Nasa Cr 142379 Detailed Test Objectives for Mercury Atlas Vehicle 67 D NASA Mission MA 2 1975066485

8/9/2019 Nasa Cr 142379 Detailed Test Objectives for Mercury Atlas Vehicle 67 D NASA Mission MA 2 1975066485

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AEROSPACE CORPORATIONPost Office Box 95085

Los Angeles U5, California

17 October I960

Recipients of Report No. AS-60-0000-00338, "Detailed TestObjectives for Mercury/Atlas Vehicle 67-D," should note that thereport is a complete revision of the currently existing DTO for67-D, issued as STL Report No. STL/OR-60-0000-69020. The STLreport should be destroyed in accordance with existing securityregulations.

B. A. HoProgram DirectorMercury Space-Booster

c *


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Technical Operating RepoBOB Approval No. 21R-138. 1 /AS-60-0000-00338J)Copy /^ of

DETAILED TEST OBJECTIVESFOR

MERCURY/ATLAS VEHICLE NO.J>7-D(NASA MISSION MA,

Prepared ByMercury Space-Booster

Program OfficeThejAero space Corporation\ El Segundo, California\ .,

^ContractAF04(647)594October I960

AIR FORCE BALLISTIC MISSION DIVISIONAIR RESEARCH AND DEVELOPMENT COMMANDUNITED STATES AIR FORCEInglewood, California


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AS-60-0000-00338Formerly: STL/OR-60-0000-6

DETAILED TEST OBJECTIVESFOR :

i

RCURY/ATIAS VEHICLE NO.-67-D(NASA MISSION MA-2)

For AEROSPACE CORPORATION

Approved B. A. HOHMANNProgram DirectorMercury Space BoosterFor AIR FORCE BALLISTIC MISSILE DIVISION:

Approved

Approved

H.B. Kuchemann, Jr.Director.Air Force Space BoostersSpace Programs

. A. CRISTADORO, JR.Colonel, USAFDirector for Atlas Programs

For NATIONAL AERONAUTICS AND SPACE ADMINISTRATION:

ConcurrenceDirector, Project Mercury


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AS-60-0000-00338Page ii

FOREWORDThis document was prepared in accordance with the specificationspresented in Reference (m) as part of the National Aeronautics andSpace Administration (NASA) Project Mercury.The flight of Mercury/Atlas vehicle 67-Dis the third of a seriesintended to develop, and eventually provide, the ability to placea manned satellite into orbit around the earth and then return itand its pilot safely to the surface.NOTE: Because of major changes made to the original flight planfor the MA-2 mission and because of the transfer of the STLMercury Project operations to the Aerospace Corporation, thisdocument supersedes STL reportSTL/OR-60-0000-69020,"Detailed Test Objectives for Mercury/Atlas Vehicle No. 67-D,"dated July 5, I960 as amended by STL TWX CM 6300. 7-332.


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AS-60-A001-00338Page iiiRevisediNovember I960CONTENTS

Section1. 0 INTRODUCTION 11.1 Scope 11. 2 Purpose 11. 3 Precedence of Reports 21. 4 Amendments . 2

2. 0 .GENERAL TEST PLAN 33. 0 TEST OBJECTIVES 4

3. 1 NASA Objectives 43. 2 Atlas Systems Objectives 44.0 DATA REQUIREMENTS 6

4. 1 External Data 64. 2 Internal Data 95. 0 TEST CONFIGURATION 17

5. 1 NASA Capsule Systems 175. 2 Atlas.Systems 175. 3 Ground Equipment 205. 4 AMR Test Support 21

6. 0 TEST SPECIFICATIONS 226. 1 General 226, 2 Abort Sensing Implementation System . 226. 3 Flight Control System 226. 4 Guidance System 236. 5 Hydraulic and Electric Systems 236, 6 Instrumentation 236. 7 Pneumatic.Systems 236.8 Propellant Loading System 246.9 Propellant Utilization System 246.10 Propulsion System 256. 11 Radiation Frequencies 256.12 Weather 25


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CONTENTS (Continued)Section

7. 0

8. 0

9. 0

TEST PROCEDURES7. 1 Preflight Procedures . . . .7. 2 Flight Procedures7. 3 Data Analysis RequirementsTRAJECTORY AND FLIGHT SEQUENCES8. 1 Trajectory.8. 2 Flight Events SequenceREFERENCES

Page2828282931313136

APPENDIX A 38


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AS-60-A001-00338Page vRevised November I96TABLES

PageI. CAPSULE INTERNAL DATA REQUIREMENTS 10II. ATLAS INTERNAL DATA REQUIREMENTS 12IIA. LANDLINE DATA 16III. RADIO FREQUENCY SUMMARY 26IV. FIRST STAGE PITCH PROGRAM 27V. ASIS THRESHOLD VALUES 27AVI. FLIGHT EVENTS SEQUENCE 32


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1.0 INTRODUCTION

1. 1 ScopeThis Detailed Test Objectives (DTO) Report designates the plan for

the launch and powered flight phase of the Mercury/Atlas Vehicle 67-DThis flight is to be conducted as the NASA MA-2 operation at the AtlanticMissile Range (AMR). The report specifies the test objectives, datarequirements, vehicle general configuration, recommended ground equip-ment and AMR support, systems operation specifications, test procedures,and trajectory necessary to conduct the flight in the prescribed manner;except for the following exclusions:

1. 1. 1 The DTO does not include detailed descriptions or detailedoperating specifications for the Mercury capsule systems. These ele-ments will be contained in References (p) and (q).

1. 1. 2 The DTO does not specify the facilities, procedures, orother support elements required for the free-flight and recovery of theMercury capsule. These elements will be specified in Reference (p).

1. 1. 3 The DTO does not specify the detailed test proceduresrequired to perform Atlas systems checkouts during prelaunch and count-down. These elements will be specified in the Flight Test Directive (FTD)document.1. 2 Purpose

The DTO was established by Reference (a) and modified to fit theProject Mercury operations to fulfill the following purposes:

1.2. 1 To serve as the primary document for the coordination of thetest plan between the NASA and AFBMD/Aerospace.

1. 2. 2 To furnish the flight planning details that are necessary forthe establishment of the detailed test procedures within the FTD and thedetailed support requirements within the OR.


