astsyshansatlauvehtocsec1and2_073007085937

8/6/2019 astsyshansatlauvehtocsec1and2_073007085937

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HANDBOOKSATURN

GEOR6E C. MARSHALL SPACE FLIGHT CENTERNATIONAL AERONAUTICS AND SPACE ADMINISTRATION

HUNTSVILLE, ALABAMA .


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ASTRlONlCS SYSTEM HANDBOOK

SATURN LAUNCH VEHICLES

1 August 1965


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Astrionics SystemI

DELETED, THE AFFECTED PAGES A N D CH AN GE DATE WILLBE INCLUDED IN THIS LISTING.

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Astrionics SystemContents

i

TABLE OF CONTENTSChapter Page1 INTRODUCTION . . . . . . . . . . . . . 1-i

1.1 Purpose of Document . . . . . 1.1-11.2 Satu rn Launch Vehicles . . . 1.2-11. 3 Sat urn V/Apollo Mission

Profile. . . . . . . . . . . . 1.3-11.4 Astrionics System . . . . . . . 1.4-11. 5 Reliability Considerat ions

(To be supplied at a la terdate) . . . .. . . . . . . . 1. 5-1

2 NAVIGATION AND GUIDANCE . . . . 2-i2.1 The Navigation, Guidance,

and Control System . . . . 2.1-12.2 Navigation . . . .. . . . . . 2.2-12. 3 Guidance . . . . . . . . . . . . . 2.3-1ATTITUDE CONTROL . . . . . . . . . 3-i3.1 Attitude Control During

Powered Flight . . . . . . . 3.1-13.2 Attitude Control During

Coast Flight . . . . . . . . . 3.2-13. 3 Control Sensors . . . . . . . 3.3-13.4 Flight Control Computer . . 3.4-13. 5 Engine Servo Actu ator s . . . 3.5-1

4 MODE AND SEQUENCE CONTROL. . 4-i4.1 Introduction . . . . . . . . . 4.1-14.2 Switch Selector . . . . . . . . 4.2-14. 3 Satu rn V Operation

Sequence . . . . . . .. . . . 4.3-15 MEASURING AND TELEMETRY. . . . 5-i

Introduction . . . . . . . . . . . 5.1-1Measuring System . . . . .. . 5.2-1Remote Automatic

Calibration System. . . .. 5.3-1Telemetry. . . . .. . .. . . . 5.4-1Multiplexers . . . . . . . . . . . 5. 5-1Telemetry Calibration

Subsystem . . . . . . . . 5.6-1Digital Data Acquisition

System. . . . . . . . . . . . . 5.7-1Television System

(Saturn V) . . . . . . . . . 5.8-1

Chapter6 RADIO COMMAND SYSTEMS . . . . . .

6.1 Introduction . .. . . . . . . . .6.2 Inst rument Unit CommandSystem. . .. . . . . . . . . .

6. 3 Sec ure Range SafetyCommand System . . . . . .

6.4 Saturn Command andCommunication System(SCCS) (To be suppli edat a late r date) . . . . . .

6. 5 Range Safety CommandSystem (AN/DRW-13) .. .

7 TRACKING SYSTEMS. . . . . . . . . . .7.1 Saturn TrackingInstrumentation . . . .. . .

7. 2 C -band Radar . . . . . . . . . .7.3 Azusa/Glotrac . . . . . . . . .7.4 ODOP Tracking Syste m. . .7. 5 S-band Tracki ng (To be

supplied at a lat er date). .8 POWER SUPPLY AND DISTRIBUTION

SYSTEM. . . . . . . . . . . . . . . . . .8.1 General Discussion . . .. . .8.2 IU Power and Distribut ion

System. . . . .. . . .. . . .8. 3 Batteries . . . . . . . . . . . .8. 4 56 Volt Power Supply . . . . .8. 5 5 Volt Measurin g Voltage

Supply . . . . . .. . . . . .9 EMERGENCY DETECTION

SYSTEM. . . . . . . . . . . . . . . . . .9.1 Crew Safety System . . . . .9.2 Emergency Detection

System. . . . . . . . . . . . .9.3 EDS Operation fo rSaturn V Vehicles . . . . .

10 LAUNCH SITE SUPPORT SYSTEMS(To be supplied at a l ate r date) . . .

11 OPERATIONAL PHASES OF THEASTRIONICS SYSTEM (To besupplied at a lat er date) . . . . . . .

Page6-i

6.1-16.2-16.3-1

6.4-16. 5-17-1

7.1-17.2-17.3-17.4-17.5-1

8-1

8.1-18.2-18. 3-18.4-18.5-1

9-1

9.1-1

9.2-19.3-1

10-1

11-1


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Chapter Page Chapter Page. . . . . . . . . . . .2 INSTRUMENTUNIT 12- i1 2 . 1 Instrument Unit . . . . . . . . . 12 . 1 - 1

1 3 ENVIRONMENTAL CONTROLSYSTEM.. . . . . . . . . . . . . . . . . 13- i

1 3 . 1 Thermal ConditioningSystem. . . . . . . . . . . . . 13 . 1 - 1

1 3 . 2 Gas Bearing SupplySystem. . . . . . . . . . . . . 13 . 2 - 1

1 4 STABILIZED PLATFORM . . . . . . . . 14- i1 4 . 1 Introduction. . . . . . . . . . . . 14 . 1 - 114 . 2 S T - 124 - M Inertial Platform

Assembly . . . . . . . . . . . 1 4 . 2 - 11 4 . 3 Gyro and AccelerometerServosystem. . . . . . . . . 1 4 . 3 - 114. 4 Gimbal Angle Multispeed

Resolvers. . . . . . . . . . . 1 4 . 4 - 114. 5 Other Platf orm System

Un i t s . . . . . . . . . . . . . . 14. 5-1

1 4 . 6 Platform ErectionSystems . . . . . . . . . . . . 14 . 6 - 1

1 4 . 7 Azimuth AlignmentSystem. . . . . . . . . . . . . 1 4 . 7 - 11 4 . 8 Gas Bearing Gyro . . . . . . . 1 4 . 8 - 11 4 . 9 Pendulous GyroAccelerometer. . . . . . . . 14. 9-1

1 4 . 1 0 Power and Gas Require-ments . . . . . . . . . . . . . 14 . 10 - 1

1 5 LAUNCH, VEHICLE DATA ADAPTERAND LAUNCH VEHICLE DIGITALCOMPUTER . . . . . . . . . . . . . . . 15- i. . . . . . . . . . .5 . 1 Introduction 15 . 1 - 1. . . . . . . .5 . 2 Physical Design. 1 5 . 2 - 1. . . . . . . . . . . .5. 3 Reliability 15. 3-11 5 . 4 Descripti on of the Launch

Vehicle DigitalComputer . . . . . . . . . . . 1 5 . 4 - 1

15. 5 Descri ption of the Launch. . .ehicle Data Adapter 15. 5-1LIST OF ILLUSTRATIONS

Number Tit le Page Number Title PageSaturn IB Launch Vehicle. . . . . . . .haracteristic Data 1 . 2 - 2Satur n V Launch Vehicle. . . . . . . .haracteristic Data 1 . 2 - 3Saturn ~/ApolloMission. . . . . . . . . . . . . . . .rofile. 1. 3-3Saturn IB Astrionics System.......Operational Vehicle) 1 . 4 - 2Saturn V Astri onics System. . . . . . .Operational Vehicle) 1 . 4 - 3Block Diag ram of Saturn V

Navigation, Guidance, andControl System . . . . . . . . . . . 2 . 1 - 2

Navigation Coordinate Systems. .. 2 . 2 - 2Navigation Flow Diagram ...... 2 . 2 - 3Coordinate System Used for

Iterative Guidance Mode(IGM). . . . . . . . . . . . . . . . . . 2 . 3 - 3

Iterative Guidance ModeEquations, Flight to Orbi t .... 2. 3-5

Iterative Guidance ModeEquations, Flight out ofOrbit . . . . . . . . . . . . . . . . . .

Launch Window Parameters. . . . .Linkage Between Vehicle and

Ground. . . . . . . . . . . . . . . . .Guidance System Input. . . . . . . . . . . . . . .equenceAltitude, A cceler ation (F/M) ,

Velocity (V,), and Aero-dynamic Pr essu re (Q ) for . .Typical Saturn Trajectory.

Variations of C P and CGDuring Flight . . . . . . . . . . . .

Typical Saturn V Frequency. . . . . . . . . . . . . . .pectrumShape of the F i r s t and Second

Bending Modes (Saturn V) . . . .Control Loop Block Diagr am . . . .S-IB Engine and Actuator............onfiguration


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Astrionics SysteContenI

Number Titl e Page Number Title PagSaturn V Engines, Actuators,

and Nozzle Arrangement . . . . .Saturn Control System Block

Diagram (Powered Flight) ....Attitude Signal Flow Diagram. ...Limit Cycle Phase Diagram. . . . .Attitude Error Signal Sources . . .Pitch Channel of the Auxiliary

Propulsion System . . . . . . . . .Pitch and Yaw Deadband for

S-IVB Auxiliary Propulsion. . . . . . . . . . . . . . . .ystem.

Roll Deadband for S-IVBAuxiliary Propulsion System . .

S-IVB Attitude ControlSystem. . . . . . . . . . . . . . . . .

Deadbands of Attitude ControlSystem. . . . . . . . . . . . . . . . .Relay Control Unit andQuad Redundant Valves . . . . . .

Cutaway View of a Self-testRate Gyro . . . . . . . . . . . . . .

Control - EDS Rate GyroPackage with Covers Removed.

Control - EDS Rate GyroBlock Diagram . . . . . . . . . . .

Control Signal ProcessorBlock Diagram . . . . . . . . . . .

Control Rate Gyro BlockDiagram . . . . . . . . . . . . . . .Accelerometer Block Diagram . . .

Simplified Diagram of theFlight Control Computerfor Powered Flight. . . . . . . . .

Simplified Diagram of FlightControl Computer APSControl . . . . . . . . . . . . . . . .

Saturn IB Flight Control ComputerBlock Diagram . . . . . . . . . . .

Typical Gain Pro gram f or ao,al, g2 Coefficients. ........Block Diagram of the Servo

Amplifier in the Flight ControlComputer (For PoweredFlight) . . . . . . . . . . . . . . . . .

Saturn V Flight Control ComputerBlock Diagram . . . . . . . . . . .

