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.. AUTHORITY...... .S.. Lt. Feb... 28,...... 197

THIS.. PAG IS UNCASSFIE

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TODISTRIBUTION STATEMENT: A

Approved for public release; Distribution Unlimited.

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AUTHORITY

SAMSO, USAF Ltr; Feb 28, 1972

THIS PAGE IS UNCLASSIFIED

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• UNCLAsSS

AA 61-0146IZ October 1961

-' • WS I07A-I FLIGfT TEST WORKING GROUP

- •IFIGHT, TEST REPORT

,.Jf'* ATLAS MISSILE 30E &. . ",s/~~~ ": : ;,UTIC$

5 October 1961 I' 1, 6'- 1961

Log No. T 61-Z7ZZ Copy No.47

AMR RANGE TEST NUMBER 1804* -00•0., 0O

0 co ASTRONAUTICS TEST NUMBER P3-501-00-30

SECURITY NOTICE

0 a 0._4 -1 This document contains information affecting the

National Defense of the United States within themeaning of the Espionage Laws. Title 18, U.S.C.

S • :Sections 793 and 7"4. The transmission or therevelation of its content@ in any manner to anunauthorised person is prohibited by law.

compiled and published at AMR by

GENERAL DYNAMICS/ASTRONAUTIlQSiGNCCRING CORRP7ONUiJftj

CONVAIR ASTRONAUTICS

POST OFFICE BOX 1128GAN IDITQ 12, CAL. 11NIA

G R ~O U P - 3e- 7 1 ' - 7 ,_

DOWNGRADED AT 12 YEARINTERVALS; NOT AUTOMATICALLY

DECLASSIFIED. DOD DIR 5Z00. 10

JNCLA I/I

Vore rl-k3 (MT 61-89184

Page No. iiAA 61-0146

This report has been prepared to present preliminary information relativeto the flight of Atlas Missile No. 30E. The information presented is basedon visual observation and data evaluation to the extent permitted by time

limitations. It should be considered as preliminary only, and the final re-ports on this flight referenced for further information. The technical contenthas been prepared and jointly agreed upon by members of the WS 107A-1Flight Test Working Group.

Prepared by: Data Operations. General Dynamics/Astronautics. AMR

peV .V. Thatcher f0 e t ný MManager, WS 107A-I ProjectOffice Cl f ofTest O0erations, AMRSpace Technology Laboratories, Inc. GD/Astronautics, A Division ofAMR, Florida General Dynamics Corporation

L. D. Solid H. B. HoepeqResponsible Representati-e Manager, Field Test EngineeringWS 107A-I General Electric (DSD)Rocketdyne, AMR, Florida AMR, Florida

W. H. Caldwell .. HriksenManager, Test Operations Test Operations Section HeadBurroughs Laboratories American Bosch ARMA Corp.Burroughs Corporation

ftJ. R. Cardso T. F. DavidsonJOI&Test OperaJrns Section Head Site Manager, AMR

AVCO Corporation Ac tica Associates

-P. R. wig~nll /Colonel, USAFiCommander

-' i

CONVAI R-AS RONAUTICS

" •Page No. iiiAA 61-0146

SUMMARY

Atlas Missile 30E. the 14th !'E" Series missile to be flight tested at AMR. waslaunched from Complex 13 at 084Z EST on 5 October 1961. The flight waasuccessful and the Mark 4 Mod IB Re-entry Vehicle impacted in the IndianOcean target area at a rango of 7, 539 nautical miles. The ejectable data cas-sette was successfully recovered.

This was the first flight for the Mark 4 Re-entry Vehi-cle to- a kzngrange target. The prime objectives of the flight, which were to evaluate re-entry vehicle performance under conuitions of maximum heating resulting fro~nthe flattened trajectory, and to demonstrate the long range capability of the "i"Series missile, were satisfied.

Performance of all missile systems was satisfactory.

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R-ASTRONAUTICS

Page No. ivAA 61-0146

TABLE OF OWNTENTS

FOREWORD . ........ ii

SUMMARY .- . . . .. iii

TABLE OF CONTENTS .... iv

FLIGHT TEST OBJECTIVES . . .

FLIGHT TRAJECTORY DATA . 3 •

SYSTEM PERFORMANCE .... 7

Ai2 frame . . . . . . . . . . a

Propulsion System . . . .. . . . 9Acoustics Propellant Utilization System . . . 14

Pneumatic System . 16

Hydraulic System . . . . . . . . . 18Missile Electrical System . . . . . .. 20

Range Safety Command System ...... 21

Asusa System . . . . . . . . . 22

Flight Control System .... . . . . 23Inertial Guidance System . 24

Mod III Range Safety and Instrumentation System . i 28Re-entry Vehicle . .a . . . . . 30

Propellant Tanking .# . . . . . . 31

Airframe Internal Instrumentation System .. .. 32

Landline Instrumentation . . . . . . 33

Mls ller Aft-end Motioa Determiuwon .... 33

CONCLUSIONS AND RECOMMENDATIONS • . 34

COUNTDOWN TIME VERSUS EVENTS . . 35

MISSILE CONFIGURATION. 41

HISTORY OF SM-65E MISSILE NO. 30E • . . 44

AF -

~II~I R-ASTRONAUTICS

Page No. vAA 61-0146

Pacge

APPENDIX . . . . . . . . .. . Ia

Fluid Chemical Analysis . . . . . . . ZaReference Documents. SaSerial Numbers of System Components . . . . 6a

Significant Dates During Testing of "A" Series FlightMissiles at AMR . . . . .. . * a

Significant Dates During Testing of "B" Series FlightMissiles at AMR . . . . . . . . 9a

Significant Dates During Testing of "C" Series Flight"Missiles at AMR . . . . . . . . l0a

Significant Dates During Testing of "ID" Series FlightMissiles at AMR .. . . . .. . Ila

Significant Dates During Testing of "E" Series FlightMissiles at AMR . .. . . . . 14a

Significant Dates During Testing of "F"41 Series FlightMissiles at AMR . . . . . .. 16a

Significant Dates During Testing of Mercury/AtlasVehicles at AMR. . . . . . .. . 17a

Significant Dates During Testing of Midas Vehiclesat AMR *. * 0 .. . . . . . 1Sa

Significant Dates During Testing of Atlas/Able LunarProbes at AMR . . . , . . . . 19a

DISTRIBUTION . . . . . . .. . Z0a

Yo A(-9

SME

CONVAIR AUTICS

Page No. I

AA 61-0146

FLIGHT TEST OBJECTIVES

The primary purposes of 4Ws flight were to evaluate re-entry vehicle per-formance under conditions of extreme heating and to demonstrate the satis-factory performance of the airframe using a flattened trajectory.

Detailed objectives are listed on the following page along with applicablecomments relative to the degree of satisfaction.

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r Page No. 3AA 61-0146

FLIGHT TRAJECTORY DATA

Atlas Missile 30E was planned for a range of approximately 7539 nauticalmiles downrange with impact in the Indian Ocean at 38. 20 S latitude and

34.00 E longitude.

A tabulation of miss-distances and a comparison of nominal flight performanceparameters from Flight Trajectory Simulation Case E - XIII, and measuredtest values from Azusa and telemetry data at significant times along the tra-jectory are presented below.

Figure I presents impact points as determined from the impact predictions.

NOTE: All times in this report are based on range zero time which occurredat 084Z:37 EST. One Inch Moticn occurred at 084Z: 37. 12 EST.

- Source Miss Distance 95% Confidence Limits

Mod rIr Impact Prediction 3.50 nm Long 13.86 nm x 6.46 nm9.40 nm Left at 95°T"

Mod III/Guidance Velocity 2. 56 nm LongComparison @ VCO 5. 86 nm Left

A susa Impact Prediction 19.63 nm Long 1.44 nm x 0.14 nm4.05 nm Left at 930 T

Item Unit Nominal Measured

Liftoff Weight lbs 266, 033 -

Pitch Plans Azimuth dog 111.9 111.9

BCO Weight lbs 68, 256 -

BCO Velocity ft/sec 8955 8935

BCO Altitude ft 195,178 199.106

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CONVAIR-ASTRONAUTICS

Page No. 4AA 61-0146

Item Unit Nominal Measured

BCO Range nm 42.Z 12.5

BCO Time sec IZ3.1 114. 8

SCO Weight lbs 11,9ZI -

SCO Velccity ft/sec Z3. 146 23, 116

SCO Altitude ft 7Z3,644 739,777

SCO Range nm 452.1 443.2

SCO Time sec 306.9 304.7

VCO Weight lbs 11.921 -

VCO Velocity ft/sec 23,146 Z3,116

VCO Altitude ft 723,644 739.777

VCO Range nm 452.1 443.2

VCC Time sec 306.9 304.7

Impact Time sec 2771.7 2786.9

Impact Range nm 7539 7543

Impact Latitude (Geodetic) dog S. 38.2O 38.05

Impact Longitude (Geodetic) deg E. 34.00 34.09

NOTE: Nominal times :re corrected foz the difference between range zeroand one inch motion. Measured velocity, altitude, and range at

• b booster, sustainer, and vernier cutoffs are from Azusa reviseddata. Impact range and co-ordinates are taken from Mod III data.Measured times are taken from telemetry recordings of discretegeneration.