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C O N F I D E N T I A LAS-60-0000-00338Page 2

1. 3 Precedence of ReportsIn case of conflict between the DTO and the test plans prepared by

the Atlas Associate Contractors, the DTO will take precedence. In casethe conflict between the DTO and the detailed test procedures within theFTD, the FTD will take precedence. In cases of conflict between theDTO and the operational procedures and support requirements for theMercury capsule within the OR; the OR will take precedence.1.4 Amendments

This DTOwill be amended as required to reflect approved changesin the over-all plan for the flight. Last minute changes will be trans-mitted by TWX to NASA, Atlas Associate Contractors, and/or agenciesaffected. Published copies of the changes will be mailed to all agenciesand personnel on the original DTO distribution list.


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AS-60-A001-00338Page 3Revised November I92. 0 GENERAL TEST PLAN

The Mercury/Atlas vehicle will propel the instrumented capsule ona ballistic trajectory which is shaped to provide the full scale conditionsof dynamics and heating that is expected for a capsule re-entry from post-stage abort. The vehicle will be guided (pitch plane azimuth of 108 ) tomatch a capsule exit trajectory at an altitude of approximately 89 nauticalmiles, at which point the capsule will be released. The trajectory ter-minates at approximately 1257 nautical miles downrange from CapeCanaveral, where the capsule is to be recovered.

The primary purpose of this test is to determine the integrity ofthe Mercury capsule structure, ablation shield, and afterbody shingles,as well as to determine heating rates and dynamic characteristics of thecapsule during re-entry. In addition, the test will provide an evaluationof on-board capsule equipment and an evaluation of the compatibility ofthe capsule escape system with the Mercury/Atlas System. Specialinstrumentation has been added to investigate the structural integrity anddynamics of the capsule/Atlas interface area in an effort to ascertain thecause of the failure of the previous flight (50-D/MA-l).

Four and one quarter (4. 25)seconds of launch hold-down time delayhas been provided in conjunction with dual rough combustion acceler-ometers and ground circuitry on the B. and B_ engines.

A Flight Readiness Firing (FRF) of the Mercury/Atlas vehicle willbe conducted.


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3. 0 TEST OBJECTIVES

3. 1 NASA Capsule Objectives3. 1. 1 Determine the integrity of the structure,

ablation shield, and afterbody shingles, fora re-entry from a critical abort.

3. 1. 2 Determine the full-scale motions and afterbodyheating rates during re-entry from a criticalabort.

3. 1. 3 Evaluate the performance of the operatingsystems during the entire flight.

3. 1. 4 Evaluate the compatibility of the escape systemwith the Mercury/Atlas System.

3. 1. 5 Establish the adequacy of the location andrecovery procedures.

3. 1. 6 Evaluate prelaunchs launch, and flight moni-toring procedures and facilities.

3. 2 Atlas Systems Objectives3. 2. 1 Determine the ability of the Atlas booster to

release the Mercury capsule at the conditionsof position, altitude, and velocity defined bythe guidance equations.

3. 2. 2 Determine the closed-loop performance of theAbort Sensing Implementation System.

Order of sObjectives

Definitions ofOrders ofObjectives are given in Reference (b).


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Order ofObjectives3. 2. 3 Evaluate the aerodynamic loading,

vibrational characteristics, andstructural integrity of the Atlas LO?boil-off valve, tank dome, capsule adapterand associated structures.

3. 2. 4 Determine the magnitude of the sustainer/vernier residual thrust after cutoff

3. 2. 5 Obtain Data on the repeatability of theperformance of all Atlas missile andground systems.

3. 2. 6 Demonstrate the suitability of equipmentand procedures used for checkout and launchingthe Mercury/Atlas vehicle.

3. 2. 7 Evaluate the Mercury/Atlas vehicle withregard to engine start and potential causesfor combustion instability.


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4. 1. 2. 3 All other RTS dataexcept for above.4. 1. 2. 4 Computer intermediate

tapes.4. 1. 2. 5 Sanborn recordings

of shutoff discretes.4. 1. 2. 6 Computer output tapes.4. 1. 2. 7 Plotter data.

4. 1. 3 AMR Radar Tracking Data4. 1. 3. 1 Metric data suitable for

determining range,range rate, lateral rateand position from liftoffuntil end of.flight or tothe limits of trackingsystem.

4. 1. 4 AMR Optical Tracking Data

Test ObjectivesData Priority

3.2. 1 (R)

3.2. 1 (M)

3.2. 1 (R), 3.2.5 (R)3. 2. 1 (D)3.2. 1 (D)

3. 2. 1 (R)

4. 1.4. 1 Engineering sequentialmotion picture data, pref-erably color, for 360coverage of the missileduring engine start andliftoff. 3. 2. 5 (M)


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4. 1. 4. 2

4. 1. 4. 3

Engineering sequentialmotion picture data, pref-erably color, for cover-age of the missile andpropulsion exhaustplumes from l i f to ffthrough staging or to thelimit of tracking system.Metric camera data suit-able for determiningmissile accelerationduring the first 300 feetof travel.

4. 1.4. 4


Metric camera datasuitable for determiningmissile position and/ velocity from l i f t o f f to

:he end of powered flightor the limit of the trackingsystem.

4. 1. 5 Range Safety Data4. 1. 5. 1 Records of all trans-

mitted RSC signals.4 . 1 . 6 Weather Data

4. 1.6. 1 Atmospheric data suitablefor determining the surfaceindex of refraction

3 . 2 . 5 (R)

3.2.5 (D)

3 . 2 . 5 (D)

3.2. 5 (M)

3. 2. 1 (M)


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4. 1. 6. 2 Weather forecasts as wellas atmospheric data asspecified in References( n ) and (p)

4. 1. 7 AMR Time4. 1. 8. 1 AMR time correlation of

data records.4. 1. 8 Propellant Loading System Data

Time event history of ConvairPLCU system, the fuel totalizer,and load cell data from initiationof tanking to launch.

4. 2 Internal Data4. 2. 1 Telemetry Reception

4. 2. 1. 1 Recording of data shownin Tables I and II fromT-30 seconds until lossof signal. 100 per centbackup is required duringpowered flight.

4. 2. 1.2 Center frequency ofeach telemetry link

4. 2. 1. 3 Signal strength at inputto each ground stationreceiver.