Connections of Control Signals. . . . . . .o Spatial Amplifiers

Block Diagram of a SpatialAmplifier in the Flight ControlComputer . . . . . . . . . . . . . . 3.4-

Redundant Cabling in the FlightControl Computer . . . . . . . . . 3.4-Hydraulic Actuator System . . . . . 3. 5-Flow-through Valve and

Actuator with MechanicalFeedback . . . . . . . . . . . . . . . 3. 5-

Switch Selector Configuration. . . . 4.2-Switch Selector Register Word

Format . . . . . . . . . . . . . . . . 4.2-1LVDC - Switch Selec tor Int er-

connection Diagram . . . . . . . . 4.2-Switch Selec tor (Mod 11). . . . . . . .implified Diagram 4.2-7Switch Selector Timing Diagram. . 4.2-9Automatic Reset Circuitry. . . . . . 4.2-1Typical Saturn V Astrionics

System Prelaunch Sequence . . . 4.3-3Measuring and Telemetry

. . . . . . . . . . . . . . . .ystem. 5.1-Typical Saturn Measuring

System. . . . . . . . . . . . . . . . . 5.2-1. . . . . . . . .ypical Tran sduc ers 5.2 -

Force-balance AccelerometerBlock Diagram . . . . . . . . . . . 5.2 -Bourdon- tube Potentiome ter

. . . .ype Pr es su re Transducer 5.2-6Strain-gage Type Pressure

Transducer. . . . . . . . . . . . . . 5.2-6Liquid Level, Discr ete,

. . . . . . . .unctional Diagram 5.2-7Liquid Level Sensor Electrical

. . . . . . . . . . . . . .chematic 5.2-7Basic Pri ncip les of a

Flowmeter . . . . . . . . . . . . . . 5.2-7. . . . . . . .ypical Bridge Circuit 5.2-9

Piezoelectric Accelerometer. . . . . . .nd Emitt er Follower 5.2-1Strain-gage Accelerometer Block

Diagram.. . . . . . . . . . . . . . . 5.2-1. . . . . . .ypical Measuring Rack 5.2-1. .llus tration of an AC Amplifie r. 5.2-1

Typical Signal Conditioning Cardfor Temperature Measure-ments . . . . . . . . . . . . . . . . . 5.2-1

i


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Astrionics SystemContentsLIST O F ILLUSTRATIONS (CONT D)

Number5.3-15.4-15.4-25.4-3

Title Page Number Title PageBlock Diagram of RACS fo r theIU.. . . . . . . . . . . . . . . . . . 5.3-1Typical S-IvB/IU Tel eme trySystem (R&D) Version ...... 5.4-5Typical S-IVB/IU TelemetrySystem (Operational Version). . 5.4-7Block Diagram of TypicalSaturn V FM/FMTelemetry. . . . . . . . . . . . . . . .ystem. 5.4-9Block Diagram of Airborne SSTelem etry Assembly. . . . . . . . 5.4-11General Block Diagram - ......CM/DDAS Assembly 5.4-13Analog-to-Digital ConverterBlock Diagram . . . . . . . . . . . 5.4-14Digital Multiplexing andFormating Logic .......... 5.4-1 5Clock Programming and TimingLogic Block Diagra m . . . . . . . 5.4-16PCM/RF Assembly BlockDiagram. . . . . . . . . . . . . . . . 5.4-18Mod 270 MultiplexerBlock Diagram . . . . . . . . . . . 5.5-2Mod 270 Multiplexer Assembly. . . . . . . . . . . . . .aveforms 5. 5-3Block Diagram of RDSM Inputsandoutputs. . . . . . . . . . . . . . 5.5-4

Remote Digital SubmultiplexerSimplified Block Diagram . . . . 5. 5-5Inputs and Outputs of Mod 41 0Multiplexer. . . . . . . . . . . . . . 5. 5-6Block Diagram of Mod 410Multiplexer. . . . . . . . . . . . . . 5. 5-7Block Dia gram of Mod 245Multiplexer . . . . . . . . . . . . . 5. 5-8Typical Telemetr y CalibrationSubsystem.. . . . . . . . . . . . . . 5.6-3Telemetry Calibrator Assembly . . 5. 6-7Computer Interface Unit - SystemInterconnection Diagram . . . . 5.7-2Computer Interface UnitBlock Diagram. . . . . . . . . . . . 5.7-3S-IC Television FunctionBlock Diagram . . . . . . . . . . . 5.8-1Television Ground ReceivingStation. . . . . . . . . . . . . . . . . 5.8-2S-IC Tentative Te levisionOptics Layout . . . . . . . . . . . . 5.8-3Saturn V IU CommandSystem. . . . . . . . . . . . . . . . . 6.2-1

Saturn IB IU Command System . . .Phase Shift Keyed Signals . . . . . .Digital Form at Showing AddressDistribution . . . . . . . . . . . . .Digital Format Showing Informa-tion Bit Groups . . . . . . . . . . .Mode Command Word Fo rmatand Coding . . . . . . . . . . . . . .Data Command Word GroupFormat and Coding for anUpdate Command . . . . . . . . . .MCR-503 Command Receiver. . . .IU Command Decoder FunctionalBlock Diagram . . . . . . . . . . .PSK Sub-bit DetectorSynchrogram . . . . . . . . . . . . .

PSK Sub-bit DetectorBlock Diagram . . . . . . . . . . .IU Command Decoder SimplifiedLogic Diagram . . . . . . . . . . .IU Command Decoder FlowDiagram. . . . . . . . . . . . . . . .Example of Wiring Between ShiftRegister and Sub-bitComparators. . . . . . . . . . . . .MAP Circuitry Block Diagram . . .Secure Range Safety Command. . . . . . . . . . . . . . . .ystem.Range Safety Ground System . . . .Downrange Station . . . . . . . . . . .. . . . . . . . . . .ode Plug Wiring.Secure Range Safety DecoderSimplified Logic Diagram ....Range Safety Command. . . . . . . . . . . . . . . .ystem.AN/FwR-2A Modulation BandwidthUsage .................AN/DRW-1 3 Receiver/Decoder . .Launch Phase TrackingStations . . . . . . . . . . . . . . . .Station Visibility fo r Saturn VPowered Flight ...........Accuracy of Position and . . . . .elocity Measurements.SST-135C Transponder System...........lock Diagram.AZUSA (MK 11) Ground Station. . . . . . . . . . . . . . . . .ayout


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Number Titl e Page Number Titl e PageAFC Loops. Ground Station..........nd TransponderAZUSA Transponder Block

Diagram . . . . . . . . . . . . . . . .. . . . .DOP System ConfigurationODOP Transponder BlockDiagram . . . . . . . . . . . . . . . .

Saturn V Instrument Unit . . . . .quipment Layout. R&DSaturn IB Instrument Unit

Equipment Layout.. . . . . . . . . . . . . .perationalSaturn V Instrument UnitEquipment Layout...............perational

Saturn IB Instrument UnitEquipment Layout. Antenna. . . . . . . . . . . . . .rientation

Saturn V Instrument UnitEquipment Layout. AntennaOrientation . . . . . . . . . . . . . .

Block Diagram of the Saturn VPower Supply and Distribution................ystems

Part ia l Schematic of the Powerand Distribution Systems . . . . .

Battery Load Profiles for...............A.IU.201Battery Load Profiles forSA-IU-205 and Subsequent. . . . . . . . . . . . . . . .ehiclesPart ia l Schematic of the IU

Power Transf er Switch ......Ground Checkout Configuration

of IU Power Derivation . . . . . .Inert ial System Power Flow . . . . .Instrumentation System Power

Distribution . . . . . . . . . . . . .IU Distribution Equipment. . . . . . . . . . . . . . . . .ayoutIU Grounding System fo r Saturn IB

and V Vehicles . . . . . . . . . . . .56 Volt Power SupplyBlock Diagram . . . . . . . . . . .5 Volt Measuring Voltage

Supply Block Diagram . . . . . . .

Thermal Conditioning Panel

. . . . . . . . . . . . . . . . . .etailsEnvironmental Control System. . . . . . .echanical DiagramSublimator Details . . . . . . . . . . .ST-124-M Inertial Platform . . . . .ystem (Saturn IB and V)Inertial Platform Subsystem

Block Diagram . . . . . . . . . . .Platform System Signal. . . . . . . . . . . . . . .nterfaceST-124-M Gimbal Configuration . .

. . . . . . .rientation of Gyro AxesOrientation of AccelerometerAxes . . . . . . . . . . . . . . . . . .Platform Gimbal Pivot Scheme . .. . . . . . . . . . . .lip-ring Capsule . . .latform Gimbal ArrangementST-124-M Inertial Platform. . . . . . . . . . . . . . .ssemblyGimbal Design . . . . . . . . . . . . . .Gyro Servoloops Block

Diagram . . . . . . . . . . . . . . . .. . . . . .ccelerometer Servoloops . . . . .imbal Servoloop HardwareGimbal Electronics BlockDiagram . . . . . . . . . . . . . . . .Accelerometer Electronics

Block Diagram . . . . . . . . . . .. . . . .T-124-M Gyro OrientationThree-axis Inertial Platform . . . .ssembly Block Diagram. . . . .hree-gimbal ConfigurationTwo-speed Resolver Schematic . .

Saturn IB IU Astrionics System. . . . . . . . . . . .ailure Modes ...rew Safety System (Saturn IB)...rew Safety System (Saturn V)Saturn IB Crit ica l Angle of . . . .ttack Versus Flight TimeSaturn IB Crit ica l Angle ofAttack Ver sus Gimbal Angle

(76 Seconds) . . . . . . . . . . . . .Saturn IB and V Instrument Unit.........hysical LocationSaturn IB Inst rument Unit . . . . .quipment Layout. R&D


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LIST O F ILLUSTRATIONS (CONT'D)Number Title page Number

Gimbal Angle Vector Diagrams . . 14.4-2Resolver Chain Schematic . . .. . . 14.4-4Resolver Chain Signal Detection

Scheme ... . . . . . . . .. . 14.4-5Platform AC Power SupplyAssembly. . . .. . . . . . . . . . 14. 5-2Generation of 3- Phase, 400-Hertz

Voltage. .. . ... . . .. .. . 14.5-3Gas Bearing Pendulum . . . .. . .. 14.6-1Platform Erection System Block

Diagram. . . .. . . . . . . . . . 14.6-2Automatic Azimuth Alignment . . . 14. 7-2Theodolite .. . . .. . . . .. .. .. . 14.7-3Optical Schematic Diagram of

SV-M2 Theodolite . . . . . . . . . 14.7-4Optical Spectrum . . . . . . . . . . . . 14.7-5Optical Gimbal Laying System . . . 14.7-6Azimuth Alignment Scheme . .. . . 14.7-7Cutaway View of a Single-axis

