 lO ~ 50 to 70 10 90 100 120 10

A al ul er v el oe te , P te ch ~12-15)(~/.) 11. . . I : I1 .J . I j !...l l a,.. 'j '-" . . . . I'" -,.....~ ~ M I ~ r : . : 1. . . . . . . I

I . .j 1 , ~ , ~o 50 D 10 10 90J 100 1101 120 il:~I ; . . . . . . . . . . . ,-..

7.1 SUMMARYSA-4 was the first Saturn vehicle to employ active

accelerometer control. Performance of the total con-trol system was satisfactory. Control gains (ao andg2 , were shaped essentially for dl'ift minimum prin-ciple in the maximum dynamic pressure region andwere close to this at other periods.

The tilt program for the ST-90 platform was gen-erated by a synchronous motor driven cam. An s en-gine tilt program, also biased for a !i(l percent Marchzonal wind, was used for the first time on this vehicle.This resulted in a decrease, compared to earlierflights, of control angles experienced in the pitch planeduring the period of maximum dynamic pressure.

Angle-of-attack measuring systems performedsatisfactory. The first flight test of the F-16 Q-ballangle-of-attack indicator, to be used as a backup forthe control accelerometers, indicated proper oper-ation. There was an indication of the possibility ;)fsmall misalignments of the indicator of about 0.6 degin pitch and O. 2 deg in yaw.

c.- Pt h lit Prop (Ill 15)(dq)1o-1

-22

o-1

-2

AttituUe measurement" from the passengerST-124Pplatform were considerably better than thoseobtained on the SA-:l flight. Some minor er-rors arestill evident in the performance of the ST-12-1P, butare due to the limitations of this prototype equipment.7.2 CO~'TROL ANALYSIS7.2. 1 PITCH PLANE

TABLE 7-1. MAXIMUM PITCH PLANE RangeCONTROL PARAMETERS Timc(Sc,-;arumeter'---- Magnitude

AttitudeAngle-of-attack (Free-stream)Angular VelocityNormal AccelerationActuator Position

O.(j (degj-5. -I (deg)-1. 0 (deg/s)-0.5 (m/s2)-1.5 (deg)543S(kg-deg ;m2

53.527.570.151".92~. -I

Angle-or-attack DynamicPressure Product

28.8

Pitch attitude deviations were small throughoutthe powered flight (see Figure 7-1,. Vehicle tilting

...... T... (HC)

A.... Piech khaCOr ... tu .... (Cl-1. Cl-2, CI-3. Cl....)(~

nGURE 7-1. PITCH AT'TITIJDE. ANGULAR VELOCITY Al\,TDACTUAL ACTUATORPOSmON

21


36/87

was initiated by the ST-90 tilt cam at 10.58 sec (seeFigure 7-2). The actual mean wind was very close to

.~~. .- ~ / . . : : - /. ~;{:-~_.x.,.. . : . ." .' + - - t . . , . r L . ~ ~ _ / ~ _ ~ _ ~ _ _ . - . : : _ _ - _ , : _ _ . > +

" ", .. . r -, .. . , . .. .

FIGURE 7-2. TILT PROORAM AND PITCHVEWCITY VECTOR ANGLE

the 50 percent zonal March wind used to bias the tilt-ing prog-am (Figure 7-2) and resulted in reduced con-trol angles for the pitch plane (see Figure 7-4). Finaltilt arrest occurred at 105. 5S sec with the vehicle tilted42.6 deg from the launch vertical.

The maximum actuator deflection of -1. 5 deg inpitch occur-red at 28." sec as a result of the wind andthe wind biased tilt program. _Theexpected engine de-

flection .at this time withthe 50 percent zonal MardI.wind was approximately 0.5 degrees. However', theactual wind was about 10 mls greater causing an addi-tiona IIdeg actuator def'lectton. This was the greatestdeviation from the wind used in designing the tilt Pl'O-gram that occur red during flight.

Shown in Figure 7-3 is a compar-ison of the pitchcomponent Winds as a function of time from threesources: rawinsonde, r=cketsonde , and angle-of-at-tack Winds. The angle-of-attack winds \\ '_I"I:' deter-mined fromonboard vehicle measurements (using theangle-of-attack measured by the local indicator) andtracking data. The angle-of-attack Winds are consid-ered questionable after 107 seconds.