3.2. 1 (R), 3.2.5 (R)

All M

3. 2. 5 (R)

As specified inTable II and III

All (R)

All (R)


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TABLE ICAPSULE INTERNAL DATA REQUIREMENTS

Data Measurement Test ObjectiveOutside skin temperature, afterbody 3. 1. 1Inside skin temperature, afterbody 3. 1. 1Ablation shield temperature 3. 1. 1Structural temperature at 81 other points in capsule 3. 1. 1Cabin air temperature 3.1.3Cabin pressure 3. 1. 3Helium high pressure, automatic control system 3. 1. 33-volt DC power supply 3. 1. 37-volt AC power supply 3. 1. 3115-volt AC voltage -3.1.324-volt DC voltage 3. 1. 3DC current 3. 1. 3Capsule ground reference 3. 1. 3AC power failure 3. 1. 3DC power failure 3. 1. 3Pitch, yaw, and roll attitudes 3. 1. 2, 3. 1. 3Pitch, yaw, and roll rates 3. 1.2,3. 1.3Oxygen supply pressure - 3. 1. 3Oxygen emergency supply pressure 3. 1. 3Reaction jet high pressure; pitch, yaw, roll 3. 1. 3Reaction jet low pressure; pitch yaw, roll 3. 1. 3Vibration levels 3. 1. 1Sound levels 3. 1. 1Accelerations s 3-axis 3.1.2Parachute compartment pressure 3. 1. 3Heat shield cavity static pressure 3. 1. 3


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TABLE I (Continued)Data MeasurementHorizon scanner outputsCommand receiver signalPeriscope retreat signalEvents

Time elapsed from launchTower releaseCapsule separationEscape rocket firedRetropack jettisonMaydayAntenna fairing releaseMain chute deployDrogue chute deployMain chute jettisonReserve chute deploy0. 05g relay actuationStandby inverter onStandby battery on

Onboard Data Recording EquipmentEarth and sky cameraPeriscope cameraTape recorderCosmic ray film pack

Test Objective3. 1. 33. 1. 33. 1. 3

3. 1. 33. 1. 33. 1. 33. 1. 33. 1. 33. 1. 33. 1. 33. 1. 33. 1. 33. 1. 33. 1. 33. 1. 33. 1. 33. 1. 3

3. 1. 23. 1.2

3. 1. 1, 3. 1.2. 3. 1. 3,3. 1. 3


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TABLE IIATLAS INTERNAL DATAREQUIREMENTS

Data MeasurementsA 622 I Th Sect Light Quad 4A 828 D -- Manhole Cover-RetropackA 842 D -- BI POD Door Displacement, Sta 960 QIVA 845 X -- Boiloff Duct to Bulkhead and Duct to ValveA 838 X -- Boiloff Duct to ValveA 833 P -- Adapter Differential Press, No. 1A 834 P -- Adapter Differential Press, No. 2A 835 P -- Adapter Differential Press, No. 3A 836 P -- Adapter Differential Press, No. 4A 827 O -- Manhole Cover LongitudinalA 826 O -- Adapter Base TangentialA 830 O -- Boiloff Valve TangentialA 832 O -- AdapterA 825 O -- Adapter Base LongitudinalA 844 S -- Differential Strain, Yaw Plane, Sta 512A 640 S -- LO- Tank Strain, B2 Side, Sta 512A 843 S -- LO2 Tank Strain, Bl Side, Sta 512D 1 V -- RSC Cutoff OutputD 7 V -- No. 1 RSC RF Input/AGCD 3 X -- RSC Destruct OutputD 120 X -- Destruct Enable#E 50 Q -- 400 Cycle AC PwrsupE 28 V -- MSL Systems InputE 51 V -- 400 Cycle AC Phase A*E 34 X -- AC Low Voltage

Test ObjectivesData Priority3. 2. 5 (D)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3.2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 3 (R)3. 2. 5 (M)3. 2. 5 (R)3. 2. 5 (M)3. 2. 5 (R)3. 2. 2 (M),3. 2.5(R)3. 2. 2 (M),3. 2.5(R)3.2.5 (R)3. 2. 2 (M)

Provides data of ASIS performance.


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AS-60-A001-00338Page 13Revised November I96TABLE II (Continued)

Data MeasurementsI P - - LO2 Tank Helium

F 3 P -- Fuel Tank Helium F 116 P -- Differential PressF 246 P -- Booster Tank, He Bottles HiF 291 P -- Sustainer Control He Bottle

*F 132 X -- LO2 Tank Pressure Switch*F 133 X -- Delta P Switch*F 152 X -- BCO-LOX Tank Pressure SwitchG 4 C -- Pulse Beacon Magnetron AverageG 82 E -- Rate Beacon RF OutputG 3 V -- Pulse BeaconAGCG 279 V -- Rate Beacon AGC No. 1G 287.V -- Decoder Pitch OutputG 288 V --Decoder Yaw OutputG 290 X -- Decoder Contacts No. 1 and No. 2H 33 P -- Booster 1 Hyd AccumulatorH 140 P -- Sust/Vernier HydPressure

*H 220 X -- Sustainer Hydraulic Pressure SwitchM 79 A -- Missile Axial Acceleration, Fine

*M 28 X -- Abort System SignalM 91 X -- Missile 8 inch MotionP 83 B -- Booster 2 Pump SpeedP 84 B -- Booster 1 Pump SpeedP 349 B ---Sustainer Pump SpeedP 528 D -- Sustainer Main Fuel ValveP 6 P --Sustainer Thrust ChamberP 28 P -- Vernier 1 Thrust Chamber

Test ObjectivesData Priority3. 2.2 (M), 3. 2.5 (M)3.2.5 (M)3.2.2 (M)3. 2.5 (R)3.2.5 (R)3. 2. 2 (M)3. 2.2 (M)3.2.2 (M)3.2. 1 (R)3.2.1 (R)3.2.1 (R)3.2.1 (R)3.2.1 (R)3.2.1 (R)3.2. 1 (R)3. 2.5 (M)3. 2,2 (R)3. 2.2 (R)3. 2.4 (M)3. 2.2 (M)3. 2.2 (M)3. 2.5 (R)3. 2,5 (R)3. 2.5 (R)3. 2.5 (D)3. 2.5 (R)3. 2.5 (R)



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TABLE II (Continued)