Integrating Gyro . . . . . . . .. . 14.8-3Exploded View of a Signal Genera tor

and Torque Generator . . .. . . 14.8-4Ele ctr ica l Schematic of an

AB5-K8 Stabil izing Gyro-scope . . . . . . . . . . . . . . . . . 14.8-5

Cutaway View of a PendulousIntegrating Gyro Accelerom-eter . .. .. . . .. . . . .. ... . 14.9-3

Pendulous Integrating GyroAccelerometer Schematic . . . . 14.9-4

Ele ctr ica l Schematic of anAB3-K8 Accelerometer. . . ... 14.9-5

System Power Requirements andHeat Dissipation . . . . . . . . . . 14.10-1

Title PageConnections Between Digital

Computer, Data Adapter, andthe Astrionics System. . .. .. . 15.1-2

Unit Logic Device Buildup . .. . . . 15.2-1Unit Logic Device PageAssembly . . . . . . . . . . . . . . . 15.2-2

Exploded View of the LaunchVehicle Data Adapter . . . . . . 15.2-3

Exploded View of the Launch VehicleDigital Computer . . . . . . . . . . 15.2-3

Computer Redundancy Con-figuration . . . . . . . . . . . . . . . 15. 3-1

Triple Modular Redundancy(TMR) . . . . . . .. . . . . . . . . 15.3-2

Computer Functional BlockDiagram. . . . . . . . . . . . . . . 15.4-7

Word Organization .. . . . . . . . . . 15.4-10Computer Timing Organization. .. 15.4-11Clock Generator Block Diagram . . 15.4-12Memory Module Arrangement . . 15.4-1 3Ferrite Core Characteristics. . . . 15.4-1 3Core Plane . . . . . . . . . . . . .. . . 15.4-14Memory Module Block Diagram . . 15.4-1 5Arithmetic Element Block

Diagram. . . . . . . . . . . .. . . . 15.4-2 5LVDA - LVDC Interconnection

Block Diagram . . . . . . . .. 15. 5-3LVDA - IU Equipment Inter -

connection Block Diagram . . . . 15. 5- 5Launch Vehicle Data Adapter

Block Diagram . . . . . . . . . . . 15. 5-13Redundant Power Supply Block

Diagram . . . . . . . . . . . . . 15.5-16


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Astrionics SystemI Content

TABLESNumber1 . 4 - 1

Title PageDifferences Between Saturn IB

and Saturn V AstrionicsSystems (To be supplied. . . . . . . . . . . .t a l a t e r d a t e ) 1 . 4 . 4

Navigation Equations ......... 2 . 2 - 4Control Accelerometer. . . . . . . . . . .haracteristics 3.3.7Input Signals to the Flight Control

Computer ............... 3.4.2Servo Actuator Ch aracteris tics

(Design Goals) . . . . . . . . . . . . 3. -4Phase. Sequence. and Time

Breakdown .............. 4.3.1Saturn V Flight Sequence. Phase 1- .. . . . . . . . . . . . .relaunch 4 3-2Saturn V Flight Sequence. Phase 2

through 5 . owered Flights andOrbital Coast ............ 4 . 3 . 5

Saturn V Flight Sequence. Phase 6and 7 . -IVB Restart andDocking ................ 4.3.13

Number of Measurements. Tr an s.ducers. and Measuring Racks. . 5 . 2 - 2

Typical Saturn V Measurements . . 5 . 2 - 3Typical Sa turn V Operational

Measurements . . . . . . . . . . . . . 5.2.4Data Categories . . . . . . . . . . . . . 5.4.1Telemetr y Systems in the Various

Saturn Vehicle Stages ....... 5.4.4IRIG Subcarrier Channels. . . . . . . . .* 7 5% Channels) 5.4 .9Mod 270 Multiplexer Assembly

Performance Characteristics . . 5. -3Saturn V Launch Vehicle Television

Characteristics . . . . . . . . . . . 5.8.2Number of Words Transmi tted for. . . . . . . .ifferent Commands 6 . 2 . 5Characteristics of the MCR-503

Receiver ............... 6.2 .8Charact eris tics of the IU. . . . . . . . . .ommand System 6.2.12Coding Scheme for Function. . . . . . . . . . . . . .haracters 6.3.2Charact eris tics of the Command

Decoder . . . . . . . . . . . . . . . . 6 . 3 . 7

Number Title6 . 5 - 1 Range Safety Command Receiver

Decoder (AN/DRW-13). . . . . . . . . . .haracteristics

7 . 1 - 1 Saturn Tracking Instrumentation . .. . . . . .. 1-2 Orbital Tracking Stations7 . 2 - 1 C-band Radar Transponder.

Model SST.135C . . . . . . . . . . .7 . 2 - 2 Radar Ground Station. . . . . . . . . . .haracteristics7.3.1 AZUSA Char acte rist ics . . . . . . . .7.4.1 ODOP System Characteristics . . .

. . . . . .. 3 - 1 IU Battery Characte ristics8.4-1 56 Volt Power Supply Electrical

Characteristics . . . . . . . . . . .8 . 5 - 1 5 Volt Measuring Voltage SupplyCharacteristics . . . . . . . . . . .

9 . 1 - 1 Important Guidelines fo r theCrew Safety System for SaturnApollo Vehicles . . . . . . . . . . .

9. -1 Abort Criteria and Ground. . . . . . . . . . . . . . . . . .ules9 . 3 - 2 Saturn IB EDS Design Ground

Rules . . . . . . . . . . . . . . . . .Slip-ring Cartridge

Characteristics . . . . . . . . . . .Angular Readout Characteristics. .Resolver Chain System

Characteristics . . . . . . . . . . .Charact eris tics of the Accelerom-

et er TM Velocity Signals . . . . .Gas Bearing Pendulum

Characteristics . . . . . . . . . . .Gyro Characteristics . . . . . . . . .Accelerometer Characteristics....Power Supply Specifications . . . .

15 . 3 - 1 TMR Computer ModuleBreakdown . . . . . . . . . . . . . .1 5 .4-1 Launch Vehicle TMR ComputerCharacteristics . . . . . . . . . . .. . . . . . . . . .5 . 4 - 2 List of Instruc tions. . . . . . . . . . . . .5 . 4 - 3 PI0 Addresses . . . .5 5-1 Data Adapter Chara cteristic s. . . . . . . . .5. -2 Signal Characte ristics

1 5. 5-3 Use of Word Locations in theDelay Line . . . . . . . . . . . . . .

Page


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Astrionics Systemt

PART IFUNCTIONAL DESCRIPTION


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A s t r io n ic s S y s t e m1

CHAPTER 1INTRODUCTION

TABLE OF CONTENTS

S e c t io n P a g e1.1 PURPOSE OF DOCUMENT . . . . . . . . . . . . . .. . . . . . . . . . . . 1.1-11.2 SATURN LAUNCH VEHICLES . . .. . . . .. . . . . . . . . . . . . . . 1.2-11.3 SATURN V/APOLLO MISSION PROFILE .. . ... . . . . .. .. . 1.3-11.4 ASTFUONICS SYSTEM .. . . . .. . .. . . . . . . . . . . . . .. . .. . 1.4-11.5 RELIABILITY CONSIDERATIONS (To be s upp lied

at a la ter date) . . .. . . . . .. . . . . .. . . .. . . . . . . .. . 1.5-1


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Astrionics System, Section 1.1

SECTION 1.1PURPOSE OF DOCUMENT

The intent of t his document i s to provide asyst em description of the Astrionics System for theSaturn IB and V Launch Vehicles. It i s not intendedto go too deeply into any given subject but rat her togive an overall p icture of the Ast rionics System fromfunctional and operational viewpoints. General missionrequirements and system capabilities are briefly dis-cussed in Chapter 1, Introduction, to provide a totalview of the Ast rio nics System. The subsequent chap-te rs of P ar t I presen t a functional description of thevar iou s sub syst ems and the involved hardware. Theoperational phases of the Astrionics System, includingpre-launch checkout, a r e discus sed in Chapter 11 andindicate how the system will be used during a typicalSaturn V Apollo mission. Astri onics hardwar e whichperfor ms seve ral different functions (e. g. , LaunchVehicle Digital Computer and Data Adapter) is describ-ed in Part 11.

/The Astrionics System description includes allof the ele ctr ica l and electronic equipment on board the

vehicle. It also includes the launch si te electronicsupport equipment. However, this par tic ula r desc ription does not cover the individual stage relay circuitrwhich controls cer tain stage functions. It does de-scribe the signal flow through the system to the pointof energizing this special circ uitry so that an overallunderstanding of syst em opera tion is presented. Likwise, stage propellant utilization s yste ms and internaengine sequencing sys te ms ar e not covered; since, fothe purpose of this description, they a r e considered apar t of the propulsion system.

Since Astrionics Syste ms of Saturn IB andSaturn V ar e very similar, this document is devotedprimarily to Saturn V. The are as in which the SaturnIB Astrionics System deviates f ro m the Saturn VAstrionics System a r e liste d in Section 1.4. Whereapplicable, these deviations ar e specified in the textof that par tic ula r chapter.


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Astrionics SystemSection 1.2

SECTION 1.2

SATURN LAUNCH VEHICLES

Figures 1.2-1 and 1.2-2 illustrate the SaturnIBVehicle and Saturn V Vehicle, respectively, and includesome character isti c vehicle data. Both vehicles havethe s am e upper stage (the S-IVB Stage) which is pro-pelled by a restartable engine to provide injection intoescape trajectory from a parking orbit. The Inst ru-ment Unit is mounted on top of the S-IVB Stage and isver y s imil ar for Saturn IB and Saturn V.

The prim ary mission of Saturn IB is to serveas a launch vehicle for the Apollo Spacecraft earthorbit al flight tests. These earth orbital flights willsimu late cert ain phases of the lunar landing missionand will provide flight tes ts f or the spacecraft and theS-IVB/IU Stage. Saturn V is the launch vehicle forthe actual Apollo lunar landing missions. The typicalprofile of a lunar landing mission is described inSection 1.3.

The pr imary mission of the Saturn Vehicles isthe successful accomplishment of the Apollo mission.

In addition, the Saturn Vehicles ar e capable of pe r-forming other types of missions which can be generallclass ified a s the insertion of heavy payloads into ear thorbi ts and escape trajectori es. This may include:

Transfer between earth orbitsRendezvous in earth orbitDirect ascent and injection into escapetrajectoryInjection into escape trajectory followingextended earth-orbit phasesExtended missions beyond injection

Saturn Vehicles a r e numbered consecutively,beginning with 201 fo r the fi rs t Saturn IB Flight Vehicand 501 for the fir st Saturn V Flight Vehicle. The fi rsfew vehicles of each ser ie s ar e considered as researcand development vehicles.