Figure 7-4 shows an estimate of the pitch angledesign crtterta requtred fOI' the Block Ivehicle whichwas based on eight engine operation on a seven enginetilt program. The garns used for establishing the de-sign crtte r , we re lor the Drift Minimum Principle.The response due to a 95 percent, uon-drrecuonal ,steady-state wind was increased by 25 percent to ac-count fOI' gusts. Variations in ae rodynamtc parame-ters we re accounted fOI' by increasing the nominal re-sponse 11 percent.

10

WindVelocity. ""X, (Positive from the nearl (m/sl I60~----'-----~-----r----~-----.r-----r-----~----'------r~~~----~---'--~----~50~ __ ~~ __ -+__+__~__~__~__+-__+-__~~-*r~__~I I I---,. II. ., "j.Angle of Attaek Winds \ _ ~J.lr\ I.~t-----+-----+-----~----~----~----~~~~----~----~----~----~~--~----~30__-+---+_-+---+-_+--.iOfII.At.",,~~i('~tWIIJ,.~_-+-II~--+--.J-')~~r~f------I20~_+--_+--_+-- / .: r i . ~ 1 f.l"~~ ~,?, 't~~, ! ' .. J J \ . RockelsondeWind

.... ~. _,,..... " '\1\0 '" -- RawinsondeWirdU' ~ :;~ :. ; Ii.. .lIPo It .,.-- I I .: lIIi;... .~ I I' '.-10~ 20 30 40 50 6~ 7~ 80'J~ " 2 \ 1 : I.100 110 130~0

Range TIme (sec I

; Telemetry Calibration Range Time (sec)

FIGURE 7-3. PITCH PLANE WThI'DCOMPONENT AND FREE-STREAMANGLE~F-ATTACK

22


37/87

- ttl - - -. ! 1 1JI'O 1\ v ' v . , I ItI" " IV I ,

0 20 10 . . 50 60 70 . . , , ~ ! . L O ' Yi J oT_ -. . . . W e . . . ttyT_ "U-ISU... ,.) I Ill . . . . . . . A I \A - , II _ . . . . . _ . . . . . . . . .. . . ~ V ' I t : ~ . I III ,. , . z o JO 60 50 60 10 10 10 I . 1101 12 0 l ! o&..~ 1'_ khetor " " t t _ _ -I. Q.. OZ... CZ-4,(..., _j , I I - I I

1\ 111l 1\ ~ I I.- - .l." ' " " " '~Jrv: ~ I [ . ., I ' II ,V I Iu I" II I, I JO ~ to 10 II to 1 ! t c 1101 uo uo

ItliL!lCl 'I"" fl loO't 'f

FIGURE 7-4. COMPAK~SONOF PITCHPARAMETERSWITHDESIGNCRITERIA

T & ..... (lIZ ISH...,

1-I

-I

- .-I-J

The solid lines in Ftgure 7-4 represent the designcriteria as a functton of time, and the points are ob-served values from the SA-4 flight. Shownas a dashedline aloe the values that would have been reached, hadthe tilt program not been biased for a 50 percent zonalMarch wind. The actual engine deflection values wereapproximately 16 percent of the design value at themaximum dynamic pressure region. The maximumwind component in pitch during this period was 45. Rmzs, All parameters were well below the design con-dition.7.2.2 YAWPLANE

TABLE 7-11. MAXIMUMYAWPLANE CONTROLPARAMETERS

ParameterRangeTime

Magnitude (sec)AttitudeAngle-of-Attack (F'ree -strcamjAngular VelocityNormal Acceleration..ctuator Position

1. 4 (deg) 65. ~-2.7 (deg] 61.41.0 (deg/~ 62.8

-1.3(m/s2 ; 61.5-3.0 (deg) 61.59788 (kg-rg; 61.4mAngle-of-Attack Dyr.micPressure Product

Larger-yaw deviations were observed on the SA-4flight (Figure 7-S} than on previous flights. These

( )

FIGURE 7-5. YAWATTITUDE. ANGULARVELOCITY ANDAVERAGEACTUATORPaiITION 23


38/87

oscillations were most predominant between 59 and 76sec and were the result of winds. Maximumactuatordeflections of 4 dge peak-to-peak were required. Theoscillations were initiated by an initial wind buildup of20 ml s with a wind gradient of 0.02 sec-1 (see Figure7-6). The resulting large amplitude slowly-dampedoscillation was the result of two (actors. First, thedesign of the bending filters created a lowdampedrigid body mode especially during the period between60 and 80 seconds. Second, the wind disturbance wasof lowmagnitude, but theoscillations in thewindspeedwere nearly in resonance with the rigid body controlfrequency of the vehicle.