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Altitude is height above launch horizontal.

Velocity is speed relative to the earth's surface.

Range is measared horizontal from the launch pad with the exception of impact

range which is measured along the surface.

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Page No. 6

AA 61-0146

) jIMPACT POINT COMPARISON

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r Page No. 7AA 61-0146

3YSTEM PERFORMANCE

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4Eeftelr-CONVAIR-ASTRONAUTICS


AIRFRAME

Missile structural integrity was satisfactorily maintained throughout powered

flight and well beyond re-entry vehicle separation. Booster separation appeared

satisfactory as indicated by measurement M 32 X, Conax Valve Command,

and Flight Control data. Re-entry vehicle separation was satisfactorily accom-

plished as recorded by measurement Y I X. Separation Signal.

I Missile 30E was the first hard stripped, maximum heat type missile to be

flight tested at AMR. The missile was stripped of equipment not absolutelynecessary for accomplishing the primary objectives and maintaining range

safety. The equipment removed from this missile included two telemetry

packages, strobe light, the vernier LOZ and fuel solo tanks and associatedplumbing, and the vernier solo accumulator. The insulation bulkhead was alsodeleted.

This missile was instrumented for measuring axial acceleration and tempera-ture in the thrust section. Measurement U 101 A, Axial Acceleration, providedvalid data. Satisfactory operation of the retro-rockets was indicated by this

measurement. The axial acceleration at BCO was 6.84 G and 6.73 G at SCO/VCO.

Measurement P 671 T, Thrust Section Ambient. Quadrant 4. provided datathat compared well with the information received on previous flights. The

temperature rose to 95°F at BCO from 82°F at liftoff. The temperature atSCO was 154°F.

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CONVAIR*ASTRONAUTICS


PROPULSION SYSTEM

Performance of the MA-3 Propulsion System was satisfactory throughout theflight. Engine thrust rise and decay were normal. Sustainer engine delaytime was 550 milliseconds.

A trailing wire umbilical was added to this missile to evaluate propulsionparameters during the first 40 feet of rise. BI and B2 L.02 dome accelerationswere 10 G's and 15 G'Is at liftoff and 29 G's and 23 G'Is at 2. 6 seconds. Sus-tainer accelerometer data was not valid. The B2 L402 pump Inlet Pressure(Measurement Number P 1003 P) indicated two transient peaks of 170 and 150pounds at 0. 3 and 0. 8 seconds prior to liftoff. Valid data for comparison hasnot been acquired from previous "E" Series flights.

BI Pump Speed (Measurement P 84 B) indicated approximately 1000 rpm belownoraml throughout engine operation. Since other parameters indicated normal

( engine operation the BI pump speed measurement was considered qualitativeonly.

There was no vernier solo operation planned on this flight and sustainer and

vernier cutoff was effected simultaneously. Vernier chamber pressuresremained nearly constant throughout the flight without the step increase thatnormally occurs when the vernier tanks are pressurised at BCO. As therewere no vernier propellant tanks aboard this missile the vernier tank fill flowrequirement was eliminated thus allowing nearly constant propellant flow ratesto the vernier engines throughout the flight. Between 171 and 208 seconds VZChamber Pressure (P 29 P) showed a gradual decrease from 378 psia to 358psia. Since the chamber pressure was the only vernier engine function moni-tored, it is not knownwhether this was an instrumentation anomaly or an actualchange in chamber pressure.

Time slices of Propulsion System Data are shown in the following tabulations.

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Page No. 14AA 61-0146

PROPELLANT UTILIZATION

Operation of the Acoustica Propellant Utilization (PU) System was satisfactory.The PU valve responded correctly to the error signal and the HeadSuppressionValve was also positioned properly. Significant PU System Parameters areshown in the following tabulation.

Fuel LOZ CorrectedSensor Sensor Error PU Valve

Station Uncov=•ina Uncoverina Time Position

1 7.35 8.50 1.15 21.50

Z 47.55 47.55 0.00 47.00/Z7.50

3 86.26 86.35 1.09 21.50

4 116.36 116.73 0.37 ZS. 0°

5 202.85 Z03.5Z 0.67 2.6.00

6 Z58.03 Z55.18 2.85 38.00

As noted above the LO and fuel sensors at Station 2 uncovered at the same time.This simultaneous sensor uncovering was interpreted by the system as a LOZonly sernsor uncovering. In response to this indication the computer set the PUvalve at the open limit until automatic monostable reset. When the monostablereset without seeing a fuel sensor uncovering the computer automatically pro-grammed the PU valve to the nominal position un*il Station 3 sensor signalswere received.

Residuals determined by various methods were as follows:

Flow Rates Used(lbs/sec) Residuals (lbs)Method LO . Fuel Loa Fuel

Head Sensing Ports uncoveringand engine acceptance data withcorrected flow rates 178.2 89.8 660 200

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Page No. 15AA 61-0146

Flow Rates Used~lbs /sec) ResidualsIlbs)Method L.40Z Fuel L0Z Fuel

Head Sensing Ports uncoveringand flow rates between PU Stations5 and 6 corrected for changes invalve angle at Station 6 185. Z 80.2 547 259

Head Sensing Ports uncoveringand flow rates between PUStation 6 and ports uncovering 186.6 94.3 581 167

PU Station 6 uncovering andengine acceptance data with cor-rected flow rates 184.3 87.8 768 346

PU Station 6 uncovering and flowrates between PU Stations 5 and6 corrected for change in valveangle at Station 6 181.6 89.2 932 404

As can be seen, some discrepancies exist between the various methods ofcomputation.

The following constants were applicable for this flight.

Acoustica Computer CA-108B (5 card) Serial No. 0114LO2 Stillwell Model No. SL-192 ("F" Series)Fuel Stillwell Model No. SF-191 ("F`" Series)

PU Valve Angles:

Lower Limit 21.4 degre.3sNominal 27.5 degreesUpper Limit 46. 5 degrees

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I Page No. 16AA 61-0146

PNEUMATIC SYSTEM

Performance of the Pneumatic System was satisfactory. Missile tank pres-sures were normal prior to engine start and were satisfactorily maintainedthroughout the flight. This was the third "E" Series missile to be flight testedwithout the insulation bulkhead between the missile tanks. There were noadverse effects noted as a result of this configuration. F and G pneumaticregulators were utilized on this flight and operated satisfactorily.

W.02 and fuel tank pressures were satisfactorily regulated within boosterphase specification of Z3 to 25. 5 psig and 57. 0 to 62. 0 psig, respectively.1.02 boiloff maintained 1.Z tank pressure at 24 psig during sustainer phaseand well beyond re-entry vehicle separation. Fuel tank pressure started todecay ifter the regulator was jettisoned with the booster section, and reached40 psig by SCO/VCO. Tank pressures and trends were similar to thoseobserved on Missiles 22E and 18Z, the two previous "E" Series flights, with-out insulation at the intermediate L0Z/fuel tank bulkhead. Fuel tank pressures3: were from 8 to 10 psi lower by SCO without bulkhead insulation as compared to"E" Series flights with bulkhead insulation.

Booster tank helium bottles Measurements F 246 P and F 247 T reflected nor-mal pressure and temperature decreases during booster phase. Bottle pres-sure at booster jettison was 650 psig.

Sustainer controls helium bottle discharge pressure remained steady duringbooster phase at 29Z5 psia and supplied adequate pressure for activation of theseparation mechanism at the time of booster section jettison. Pressureremained at approximately Z840 psia during sustainer phase. Since there ./.no vernier engine solo operation on this flight, bottle pressure did not dropafter sustainer engine cutoff.