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strionics SystemSection 1 . 2


17/42

Astrionics System, Section 1 . 2


18/42


/

SECTION 1.3SATURN V/APOLLO MIS SIO N PROFILE

The mission of the Apollo Project is to land2 Astronauts on the moon and retu rn the total crew of3 Astronauts safely to earth.

Th e overa ll Apollo Space Vehicle, composedof the Sa turn V Launch Vehicle and the Apollo Space-craft, i s shown in Figure 1.2-2. The Sa turn V con-si st s of t hr ee propulsion stage s and the InstrumentUnit. The IU contains the navigation, guidance andcontrol, communication, and power supply equipmentcommon to the main propulsion stages.

The following i s a brief description of theSaturn V Launch Vehicle mission only. The profileof the mission is illustrated in Figure 1. 3-1. TheS-IC Stage boosts the vehicle through the atmosphericfligh t phase. Cutoff of the engines is initiated closeto fuel depletion and occurs at an alti tude of approx-imate ly 62 kilome ter s (34 nautical miles ). Afterseparat ion from the fir st stage, second stage (S-11)boost follows immediately; engine cutoff is executeda s in the fi rs t stage. Both stages drop to earth in aballi stic flight path. After sep ara tio n of the S-I1 Stage,the S-IVB Stage engine i s ignited. The engine i s cutoff

when the vehicle has achieved the necessar y orbit alvelocity. The vehicle, consis ting now of the S-IvB/IUStage and the Apollo Spacecraft, orb its the e art h at analtitude of approximately 200 kilomete rs (108 nauticalmiles ) for a maximum of 3 orbit s. An orb ita l launchwindow exi sts once in each orbit . When the selecte dorb ita l launch window occurs, the S-IVB engine isignited a second time to provide the thrust for injectiointo the translunar trajectory. The engine is cutoffwhen the required esca pe velocity is achieved. In thecoas t perio d following injection, the transposi tionmaneuver is performed. In this maneuver, the Serv-ic e and Command Modules of the spa cecraf t move awafrom the Saturn S-IVB/IU Stage, tu rn around, and dockwith the Lunar Excursion Module still attached to theS-IVB/IU Stage to achiev e the. pro per spac ecraft con-figuration for the lunar landing operation. TheS--IVB/IU Stage is then separat ed fro m the ApolloSpacecraft. The launch vehicle mission ends withthe separation fro m the spacecraft a t approximatelyone hour after injection (maximum time of 2 hoursaft er injection). The spacecra ft continues its coastflight toward the moon.


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Astrionics SystemSection 1. 3

f185 KILOMETERS (100 NAUT. MILES)

S- l STAGE CUTOFF 8 INSERTION INTOJETTISON S-11 JETTISON- -IV B PARKING ORBIT 8AFT INTERSTAGE STAGE IGNITION S-IVB ENGINE CUTOFF COAST S-IV B RE-IGNITION

S-IVB ENGINES-ll IGNITION 8LAUNCH ESCAPESYSTEM JETTISON

BEGINNING OF

S-IC STAGE CUTOFF8 JETTISON

S-IV B SEPARATIONS-IC STAGE IGNITION8 VEHICLE LAUNCH

S A T U R N V / A P O L L O

Figure 1.3-1 Saturn V/Apollo Mission Profil1.3-3/1.3-4


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SECTION 1.4

ASTRlONlCS SYSTEMThe overall Astrion ics Sys tems of the SaturnIB

and V Launch Vehicles a r e shown in the simplifiedblock diagrams, Figu res 1.4-1 and 1.4-2, respectively.The ma jor portion of th e Astrio nics equipment is locat-ed in the IU, which is mounted on top of the S-IVBStage. During flight, the Astrionics System perf orms ,o r i s involved in, the following main functions:

Navigation, guidance, and control of thevehicleMeasurem ent of vehicle par ame ter sData transmission between vehicle andground stations (up and down)Trac king of the launch vehicleCheckout and monitoring of vehicle func-tions in orbitDetection of emergency situationsGeneration of el ectr ical power for s yste moperation

The operat iona l lifet ime of the S-IVB/IU Astr ioni csSystem is 4-1/2 hour s for Saturn IB and 7 hours forSaturn V. The operational lifetime is limited only bythe cap acity of the power supply (ba tteri es) and thewate r supply of the environmental control sy ste mwhich is sufficient to complete the presentl y definedlaunch vehicle missions. With incre ased power andwate r supply capacity, the operational lifetime of theAstrionics System can be extended for longer durationmissions if required.

NAVIGATION, GUIDANCE, AND CONTROL

and finally the propulsion engine actuato rs and theauxiliary propulsion system. The Saturn inertialnavigation and guidance system can be updated by dattransmission from ground station s through the IUcommand system. The Inertial Platform Assembly'car rie s three integrating accelerometers which measur e the thrust acceleration in a space-fixed referencefram e. In addition, the platform gimbal angles indi-cate the attitude of the vehicle in the platform re fe r-ence frame. The LVDA se rv es a s the input/outputdevice for the LVDC and als o per for ms the nec ess arydata processing.

The LVDC per for ms computations fo r navigation, guidance, and cont rol functions. The positionand velocity of the vehicle is obtained by combiningacceleromete r measureme nts with computed gravi-tational acceleration . This information is the inputto the guidance computations which d eter mine therequired thr ust vector orientation and engine cutofftime according to the guidance scheme sto red in thememory of the LVDC.

Attitude contr ol during powered flight is ac-complished through swivelling of propulsion enginesby means of hydraul ic actua tors t o obtain the prope rthrust vector orientation. The actuator commandsa re generated i n the Flight Control Computer. TheFlight Control Computer combines attitude e rr orsignals from the LVDA and angular ra te signals fro mRate Gyros to provide stabl e attitude control of th evehicle. The attitude e r r o r signal is generated inthe LVDC by comparing the required thrust vectororientation (from guidance computations) with th eactual vehicle attitude (obtained fro m platform gimbangles).

The Saturn Astrio nics System provides naviga- During coast flight periods, attitude controltion, guidance, and cont rol of the vehicle from launch achieved by the auxiliary propulsion syst em. Thisuntil sep ara tio n of the S-1VB/IU f rom the spacecr aft . system consists of 6 nozzles which ar e arran ged inThe equipment involved in the se functions a r e the 2 modules and mounted on the aft end of the S-IVBST-124-M Inert ial Plat form Assembly, the Launch Stage. T he auxiliary propulsion sys tem is als o conVehicle Digital Computer and Launch Vehicle Data troll ed by the Flight Control Compute r located in thAdapter, the Flight Control Compute r, the Rate Gyros, IU.


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Figure 1.4-1 Saturn IB Astrionics System ( Operational Vehicle )

Abort Decision Alter nate Steering Mode Command StatusA Commands A SPACECRAFT- - - - -= 1-1- - - - -1 - - -1 -11- -11--11--11 ---------------.

-) EDSDistributor-

Control -EDS Launch Vehicl e Stabil ized C-band RadarRate Gyros Digi tal Computer platform ~ p '

4 - - -ccelerometers Control , IU Command IU Command

Computer I Decoder Receiver4 -I

Control

-EDS-.---

EDS '

RACS I DDASPower (Supply)& Distribution

- --I----------------------------Switch Selector-

-witch Selector- -Measuring VHF TelemetryS-IV B STAGE

-----------------11-----0-------- -

I 8 and Telemetry Transmitter

RACS I DDASMeasuring

and Telemetry

I.--------I----.------------

IF t " - { T jAccelerometers

& Distribution Fl-yS-IB STAGE

ISM Bl35

*+ SE

4 I u

t C )

VHFTelemetry

Transmi ter

-1---1 I-- --I--------------Aux Propulsion

System

Actuators

*Switch Selector

RACS I DDASMeasuring

and Telemetry-& Distribution LEIower (Supply) Propulsion

.)+ SEVHF

TelemetryTransmitter -


22/42

Abort Decision Alter nate Steering Mode Command StatusCommand: I 1 t SPACECRAFT,----.t-------------- --------------.------------------------------------.

EDS Con trol -EDS-) Distribu tor Rate Gyros

1control f

Computer.

Launch Vehicle + Stab ilize d C-band RadarDig ital Computer PlatformIIU CommandLaunch Vehicle Decoderr----'I?

IPower (Supply) + .-Switch Selector& DistributionFnS II - RACS I DDAS I- I I SCCS S-Band IMeasuring d SEand Telemetry I UHF and VHF TelemetryTransmittersAux Propulsion Switch Selector

Power (Supply)~~sb-1 Propulsion4- & Distribution

EDS4----

EDSC-

Switch Selector

L H& Distribution--------1111-----11-1------.

Switch SelectorP& Distribution

L - ) Telemetry 1Measuring Transmitterand Telemetry d SEI S-IVB STAGE-11 1111 1 1

RACS I DDAS- VHFTelemetryMeasuring Transmitterand Telemetry + SE

I IS -l l STAGE

-----11---1------11------1--------1----.

-RACS I DDAS VHF- ) TelemetryMeasuring Transmitterand Telemetry,SE p T - ] - Y S- IC STAGE

IBM B 1 3 1 IFigure 1.4-2 Saturn V Astrionics System ( Operational Vehicle )


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Flight sequence control (e. g., vehicle staging,engine ignition and cutoff) is performed by the LVDC.The flight program, sto red in the LVDC memory,generates the ne cess ary flight sequence commandswhich a r e transm itt ed through the LVDA and SwitchSelector to the proper circuit in the particular vehiclestage.MEASUREMENTS AND DATA TRANSMISSION

Each vehicle stage is equipped with a completemeasuring and tele metry system, including RF trans-mit ter and antennas. Fo r efficient utilization of avail-able bandwidth and to obtain the r equir ed accur acy,three different modulation techniques are used in eachstage telemetry system. These three are: frequencymodulation/frequency modulation, pulse code modula-tion/frequency modulation, and single sideband/fre -quency modulation (employed in re se ar ch and develop-ment only).

In Saturn IB vehicles, telemetry data is radiat-ed fr om the vehicle t o ground stations i n the VHF band(225-260 MHz). The PCM/FM sys te m of the S-IVBStage and the IU a r e interconnected to provide a redun-dant transmission path and to make S-IVB measure-ment s availa ble to the LVDA. All flight control datais transmitted through the PCM/FM system.

In Saturn V Vehicles, the PCM/FM te leme trydat a of the S-IVB and IU i s transm itt ed in VHF band(225-260 MHz) and in the UHF band (2200-2300 MHz).The UHF-band transmission is provided primaril y fortrans miss ion over the longer ranges a fter the vehiclehas left the parking orbit. In addition, the PCM/FMdata can be tra nsmi tted through the communicationand command syste m transponder. This arrangementprovides high rel iabil ity through redundancy in tra ns-miss ion path.