The design of the bending filters for 8A-4 was

\\'ibd Vf"IOl'ity III. (Po_tUw fr_tt. &.H)(a/.)

complicated bythe non-optimum locationof the controlaccelerometers which were used for system stabi-lization. Thebendingfilters were required toattenuatea large range of frequencies from the accelerometersin order to insure system stability. The result was asystem which was relatively insensitive to dynamicsignals from the accelerometer except at extremelylowfrequencies. This condition was noted in preflightclosed loop transient responses. Saturn I - Block IIvehicles will have different accelerometer locationsand other characteristics which will allow better re-sponse to winds.

An analog simulation of the vehicle response inthe yaw plane using angle-of-attack winds as input isshownin Figure 7-7. Theagreement with telemetered

FIGURE 7-6. YAW PLA~"E WIND COMPONENT AND FREE-5TREAM ANGLE-OF-ATTACK

24FIGURE 7-7. COMPARISON OF YAW PLANE SIMULATION RESULTS WI11I TELEMETERED (35-95 sec;


39/87


40/87

TABLE 7-IV. ROLL MOMENT

Vehicle

Prior toEngine 5Cutoff(kg-:n)

Prior to OECO(kg-m)

Prior to IECO(kg-m)

SA-4SA-3SA-2SA-l

964928713672

16S8 1486155321401490

Very similar patterns are exhibited on all fourflights. Even though there is a systematic deviation,it had no serious consequences.

7.2.4 ALTl1 UDEANDCONTROL AFTER CUTOFFSince the effect of the thrust vector angularity

of the engine during thruRt decay is of interest in future

Pitch Control Accelerometer ('To/s2\

1 00Io

-1I n iOutboard En,: n('Jrust &cay Peu

YawCon' -11 Accelerometer (m/s2)

design for separation, values have been obtained forall Block I vehicles. The largest value obtained oc-curred on the SA-I flightand was 0.38 deg during thelOpe rcent to 0pe rcent thrust decay period. The valuesobtained for the SA-4 flight are listed in Table 7-V.All values are well within the design angularity "f onedegree allowed for inthe S-IV stage separation design.TABLE 7-V. THRUSTVECTORANGULARI1Y

DURINGCUTOFF DECAYAveraging Period Pitch (deg) Yaw (deg)100 to 10 Percent Thrust 0.0310 to 0 Percent Thrust O. 19100 to 0 Percent Thrust 0.01

0.010.110.03

A large degree of unceetatnty (estimated to be 0.75deg) exists in the measurements presented in Table7-V due to the small deviations being analyzed.

As on SA-3, cutoff of the outboard engines ex-cited vehicle bending in both pitch and yaw. Figure7-10 shows the oscillations as measured by the con-trol accelerometers located inthe Instrument canister.

llino Plre PeriOd IH y d ra ul iC : P r es s u re~--~~-c-~~P~e-ri-oo~--------------~~

o

-1

12624 12 ,

nGUU 7-10. PITCH ANDYAWCONTROL ACCELEROMETER AFTER OECO


41/87

Again. a8 iudicated OIIEA -3. ther~ it; a difference if!the damping btm\'epfJpik.ll and yaw. "nIe diffcrtmce iillstabiUtylJet:ween o1tObtainUJi8ft),(;1 rJlte,,&navet'lt4Jemi..'~(Jlaf.l ,foart'etort"OCkett!ofO.1l>deg perpen-dicular 10 the cant .angle piatle it; tleqllired.

nre ' 1 " 0 1 1 ,,"ate Yf'O 0IItpUl eontinued to mcreawe_ightly .Ilftct-!Mdt; tlme~ madli1lgama1{iftlltlll m %.1,deg!.at3ao tleOOhCtll. rtd....... !; < 1 " f T.~vlCl -iI!!~ua 1 45tI iV ( .. tr) ~~01_ . .-

. .,__ .. l"""O~.(_),. ..... rtfi!!iM)

FIQJIlI: :i-I%. AWGI...ES-Of'-A1'TACK 00 ...._Dlmt___ ~"'1IIe alOeler ...... O IM I III .. f'............................. mFpR3 -u .t'IIe .. ..-,e ..., e........... ;8ft!__ rtf.." a III .eIIect:s a(die ftiIIide4*tile .. elena' WIle -.I~__ .a-4' teIe.. .." ''' .. ..ct die ........ ~_IZ.!,..a.