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4GR~CONVAIR-ASTRONAUTICS Page No. 17

AA 61-0146

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Page No. 18AA 61-0146

HYDRAULIC SYSTEMS

Performance of the Hydraulic Systems was satisfactory. Booster and sus-tainer/vernier system pressure were properly maintained throughout flight.There was no vernier solo accumulator on this missile as no vernier enginesolo operation was required. All high and low pressure system instrumenta-tion yielded valid data and indicated satisfactory operation.

Booster and sustainer hydraulic pressures rose from ground levels of 1960psia and 1955 psia at engine start to stable flight pressures of 3080 psia and3000 psia respectively. The booster and sustainer low pressure systemsremained between 90 and 100 psia throughout:flight.

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Page No. 20AA 61-0146

MISSILE ELECTRICAL SYSTEM

Operation of the Missile Electrical System was satisfactory. All parametersremained within specification throughout flight. Normal transients, reflect-ing changing load conditions, were evident during flight.

Main missile battery voltage remained between Z7.9 and Z9. Z vdc. Theinverter frequency remained between 399.4 and 400.0 cps. Phase A andPhase C voltages varied between 114.3 and 115. 1 vac and between 113.4 and114.8 vac rtspectively.

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CONVAIR-ASTRONAUTICSr

Page No. 21AA 61-0146

RANGE SAFETY COMMAND SYSTEM

Operation of the Range Safety Command System was satisfactory. TelemeteredRF input/AGC data indicated adequate signal strength to ensure proper opera-tion of the system until well beyond nose cone separation. The AutomaticSustainer Cutoff (ASCO) signal and the Manual Fuel Cutoff (MFCO) signal weresatisfactorily generated and transmitted by the ground equipment and properlyreceived and decoded at the missile. It could not be determined whether theguidance sustainer cutoff discrete or the ASCO signal initiated cutoff. TheASCO signal was decoded at 304.808 0. 10 seconds. The MFCO signal hadbeen requested for 350 seconds and was decoded at 351.234 40. 10 seconds.

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"Page No. 22AA 61-0146

AZUSA SYSTEM

Operation of the Azusa System was satisfactory. The range parameter was inthe fine mode at liftoff and the angle cosines, which were in the intermediatemode at liftoff, had switched into the fine mode by 4. 4 seconds. Automatictrack was established at 4.4 seconds. One ambiguity was resolved from theSangle cosine at Z4 seconds and an ambiguity was resolved from the ,.*Vangle cosine at V7. 5 seconds. The estimate for reducible data was from 12to 408 seconds.

Telemetered Input/AGC data and ground receiver AGC data indicated adequatesignal strength to ensure proper operation of the system. The ground receiverAGC data experienced a 10 db drop from -100 dbw to -I10 dbw during the periodfrom 130 to 139 seconds. The signal did not recover from this drop. Thisreceived signal strength drop, seen only at the ground station, was believed tobe the result of a change in look angle caused by the 11 degree pitch down ofthe missile after guidance steering enable. The system switched to the highpower mode of operation at 100 seconds and loss of signal was at 434 seconds.

The system experienced loss of phase lock 56 times between 0 and 300 seconds.These intermittent occurrencea involved from 1 to 3 data points each. A simi-lar occurrence was observed on another range operation and this condition isapparently associated with the ground equipment.

The 7090 Computer generated an open loop Automatic Sustainer Cutoff (ASCO)signal based on Azusa data at 304. 590 seconds. This was 236 millisecondsbefore Mod Ill generated the closed loop ASCO function.

Azusa data was utilized by the 7090 Computer for I.P. information during theperiods from 50.95 to 54. 15 seconds, 55. 15 to 208.35 seconds, and Z08.75 to375. Z5 seconds. The following 1. P. was derived from the Azusa System.

Miss Distance 9S% Confidence Limits

Downrange 19.63 am 1.44 nm x 0.14 nm

Crossrange 4.05 am at 930 T

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Page No. 23AA 61-0146

FLIGHT CONTROL SYSTEM

Performance of the Flight Control System was satisfactory. Data showed thatresponse to the guidance roll maneuver was satisfactory and that the pitchprogram was accomplished satisfactorily. All guidance discrete commandswere acted upon properly and response to guidance steering commands wassatisfactory. Data indicated that all programmer switch functions occurredproperly. There were no missile bending mode buildups during the flight andpropellant slosh during booster phase was small. Engine displacements atengine start were within the allowable tolerance of 0 0. 6 degrees. Liftoff andstaging transients appeared normal as did the staging sequence. The missileharness was modified such that SCO, VCO and pre-arm were effected simul-taneously.

A 0.9 volt drop of Measurement S 3Q1 X (Thor Retro-rocket Command) forone segment at 32Z. 57 seconds indicated that the retro-rocket igniter squibwiring probably shorted as usual following retro-rockets firing. However.the programmer was protected from burnout by the switch current limiter.

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Page No. 24AA 61-0146

INERTIAL GUIDANCE SYSTEM

Performance of the Inertial Guidance System was satisfactory. The roll man-

euver and the steering signals were generated properly. The discrete com-mands were issued properly. No target spotting offsets were required onthis flight.

Inerti*.l Mode Start occurred at 0847:31. 000 EST.

Missile Behavior

Missile axial acceleration at sustainer cutoff was as follows:

Nominal Actual

Thrust Acceleration 6.76 6.8

t Net Acceleration 6.47 6.5

A comparison of the telemetered velocities and positions with those listed inTrajectory Xmn at the approximate time of guidance enable was as follows:

Conditions at Guidance Enable

3 SigmaFunction Units Actual Nominal Difference* Limits

Time** sec 136.5625 134.0 42.56 M6.5

xft/sec 9411.25 9412.5 -1.25 170

Yft/sec 2064.Z5 2057.75 06.50 0600

± ft/sec 4018.50 3947.5 471.0 1900

X ft 464,000 450, 11 Z ,13,888 419,500

Y ft 147,264 142,848 04,416 025.500

Z ft 222,720 Z13,632 49,088 n_8,500

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"Page No. 25AA 61-0146

3 Sigma

Function Units Actual Nominal Difference* Limits

CEF rad -0.000213 -0.00027 0.000057

REF rad 4.587645 4.6003 -0.012655

* Actual minus nominal** Referenced from Inertial Mode Start

Abnormalities -

A minor disturbance at 43 seconds on the azimuth resolver signal was indi-cated by a 0. 80 step, possibly indicating left roll of the missile. The rolldisplacement gyro data also showed a minor disturbance at the same time.The cause of these disturbances is unknown.

Residual Missile Thrust at SCO/VCO-

Telemetered guidance velocity data yielded the following approximate valuesfor residual thrust. Decay from 6. 8 g's to 0. 3 g's occurred within 0.2 sec-onds. During the next 3 seconds, the decay from 0. 3 to 0 g's appeared linear.The total thrust velocity gained between SCO plus 0. 4 seconds and SCO plus 3seconds was approximately 12 ft/sec.

Platform and Control

The roll maneuver, as indicated on the azimuth resolver data, was executedprope.-Iy from 2 to 19 seconds.

The pitch resolver came into the instrumented range at guidance enable plus. seconds and settled at zero 21 seconds later.

Servo errors were normal and less than 0.5 minutes deviation. Azimuthservo was not instrumented on this flight.

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-) Page No. 26

AA 61-0146

Gyro drifts measured prior to launch were:

Gross Azimuth 4 2.07 deg/hr Precountdown

Roll Fixed - 0. 066 deg/hr X-1 Day

Gross Pitch - 0. 399 deg/hr X-1 Day

These values were consistent with previous measurements. Redundant gyrotorquing currents were normal. Maximum amplitude oscillation occurred at80 seconds and was 80 deg/hr peak to peak.

Gyro temperatures were not instrumented for this flight.

Accelerometer scale factors measured during the precount and countdownwere as follows:

X Y Y

Z.000Z6 1.99858 Z.00023

These values were consistent with previous measurements.

Computer

Computer operation was satisfactory. All discretes and yaw steering wereproperly generated. Pre-arm was not issued, demonstrating the azimuthresolver inhibit capability. SCO was issued prior to Mod III ASCO.

Yaw steering was saturated left polarity at liftoff, diminishing slowly in ampli-tude to a crossover at 57 seconds, and then continuing to right saturation atBCO. At guidance enable a right command caused the missile to yaw 7. 60right and to stabilize on this heading at guidance enable plus 26 seconds.

The tests of the Range tape recorded during the countdown indicated that, byYisual inspection, the quality of the telemetered MGS signals was satisfactory.A normal data checker run could not be made because the accelerometerstrings were not instrumented on telemetry.