The telemetry system of each stage has asep ara te output via coaxial cable to the electronicsup por t equipment, which is used with the digitaldata acquisit ion syst em for vehicle checkout beforelaunch.

The Instrument Unit command system permitsdata tra nsm iss ion from ground stations to the IU forinsertion into the LVDC.TRACKING

The Saturn Vehicles ca rr y se veral trackingtransponders. The ODOP Transponder is located inthe fi rs t sta ge of S aturn IB and V Launch Vehicles.

The Instrument Unit is equipped with two C-bandRadar Transponders, an AZUSA Tran sponde r, andthe CCS Transponder (S-band tracking).EMERGENCY DETECTION SYSTEM

The emergency detection sy stem collectsspecial m easurements fro m each stage of the launchvehicle. Based on these measureme nts, cri tic alsta tes of the vehicle which may requ ire mission abor ta r e detected, and the information i s sent to the spa ce-cra ft for display and/or initiation of automatic abort.SPACECRAFT INTERFACE

Several lines cr oss the IU/spacecraft interfacefor exchange of signals. Alternate steer ing commandsfrom the s pace craf t navigation and guidance syst emmay be used to control the launch vehicle during S-I1and S-IVB powered flight phase s. This type of op era -tion is cons idere d as backup in ca se of a f ailur e of theIU navigation and guidance sys tem. During coastflight, the Astronaut may control the attitude of thevehicle through manually generated commands. Inany case, a mode command must be sent fir st fr omthe spac ecra ft to the LVDA to perform t he neces saryswitching before the IU Flight Control Computer canaccept the steering signals from the spacecraft. Toindicate the sta te of the launch vehicle, cert ain measurements a re sent to the spacecraft and displayed tothe Astronaut.

Before launch, automatic checkout of the ve-hicle system is controlled by the launch computercomplex and the electronic support equipment. Thissyst em als o includes the digital data acquisition sys -tem.

Table 1.4- 1 indicates the differences in theSaturn IB and Saturn V Astrionic s Systems.

Table 1.4-1 Differences Between Saturn IBand Saturn V Astrionics Systems


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Astr ion ics SystemI Section 1. 5

SECTION 1.5RELIABILITY CONSIDERATIONS

( To be supplied at a later date )


25/42

A s t r i o n i c s S y s t e mI

CHAPTER 2NAVIGATION AND GUIDANCE

TABLE OF CONTENTS

S e c t i o n P a g e2.1 TH E NAVIGATION, G UIDANCE, AND

CONTROL SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1-12.2 NAVIGATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2- 12.3 GUIDANCE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.3-1

2.3.1 G e n e r a l C o n s i d e r a t i o ns . . . . . . . . . . . . . . . . . . . . 2. 3-12.3.2 I t e r a t i v e G u i d a n c e M o de . . . . . . . . . . . . . . . . . . . . 2.3-2


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SECTION 2.1THE NAVIGATION, GUIDANCE, AND

CONTROL SYSTEMThe problem of directing a balli stic miss ile

or space vehicle, to accomplish a given mission, iscustomarily discussed in te rm s of three s eparatefunct ions: navigation, guidance, and control. Theboundaries between t hese 3 area s a re to some extentar bi tr ar y and conventional. The 3 terms, navigation,guidance, and control, will be used in this text accord-ing to the following definitio ns:

Navigation is the deter minat ion of positionand veloci ty of the vehicle fr om m easu re-ments made onboard the vehicle.Guidance i s the computation of maneuv ersnece ssar y to achieve the desire d end con-ditio ns of a tra je ctory (e. g. , an insertioninto orbit).Control is the execution of th e maneuver(determined from the guidance scheme)by controlling the proper hardware.

A block dia gra m of the ove rall Saturn V navi-gation, guidance, and contro l sys tem is shown inFigure 2. 1-1. (This figure is also true for theSaturn IB Vehic le if th e S - I1 Stage Switch Selectorand engine actuator blocks ar e omitted.) The3-gimbal stabil ized platform (ST124-M) provid es aspace-fixed coordinate referen ce fra me for attitudecontrol and for navigation (acceleration) meas ure-ments. Thr ee integrating acceler omete rs, mountedon the gyro-sta biliz ed inner gimbal of th e platform,measure the 3 components of velocity resulti ng fro mvehicle propulsion. The acceler omete r measur e-ments a r e sen t through the LVDA to th e LVDC. Inthe computer, the accelerometer measurements arecombined with the computed gravitational accelera-tion to obtain velocity and position of the vehicle.

The LVDA is the input/output device fo r theLVDC. It per for ms the nece ssar y process ing ofsignals, from different sources, to make these sig-nals acceptable to the computer.

According to the guidance sch eme (programmedinto the computer), the maneuver required to achievethe desired end conditions is determined by the LVDCThe instantaneous position and velocity of the vehiclear e used as inputs. The result is the required thrustdirect ion (guidance command) and the time of enginecutoff.

Guidance information st ored in the LVDC(e. g., position, velocity) can be updated through theIU command syst em by data transmissi on fr omground stations. The IU command system providesthe general capability of changing or inser tin g infor-mation in the LVDC.

Control of t he launch vehicle can be dividedinto attitude control and di scre te control functions.For attitude control, the instantaneous attitud e of thevehicle i s compared with the desi red vehicle attitude(computed according to the guidance scheme) . Thi scomparison i s perform ed in the LVDC. Attitude correction signals ar e derived from the difference be-tween the existing attitude angles (gimbal angles) andthe desir ed attitude angles. In the Control Computerthes e attitude correct ion signals a r e combined withsignals fr om control sen sors to generate the controlcommand fo r the engine actuators. The requir edthrust direction i s obtained by swivelling the enginesin the propelling stage and thus changing the thrustdirect ion of the vehicle. Since the S-IVB Stage ha sonly 1 engine, an auxiliary propulsion sys te m isused fo r roll control during powered flight. Theauxiliary propulsion s ystem provid es complete atti-tude cont rol during coast flight of t he S-IV B/IuStage.


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Commands for flight sequence control are Attitude and sequence cont rol of t he launchgenerated in the LVDC according to a s tor ed program. vehicle is described in Chapters 3 and 4, respectively.These commands a r e transf erre d through the LVDA The stabil ized plat form, LVDA and LVDC, which ar eto the Switch Selector of th e corresponding veh icle involved in navigation, guidance, and cont rol opera-stage. Examples of flight sequence con trol a r e engine tions, a r e desc ribed in Chapters 14 and 15.ignition, cutoff, and sta ge separation. The SwitchSelector in the addre ssed sta ge activates the neces-sary circuit to perform the commanded function.

IU CommandReceiverDecoder

LVDC Attitu de Control Signal S-IC Stager------- 1 from Spacecraft Engine ActuatorsI STABILIZED 1I PLATFORMI ! AttitudeI _ I Angles ToS - l l Stage

Y -IVB Auxiliary To ~~~~l~~Propulsion SystemI-

II

S-IC StageSwitch Selector

S -1 1 StageSwitch Selector

Flight SequenceEl-

I - IL- - - - - -JControl . S -IVB Stage

Engine ActuatorsSensors

lntegrating4 I Engine Actua tors Engines

I LCommands t -

Accelerometers

S-IVB Stage

-~-

Switch Selector

IBM ~ 6 3Fig ure 2.1-1 Block Diag ram of Satur n V Navigation, Guidance, and Control System


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Astrionics SystemI Section 2. 2

SECTION 2.2NAVIGAT ION

Th e function of navigation includes the de te r-mination of position, velocity, and thr ust accele rationof the vehicle fr om acceleromete r mea surements.Thes e quantities a re required input data for guidancecomputations.

Th e velocity of the veh icl t :.led by inte-gration of acceleration. T hree integrac- ?lero-me te rs a r e mounted on the stabilized inner g,- Cthe platform. The mutually orthogonal sens itive L.of the 3 accelerometers define the measuring or ac-cel ero met er coordinate syste m (xIyIzI). The platformis aligned before launch s o that the y~ axis is parallelto th e vert ical at the launch sit e and pointing upward,the XI axi s points in the directi on of t he flight azimuth(Az), and th e ZI axis completes the right-handed co-ordi nate syst em. The origin of the XIYIZI coordi-nate system is at the stabilized platform. For navi-gation computations, t he xsyszs coordinate sys tem isused. It has its origin at the center of the earth . TheXIYIZI syste m and the xsyszs system ar e parallel. Be-fo re launch, both system s ar e earth-fixed (rotate withthe ear th), but at the moment of platform re le as e (5seco nds befo re launch), both sys tem s become space-direction fixed (see Figure 2.2-1). The total accel-eration i: ( x y z ) of the vehicle in the xsyszs sy st emis given by:

and the velocity is

where F/M is the thrust acceleration, g ( r ) i s thegravitation al acceleration , and Co is the initialvelocity of th e vehicle a t launch (caused by ear throtation).

During flight, the integrating accelerometersdo not respond t o gravitational acceleration so theiroutputs a r e the velocity components (kIGIid result ingonly from thrust acceleration. To obtain the tota lvelocity of the vehicle, the velocity componentslxqj'gkg) caused by gravity must be added to the ac-

.rometer readings. The gravitational accelera-tl, ' r ) , which i s acting on the vehicle, is a functioof vt.. qosition and i s computed in the gravitat ionloop. I. -ust acceleration (F/M) is computedfrom the d~ +iated accelerometer output accord-ing to the equa~

A flow diagr am of the navigation computa-tions is shown in Figu re 2.2-2. Accel erome terreadings, initial velocity, and velocity (vg)(caused bygravitation) a re added to obtain the vehicle velocity( i ) n the space-fixed coordinate system (xsyszs).The velocity ( E ) is integrated and the in itial positionis added to yield the vehicle position ( r ). This posi-tion data is used t o compute gravitational accelerationand velocity which i s then added to the accel erome terreadings.

The gravitational acceleration acting on thevehicle is derived fro m the gravitational potentialfunction of the earth. The expression for the grav ita-tional potential (a ) f the ea rth (used for Saturnnavigation) is based on the Fisher ellipsoid and isgiven by:


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Ast r i on i c s Sys t e mSect ion 2 .2RE = equator ia l radiu s of the ear thr = 1 1 = di s ta nc e f ro m e a r t h ' s c e n t e r t o t hevehicleA = angle between r and the equa tor ia lplane

5 r 3-t J, D, H = constants

D R E ~+ ( 3 - 30 sin2 A + 35 sin4 h. )1 In or de r to combine the components of g rav i-35r4 ta t iona l acce le ra t ion wi th the m easur ed compo nentsof thru st accelerat ion , the gravi tat ional potent ial iswhere (2.2-4) f i r s t e xpre s se d i n a rec tangular coordina te sys tem(uvw ) which is then "rotated" to be pa ral lel with theG = univ ersa l gravi ta t iona l cons tant acce le rometer coordina te sys tem (xIy l z j a nd t heME = m as s of the ear th xsy szs coordina te sys tem.