Data MeasurementsP 29 P -- Vernier 2 Thrust ChamberP 55 P -- Sustainer Fuel Pump InletP 56 P -- Sustainer LCXPump InletP 59 P -- Booster 2 Thrust ChamberP 60 P -- Booster 1Thrust ChamberP 100 P -- Booster Gas Generator CombustionChmP 339 P -- Sustainer Gas Generator DischargeP 14 T -- Engine Compartment Ambient TemperatureP 72 X -- Booster Cutoff RelayP 77 X -- Vernier Cutoff Relay

*P 300 X -- Booster 1 Injector Manifold P Sw*P 313 X -- Booster 2 Injector Manifold P Sw*P 574 X -- Sustainer Injector Manifold P SwS 61 D -- Roll Displacement Gyro SignalS 62 D -- Pitch Displacement Gyro SignalS 63 D -- Yaw Displacement Gyro SignalS 252 D -- Bl Yaw RollS 254 D -- Booster 1 PitchS 256 D -- Sustainer YawS 257 D -- Sustainer Pitch

*S 52 R -- Roll Rate Gyro Signal*S 53 R -- Pitch Rate Gyro Signal*S 54 R -- Yaw Rate Gyro Signal

Test ObjectivesData Priority3. 2. 5 (R)3. 2. 5 (R)3. 2. 5 (R)3. 2. 2 (M), 3. 2. 5 (R)3. 2. 2 (M), 3. 2.5 (R)3. 2. 5 (R)3. 2. 5 (R)3. 2 ; 5 < ( R )3. 2. 5 (R)3.2. 5 (R)3. 2.2 (M)3. 2. 2 (M)3. 2.2 (M)3.2. 1 (R),3.2.5 (R)3.2. 1 (R), 3. 2. 5 (R)3.2. 1 (R), 3.2.5 (R)3. 2. 1 (R), 3. 2. 5 (R)3.2. 1 (R), 3.2.5 (R)3.2. 1 (R), 3.2.5 (R)3.2. 1 (R), 3.2.5 (R)3.2. 1 (M), 3. 2. 2(M),3.2.5 (M)3.2. 1 ( M ) , 3. 2. 2 (M),3.2.5 (M)3.2. 1 (M),3. 2.2(M),3. 2. 5(M)



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TABLE II ALANDLINE DATA

Data MeasurementsP 1439 O -- S NAA RCC AccelP 1452 O -- BI NAA RCC AccelP 1453O -- B2 NAA RCC AccelP 1893 O -- BI NAA Accel Back-upP 1894 O -- B- NAAAccel Back-upP 1003 P -- B- LO Pump InletC* CtP 1004 P -- B Fuel Pump Inlet

L*P 1006P -- S Thrust ChamberP 1059 P -- B, Thrust ChamberL*P 1060 P -- B Thrust ChamberP 1091P -- BI LO2 Inj ManifoldP 1092 P -- B2 LO2 Inj ManifoldP 1093P -- B Fuel Inj ManifoldP 1094 P -- B2 Fuel Inj ManifoldP 1054 T -- B LO, Pump Inlet TempP 1437 W -- S RCC Binary CounterP 1454 W -- B RCC Binary CounterP 1455 W -- B2 RCC Binary CounterP 1897 W -- B RCC Binary Counter Back-upP 1898 W -- BZ RCC Binary Counter Back-upP 1192 X -- Bj Rough Comb CutoffP 1193 X -- B2 Rough Comb CutoffP 1438 X -- S Rough Comb CutoffP 1975 X -- Bj RCC Back-up RelayP 1976 X -- B RCC Back-up Relay

L*

Test ObjectivesData Priority3. 2. 7 (R)3. 2. 7 (R)-i-l of 23. 2. 7 (R)-i3. 2. 7 (R)J3. 2. 7 (R)J3.2. 7 (R)3.2.7 (R)3.2.7 (R)3. 2. 7 (R)3.2. 7 (R)3.2. 7 (R)3.2.7 (R)3. 2. 7 (R)3.2. 7 (R)3.2. 7 (R)3.2.7 (R)3. 2. 7 (R)3. 2. 7 (R)3. 2. 7 (R)3.2.7 (R)3. 2. 7 (R)3.2. 7 (R)3.2.7 (R)3. 2. 7 (R)3. 2. 7 (R)

( M )-1 of 2( M )


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5.0 TEST CONFIGURATION

5. 1 NASA Capsule SystemsA full-scale production Mercury capsule (No. 6), with an active

escape system, will be used for MA-2. The capsule will also have activeposigrade rockets but the retrograde rockets will be dummies. A detaileddescription of the individual capsule systems are presented in Reference(o), (p) and (q).5. 2 Atlas Systems

The systems installed are identified briefly below. Detaileddescriptions of the individual systems will be presented in the FTD.

5. 2. 1 AirframeConvair Series D missile system with following modifications:a. Mercury capsule-adapter replaces the Convair

re-entry vehicle adapter. (The adapter has beenmodified to provide increased strength characteristicsover the adapter provided formerly for MA-1.)

b. Pod-mounted retrorockets are deleted.c. Skullcap insulation added to top of liquid oxygen tank

dome, boiloff valve and ducting.d. "C" type peacock boiloff valve in lieu of standard "D"

series. (The valve support structure has been changedto provide stronger support members and to eliminatethe attachment of the members to capsule adapterstructure.)

e. 117L rate gyro pod installed.f. The APSjfairing has been deleted (Replaced by production

type fairing).


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NOTE: (To allow for future identification of photographicrecords', the number '%7-D'1 will be painted in14 inch high letters at the base of the fuel tank onthe centerline of Quad I-II and Quad III-IV.)

SystemPriority

5. 2. 2 Electrical PConvair remotely activated battery, rotaryinverter system.

5. 2. 3 Flight Control PConvair Series D "Round" autopilot system,modified to provide the special programmersequence required for this flight. Autopilotcontrol gyros are located at Station 675.

5. 2. 4 GuidanceGeneral Electric Mod III radio guidance system P

, with traveling wave-type antennas in place of theGE slotted antennas to provide improved "lookangle" for the guidance tracking radar.

5.2.5 HydraulicConvair system. (Vernier solo hydraulic package Phas been deleted. )

5. 2. 6 Instrumentation *5. 2. 6. 1 Two FM/FM telemeters for Atlas

systems data.5. 2. 7 Pneumatic P

Convair system.

The priority of the instrumentation will be dictated by the priority ofdata it supports as shown in Section 4. 0.


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SystemPriority

5. 2. 8 Propellant Loading PConvair PLCU system.