Computer voltages were not instrumented for this flight.

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( Page No. V7

AA 61-0146

Alignment Countdown Set

The Alignment Countdown Set (A-CS) operation was satisfactory. Accelero-meter zeros were within the specified tolerances before launch, as measured

with the A-CS. inaicating proper operation of the zeroing loops.

Function Nominal Computed Measured Error

X Offset 0.667 .638885 .640800 40.001915

X Zero 1.000 1.00067689 p0.000677

Y Zero 1.000 1.00149338 40.0014 9

Z Zero 65. Z5407 65. 26227 65. 26007 -0. 0022

(in. Instrumentation

All instrumented channels of the Analog Signal Converter (ASC) operated satis-factorily. ASC temperature was not instrumented for this flight.

The Digital Signal Converter performance was satisfactory.

Telemetry quality on this flight was good, with the normal dropout due tobooster separation. Only one telemetry link was flown on this flight.

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CONVAI R-ASTRONAUTICS

Page No. Z8

AA 61-0146

MOD III RANGE SAFETY AND INSTRUMENTATION SYSTEM

Performance of the Mod III Range Safety and Instrumentation System was sat-isfactory; however, monopulse tracking errors after 60 seconds were largerthan were expected. Also, the noise in the laterai rate data after BCO washigher than anticipated. These unexpected conditions are being investigated.The Mod III System provided range safety and impact prediction functions andwas primary for the generation of the Automatic Sustainer Cutoff (ASGO) sig-nal.

A good UIP plot was presented to the Range Safety Officer from approximatelyliftoff through SCO/VCO. The ASCO Inhibit Switch remained in the "Off"position for the entire flight and the ASCO signal was generated properly at304. 826 seconds.

Missile position data were available from liftoff to 393. 6 seconds. Rate datawere available from liftoff until 4Z5.1 seconds. Track and Rate Subsystemdata were interrupted momentarily after booster staging.

Performance of the individual subsystems was as follows:

Track Subsystem

Performance of the Track Subsystem was satisfactory; however, larger thannormal monopulse tracking errors were encountered from 60 seconds for theremainder of the flight. The cause of the large tracking errors is still underinvestigation.

The Track Subsystem was in monopulse mode from liftoff until all signal waslost, excluding two seconds of selected conical track near booster staging.Lock was intentionally broken at 393. 6 seconds in order to reposition the rangegate back in an attempt to reacquire the missile beyond the normal range limitof the equipment. Reacquisition was unsuccessful.

Tracking errors following the major portion of the pitch program and just priorto booster cutoff were 0. 20 mils peak to peak. The track console operatorselected conical hold mode for two seconds near booster staging to confirm thereliability of the monopulse lock. There was no change in tracking. Mono-

* • pulse tracking errors were 0.25 mile peak to peak through sustainer phase.

Signal strength during sustainer phase averaged -60 dbm. The signal strengthaveraged -68 dbm when lock was broken at 393. 6 seconds.

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Page No. 29AA 61-0146

Rate Subsystem

Performance of the Rate Subsystem was satisfactory. Intermittent lock wasencountered following liftoff for the first 35 seconds. Thereafter, lock wascontinuous until 425. 2 seconds except for the usual loss at staging.

The rate received signal strength during the sustainer phase averaged -95dbm. All signal was lost at 4Z5. 2 seconds, 32 seconds after the Track Sub-system began the missile reacquisition procedure.

Noise in the lateral rate data after booster engine cutoff was more severe thanwas anticipated. The cause of this phenomenon is under investigation. Post-flight checks revealed no equipment malfunction.

Mod IlI Computer

The Mod 1II Computer operated satisfactorily during the countdown and ensu-ing flight with no malfunctions observed. The simulated flight re-run was made

0 without deviations from the realtime results.

Data obtained near SCO/VCO were extremely noisy on this flight. The follow-ing impact point was calculated from data gathered between SCO/VCO andretro-rockets firing and reflects the noise encountered.

Mean-Miss Standard Deviation

Distance Deviation of the Mean

Crossrange 9.40 Left tO 7.2Snm 41.3Znm

Downrange 3.50 Long 415.52 nm I02.83nm

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Page No. 30AA 61-0146

RLE-ENTRY VEHICLE

A Mark 4 Mod 1B was flown on Missile 30E.

Re-entry vehicle systems operation was satisfactory. Separation appearednormal and the vehicle satisfactorily survived re-entry. Both the C-bandradar beacon and the telemetry system functioned normally. The data cassettewas ejected as programmed and recovered.

Re-entry vehicle functions during the flight were as follows:

Function Time

Range Zero 0

Lockout Switch No. 2 53.6

Lockout Switch No. 1 58.2

Guidance Pro-arm 304.8

PSA Battery No. 1 305.7

PSA Battery No. 2 305.7

Separation 320.9

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r Page No. 31AA 61-0146

PROPELLANT TANKING

Fuel was tanked to flight level Qfn X-2 Day. Flight level was obtained by tank-ing to the overfill PLCU probe., Replacement of a propulsion line on X -I Daynecessitated the drainage of an unknown quantity of fuel from the engine fuelplumbing. Fuel was retopped to the overfill PLCU probe. As a result of thisunknown quantity of fuel drainage, an ignition fuel density cannot be determined.

Based on the fuel density as determined during X-Z Day tanking there wereapproximately 76, 200 pounds of fuel aboard the missile.

Fuel API Rating 43. 6 DegreesFuel Temperature 87. 0 dgfFuel Density on X-Z Day 49. 68 lbs/ft3

L02 tanking was concluded with a satisfactory 1.02 slug transfer of 4Z. 9 sec-onds duration as measured from the activation of the topping high probe to theactivation of the 100 per cent slug cutoff probe. Slug discharge pressurepeaked at 260.9 psig. decayed to 226.1I within 6 seconds and returned to 260. 3psig prior to activation of the 100 per cent slug cutoff probe.

The 100 per cent slug cutoff probe uncovered 1. 36 seconds after ignition.This indicated that there were approximately 850 pounds of 1.02 above the 100per cent probe at ignition.

Based on tank volume and an igni~tion 1.02 density of 70. 32 lbs /ft 3 there were

approximately 175, 450 pounds of 1L02 aboard the missile at ignition

Weather Data

Fuel Tankitz- Ignition

Barometric Pressure 29. 850 Inches of Hg 30. 020 Inches of Hg

Ambient Temperature 81. 30r 73. 30 F

Relative Humidity 64 Per cent 91 Per cent

Wind - Velocity and Direction 3 Knots - West 4 Knots - East

Cloud Cover 1/10 8/10

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Page No. 3Z

AA 61-0146

AIRFRAME INTERNAL INSTRUMENTATION

Operation of the Telemetry System was satisfactory. Good signals werereceived at the AMR Ground Station well beyond powered flight. The normaldropout of telemetry signal occurred after staging and lasted for 0. 6 seconds.The signal for D I V, RSC Cutoff. was not received on Channel I - 6, althoughit activated satisfactorily on the commutated segment.

There were three measurements that did not operate properly:

P 709 T SGG Combustor Remained at zero during flight.

P 330 P S Fuel Pump Dlsch Remained at zero during flight.

P 84 B B2 Pump Speed Qualitative only.

* ~. Missile 30E contained one Bendix Mod 7 FM Telemetry package. Basic telem-etry channel assignment is given in General Dynamics/Astronautics reportACZ-27-059-30.

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Page No. 33AA 61-0146

LANDLINE INSTRUMENTA T0N

Landline measurements on Brown recorders and E-A sequence recordersprovided satisfactory data. The following minor discrepancies were noted:

Measure-ment No. Description Discrevancy

r 1047 P PCU Fuel Sensor Line Out of ink at liftoff.

P 1712 T BZ Nacelle Ambient Out of ink at liftoff.

P 1325 T Engine Compartment Amb Timing Pen out of ink at liftoff.

P 1710 T Sustainer EngEnvironment Timing pen out of ink at liftoff.

0P 1020 T Bi LOZ Pump Inlet Irregular paper motion at liftoff.

P 1054 T BZ LOZ Pump Inlet Irregular paper motion at Uftoff.

H 1360 P HPU Pressure Provided qualitative data only.The transducer will be replaced.

A trailing wire umbilical provided data during the first forty feet of motion.Out of twelve measurements, ten provided good data. P 1206 0, SustainerEngine LOZ Dome. and P 1001 P. BI LO0 Pump Inlet.were invalid.