NorthGround

Coordinate Systems:

UVW: Earth-fixed rectangular system used forgrav ity computations. Ori gin at the centerof the earth; W axis collinear with theearth's spin axis and pointing south; V axisi n the equatorial plane intersecting theequator at the longitude of the launch site;

Coordinate Systems:

xyz Accelerometer coordinate system, parall el to .the xsys5 system, orig in at the vehicl e platformU axis completes right-handed orthogonalsystem.

4 L: Geodetic latitude of the launch site.XsYsZs: Space-fixed rectangular coordinate system A Launch azimuth, measured in clockwisefor navigation computations. Ori gin at the direction from north.earth's center; the Y axis i s ~ a r a l l e lo the

vertical at the launch site at the instant oftake-off; the X-Y plane contains the fl igh t r : Radius vector from the earth's center to theI lane; the Z axis completes the right-handed vehicle.system.

Figu re 2 .2-1 Naviga tion Coordina te Sys tems2.2-2


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Astrionics System, Section 2. 2The potential function (2.2-2) i s expres sed in

polar coordinates and may be written in a rectangularearth-fixed coordinate system (uvw) with the originat the center of the earth and oriented as shown inFigu re 2.2-1 by using the relationship:

and w = - r sinkThe components of gravit ational accelerat ion

(ugvgwg) along the uvw coordinate axis ar e given bythe parti al derivatives of (u v w)

The expressions for Q and P ar e given in Table2. 2-1.

The components of gravitational a cceleration.. .. ..(ugvgwg) must be tran sform ed now into the accel ero-mete r coordinate system (xIyIz$(or xsy szs system).

The relationship between the uv w system and thexsyszs syste m may be expresse d in matrix form:

the rotational transformation matri x b] iss in AZ -sin ~ L C O S A , - C O S

co s 4 L - si n 4 Lcos A, si n 4 si n AZ si n Az CO S 4 L

where 4 L is the geodetic latitude of the launch sit eand Az is the launch azimuth me asure d clockwise fromnorth. The matr ix [A] corresponds to two successiverotations of the uv w system: fi rs t about the u axisby an angle + L and second, about the new v axis(now para ll el t o the ys axis ) by an angle of ( 90 " - AZ)(See Figure 2.2-1. )

Fig ure 2. 2-2 Navigation Flow Diag ram2 . 2 - 3

In i t ia l Veloci tytc Engine- utoffGuidanceComputations. a .Xo Yo zo rIntegratingAccelerometer ti Engine

Ignit ion-t. . . .; ; i '(xs Ys zs) - r(xsYs zs) ;-)I I

FM+' - - +

7X xXX z

--

RequiredAtti tudeAngles

GravityComputation

9 ( r )

A t7

7-- .

7

AttitudeCorrectionComputation

*d t

To' 'Y 'R+ ControlAttitude ComputerCommandI Computation.. .. ..x l Y ~ Z ~ ex ey e zF/M PlatformGimbal Angles

IBM B6 5


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Astr ionics SystemSection 2. 2

The gravitat ional accelerat ion components in The quanti t ies and equations requ ired for navi-the xsyszs coordinate sys tem are: gat ional computations a r e l isted in Table 2.2-1.

ig ysQ - P s in $L-.zg = zsQ + P cos 9~ si n A, (2.2-9)The quanti t ies sin 6L, os $L , ioiOio,xoyozo ar e constants for a given launch s i te and a resto red in the LVDC. The ini t ial condit ions xoyozorep re sen t the posi tion of th e launch si te in the xsyszscoordina te sys tem while &oioios the velo city of th elaunch si t e (equals ini t ial vehicle velocity) caused bythe rota t ion of th e earth. For ce r t a i n mi s s ions , t helaunch azimuth A, va ri es with time. In thi s case , thequantity cos AZ stored in the LVDC is continuouslyupdated.

Navigation is perfo rme d continuously through-out the mission. During coast fl ight per iod s (in orb i tand af te r t rans lun ar injec tion), no acce lerom eterreadings a r e obta ined. Posi t ion and ve loc ity a reobtained fro m grav ity computat ions alone; i. e. , bysolvin g the eq uation s of m otion.

Actually, a sma ll thru st is applied duringcoast fl ight which is th e res ul t of the venting of th eS-IVB hydrogen tank. Whether the accelera t ioncaused by vent ing can be m easured, considered incomputat ions, o r neglected, is under investigation.

Table 2.2-1 Navigation Equations

Velocity: Positio n (Displacemen t)xs = i I + J x g d t + i o x s = !is dt + xoh = h + J a d t + y o YS = J js dt + YOzs = iI + J E ~ t + io z s = J i g dt + zoThru st A ccelera t ion:

F --J'm+ -x0yozo and ko.j'oio a r e initia l conditions a t lift-offiIjTIiIAccelerometer outputGravitat ional Accelerat ion:..xg = xsQ - P cos $L cos Az 4 L : Latitude of launch site

fg = ysQ - P s i n $ L AZ : Launch AzimuthE g = z s ~ + po s d ~ ~ s i n,

Q = - % ~ + J ( " ) ~3 (1-52)+~(~)3(3(3-7)+~(%7( 4 r2 ?)I4P = - ( ) +2 - ( 1 3 ) + ( ( 2 2 f ]


32/42


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Astrionics SystemSection 2. 3TRAJECTORY CONSTRAINTS AND LAUNCHWINDOW

A lar ge number of c onst rain ts apply to thetra jec tor y of a launch vehicle. These constraintsresu lt f rom m ission rule s; environmental conditionssuch as the at mos phe re and location of launch point;operational requirements such as safety restrictionsof launch azimuth and tracking requi rements; andhardware limitations such a s structural load limitsand avai lable propulsion means. The dominatingand most sev ere requirement applying to the choiceof trajectories is optimal propellant utilization.

A combination of the various const rain tsgenerates a limited tim e period f or launch; i. e. , alaunch window to meet mission requirements. Therea r e 2 types of launch windows fo r the launch vehiclein the Saturn ~/Ap ol lomission: the ground launchwindow for as cen t into parking orbi t and the orbitallaunch window for tr ansl unar injection.

The launch azimuth at the ground and theorie ntat ion (inclinat ion and descending node) of theparking orbit a r e varying with time. A brief ex-planation of the r easo ns f or time varie nt launchazimuth, inclination, and descending node se em sadvisable a t this point. At any instant of tim e aplane can be defined which contains the launch site,the ce nte r of the eart h, and the moon at the desiredtime of arri val . In or de r to maintain this plane it isnecessary to vary the launch azimuth on the groundcontinuously, sinc e the launch site is moving with therotating e art h and the moon is moving. When rangesafety constra ints a r e introduced, the launch azimuthwill be limited to a certain band of values. Forlaunches fr om Cape Kennedy, the band is approxi-mately 45 to 110 degre es meas ured e as t of north andwill be encountered twice each day. Within thi s band,a range of azim uths of 26 deg rees will guaran tee atleas t a 2-1/2 hour launch window and normally thisvar iat ion will be applied to the portion of the bandwher e launch azimuth var ie s line arly with the time oflaunch. Of the two daily launch peri ods, one willgenerally lead to a sho rte r coasting a r c in orbit thanthe other . Since the geographic position of the launchsi te is fixed, it follows that one of the launch per iod slea ds to time of coas t in parking orbi t such that igni-tion in orbit occu rs within a range close to thelaunch site-sometimes ref err ed to a s the "Atlanticopportunity". The second launch period ca lls fo rcoasting a r c s of gr ea te r length-commonly re fe rr edto a s "Pacific opportunities".

Since launch vehicle payload capability isdegraded by powered plane changes, it i s assum edthat in the nominal case there will be no such ma-neuvers. If it is des ire d to go into an orbi t with aninclination to the ear th's equator (g rea ter than thelatitude of the launch si te) , then it i s possible tolaunch directly into the properly selecte d launchazimuth. There a r e other const raints which mightbe present which affec t the launch window problem;e. g., launch during daylight, proper lighting of thelaunch site, and lighting of the re tur n landing si te s.These, however, do not affect the basic geometricalconsiderations.2.3.2 ITERATIVE GUIDANCE MODE

The iterative guidance mode was developed tomeet the mission flexibility requireme nt of la rgespace vehicles with minimum propellant consumption.The scheme is based on optimizing techniques usingcalculus of vari atio ns to dete rmine a minimum pro -pellant flight path which satisfies the mission require-ments. Experience with hundreds of minimumpropellant trajectories for various orbital injectionmissions has demonstrated that the optimum thrustdirection, relative to the local vertical, is verynear ly a l in ea r function of t ime du ring vacuum flight.Moreover, the si ze of the angle between the opti-mum thrust direction and the local horizon is neververy large. These observations show a remarkableagreement with the mathematical results obtainedfro m the calculus of varia tions when a flat eart hmodel having a constant gravitational field is used,and position and velocity constraints a r e imposed atcutoff. A clos ed solution can be obtained with thi smathematical model and yields an explicit equationfor the optimum thrust direction. This equation hasthe form:

X p = a rc tan (A + Bt)whereXp i s the optimum thrust direction for mini-mum propellant consumption and t i s the time.Constants A and B ar e determi ned by the spe cifiedcutoff velocity and position, the init ial values of th estat e variables, the vehicle thrust acceleration, andthe engine specific impulse. The comp aris on of thisequation with the res ult s of tra jec tor y studie s sug-ge st s the us e of the approximation:

A rectangular guidance coordinate system( r , q , 5 ) (Figure 2. 3-1) is establis hed with the origin


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Astrionics System, Section 2. 3a t the cent er of the e ar th and with the 7 axis lyingalong the vertical which int ers ect s the calculated cut-off position of the vehicle. Simplified equations ofmotions a r e derived to approximate the motion overan oblate earth with a realistic gravitational field.These equations of motion a r e solved during flight todetermine the instantaneous range angle to cutoff, thetime-to-go to cutoff, and the gravitational effectsoccurring over the remaining flight time. This in-formation i s used to compute values for A and B con-tinuously during flight. In pract ice, only the value ofA need be calc ulate d when computation ra te s on theord er of one or mor e pe r second ar e used (exceptduring the la st few seconds before cutoff when bothA and B a r e computed and held constant over the re-maining flight time). This is necessar y because theequation gives an indeter minate command angle atthe cutoff point.