42/87

7. 3. 1. 1 CONTROLACCELEROMETERSTwo Statham control accelerometers (pitch

and yaw. were flown "closed loop" for U Ie first timeon SA-4. They were used in place 01the local angIe-of-aUac'kmeters to provide the artificial stabilizationterm for the control svstem, The telemetered ac-celerations are compared to calculated accelerationsfrom the telemeteredactuat: 1 " deflections and angles-

'0 PItch "n~i acce l( ,f"at 1= (-.1.2)

II, . . . I1~',.. . I

~t.. I\ IL . . . . . . x-, ~ , , . .\! (. . . - \'110

-1 0

- - - - - Calcul.tacI "fr_ ..,_ of,,,,le!> of *u.dr. u.tt_ ...

T.... lIIomal kCIPlenUOtl .,.; II I,0 " " = = ' I~ . . - f~- I (I01

FIGVII 7-13. 1'LE)(ETERED D'"D CAU::ULATEDNOlDIAL ACCELERATIONS OOBING REENTRY

of-attack in Figul-e 7-14. '(be agreement is withinO. 1 ml s2. The telemetered accelerations were nu-merically filtered with a 10.' pass filter to removebending oscillations. A considerable amount of firstbendiag oscil....ion (2 to 3 CPS) was seen on these ac-celerometers throughout the fiigbL Theydid not effectthe control system due to theaccelerometer loopfiltersin the control computer.

7.3.1.2 ANGLE-QF-ATTACKMETERSFour local angle-of-attack meters (U.S.

Science) anda model F16Q-baUangle-of-attack meterwere fto.'O fur measuring purposes.

The two local angle-of-attack measurements ineachplane.'ereaveraged and corrected for an upwashfactor. The resulting free-stream angle-of-attack isshown in Figures 7-3 and 7-7. From the comparisonof the calculated angles-of-attack (rom ra,,-insoOOe(square points. and rocketsonde ,,1nd data (circledpoints.. it can be concluded that the local angle-of-attack meters functioned property to approximately107 S~. corresponding to a dynamic pressure of 340kg/m2

Sho.:n inFigures 1-3and 7-7 as dotted lines arethe angIes-of-attad determined from the Q-baU indi-calor. '(be agreement with the local measllred andcalcuia1ed aagIes-of-auack is good throughout flighLRo..ever. this agreement was obtained by assumingthat some portion of the Q-baU indicator.'aS slightlymisaligned. Without this assumption. the agreementprior to 50 and after 80 sec deterionled. 'DIe---~

,I . :~~1__~'~1.~_2~"_ ~~~~~~~5~. __ ~~~'_ ~7~n_ ~~~~~~ __ ~100~_'~10~;~~~~~!~~TbW._'1

II'II

- - -FIGUIlEV-I4.. P'I1CIIAJID YAW CONTROL ACCEtERATmNS


43/87

misalignment, calculated bycompariSOllwith the localmeasureme_, was 0.6


44/87

The maximumdemands on the actuators occurredbetween 58 and 70 sec when all actuators experienceddemands of 1.5 to 2.0 deg!s. The nominal level ofdemand was less than 1deg/s, The demands on thepitch actuators were lessened due to the modified tiltprogram.7.3." ST-124P STABIUZED PLATFORMATn-

TUDECOMPARISONA compartson of the attitude measurements

from the two platforms (ST-90 and ST-124P) showssome difference in all three axes (see Figure 7-171.The largest deviation occurs in the pitch plane duringthe period 0( vehicle tilting. Shownas a dashed linein the upper graph of Figure 7-17 is the error due toimpedance mismatch. The relM;, ,ng difference isprimarily due to the less-than-optimum gain in theservo systems and some backlash in the gear trainsof 1he ST-124P.