5.2.9 Propellant Utilization PConvair Manometer system

5. 2. 10 Propulsion5.2. 10. 1 Rocketdyne MA-2 system consisting of

Booster No. 112105, Sustainer No. 222105,and Vernier engine assemblies.NOTE: (Vernier solo refill orifices

normally used to refill the vernierstart tanks, have been plugged.)

5. 2. 10, 2 RC filter networks have been added to thebooster and sustainer vernier cutoff relaycircuits to reduce the possibility ofinadvertent engine c u t o f f resulting fromspurious transients.

5. 2. 10. 3 A wet start will be conducted.5.2. 10.4 Two independent parallel rough combustion

cutoff (RCC)systems for each boosterengine and one RCC system on the sustainerare installed. For instrumentation purposes,the backup RCC accelerometers will be con-nected through 42"' motion.

5. 2. 11 Range Safety5.2. 11. 1 Range Safety Command (RSC)system P

consisting of two ARW-62 receivers,decoders and associated circuitry which


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SystemPriorityhas been modified to provide: (1) a3-second time delay between enablingof AFCO and initiation of destruction;(Z ) increase the linearity and decreasethe gain of the tone amplifier.

5.2. 11.2 AZUSA impact predictor type B coherentcarrier transponder.

5. 2. 12 Abort Sensing and Implementation System( ASIS)The system is comprised of instrumentation withinthe Atlas booster to sense impending malfunctionand to activate the capsule escape system (seeSection 7. 2. 4). Critical parameters beingmonitored are;(1 ) A. C. Voltage (2) Attitude Rates (3) L C > 2 TankPressure (4) LO_ Fuel tank differential press.( 5 ) Booster engine fuel injector pressure( 6 ) Sustainer engine fuel injector pressure( 7 ) Sustainer hydraulic pressure (8) Capsule/-booster inter-face.This will be a closed loop flight of the system.

5. 3 Ground Equipment5. 3. 1 Launch Complex

5. 3. 1. 1 AMR Complex 14 (2829'27" N,8032'50" W).

5. 3. 2 Test Support and Ground Support Equipment5. 3. 2. 1 TSE and GSE as specified in the FTD.


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C O N F I D E N T I A L A s _ t o . A 0 0 1 . 0 0 3 3 8Page 21RevisedNovember I960SystemPriority

5.3.3 Ground Guidance Station5.3.3. 1 GE/Burroughs Mod IU System at

Cape Canaveral.5.3.4 Impact Predictor System

5.3.4.1 AZUSA AN/FRW-1 (XW-1) system atCape Canaveral.

5. 4 AMR Test SupportThe following AMR equipment is recommended to be furnishedto support the fulfillment of the data requirements specified inSection 4.0 of this document. The recommended equipment isidentified with the data requirement it supports.5.4.1 Radar Tracking Systems

5. 4. 1. 1 FPS-16 radars tracking from Station 1, 3,and Station 5 (4. 1. 3. 1).

5.4. 1.2 Mod III and Mod IV radars (4. 1.3.1).5. 4. 1. 3 AZUSA system tracking from Station 1

(4.1.3.1).5.4.2 Optical Tracking Systems

5. 4. 2. 1 Mitchell and Fastax Cameras (4. 1. 4. 1).5.4.2.2 ROTI and IGOR cameras (4. 1. 4. 2).5. 4. 2. 3 Ribbon Frame cameras (4. 1. 4. 3).5. 4. 2. 4 Cine-Theodolites (4. 1.4. 4).

5.4.3 Telemetry Receiving Stations5.4. 4. 1 Stations 1, 3, 5 and 7 (.4. 2. 1.1).5. 4. 4. 2 Station 9A, using 60-foot dish (4.,2. 1.1).


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AS-60-A001-00338Page 22Revised November I966.0 TESTSPECIFICATIONS

6. 1 GeneralThe specifications presented below are of particular importance for

the achievement of the test objectives relative to the Atlas booster.Specifications (redline values) required for the launching of this missilewill be defined by the Flight Test Working Group (FTWG) at AMR but willbe consistent with the corresponding items listed below.6. 2 Abort Sensing Implementation System

The threshold values of the ASIS sensors will be as specified inTable IV-A.6. 3 Flight Control System

6. 3. 1 Roll ProgramThe vehicle will be rolled during vertical ascent from thelaunch azimuth clockwise to the pitch plane azimuth of 108 .The tolerance of the roll program will be 1.5 per cent or0. 3 whichever is greater.

6. 3. 2 Pitch ProgramThe first stage pitch program will be shown in Table IV.The pitch program is not to exceed 1.5 of nominal. (Thesecond stage pitch program will be set at 2 /second).

6. 3. 3 Programmer SettingsThe programmer events will be as specified in Table V.

6.3.4 Flight Control System Gains6. 3. 4. 1 Reduced pitch and yaw position gains during

launch +85 seconds to booster cutoff.6. 3. 4. 2 Incorporation of a 4 + 8 CPS lag for the stabiliza-

tion filter from launch to +85 seconds and frombooster cutoff to the end of powered flight.


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AS-60-0000-00338Page 23

6. 3. 4. 3 Incorporation of a 4 + 3 GPS lag for thestabilization filter from +85 seconds to BECO.

6. 4 Guidance System6.4. 1 Nominal times of occurrence of guidance discrete commands

will be shown in Table V.6. 4. 2 The following octal values of manual constants are to be

used in the Mod III guidance computer at Cape Canaveral.Octal Value^anual Constant

Hydraulic

J SixJ SevenJ EightJ NineJ Ten

and Electric Systems

Cell11251126112711301131

Two Zero Zero Six ( 2 0 0 6 )One Five Three Seven ( 1 5 3 7 )Zero Five Six Five ( 0 5 6 5 )Three Zero Six Two ( 3 0 6 2 )Four Two Two Zero ( 4 2 2 0 )

No special specifications,6. 6 Instrumentation

6. 6. 1 Telemetry channel assignments for the Atlas telemetersare contained in Reference (e).