Aft-end Motion Determination

Satisfactory photographic data were obtained from the lights mounted on themissile skin for determination of aft-end motion during the first 25 feet ofrise.

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CONYAIR-ASTRONAUTICS

Page No. 34AA 61-0146

CONCLUSIONS AND RECOMMENDATIONS

Conclusions

1. The flight was successful.

2. All systems performed satisfactorily.

Recommendations

1. Resolve discrepancies in flow rates used to compute propellant residuals.

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Page No. 35AA 61-0146

COUNTDOWN TIME VERSUS EVENTS

This test was scheduled for a 150 minute countdown and began on time at0530 EST. There were two unscheduled holds resulting in 4Z minutes holdtime. Total countdown time was 19Z minutes. The holds were as follows:

1. At -35 minutes (0725 EST) for 41 minutes due to poor weather conditions.Rain in the area was adversely affecting camera coverage. As soon asthe camera lenses were wiped dry the count was resumed at -35 minutesat 0806 EST.

2. At -3:30 minutes (0837 EST) for approximately 1 minute to completeliquid oxygen topping. Count was resukaed at -3:30 minutes at 0838 EST.

The following notations were made by an observer in ths blockhouse.

Countdown Countdown

EST Time Proceduen

0530 T-130 T-150 Telemetry Radiation Test.

Computer Warmup Test Started.

0535 T-145 T-150 Acoustic& Sin.sor Response Check Completed.

0536 T-144 T-144 Telemetry Br.ctery Activation Started.

0538 T-142 T-145 ALL GAP Test Support Systems to Channel 5.

0539 T-141 T-147 Telemetry Coming Up for GAP Test.

0547 T-133 T-144 GAP Test Completed - "Go".

Telemetry Off.

0548 T-132 Securing Hydraulics.

0550 T-130 T*144 Telemetry Internal Power Checks.

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Page No. 36AA 61-0146

Countdown CountdownEST Time Procedure Event

0553 T-127 T-135 RSC Tests Started.

0604 T-116 T-125 RSC Tests Completed.

Begin Red Box Installation.

0613 T-107 T-125 Red Box and Retro-rockets are Installed.

0616 T-104 T-115 Start Nose Cone Arming and Fusing Checks.

0619 T-101 T. 100 Flight Control Systems Test Started.

Activate Main Missile Battery.

0621 T-99 T-105 Remove Landlines.

0622 T 98 Arming and Fusing Checks Completed.

0630 T-90 F~ight Control Tests Completed.

Begin Gyro Drift Test.

0644 T-76 Fuel Level Between 100 Per cent and

Overfill.

0654 T-66 Started GE Beacon Tests.

0656 T-64 T-65 Started Landline Calibrations.

Weather Report.

0659 T-61 T-70 Helium Storage Started.

0702 T-58 T-55 Begin GAP Test Preparations.

T-55 Started Telemetry Warmup.

0704 T-56 T-52 Readiness Callout by Flight Control - All "Go".

T-5Z Begin GAP Test.

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Page No. 37AA 61-0146


0711 T-49 T-52 GAP Test Completed - IGo".

Landline Calibrations Completed.

All Personnel Clear Test Stand Area.

Remove All Private Vehicles.

L402 Tankine Preparation Started.

0719 T-41 LO2 Tanking Preparation Completed.

0725 T-35H Holding for Camera Coverage - EstimatedTwenty Minutes.

0731 T-35H Securing Helium Storage for Duration ofHold.

0735 T-35H GE Reports ARMA Liftoff Pulse is NowCorrect and GE Is "Go".

0806 T-35C Resume Helium Storage and LN2 Flow.

Begin L402 Tanking.

Begin Final Guidance Computer Check.

0810 T-31 T-30 BeSin Flight Control Systems Final Checks.

0818 T-23 T-25 Begin Final Azusa and Range Safety Comma.-.dChecks.

0821 T-ZC Retract Hooks.

Start Telemetry Final Warmup.

0828 T-13 Station 3 IGOR Estimates 0 Per cent 'overage.

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,) Page No. 38AA 61-0146

Countdown Countdown

EST Time Procedure Event

0829 T-1Z T-I2 Telemetry and Nose Cone to External Power.

Azusa is "Go".

Begin Telemetry/RSC AGC Check.

0831 T-10 T-10 Flight Control System Final Checks Com-pleted - "Go".

Final Acoustica Check.

0832 T-9 RSC Final Check Completed - "Go".

0833 T-8 Acoustica Final Check Completed - "Go".

Final Range Clearance.

Started Guidance Final Checks.

0835 T-6 T-6 All Personnel Remain in Present Position.

0836 T-5 T-5 All Communications to Channel 1.

0836 T-4:30 T-4:30 Squibs Disarm Switch to "Off".

0837 T-3:50 T-3:50 Status Check - All "Go".

0837 T-3:30H Holding for LOZ Tanking.

0838 T-3:30H Status Check - "Go".

0839 T-3:30C T-3:30 Resume Count.

Telemetry to Internal.

0839 T-3:00 T-3:00 Timer "ýOff" - Ready Switch to Ready.

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Page No. 39AA 61-0146

Countdown Countdown

EST Time Procedure Event

0839 T-Z:4S T-Z:45 Shut Down Power Switch to Arm.

0840 T-Z:05 T-Z:05 Commands to Internal.

0840 T-Z:00 T-Z:00 Nose Cone Beacon and Telemetry to Internal.

Nose Cone to Ready.

0840 T-l:55 T-1:55 Autopilot to Arm.

0840 T-1:50 T-1:50 Turning Water Systems "On".

0840 T-1:45 T-1:45 Commands to Arm.

0840 T-1:40 T-1:40 Range Ready Switch "On".

"0840 T-1:35H T-1:35H T-1:35 and Holding - All Pre-start Lightsare Green.

Slug Start.

0841 T-1:35C T-1:35 T-1:35 and Counting.

Proceeding to Flight Pressurization.

0841 T-1:10 T-1:10 Missile to Internal Power.

0841 T-1:05 T-1:05 Missile Helium to Internal.

0842 T-0:60 T-0:60 T-60 Seconds and Holding

T-60 Seconds and Counting.

0842 T-0:50 T-0:50 Water Full Flow.

0842 T-0:35 T-0:35 Status Check - All "Go".

A A L i~mL L . - HI W'L

8• Page No. 40AA 61-0146


084Z T-0:ZO T-O:ZO All Launch Commit Lights Are Correct.

084Z T-0:05H T-0:05H T-5 Seconds and Holding.

All Recorders to Fast.

084Z T-0:05C T-0:05C T-5 Seconds and Counting.

084Z:37 T-0 T-0 Liftoff

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Page No. 41C" AA 61-0146

MISSILE CON FIGU RATION

Airframe

The SM-65E Missile is approximately 71 feet long from the re-entry vehicleinterface to the aft surface of the thrust chambers. The missile structureis comprised of the booster structure and the main propellant tank structure.With the re-entry vehicle attached, the complete missile is approximately 81feet long. The insulation bulkhead was removed for this flight.

Azusa System

Type B-IA coherent carrier transponder in conjunction with the Mark IIGround Station. Utilized a tripod-mounted, elliptical horn antenna mountedin missile Quad IV.

Electrical System

C Remotely activated, primary type, Eagle-Picher main missile battery and

Leland rotary inverter.

Flight Control System

Square canister configuration with forwr.rd rate gyro canister containingpitch and yaw rate gyros. This was the second flight using the 27-41000-831"Transistors" Servo Canister (previously flown on Missile 25E). The -831"Transistors" Servo Canister was functionally the same as the -831 ServoCanister previously flown on Missile ZF, but incorporated changes to certaintransistors (manufacturer only). This was the third "E" Series flight andthe fourth flight using the 27-4520Z-801 Gyro Canister and the 27-45045-5Forward Rate Gyro Canister which incorporated the Spin Motor RotationDetector (SMRD) System (previously flown on Missiles ZF, Z5E, and 26E).This was the fifth "E" Series flight using the Z7-41001-929 Programmer Can-ister (previously flown on Missiles 12E, 17E, 18K, and ZZE). Missile harnesswas modified to provide simultaneous SCO, VCO, and pre-arm functions.

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Page No. 42.1 AA 61-0146

Guidance System

ARMA Lot IV Missile Guidance Set in conjunction with Lot In GroundEquipment.

Hydraulics Systems

Comprised of two independient systems which provide hydraulic pressurefor booster operation and sustainer/vernier operation. Hydraulic tubing wasof aluminum except that directly associated with the sustainer engine, whichwas of corrosion resistant steel. A Z5 cubic inch accumulator for vernier Molooperation was not required for this flight.