The iterative guidance scheme, which is aquasi-explicit scheme, will be activated aft er jetti-son of the launch es cape s ystem in the S-I1 Stage

ytLaunch

Point-B

Situation Shown in the Flight Plane

Xs Ys z~ - Space -f ixed Coordinate Systemfor Navigation

& = r ] { - Space -f ixe d Coordinate Systemfor Guidance (IGM)

*T - Range AngleIBM B66

Figur e 2.3-1 Coordinate System Used fo rIterative Guidance Mode (IGM)

burn period and will continue in operation to inser tioninto parking orbit by the S-IVB Stage and subse-quently to lunar tr ansi t injection with the second burnto the S-IVB Stage.

The iter ative guidance mode equations forascent into parking orbit and for powered flight out oforbit ar e shown in Figures 2.3-2 and 2.3-3, respec -tively. The guidance scheme gener ates commands fo rthe pitch and yaw angle of t he th ru st di rection and thecutoff velocity.

The inputs required are divided into 2categories: ( 1 vehicle dependent inputs and ( 2 )mission dependent inputs. Figur e 2.3-2 shows the seinputs require d fo r the ascent-to-orbit phase and theguidance equations-the solution of which genera te sthe steer ing commands in pitch and yaw. Fig ure2.3-3 shows the additional input for flight-out-of-orbi t and the additional computation. Although it isseen that the required inputs for the 2 phases differsomewhat, empha sis should be placed on the fact thatthe guidance equations ar e identical for both phases ;i. e. , hrough built-in logic on the LVDC, the multi-stage equations necessary fo r ascent to parking orbitar e reduced to single stage equations for flight-out-of-orbit by means of nulling the prope r par am eter s.This task can be accomplished entirel y onboardduring the time in parking orbit.

Figure 2.3-3 shows that there a r e 10 inputsdependent on the physical cha ra cter is ti cs of the vehi-cle and 33 inputs dependent on the mission. It shouldbe noted that the inclination of the parking orb it cutoffplane, the decending node, and the launch azimuth a r efunct ions of tim e of launch. During th e las t few min-utes pri or t o lift-off, these quantities will be computein the launch ground computer and the r esul ts will beput into the LVDA. At the sam e tim e, the StabilizedPlatf orm will be turned in the direction of the d esi redlaunch azimuth. Fig ure 2.3-4 shows how the quanti-tie s, dependent on tim e of launch, va ry fo r a typic allaunch day. Very simple repres entat ions of. the securves, as a function of t im e of launch, can beobtained.

When the inputs a r e considered fo r the out-of-orbit case, it is convenient that the vehicle characterist ics of the S-IVB Stage ar e alre ady in the LVDC fr omit s fi rs t burn into parking orbit. It has been statedpreviously that the multistage equations are reducedto the single stage by properly storing ze ro s into thepertinent paramete rs while the vehicle is in parkingorbit. Since this featu re is an outgrowth of p rog ram -med logic, it is questionable whether these par am et er


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Astr ionics Sys temSection 2. 3should be considere d bona-fide inputs . As a r e s u l t ,only estim ate d cutoff tim e for the second burn of theS-IVB is shown in Figure 2.3-3 a s an input. Th erea r e 27 inputs dependent on the mission. Four quan-t i t ie s a r e t ime dependent and can be represented bypolynominals a s shown in Fig ure 2.3-3: the aimvecto r, aim vecto r magnitude, nom inal ecce ntricityof t he outgoing elli pse , and cutoff e nergy. If the 2-body proble m is cons idered such that Kepler ianmechan ics holds, the a im vector might be def ined asclose to the ear th-moon l ine (a l ine which is betweenthe ce nte r of the e ar th and the ce nte r of th e moon).I ts d i rect ion is, of co urse, t im e var ian t s ince theear th and moon a re moving re la t ive to each o ther .The ac tua l a im vec to r s a r e t aken f rom fu l ly op timizedtra je cto r ies which a t ta in a specif ied per iselenum,flight t ime, and a specified inclination to the lunarequatorial plane. In the nominal case , the f l ight p laneis chosen such that no powered p lane changes ar e re -quired. If parking orbi t per turbat ions occur , exper-ience has shown that rotating the cutoff plane about theaim vector re sul ts in s ignificant payload gains . Thecutoff pla ne is defined by the aim vecto r and perigeevec to r as shown in Fig ure 2.3-3. Ene rgy input con-s t r a in s f l igh t t ime and the re fo re de te rmines the s emi -ma jor ax is of the outgoing ellipse . The magnitude ofthe a i m vector det erm ines the t ru e anomaly of thea im vec to r . In shor t , wi th the e l l ip t ical par am ete rsM, sem i-ma jor ax is a , and eccentr ic i ty e , the shapeof the outgoing ellipse is defined. The perige e vecto rS det erm ine s the orientation of the e llipse, and the-per ig ee vector and aim vector determine the p lane.S ince the i ter a t ive guidance equations predict range-to-cutoff updated a t d isc rete in tervals , the t rueanom aly of th e cutoff point can be com puted and hencethe cutof f param ete rs T T , VT, and OT which fe ed intothe i tera t ive scheme. I t mus t be emphas ized thatthe se quant i t ies change under per turbat ions such thatthe S-IVB cuts off possibly at a different point on theel l ipse for a thrus t per turbat ion for example, in ord erto maintain the desire d conditions at the moon. Theignit ion c r i ter ion shown here is that ignition willoccur when the orbiting vehicle is a f ixed angle f romthe aim vector. The evaluation will begin upon injec-tion into orbit and tes ts for th e number of changes ofs ign wil l be made unt i l the corr ect number is attained.If fo r some reason (e. g . , the re is not enough time t oper fo rm a l l checkout p rocedures be fo re the t ime toignite out of the f i rs t orbit) i t is nece ssary to go outof the second orbit , the sa m e type of repre senta tionsof ai m vecto r, energy, and ecce ntricity can be readinto the onboard computer and it is even conceivablethat the sam e coefficients may be used. Open ques-t ions still exis t as to whether to update the energysuch that the moon is r eached a t a f ixed tim e of a rr i-

val- for example, a l te r the energy such that the re -sulting tim e of flight is reduced by the additional t imein parking orbit- o r to execute a plane change and flythe s am e amount of t im e a s in the c as e of going out ofthe f i r s t orb i t .Exa min ation of th e loop shown in Fig ur e 2.3- 5

between vehicle tracking and tele m etry of the LVDCand the s tabiliz ed platform, along with transm issio nof data in rea l t im e to the IMCC, the vari ous possib il-iti es of updating information to the compu ter, of ov er-riding comm ands such a s the ignition equation, and ofperforming an a l te rnate miss ion can be seen . In theca se of abort (e. g., fail ure so mew here in the S-11 tothe nominal mission), no change in input is required.The solution of time-to-go and other para m ete rs inthe i tera t ive scheme se t t le out qui te rapid ly (af ter 2o r 3 cycles ). This s ta tement a lso holds for changesin desir ed term ina l conditions which might be builtinto the computer. Nei ther the accuracy nor the nea roptim ality of the sche me is degraded.

F igure 2. 3-6 shows a typical re la t iv e t imeseque nce of e vents: when the vehic le dat a is supplied,when the mission is defined, when the launch day isgiven, and when the inputs a r e loaded into the launchcomp uter to be transm itted to the LVDC. It would beassum ed that one repres entatio n of the time-vary ingquantities would be made available over an enti remonth which contained a potential launch date. If 1month before scheduled launch, a launch date in adiffere nt month is selected , approximately 1 weekwould be required t o obtain new rep resen tations oflaunch azimu th, inc lination, desc ending node, a imvecto r, e ccen tricity, and energy.

In summ ary, the ef for t has been d irected to-ward s gene rality of guidance equations with a mini-mum am oun t of input - oth for the nominal m iss ionand for abor t and al terna te miss ions . S ince a mini-mum amount of "tampering" with the main prog ramin the LVDC is necessary , ef f ic iency is at a highleve l, and checkout is greatly s implif ied. A g r e a tdeal of consideration was given to the question: Isit better to have one se t of guidance equ ations withthe extraneous computation of the quantities whichar e nu l led fo r a s pec ia l cas e , o r to have s ep ara teequations fo r the ascent-into-orbit and the f light-out-of-orb it? Close coordination with perso nnel whowould actually be involved with the flight progr am onthe LVDC resolv ed the answer. Extraneous computa-tion, with the m inimum amount of c hange to th e flightp rog ram, is more des irable . The t ime required forprogramming a com puter f rom one f l ight to the nextis reduced by thi s method. In fact , theoret ical ly it is


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Astrionics Syst' Section

Figure 2. 3-2 Iterati ve Guidance Mode Equations, Flight to O2 .3 -5 /2 .

CO O RDI NATE RO TAT I O NS

[KIC+d[GI

I NPUTS FO R ASCENTVehtcle Dependent Inputs (9)

T = estimated first stage burning timeTz = estimated second stage burning timeTs :Coast time between second and thrrd stagesVex, 'go Ispl

M ATRI X TO TRANSFO RM FRO M PL ATFO RMCO O RDI NATES TO G UI DANCE CO O RDI NATE S0 coo AI

O Ico s A 0 s in Az

T I M E TO CUTO FFCAL CUL AT I O NA t * = i T - c e ,~*A = ,jr-T= - i - f , ~ X

( A k * ) z + ( ~ + * ) 2 + ( ~ i * ) z

vex2 'go Isp2Vex3 =901sp3rz =nom~nalmm7 3 = n or n ln a l m 3/ m3TI =estimated hird stoge burnlng tameto porking orbtt

Mission Dependent Inputs (7)sr = radius at cutoffVT = v e lo c i ty a t c u t o ff81 =path angle of cutof fK, Emission constonti =incl~not ion f cutof f plone8 =descending node of cutoff planeAZ = launch azimuth

TERM I NAL RANG E ANG L E CAL CUL AT I ONI N O R B I T P L A N EL =ln[t2/(*,-Tz)]J = Ve Xz ( ~ 2 L 1 - T2 )S =J T Vex L9 = vexz~: r2S

+= arctan( r fp/yfp)+ 1 8,+82)?4

I0 s i n coa+

c o r e - s in 8

0 sini[ D l = [ -siny: 0 cosi I[ ~ l = [ ~ lc l [ e l [A ]

AT = ~ ( r , T1) /vT = T ' t A TT* - T * + A T

i T = VTsin BT velocity along r ) axis I

r, = V F M)Ll = ln[rl / k l -T I]JI = Va x 1 ~ I L I - TS = JI- T V L I9, : T 2 + r 1 S oL+= V Ll + V L-181: SI-JP+L*(TZ+TI)TC +T2+Tc