6.6.2 Telemetry channel assignments for the capsule telemeterswill be defined by NASA in Reference (p)

6. 7 Pneumatic Systems6.7.1 Regulator Settings

Engine tank pressure regulator: 660psigBooster pneumatic regulator: 750psig


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AS-60-0000-00338Page 24

NOTE: (The above values are the adjustment settings forthe regulators and as such are specific values. Operationalredline values, which indicate maximum and minimumpermissible operating limits and if exceeded could result inserious degradation of vehicle performance, are 25 psigof the set values. )

6. 7. 2 Tank pressure regulators will be set to regulate in flightwithin the following limits:

Fuel tank: 55. 3 to 59. 9 psigLO tank: 23. 6 to 26. 0psigL *

NOTE: (The above values are regulator operating limits andare not redline values. )

6. 8 Propellant Loading SystemLiquid oxygen will be tanked utilizing the Convair 95 per cent tankprobe at sequence III pressure to obtain a known volume and a loadcell reference weight. LO7 is then tanked at sequence II pressureL*by load cells to a level equivalent to approximately 8700 poundsabove the 95 per cent reference level.RP-1 will be tanked to the Convair overfill probe at sequence Ipressure. (Load cells will be used as the standard for determiningthe weight. )

6. 9 Propellant Utilization System6. 9. 1 The Convair PU computer will be nulled for a liftoff weight

ratio of 2. 28:1 nominal.6. 9. 2 Three values of the PU valve position versus sustainer

engine mixture ratio are listed below:Mixture Ratio PU Valve Position

1.93 (-15%) 47.22 .27 28.52.61 (+15%) 22.5


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AS-60-0000-00338Page 25

6. 10 Propulsion System6. 10. 1 Booster LO_ Reference Regulator: 552 psig

Sustainer LO2 Reference Regulator: ' 793 psigNOTE: (The above values are the adjustment settings forthe regulators and as such are specific values. Operatingredline values which indicate maximum and minimum per-missible operating limits and if exceeded could result inserious degradation of missile performance, are +25 psigof the set values. )

6. 10. 2 The booster engines are orificed for a sea level mixtureratio of 2. 29:1 based on the standard inlet conditions.

6. 10. 3 4. 25 seconds of holddown time relay and RCC activationis specified for the launch.

6. 10. 4 Booster RCC setting: 30 g with 40 millisecond count.Sustainer RCC setting: 60 g with 20 millisecond count.

6.11 Radiation FrequenciesThe frequencies of the various electronics systems to be used duringthis test are listed in Table III.

6.12 WeatherThe weather conditions for this test must be within the minimum asdefined by range safety regulations and be consistent with the testobjectives defined herein. Detailed weather conditions for launchand sea conditions for the recovery of the capsule are specified inReference (p).


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AS-60-A001-00338Page 26Revised November I96TABLE III

RADIO FREQUENCY SUMMARYMissile SystemGuidance (GE Mod III)

Rate Beacon - InterrogationReplyReplyPulse Beacon- InterrogationReplyTelemetry

Atlas Telemeter No. 1Atlas Telemeter No. 2Range Safety Command - PrimaryImpact Predictor (AZUSA) - InterrogationReplyGround Radar

Mod II(FPS-lo)Mod IV

Frequency (me)

8387.58462.5842592209310

229.9227. 741450605000

290054808600


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TABLE IVFIRST STAGE PITCH PROGRAM (MA-2ONLY)

Time ~"p( s e c ) (deg/sec)

0 015 017.8 1.12 9 . 8 1.132. 0 0.53 6 . 0 0.542. 9 0.6

5 9 . 5 0.675.0 0.791.8 0.5

1 1 5 . 6 0.2741 3 6 . 0 0.1831 4 6 . 0 0. 150

Where negative pitch rate is a nose-down command,-w is clamped to 0. 150 deg/sec from T + 146 secto staging, if staging has not already occurred.


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7. 0 TEST PROCEDURES

AS-60-A001-00338Page 28Revised November I9

7. 1 Preflight ProceduresRequirements for Atlas hangar tests, system checkout tests, count-

down integration, etc. , will be defined in detail by the FTWG within theFTD and will include the following special procedures:

7. 1. 1 A Flight Readiness Firing (FRF) will be conducted.(Duration - 20 seconds)

7. 2 Flight ProceduresRequirements for the conduct of this flight test will be defined indetail by the FTWG within the FTD and will include the followingspecial procedures.7.2.1 NASA Capsule Systems

These special procedures pertaining to the Mercury capsulewill be detailed in Reference (o), (p), and (q).

7.2.2 Range Safety7.2.2. 1 Range Safety requirements for the launching

and flight of this vehicle will be established by,and be the responsibility of, the AMR RangeSafety Officer.

7. 2. 3 Propulsion SystemThe wet start technique will be used for ignition of thethrust chambers.


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AS-60-0000-00338Page 29

7.2.4 Abort: SystemThe escape rocket on the capsule escape tower provides ameans of separating the capsule from the Atlas in the eventan abort is required during the prestaging phase of the flight.The tower is armed immediately prior to LO, tanking and isshimmed to separate the capsule on an azimuth heading ofapproximately 225 with the M/A vehicle on the launch com-plex.Abort :Prior to Capsule Umbilical Disconnect - An abort canbe initiated by: (1) signal from blockhouse (test conductor)through Atlas 8-inch umbilical to capsule, and (2) signal fromblockhouse (test conductor) through capsule umbilical.Abort Prior to Lift-off; After Capsule Umbilical Disconnect -An abort can be initiated by: (1) signal from blockhouse (testconductor) through Atlas 8-inch umbilical to capsule, and (2)signal from blockhouse, (test conductor) and/or NASA ControlCenter through ground command RF link.Abort. After Lift-off; Prior to SECO/VECO - An abort can beinitiated by: (1) signal from blockhouse (test conductor) andNASA Control Center through ground .command RF link, (2)signal from RSO through RSC link, and (3) signal from ASIS

within Atlas (Note: ability of ASIS to shut down Atlas enginesis deactivated from lift-off until T + 30 seconds).

7. 3 Data Analysis Requirements7. 3. 1 Quick Look Data Evaluation (Atlas Booster)

A preliminary evaluation analysis will be made at AMRwhich will be conducted on the basis of quick look data aswell as other data which may be necessary to determine the


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AS-60-0000-00338Page 31

8. 0 TRAJECTORY AND FLIGHT SEQUENCES

8. 1 TrajectoryThe simulation for the nominal trajectory for this test is presented

in Appendix "A."8. 2 Flight Events Sequence

The sequence of anticipated flight events is presented in Table V.