1be mean differences in yaw and roll result prt-madly from the initial azimuth misalignment of theST-124P.

1 AUhude Pilch ST-~OMinus Attitude Pitch ~T-124 (deg)-

7.4 PROPELL"NT SLOSHINGSloshing was measured, as has been done on all

previous Saturn flights, by means of differential pres-sure (~P) measurements in three of the nine propel-lant tanks (center WX tank, WX tank 04, and fueltank F2,. Theconfiguration ofthe tanks and measure-ment locations are the same as on S"\-2 and S.-\-3andcanbe seen in Volume IIof the SA-3Evaluation Report(MPR-SAT-63-1)

The propellant slosh heights in the pitch plane.as converted from the ~p measurements. during thelatter portion o f powered flight are shown in F;gure7-1". Two methods were used for eonvertmg the te-lemetered pressures to fluid heights. One was basedon a theoretical conver'ston factor and the other wasbased on some ground test calibration results,

As longas the liquid surface was within the regionof 1he slosh baffles, the sloshing was damped. When.the liquid surface went below the baffles, the dampingprovided by the smooth walls was insufficient to pre-vent a small buildup in sloc,hing amplitude. The

tECO OECO

-1 10 20Error DueTo Impedance --''Mismatch In ST-1:!4P Resoh:~,.Chain 1

40 50 60 70 ~~~0__ ~+-+___.l?O~ __ 11_0--,-1_ l~2:-0_:---::13(lnan&,!, Time (sec)

.. T~lemetry ,Calibrations,.....Meatllll'ed f'!'om Launch SpacE-Fn..~ \'~'rticai

FIGURE7-17. AT'IlTUDE:mFFERENCES BET\\'EEN ST-90 ANDST-124P


45/87

maximum slosh amplitude appeared to be about 7 emin the pitch direction; the largest since SA-t. Theinitial excitation on SA-4 may have been caused by theearly cutoff of engine 5 and the termination of the tiltprogram. No significant sloshing was observed in theyaw plane measurements.

The upper portion of Figure 7-19 compares theobserved sloshing frequency (symbols) with the pre-dicted sloshing frequency (solid line'.

The observed frequencies were obtained by digitalfil1l>ring of the telemetered pitch actuator deflectionsand normal acceleration. The increase in actual fre-quency, compared to predicted, near the end of flightis probably primarily a result of coupling with the con-trol system.

A brief parameter varfattoe study was made of thesloshing mode for the time period when the propellantsurfaces are below the baffles in the bottom of thetanks. Results are shown in ttl .Diddle graph of Fig-ure 7-19. Lines of constant system damping are shownas functions of propellant damping and total sloshingmass for the period when the propellant surfaces are

FIGURE 7-18. SUlSHING AMPLITUDES AFTER 95SECONDS

. .below the baffles. The observed SA-4 rate of damping(0-) around 110 sec was 0.2. The point representingpredicted values of sloshing mass and propellant damp-ing is also shown in this graph. This indicates that thepredicted values underestimate the instability duringthis period.

In order for the sloshing model to give resultswhich agreed with the SA-4 flight. it would generallytake an increase in sloshing mass and a decrease inthe propellant damping. Experimental investigationsof sloshing in tanks with sphericd type bottoms haveshown that the smooth wall damping is very small(gs "" 0.002). Therefore. itseems that the dampingfor the slosh model when the propellant is in the bottomof the tank should definitely be reduced.

The response of the vehicle center line (combinedtranslation. rotation and bending) to the sloshing isshown in the bottom graph of Figure 7-19. The solidline is predicted and the circled points were ohtainedfrom telemetered accelerometer values at severaltime points during the period of sloshing. The curvebas been normalized to avalue of 1 at the vehicle nose.

lSi_hutt:Frequ.,n":.(J\ ~ 0r.o .. ~ .. -0

o Act~tOf "fi~ctlOtl~} ... _.-o lIIo~l Accele-ra'tlGn " ~ u e P t C t . ~

~~ . Roit_~of Da~:-'':

f:.e>r-...... l . l " . . . . , .c Tril""_ ~~1.' '' '.8 ... ..-p :" . ..~ "J:IVehlCh" ceeaer : 1"'~ T ~-, '_'"1.....: _ 0 1

-- Preclcr ..d

Di.... eft... .~CiaNI ra;., ,.)

nCURE 7-19. SA... RESPONSE 1'0 SL


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SECTION vm . GUIDANCE

8.1 SUMMARYThe SaturnSA-4 vehicle was flown without active

path guidance or velocity cutoff. However, passengerhardware for both ST-90 and ST-124P (Prototype)guidance systems was onboard toestablish operationalcapabilities of the guidance equipment in the Saturnflight environment. The telemetered data, as well astrajectory comparisons, confirms satisfactory per-formance of both platform systems.