Impact Predictors

Azuca System and Mod IU Instrumentation and Range Safety System.4p,

Pneumatic System

Basic Astronautics system of five shrouded main propellant tank pressuriza-tion titanium helium bottles, and one unshrouded bottle for booster jettison andvernier control. Utilized F and 0 fuel and LOZtank pressurization regulators.

Propulsion System

Rocketdyne MA-3 Liquid Engine Propulsion System. Vernier - solo capabilitywas deleted by removal of solo tanks and associated propellant fill and pres-surization hardware.

Propellant Utilization System

The Acoustics Propellant Utilization system was utilized on this missile. Thissystem uses a 400 cps signal for excitation of the PU valve position feedbacktransducer and a 5-C,%rd computer. No specific gravity 'iases were used.The stillwells utilized were "F'" Series types.

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(-Page No. 43AA 61-0146

Range Safety Command System

Standard system employed two ARW-6Z receivers, a power and signal controlunit, arming switch, and destruct package. Two manually activated, second-ary type, Yardney batteries were utilized.

Telemetry System

The airframe telemetry canister utilized a remotely activated, primary type,Eagle-Picher battery. One airframe telemetry link was operational at ZZ7.7mc.

Mod III Instrumentation and Range Safety Systemn

Mod III E Instrumentation Beacon System in conjunction with the Mod III GroundSystem utilized the "twisted" missileborne antenna (Part No. Z7-136010-3)mounted on the tripod boom in missile Quad IV.

Propellant Tanking

Astronautics "Ell Series Propellant Tanking System incorporated four ultra-sonic fuel sensors, four LOZ/GOZ detectors, a Propellant Loading ControlUnit (PLCU) in the blockhouse, and 200-400 gallon LOZ slug.

Re-entry Vehicle

Mark 4 Mod 1B Re-entry Vehicle contained a C-band tracking beacon. a non-ejectable four pound SOFAR bomb and a recoverable data cassette. The telem-etry system utilized one link, transmitting at ZSZ. 4 mc.

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Page No. 44AA 61-0146

HISTORY OF SM-65E MISSILE NO. 30

Atlas Missile 30E arrived at AMR by air on ZZ August 1961 and placed in Han-gar "K" to receive a minimum of hangar checkout. On 11 September Missile30E was weighed. Missile erection was accomplished on Complex 13 on12 September. Pre-flight testing was accomplished in accordance with plan-ning documented in report AA 61-0117, Flight Test Directive, Atlas Missile 30E.

Significant events concerning Atlas Missile 30E from arrival at AMR to launchare delineated below:

Date Event

22 August 1961 Arrived at AMR.

11 September 1961 Missile Weighing.

4 12 September 1961 Transferred to Complex 13 and erected.

21 September 1961 Successful propellant tanking.

V7 September 1961 Successful Flight Acceptance Composite Test.

5 October 1961 Flight.

A brief description of significant difficulties encountered in system preparationand testing accomplished at AMR is as follows:

Landline Instrumentation

Considerable delay was encountered in checking out and calibrating the engineL4Z dome accelerometers due to unavailability of calibrated transducers.These transducers must be mated with amplifiers and calibrated as matchedsets.

Missile Electrical

No significant problems were encountered.

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Page No. 45C AA 61-0146

Range Safety Command

Upon attempting to electrically connect the Range Safety Command batteries,it was found that one battery had a loose connector at the connector mounting.The problem was corrected by replacement of the battery.

Azusa


Telemetry

The RF Canister. Serial No. 011-0006 and the Accessory Canister, Serial No.103-0017, were returned to San Diego for a transistor check, prior to anytests being performed on the system. After the return of the canisters, testswere satisfactory.

IR 662923 was written on the Pod I Telemetry/RSC antenna because the bondingfingers between the top of the antenna and the tapered fairing above the antennadid not make good RF contact. Bonding fingers were re-shaped and the antennara-installed per the IR. This is a recurring problem and should be correctedin missile assembly at the factory.

Due to Missile 30E configuration containing only one RF canister, TPS 13-1464was written to confirm that none of the deleted wires in the telemetry power andcontrol circuitry were grounded.

The transducers, 27-11956-1, for the booster pump speeds, MeasurementsP 83 B and P 84 B. were SPACE transducers. IR's 645665 and 645666 werewritten to replace these transducers with flight certified articles.

Routine inspection of the missile revealed that the separation lanyard attachedto Plug I00PII on harness Z7-12649-1 was too short to allow proper installa-tion of the plug. IR 645680 was written to relocate the anchor holes for thelanyard so as to acquire sufficient slack to allow proper installation of Plug100Pil.

During routine inspection it was revealed that the Landline InstrumentationPlug ZOIMTI341PI was connected to the sustainer lube oil manifold pressuretransducer and Plug Z01MT341PI was capped and tied back. Since the require-ment for Measurement P 1341 P was deleted for 3E and ON and the requirement

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Page No. 46AA 61-0146

existed for Measurement P 341 P on Missile 30E, the Plug 201MT1341 PIwas removed, capped and tied back and the Plug Z01MT341PI was installedon the transducer. TPS 13-1493 verified that Plug Z01MT341PI was connectedto the telemetry harness.

Due to the lack of potting compound at AMR. the installation of P 830 D couldnot be accomplished per planning card Z7-12646-9001-C. TVA A32608B wasreleased locally to substitute GD/A 894531 high temp wire and pro seal 777potting compound for this installation.

During hydraulic checks, Measurement H 224 P transducer developed a leakand was replaced per IR 653411. The replacement transducer was too shortfor the mounting bracket, and problem was corrected per IR 653531 and TVAA 3Z609.

TVA A3Z655 was released locally on 29 September to add Measurement U 134 X,AA Time Shared Oscillation Output. The operation of this measurement wasverified by TPS 13-1515.

TVA A35982 was released locally to delete the Guidance Measurement I 5Z7 X,pre-arm relay and added I 525 X, Pre-arm Signal No. 1.

Flight Control

Vernier I yaw engine went to a 000 position upon initial application of auto-pilot power with hydraulics at 3000 pounds. The VIY engine position wire wasbroken on the feedback transducer plug P605-B. The defective harness wasreplaced.

The sustainer yaw actuator was replaced due to hydraulic fluid seepage.

Mod III E Instrumentation Beacon


Hydraulics

During normal Booster Hydraulic System testing, the booster return pressuretransducer (H Z24 P Measurement) 27-01241-11 was leaking hydraulic fluidaround the indexing pin. A new -11 transducer manufactured by Colvin Labswas sent to AMR as a replacement. TVA 32609 "B" was initiated to make anew transducer bracket as the one installed (P/N 27-11854) could not be

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Page No. 47

AA 61-0146

utilized because the Colvin Labs -11 transducers have a much shorter bodylength than Bourns transducers.

Missile 30E arrived at AMR with three vernier actuators, manufactured byBohannan. These actuators were replaced with reworked Clemco actuators.

Sustainer yaw actuator Z7-85314-817 was replaced due to a hydraulic leakagebetween actuator body and end cap retaining rings.

Propellant Utilization

On X-Z Day, the PU valve mechanical stop was reset from 20.4 degrees to21.4 degrees per Rocketdyne's request. This entailed re-running 27-90480,Acoustica PU System Calibration during which the MO-Z00 error time generatormalfunctioned. The error time generator was BOI'd from Complex II forcompletion of test.

As a result of Missile 25E flight, special attention was given to the PropellantUtilizaLion System at San Diego's request. The PU harnesses were visuallyinspecte- for possible damage and poor workmanship and all PU measurementswere moaitored on telemetry while harnesses were moved and shaken. Nodiscrepancies were detected.

Propul3 ton

GMA 12907, Engine Electrical 3ystem Modification was sold, but had not beencompleted. GMA should have replaced a 500124 mount bracket with 500126bracket which accommodates at, additional electrical plug. Installation wasIR'd and correct bracket was installed.( Ref. IR 66Z9Z7)

Z7-02403-801 VE Fuel SApply Line was replaced for a blowing leak. A newline had to be removed and dimensionally checked per Inspection Surv,.y 54.Applicable leak check was satisfactorily re-run.

One LOZ topping system Marmon clamp was replaced for distortion (Refer-once: IR 653408)when T-bolts were inspected per SFTO- 184-185 dated9-14-61. System was re-leak checked.