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TABLE V (Continued)


Nominal TimeIn Seconds126

1 3 C . O

T +

Subroutine

Staging

Staging+0.1second

Staging+3.0secondsStaging+3.7seconds

Staging+5.5 0. 1secondsStaging+6 +1seconds

Function and/or ActionStaging subroutine com-mand is enabled.Staging subroutine isinitiated. R = 395,244 fta. Cut off booster engine.b. Stop programmer pitchprogram.c. Null boster engines inpitch, yaw and roll.d. Activate sustainer enginein pitch andyaw.e. Activate vernier enginesin pitch andyaw.f. Change autopilot gyrogains and integrator limitsin pitch andyaw.g. Change filter in pitch andyaw.Sustainer engine is nulledand booster section jetti-soning commences.Sufficient clearancebetweensustainer and jettisoningbooster section attained;sustainer engine is releasedto gimbal in pitch andyaw.Initiate second stage pitchprogram. (Rate = 2/sec)

Enable guidance pitch andyaw steering

Initiated ByAFCSProgrammerGuidance

AFCSProgrammer

AFCSProgrammer

AFCSProgrammer

AFCSProgrammer

AFCSProgrammer


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AS-60rA001-00338Page 34Revised: November I96TABLE V (Continued)

Nominal TimeIn Seconds

Staging + 25140

170 5

T + 248. 2

SubroutineStaging +6.7seconds

Function and/or ActionVernier engines are nulledin pitch and yaw but con-tinue to gimbal in roll. Theyaw null is 50 away frommissile centerline.

Staging + 23 Jettison capsule escapesecondsStaging+ 22.850. 1seconds

towerSecond stage pitch pro-gram is stopped

Guidance steeringBackup staging command

Sustainer cutoff (SECO)command is enabled.SECO/VECO ASIS is disabled. SECOsubroutine is initiated;sustainer and vernierengines are cut off.

Initiated ByAFCSProgrammer

CapsuleProgrammerAFCSProgrammer

GuidanceAFCSProgrammerAFCSProgrammerGuidance


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AS-60-0000-00338Page 35

Capsule Events Subsequent to SECO/VECOWhen the Atlas thrust decays, to a point where the longitudinal

acceleration is 0. 20gs as sensed by the capsule, the capsule-adapterclamp-ring explosive bolts will be fired to release the clamp and thenthe posigrade rockets will fire to separate the capsule from the Atlas.After separation, the capsule Automatic Stabilization and Control System(ASCS) will call for five seconds of rate damping and then will yaw andpitch the capsule to the retroattitude. While the capsule is in the retro-attitude, a signal from either the ground control center or the airborneinterim clock will initiate the retrofire command to enable the jettisoningof the retropackage (dummy retrorocket motors are used). Ninety sec-onds later, the retropackage will be jettisoned and the ASCS will pitch thecapsule down to the re-entry attitude. When the axial load factor buildsup to 0. 05g, the ASCS will switch from re-entry attitude hold mode to arate damping mode with a constant roll rate of about ten degrees persecond. Upon descent to 42, 000 feet, a barostat circuit initiates deploy-ment of the drogue parachute. At 10,000 feet, the antenna fairing will bejettisoned to deploy the main parachute. At touchdown, an impact switchwill disconnect the main parachute, jettison the pilot and reserve para-chute, activate the recovery aids and turn off unnecessary equipment.


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AS-60-0000-00338Page 36 .

9. 0 REFERENCES

(a) "Flight Test Responsibilities, Procedures and Organization atAFMTC," from WDTT dated 15 October 1956.(b) Aerospace Report AS-60-0000-00336, "General Flight Plan-AtlasBoosters for Project Mercury," dated September I960.(c) STLReport GM-TR-0165-00475, "WS 107A-1 Program Requirementsat AFMTC," dated 12 March 1959.( d ) STL Report TR-59-0000-00673, "Atlas Man-in-Space ProgramRequirements at AFMTC," revised July 1959.(e ) Convair Report "Instrumentation Configuration Series D,Article 50, at AFMTC," AZC-27-066-50, dated 25 January I960.(f ) ARDC Air Force Surveys in Geophysics, No. 57S "Wind, SpeedProfile, Wind Shear and Gusts for Design of Guidance Systems for -Vertical Rising Air Vehicles," dated November 1956.(g ) Convair Report, "Flight Test Plan Missile No. 67D Mercury/Atlas,"AZC-27-073.(h) Convair Report, ZA-7-166, "Performance Summary of the

XSM-65D Missile."(i) North American Aviation Report R 550 S, "Model SpecificationRocket Engine Propulsion System Modll MA-2, for the WS 107A-1Missile."( j ) General Electric Report, "Guidance System Test Plan,Missile 67D."( k ) Convair Report 27-04203, "Subsystem Autopilot, XSM-65D.Missileborne, Specifications for," Revision B, 8 May 1959.(1 ) Convair Report, "Range Safety and Dispersion Information for theXSM-65-D-67 Missile." (to be released)(m) ARDC Report, "Space System Development Plan, Research for andSupport of Project Mercury," 12 June 1959.(n ) AFMTC Document No. MT58-21776, "Operations RequirementsNo. 1003, XSM-65D Atlas Launch," 1 March 1959.


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AS-60-0000-00338Page 37

(o) NASA Document, "Project Mercury General Systems InformationDocument," NASA Project Mercury Working Paper No. 118.(p) AFMTC Document No. MT60-24978 "Operations RequirementsNo. 19.03, MA-2 Mercury Operation," 15 September I960.(q) NASA Document, "Mercury/Atlas MA-2 Mission Directive,NASA Working Paper No. 140.


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AS-60-0000-00338

APPENDIX A


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AS-60-0000-00338Page 38

APPENDIX ATRAJECTORY SIMULATION

This third Atlas launching on Project Mercury, identified as MA-2,is a repeat of the unsuccessful MA-1 (50-D) test. As before, the launchvehicle is guided to NASA-assigned burnout conditions. This subjects thecapsule to the most rigorous re-entry environment it is ever likely toiexperience.

Noteworthy features of the trajectory presented here include: neworbital guidance equations which are being used/for the first time, revisedsustainer engine parameters which reflect recent findings, and all geodeticdata which used here is consistent with the newly formulated DOD/WGS-60geophysical systern.