The last alignrnent check before liftoff showed theST-90 cross-range accelerometer biased 0.023 degabout the X-axis. This misalignment is substantiatedby an error of approximately 1.0mls in cross-rangevelocity at outboard engine cutoff.

SaturnSA-4experienceda high roll rate after ig-nition of the retro rockets, as did the SA-3 vehicle.As a result, the ST-90and ST-124Pplatforms reachedtheir mechanical limits at 133.2 sec and 135.1 sec,respectively.8.2 DESCRIPTIONOF THE GUIDANCESYSTEM8.2.1 ST-90 GUIDANCESYSTEM

The ST-90 guidance hardware included threependulous integrating gyro accelerometers (AMAB-3)mounted on the ST-90 stabilized platform. These ac-celerometers were mounted to sense forces along aset of inertial axes oriented with respect to the stableplatform X, Y, Z axes. The slant range axis was di-rected downrangeand elevated 41degupfrom the localhorizontal plane at launch; the slant altitude axis wasdirected up and normal to the slant range axis in theflight plane; the cross-range axis was parallel to thelocal horizontal plane at launchandnormal to the flightplane. thus completing a right-handed cartesian co-ordinate system. The platform was stabilized by threeA8-7 air bearing gyros. Guidance computers werenot carrier, and the velocity signals were used ""'''measuring purposes only.

Velocityenc~rs mountedon the accelerometersare driven by the precession of the accelerometermeasuring head. Each encoder generates two vari-able frequency output signals. The zero crossfng ofeach signal indicates an incremental velocity changeof o. 1mls, The polarity of the velocity increment isdetermined by the phase relation of the two signals.

Aguidance signal processor repeater accepts thetwo sine wave signals from the encoder and converts

them into a polarized incremental stairstep format fortelemetering purposes.

8.2.2 ST-124P GUIDANCESYSTFMThe ST-124P guidance system was the second

prototype, or engineering test model, of the SaturnBlock IIguidance system. The four-gimbal systemutilized twopendulous integrating gyro accelerometers(AMA8-3) mounted on the stable element. Platformorientation was maintained by three A8-5 air bearinggyros.

The altitude and cross-range accelerometerswere mounted to sense forces acting along inertialaxes oriented parallel to the Y and Z platform axes.The altitude axis was directed up and parallel to thelocal horizontal plane at launch and normal to theflight plane.

Aguidance signal processor repeater and two ve-locity encoders, similar to those uced with the ST-90system, were included in the ST-124P system.8.3 PLATFORM AND ACCELEROMETER ALIGN-

MENTSThe ST-90 stable platform was optically aligned

along an azimuth of 100.038deg E of N, while the ve-hicle Fin 1- Fin III plane was aligned 100.325 deg Eof N. The ST-124P platform was aligned approxi-mately along the Fin I-Fin m plane.

The flight plane was defined bythe ST-90 platformand the vehicle flew with offset axes as shown in Fig-ure 8-1. Both platform roll attitude indications were

-- Vehicle Axe.------- ST-90 Alipment

.... m

. . , .. . . . .FIGURE 8-1. ST-90 PLATFORMAUGNMENTRE LATIVE TO VEHICLE AXES

32


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-ulled to zero before launch. The view shown is fromthe rear with the solid lines and circles representingthe vehicle axes and engine positions. Dashed linesandctrcles represent the positions of the vehicle axesand engines if aligned at launch with the ST-90 plat-form.

Both platforms performed properly throughoutpowered flight. However, an excessive roll anglebuildup, due to retro rocket misalignment. force

Documents

Results of the Fourth Saturn Launch Vehicle Test Flight SA-4