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Page No. 48AA 61-0146

V7- 0Z404-803 VE LOZ Supply Line was IR'd (Reference: IR 653430) for blow-ing leak and was replaced with Z7-02427-1 Line during LO2 self-referencingregulator modification. After modification was complete, the 27-Z4011-69LOZ Regulator Bleed Line interfered with the sustainer low pressure fuel duct.(Reference: IR 65353Z) The structural integrity of the modification was qjues-tionable and system was returned to original configuration.

V-I engine was replaced for excessive hydraulic leakage at the gimbal shaftbleed port. (Reference: IR 645689) The replacement engine also developed aflight hydraulic seep, but it was within ailowable limits for launch. (Reference:IR 653540)

Following readiness day fuel tanking, a very slight fuel dampness was found atthe vernier fuel manifold to fuel tank flange. (Reference: IR 653539) Flangebolts were re-torqued, but seepage continued. Although dampness was notexcessive and was cleared for flight, any leak at this flange is a serious prob-lem. The seal cannot be replaced without removing the sustainer engine andapex to gain access to the inside of the fuel tank in order to disconnect solofuel tank fill and bleed line so vernier fuel manifold can be removed.

B-Z gas genarator fuel supply line was found rubbing on the B-2 turbineexhaust duct (Reference: IR 669811). Line mounting flange was loosened andline was rotated to provide clearance. This necessitated re-leak checkingB-Z chamber.

Complex Mechanical

While performing LOZ tanking preps the seal in the slug unit L02 manual drainvalve was blown. The valve was removed from the system and the line pluggedwith a pressure rated LOZ clean plug. LOZ Tanking Test was satisfactorilyaccomplished with the slug unit in this configuration. This valve failure is arecurring problem. Reference reliability action report 9B-56-603. Thisvalve was not replaced for laknch; however, the system performed satisfactorily.

During readiness fuel tanking, fuel transfer unit valve FB-l would not activate.Tanking was accomplished with Pump FB on 150 gpm flow and pump FA at fullflow (500 gpm). The failure was traced to a grounding of solenoid actuationvoltage which resulted from a faulty conduit installation during solenoid valvereplacement per EOZ35739 on 7-29-038-B/N 3A.

. I N W 11 1 IL I I"L

Page No. 49AA 61-0146

It was decided, with San Diego concurrence to flow 10, 000 gpm in the flamebucket on the launch of 30E. During the test which was run to obtain thisconfiguration it was found that there was no flow in the top jacket of the flamebucket. Upon investigation it wrs found that the TWD manifolds which controlthe flow had been restricted to an almost full closed position. This had beendone to restrict the flow to an acceptable condition while the flow rate was27. 000 4 gpm. These manifolds were opened to approximately 1/4 open posi-tion to give an acceptable flow coverage of the jacket area.

Complex Electrical

Elimination of telemetry cans 2 and 3 from Missile 30E necessitated circuitrevisions at the blockhouse to enable RF ready signal to be received b7 theTest Conductor.

Due to hurricane alert on 9-18-61, preparation was made for removal of themissile from the complex. Umbilicals, lateral motion lights and night tubewere remcved from the missile and then replaced when the alert was termin-ated.

Guidance pod air conditioner failed to operate on 9-18-61 because of a burned-out fan motor which was then replaced by Facility Maintenance.

During fuel tanking, the fuel totalizer failed to operate. Correction was madeby subsequent test prep sheet 13-1428.

During FAC Test, umbilical 600U8 failed to eject electrically. UmbilicalAdapter 600U4 did eject but exhibited high moisture content. UmbilicalAdapter 600U8 was replaced by IR 653446 and subsequent investigation by Relia-bility of the rejected adapter indicated corrosion on the solenoid retainingcollett prevented mechanical ejection.

During performance of Complex Electrical Pre-Readiness Test. BI and SustainerSPGG heatey dc voltages were found to be below specifications. GN2 purgebrought sustainer SPGG heater voltages within specifications indicating moisturein the umbilical adapter had created leakage paths. GN2 purge did not correctBI SPGG heater voltage problem which was caused by two shorted pins inUmbilical Adapter 600U4. Adapter 600U4 was replaced by IR 653549.C

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"Page No. 30AA 61-01(6

During Pre-readiness Fuel Tanking, a shorted wire was found in FPI solenoid.valve in the fuel area. This fault was corrected after completion of fuel tarn-inl,.

Propellant Tanking


Airframe

The vernier fairings were modified to close the excessive gap around theengine gimbal housings to prevent aerodynamic heating inside the fairing. Taevernier pitch actuator clamshell fairings were modified by closing tae openingaround the gimbal housing.

When Missile 30E arrived at AMR, the VE LOZ solo tank bleed extension .-ýuct4 b attached to the fireshield was installed and the L02 bleed line was removce for

weight reduction. EO 235807, (CIC 95353) was written locally to remove t.1Csolo tank bleed extension duct and install a 27-79090-19 plate.

EO Z16448 (CIC 95353 was written on 27-79060 to plug two holea in booste;-thrust section to prevent aerodynamic heating. The holes were left by re,.-ovalof booster pneumatic control lines for booster shut-down kit. The lines we-eremoved for weight reduction.

Pneumatics

The ullage bottles were reject-d for rusty inside surfaces after they ha,. bceereturned from cleaning. It was decided to use the bottles for testing au tocheck a launcher duct in each system to determine if particles had migra.e.:.The ducts were found to be within acceptable cleanliness specificationi.

The Haskel compressor charge panel and several associatedi tubez towa,' ;..ePCU had to be replaced due to rusty sleeves.

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Page No. 51AA 61-0146

A nut on a launcher LOZ pressurization duct assembly Part No. 27-80185-127,was found to be cracked, immediately prior to missile erection. This N.A.S.nut appears to be of insufficient thickness to withstand a launch environmentin which LOZ, LNZ and extremely high temperature could exist. The tubeassembly was located below the Quad IHI and IV riseoff disconnect panel.

Two ullage pressure hoses were IR'd for being frayed.

Inertial Guidance

On 15 September 1961, commenced System Test (CTP-3). While scaling Xin a minus G field, the XFZ string dropped out. Since a heavy rain shower haQoccurred prior to this test, the problem was believed to be in the umbilicalconnector. The umbilical was pulled and purged. Umbilical reinserted and theXFZ problem appeared to be corrected. The platform would not align in scaleX - IG field configuration. Also the pitch axis was drifting.

On 18 September 1961, continued System Testing (CTP-4ZB). Trouble shot4th pendulum problem. Found Sandwich Plug PZ4011 not installed on 3AZ bo.-.The same was repaired. Pod cooling vnit failed. All testing stopped onComplex to facilitate Phase mI hurricane warning.

On 19 September 1961, Phase Ill hurricane alert lifted. Continued System Tests(CTP-4ZB). Umbilicals re-installed into missile and pod cooling failure wasrepaired. Ran three computer problems with "No-Go" results. After theinitial three, all remaining problems were "Go". The thrEe No-Go's" wereattributed to the heavy raias received for four days and that thu pod air condi-tioner was off for approximately 19 hours due to hurricane warnings.

While scaling, the XTV. string dropped out. During preparations for troubleshooting, the string came back in. The platform was not holding in roll.

On 20 September 1961, continued System Checks (CTP.-42B). Ran ComputerProblem No. 120 and received a "No-Go". Re-ran the problem and receiveda "Go''.

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Page N_. 5ZAA 61-0146

Reached proper temperature and went into alignment. The XFZ strinb drop-ped out intermittently.

On Z0 September 1961, checked all pod wirir.g associated with the XFZ stringand found it to be satisfactory. It was decided to replace the Platform. Plat-form Serial No. 7210050 was removed and Platform Serial No. 7210008 wasinstalled. This failure later confirmed in Hangar N.

On 21 September 1961, commenced checkout of new platform. The pitch gimbalwould not travel past -129 . Platform Serial No. 7Z10008 was removed andreturned to ARMA, Hangar N, Component Checkout Lab. Interference wasfound to be a slipped gimbal weight. Repaired same and re-installed. Plat-form Serial No. 7210008 into the missile.

On 2Z September 1961, completed a satisfactory MGS System Test.

On 25 - 26 Septembez 1961. conducted eleven A,-muth Gyro Drift Tests in aneffort to establish cause of n'on-repeatable Azir, ut:, gyro drift measurements.

On 27 September 1961. replaced Computer Seri.l No. 7230015 with ComputerSerial No. 7230052 at the completion of the FAC Test. Computer Serial No.7230015 was out of spec on reset line impedance measurement.