Significant trajectory results are indicated on Figures A-l throughA-5 and summarized below:

1. Summary of Trajectory Data

Event

LiftoffBECOBooster JettisonTower JettisonSECO/VECOPosigradeApogee (approx)Re-entry (approx)Recovery Starts

Time(sec)

0136.1138.9158.9248.2249.2320530664

Altitude( f t )

-41240,176252,932342,589567,987568,955601,906300,00044,000

Inertia! Inertia! Flight w . ,Velocity Path Angle nl f(ft/sec) (deg) (lb)

1,34211,33511,42312,22018,92318,94518,89119,3901,672

90.065.866.169.787.087.190.098.7

110.8

257,61256,23948,50542,33519,1392,5262,4212,4212,421

SurfaceRange(naut mi)0

505486

277280479

10771256(approx)


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AS-60-0000-00338Page 39

4.

LaunchStandGeodetic latitudeLongitudeMissile antenna above MSLPad azimuthPitch plane azimuthWeights and LoadingsLiftoffAvailable oxidizerAvailable fuelGround runJettisoned boosterJettisoned escape towerCapsule at separationCapsule at re-entryPropulsion -MAZBooster sea level, IspBooster mixture ratioSustainer sea level, IspLiftoff thrustLiftoff flow rateLiftoff thrust acceleration

1428.4913N80.5474W60 ft104.977T108T

257,612 Ib167,407 Ib at70.25 lb/ft372,996 Ib at49.90 lb/ft36.25 sec7025 Ib1012 Ib2533 Ib2421 Ib

252 sec2.29220 sec364,567 Ib1491 Ib/sec1.42 g's


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AutopilotPitch program, First Stage:

Time (sec)


c o (deg/sec)0-1517. 829. 832. 036.042.959.575. 091.8115.6136.0145.0

01 . 11 . 10.50.50.60.60.70.50.2740. 1830. 15

Second Stage: 2 deg/sec, BECO + 5. 5 toBECO + 22. 85Roll programYaw programMaximum dynamic pressureStagingBECO timeTotal acceleration just before BECONet positive suction head (NPSH)Inertial attitude angle

3.023 degNone896 lb/ft2 at 60 sec

136. 1 sec7. 8g's24 ft66.3 deg


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AS-60-0000-00338Page 42

10. RecoveryParachute decent begins approximately

AltitudeTime from liftoffGeodetic latitudeLongitudeSurface range

44,000 ft664 sec20.1038N59.4663W1256 naut mi


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Printout KeyI

Project Mercury Trajectories

AS-60-0000-00338Page 94

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t Time from 2" motion, sech Altitude above (DODWGS-60) reference ellipsoid

along radius vector, ftF-Da Thrust acceleration ( = -= - ) during powered flight orD 2atmospheric deceleration ( = r - = ) during free flight, ft/sec

W Weight, Ib(3 . Inertial flight path angle, radius vector (extended) to

inertial velocity vector, dege Pitch angle, radius vector (extended) to roll axis, deg( 3 Flight path angle, radius vector (extended) to aerodynamic

velocity vector, degA V. Azimuth of inertial velocity vector from North, degZ 1e Inertial pitch angle, launcher radius vector (extended)

at liftoff to roll axis, degW Propellant weight flow rate, Ib/secW. , W . p , Weight of expendable LOX and fuel remaining, IbF Thrust, Ib|i Mixture ratio, LOX flow rate to fuel flow rateNPSH Net positive suction head at sustainer LOXpump inlet.ftFNT Total normal force, IbD Drag force, IbM Mach numbera Angle of attack, roll axis to aerodynamic velocity:

vector, degq Dynamic pressure, Ib/ftr Radius vector magnitude, distance to center of earth,ft


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AS-60-0000-00338Page 96

x, y, z Position in inertial, rectangular coordinates (com-ponents of r); origin at center of earth, x-y planecontains equator, x-z plane contains polar axis andlauncher at t = 0, x-positive at launch, y-positiveeastward, z-positive northward, ft

V. Inertial velocity magnitude, referred to EC! (earth-centered-inertial, x-y-z) coordinate system, ft/sec

x, y, z Inertial velocity components, ft/seca Acceleration magnitude referred to ECI coordinate

system, ft/secx, y, z ' Acceleration components, ft/secR Distance from launcher to projection of missile on. s spherical earth (approximate surface range), naut mius v, w Position in rectangular, earth-fixed coordinates; origin

at launcher, u- v plane normal to local gravity vertical,u-w plane contains vertical and launch azimuth,u-positive downrange, v-positive left looking downrange,w-positive up, ft

V. Aerodynamic velocity magnitude,, referred to air mass,J ft/secu, v, w Components of velocity with respect to earth, ft/secLA Pitch look angle, rearward roll axis to pitch plane

projection of radar line-of-sight, clockwise fromright, deg

LA Yaw.look angle, rearward roll axis to yaw plane pro-jection of radar line-of-sight, clockwise from below,deg

c o , w Pitchand yaw angular velocity, pitch-positive down-ward, yaw-positive right looking downrange, rad/sec


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AS-60-0000-00338Page 97

R, A, E Radar position coordinates; slant range, true azimuth,elevation angle from vertical; guidance radar-subscript1, secondary radar-subscript 2, ft, deg,

R, P, Q Velocity in radar coordinates; range rate, lateral ratesand longitudinal rate with respect to Doppler ratestations at radar site, ft/sec

X Longitude from launcher meridian, positive eastward,deg

0 Geocentric latitude, deg0 Geodetic latitude, on (DODWGS-60) referenceO ellipsoid, degR.j Distance from launcher to predicted (Analytic)

instantaneous impact point, naut miT Time-to-go before guidance shutoff, secc o , co Steering commands, computed this cycle for trans-

mission during the next cycle, rad/secc o , Pitch rate required to cause proper radius vector

magnitude at burnout, rad/secr Radius vector magnitude, converted from radar inputs,

multiply by 20 x 10 to get ft4 , ty , Position, converted from radar inputs; similar to

x-y-z system but with 4 - plane re-positioned eachguidance computer cycle (0. 5 sec) to contain themeridian of the guidance radar at the instant ofmeasurement, multiply by 20 x 10 to get ft

V Inertial velocity magnitude, converted from radar inputs,multiply by 26, 530 to get ft/sec

rj Inertial velocity co

Nasa Cr 142379 Detailed Test Objectives for Mercury Atlas Vehicle 67 D NASA Mission MA 2 1975066485

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