On 29 September 1961, commenced Readiness Checks (CTP-39). Ran severalAzimuth Gyro Drift Tests. After Platform was slewed into stops, a changewas noted in drift readings. It was decided to change the Platform. PlatfornmSerial No. 7210008 was replaced by Platform Serial No. 7A110007.

0- 2 October 1961, commenced MGS System Tests (CTP-42 -nd CTP-39) toverify new platform operation. System Test complete satisfactorily.

On 3 October 1961, commenced Pre-readiness Checks (CTP-42 and CTP-39).During the test, a I degree drift difference occurred in the Azimuth GyroDrift. Ran additional Drift Tests at various inductosyn settings. SlewedPlatform into stops and ran additional drifts. One degree change occurredagain. Pulled umbilical connector and purged for possible dampness. Stilreceived shifting drift metasurements. Checked indui tosyra and drift circuitsand they were all satisfactory.

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Page No. 53

AA 61-0146

On 4 October 1961, continued Azimuth Gyro Drift measurements. Still

received shifting Azimuth Gyro Drift measurements. Completed combination

Pre-readiness and Readiness Test. It was decided that despite shifting drift

measurements, the Platform was suitable for flight.

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Page No. laAA 61-0146

APPENDIX

Page No. ZaAA 61-0146

FLUID CHEMICAL ANALYSIS

Unit Sample Specification

Liquid Oxygen

Purity Per cent 99.5 99. 2 Minimum

HydrocarbonsAs Methane ppm by vol. 18 66.7 Total Maximum

As Acetylene ppm byvol. None 0.62 Maximum

Particle Count50 - 175 Microns 95 No solid particles

175 j Microns 0 greater than 175

Fibers Microns 5 microns. (Fibersnot acceptable.)

This item is acceptable.

( Gaseous Helium

Purity Per cent 99.9 99.94 Minimum

HydrocarbonsAs Methane ppm by vol. 4 58.3 Total Maximum

As Acetylene ppm by vol. None 0.5 Maximum


Gaseous Nitro-Zen

Purity Per cent 98.72* 99. 5 Minimum

Hydrocarbons ppm by vol. None Methane - 58.3 MaximumAcetylene - 0. 62 Maximum

* This item is out of specification.

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Page No. 3a4 AA 61-0146

Unit Sample Specification

Lubricating Oil

Viscosity at 100°F Centistokes 12.2 11. 0 Minimum

Viscosity at 210°F Centistokes 3. Z 3. 0 Minimum

Flash Point 0 417 400.0 Minimum

Viscosity Index 146 80.0 Minimum

Appearance Pass


Tricloroethylene

Appearance Pass Clear and Free

Color Pass Not Red, Blue, Green

or Purple Dyed.

Odor Pass Characteristic

Specific Gravity @680 /68OF 1.466 1. 454 to 1. 476

Distillation °C 86.5 - 87.5 86.5 to 87.5

Water Content @A14. 0°F Pass Cloudless

Non-volatile Per cent .0007 0. 002 Maximum


Fuel, RP-l

Initial Boiling OF 362 Report

10 Per cent OF 390 365 - 410

50 Per cent OF 4Z2 Report

90 Per cent OF 460 ReportEnd Point OF 490 525 Maximum

Residue Per cent 1.0 1.5 Maximum

Loss Per cent 1.0 1. 5 Maximum

Flas4 Point 0 F 138 110 Minimum

Gravity OAPI 43.5 42.0 to 45.0

Particle Countlo - 20 Microns 7.920 No solid particles

20 - 40 Microns 3, 360 greater than 175

40 - 80 Microns 360 microns. (Fibers

175 $ Microns 0 not defined.)

Fiber Count 3

Moisture Content Per cent None

This item is acceptable

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o Page No. 4aAA 61-0146

unit Sample Specification

Hydraulic Fluid - Booster

Flash Point 8F .14 ZOO.0 MinimumViscosity Centistokes 8.8 8.5 Minimum

Color Clear

Dye Red

Particle Count10 - 25 Microns 2,940 5, 500 Maximum

26 - 50 Microns 120 2, 400 Maximum

51 - 100 Microns 60 300 Maximum

Over 100 Microns 1 20 Maximum

Fibers Microns 2 z0 Maximum


Hydraulic Fluid - Sustainer

Flash Point OF 208 200.0 Minimum

Viscosity Centistokes 8.7 8. 5 Minimum

Color Clear

Dye Red

Particle Count10 - 25 Microns 2. 280 5, 500 Maximum

26 - 50 Microns 240 2, 400 Maximum

51 - 100 Microns 60 300 Maximum

Over 100 Microns 1 20 Maximum

Fibers Microns 3 zo Maximum


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Page No. 5a

AA 61-0146

REFERENCE DOCUMENTS

Flight Test Plan - Missile No. 30E AE 60-0801

Flight Test Program - SM-65 Series E.R & D Missiles AZC-Z7-005

Detailed Test Requirements (AFBMD/STL) STL/OR-60-0000-190Z8

Flight Test Directive (FTWG) AA 61-0117

Additional reports which may be referenced for further information regardingthis missile are listed below:

Reprts Approximate Issue Date

(time after test)

Acoustica Associates, Inc., Los Angeles, Calif.

Flight Test Evaluation Report 30 days

General Dynamics/Astronautics, San Diego, Calif.


AFBMD/STL. Inglewood, Calif.

Flight Summary Report 8 - IZ weeks

ARMA, CCO

CCO Quick Look Report 7 - 10 days

American Bosch ARMA Co., Garden City, N. Y.


AVCO, Wilmington, Mass.

Final Flight Report 30 days

General Electric, Syracuse, N. Y.

Evaluation Report of Mod III InstrumentationSystem with Missile 30E 6 - 10 weeks

CONVAIR*ASTROWATICS

Page No. 6aAA 61-0142

SERIAL NUMBERS OF SYSTEM COMPONENTS

Aussa Transponder

Canister, Serial No. 731-0080

Re-entry Vehicle, Mark 4 Mod IB - 12

Range Safetr Command System

Range Safety Command Receiver No. 1, Serial No. AF 60-151Range Safety Command Receiver No. 2, Serial No. AF 60-152Range Safety Command Receiver No. 1, Battery, Serial No. 009-0057Range Safety Command Receiver No. Z, Battery, Serial No. 009-0045Range Safety Command Power and Signal Control Unit, Serial No. 101-0244

Propulsion System

Sustainer, Serial No. 222733Booster No. 1, Serial No. 112781Booster No. 2, Serial No. 112840Vernier No. 1, Serial No. 332832Vernier No. 2, Serial No. 332735

Electrical System

Main Missile Battery, Serial No. 107-0474Inverter, Serial No. 002-0060Power Changeover Switch, Serial No. 007-0058

Mod III E Instrumentation Beacon System

Rate Beacon, Serial No. 638027Pulse Beacon, Serial No. 6E1022

Telemetry System

Telemeter RF, Serial No. 107-0023Telemeter IF, Battery, Serial No. 104-0559(2 Accessory Canister, Serial No. 103-001)7

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S Page No. 7aAA 61-0142

Flight Control System

Gyro Canister, Serial No. 108-0185 (217)Forward Rate Gyro, Serial No. 107-0089 (86)Servo Canister, Serial No. 107-0059 (Z04)Programmer Canister, Serial No. 108-0110 (213)

Propellant Utilization System

Computer, %,Ldel CA-108B

Pneumatics System

LO2 Tank Pressure Regulator, Serial No. 905-0298Fuel Tank Pressure Regulator', Serial No. 905-031Z

Inertial Guidance System

Platform, Serial No. 7210007Control, Serial No. 7220063Computer, Serial No. 7230052Analog Signal Converter, Serial No. 021Digital Signal Converter, Serial No. 714G083

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JEeUI%CON VAl R ASTRONAUTICS

Page No. 8aAA 61-0146

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SIONIFICANT DATES DURING TESTING Or "B" SERIS FLIGHT MULES AT AM&

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SIGM4IFCA1Cr DATES PURIMG TZSTIZ!o Or "V- SCRIES FLIGIff M13BILES AT IM tCMr'dI)

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CONVAIR-ASTRO NAUTICS AA 61-0146

SIGNIFICANT IVATV3 ZrIIRING TESTING OF "E" SERIES FLIGHT MLS8ILZS AT AbR

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in " UNCLASSIFIEDCONVAIR-ASTRONAUTICS Page No. 16a

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SIGNIFICANT DATES DURING TESTING O MERCUIYtALAS EMRICLES AT AMA

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