4' v~ (T,L'-T')L' = ln[r,/(r,-~')lTI = Ti TCd = L* tvL8 VT*-J'+A'T1-K6(r3-T' ) (A'+V-VTIL'

i, = V COO 9, velocity along E axisA

a Y z, i, , i F/m. P

7I C U T O F F W H E N V S V T IIBM

i

PI TCH AND YAW ANG L E CAL CUL AT I O NSM * c o s zy + K3 s inPyN = K4 sin xYAp = MAy - NBypI =-~T+~+TIJIP2 = -+T2 Vex2+ Jzlrz+2T1)P 3 = - + T ~ ~ V ~ ~ , +3( r3+ 2Tlc)

a t = A~ " -%ATA A~ATat = at* t g a 7Zy = o r s t o n [ ~ j l ( ~ ~ + ~ ~X p = o r c t a n ( A j 1 ~ i )Vex3L3 V1x3L'+G

B p = M B ~ - N ( T ~ V ~ X , L ~ ++T Vex3L3++ q + P z + P 3 )

E ~ =-,T+. sT *+I..s g T* 2 -- MC' - ND'I s in zpc P = ( M C' - ND ' ) COSI,Q' = { Vex+ r3S3Ul = ~ T ? v ~ ~ ~ +l01U2 . T ~ ~ v ~ ~ ~ +2(r2+2TI)U 3 = ~ T ~ V ~ , , + Q ' ( T ~ + ~ T ~ ~ )Dp {MDI- N[UI + u ~ + u ~ + ~ I ~ s ~ +

+ T ~ ~ s ~ - ( T ~ + T ~ ) ( P ~ + P ~ ++ T?VC~LL) P I T Z ] )C O ~ Y ~

KI = Bp Ep I ( Ap Dp - B p C p lx ;= I p - K l

Ay * La+ Vex3L3J = J'+ T ~ GC' = J3- AyT3- 81Cy = C c os jiyE~ = C + ~ ~ * + t i g ~ * 2 - ~ ' s i n z83 . I + J ~ + T l v e x ~ L tBy J3 + T c V~ X~ L+ 8 384 ' 01+02-TZJI+TISI8 ' (T3+ Tc)s3 ' 3 - T3vex3 L3O3 a i~ i~~( '~+TIc)D' . 4-8s+Q3oY D ' C O S ~ ) .K3 = By EY/ ( Ay Dy - By Cy )x = Py- '(3


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Astrionics SysSection

I N P U T S F O R O U T O F O R B I T

( parking orbi t 1Mission Dependent Inputs (28 )

hf; = Mi 0 + M, , T + M i 2 T 2 + ~ , , T 3uni t oim vector

M = R O+ R , T + R ~ T ~ + R ~ T ~aim vector magni tude

e = e o+ e , T + e 2 T 2 + e 3T 3nominoi eccentricity

C3 = C o + ~ , T + ~ 2 ~ 2 + C 3 T 3cutoff enery

rN =nominal radws ot t ime of igni t ionK 8 = m i s s i o n c o n s l o n tCOSA = cosine of angle before M_ to igniteTlGA =cosine of ongle before M to begin

ch~l l own

I AFTER PARKING ORBIT INJECTIONCOAST UNTIL II co s A = ++TIGAIGNITION OCCURS IN TIG SECONDS t

M A T R I X T O T R A N S F O R M F R O M P L A T F O R MC O O R D I N A T E S T O G U I D A N C E C O O R D I N A T E Ss in A 0I A ] = [ -coA r : ] I

T E R M I N A L R A N G E A N G L E C A L C U L AT I O NIN O R B I T P L A N EL 2 = I n [ r 2 / ( r 2 - T 2 ) ]J 2 = V e x 2 ~ 22-T21S2 'J2-TZ Vex2 LZQ2 = 4 vex2T: +4 S2I, = V . ( F M)L~= ~ n [ ? ( r , - ~ , ) ]J : ax1 r tLi -Tt IS = J I- T V nX L I0, +vex, ,'+T, S,L'= V L l + Vex2 L t8 = -S1 -C + LT2 +T IT C . T , +T 2 + TcL' = l n [ r 3 / ( r 3 - ~ ' ) ]J' = v ( T ~ L ' - T ' )T* T 't TICA = L*tV..3~ 'B2 = V T * - J ' + A ' T ' - K ~ ( ~ ~ - T ' I ( ~ + V - V ~ ) CE:jG [i]41= a r c t a n X F P / Y ~ P ) + 8 , + %1

-

A4 = +,+a t r u e a n o mo l )= a r c t a n [* p a t h a n g l e

p / ( l + e c o s A O I r a d i usVT ~ ~ ( l + e ~ + 2 eo s v e l o c i t yi T = V T si n BT v e l o c i t y a l o n g T a r t siTVT coo 0 v e l o c i t y a l o n g E a x i sIC OMP U T E A T S E C ON D S - I V B I GN I T I ON

unit aim vectorM aim vector magnitudeeN nominal eccentricityC c u t o f f e n e rg ye =e + l eccentr ici ty of cutof f el l ipsep =$3(e' - i l semi lotus rectumc o r e =i (b ) t rue anorndy of aim vectorK5 =m i constantN =L XY normal to parking orbi t planeD D: (N .M I? IN XM ?+ I c o ns t on tFF = N. M / I NXM I ~ constontS =[OD M-FF ~+ f i ( * -D~ )+ ]c o s + *p e r ig e e vector

M - S c o s YSL -CoS2+X vector perpendicular to perigee vectorC_, =SXS, normal lo cutof f planeg = c o s A c o s & i + s i n +o 1- in A* cos +0 Ir earth spin vectorcos i =g GI inclination of cutoff pione1, = - co s A s in 8, i+ oo 8, 1

+ s in A s in 8. k vector to launch siteDN = GIX Q cmstont vector = orc tan '$ E1~ dercendinq node0 = a r c a n (W-l)ngle from S to descending node

C O O R D I N A T E R O T A T I O N S

C A L C U L A T I O NA +* ' .,,'i.J*

C U T O F F W H E N

P I T C H A N D Y A W A N G L E C A L C U L A T I O N S

Figure 2. 3- 3 Iterat ive Guidance Mode Equations, Flight out of O

M = cos " + K3 sin gyN = K a s i z y r yA p = M A y - NB yp I = - ~ T ~ ~ V ~ , , + ~ ~ J ~P2 = - 4 T z 2 V e x 2 + J 2 ( r 2 + 2 T , IP3 = - $ T ~ ~ v ~ , ~ +3 ( r 3 + 2 T l c lB P = M B ~ - N ( T ~ ~e x 2 L z +

+T Vex3Ln ++ F / + P 2 + p , l

E p = , - , , T ++ ~ " + + i i g ~ * 2 -- MC ' - NO' ) sln ypCp = ( MC ' -ND'Ios gp0' = ~ T ~ ~ v ~ ~ ~ +3 ~ 3UI = $ T ~ ' V ~ ,~ + 1 Q 1U2 = i ~ ~ ~ ~ ~ ~ ~ + 0 ~ ( r ~ +~ ~ 1u3 $ T ~ V ~ ~ , + O ' ( ~ ~ + ~ T ~ ~ IDp. { M D ' - N [ U ~ + U ~ + U ~ + ~ ~ S Z +

+ T ~ ~ S ~ - ( T ~ + T , I ( P ~ + P ~ ++ ~ 1 ~ ~ e x 2 ~ 2 ) - ~ 1 ~ 2 ] } c 0 s j i p

A( = A(* -GATA' - A,"-igAT? -AC = A C * - ~ ~ A Tgy OK IO" [A~/(A~Z+A+~I"Zx p = o r c t o n I j / A C lVex3L3 = Vex3LS+A = L*+ VexJ~ = J '+ T ~ G3L3C' = J 3 - A y T 3 - 8 ,Cy = C' cos gyEy : + ~ ~ * + ~ ~ g ~ " 2 - ~ ' sB3 = J I + J 2 + T I V e x 2 L 2By = J3 + T1cVex3L3+ Bn8,; O l + Q 2 - T z J ~ + T ~ S pB5 = B 3 1 T 3 + T c lSa = Jn - T3Vex3 '-303 * ~ T ~ ~ v ~ ~ ~ + SD' = B a - B 5 + Q3D = D ' C O S Y ~


38/42

Astrionics System, Section 2.

Figure 2.3-4 Launch Window Parameters2.3-

Inclination of Earth Parking Orbit (degrees)A35

72.0Inc lin ati on of Earth Parking Or bi t

VersusTime of Launch 30

for a Typical Launch Window LaunchAzimuth(degrees) 98.0rt +25 30 35Time of Launch from Midnight (lo3 seconds)

Decending Node (degrees)A150 -

125

Time of Launch Venus Descending Node 100-for a Typical Launch Window75

25 30 35Time of Launch from Midnight (lo3 seconds)

Launch Azimuth (degrees)90

80Time o f Launch Versus Launch Azim uthfor a Typical Launch Window

70

24 26 28 30 32 343Time of Launch from Midnight (10 seconds)

IBM ~ 6

I +

72.> 82.8LaunchAzimuth(degrees)

%*

-98.0


39/42

Astrionics SystemSection 2. 3possibl e to prog ram the computer, change the vehicleconfiguration, change the mission, and never perfor manother checkout of the LVDC. Pra ct ica lly , of course,a high confidence level is necessary for crew safetyand fo r accomplishing the desired mission - ocheckouts a r e perf orme d again but of a much mo relimited nature than if a completely new computer pr o-gram we re required.

The guidance mode, qualitatively describedhere, is not necess aril y a final recommendation; itsimplementation is mere ly one way to accomplish thetas k of success fully performing the Apollo missionand many othe r orb ital and, indeed, interplaneta rymissions. The deta ils of the cutoff surf ace must beanalyzed carefully with resp ect to interface and sub-sequent consistency with the spacecr aft guidancemode.

The iterativ e guidance equations requ ire m oreLVDC capacit y than most of the other guidancesche mes optimized for minimum propellant consump-tion, but conside rable flexibility is gained. The sam ese t of guidance command equations is applicable toalmost a ll orbita l missions and can be formulated foruse with any number of high th rus t stages. The smallnumber of pre sett ings that must be calcula ted for a

flight repr esen ts physical quantities such as vehicleexhaust velocity, nominal cutoff time , and desir edcutoff position and velocity. This is an importahtcharacte ristic of the sche me since thes e presettin gsmay be determined without resortin g to time-consum-ing statistica l methods. The accurac y and propellanteconomy with the scheme a r e excellent. The fuelrequire d to attain the des ired cutoff conditions (atorbit insertion) is within 5 kilograms (11.1 pounds)of that re qui red using exact minimum propellant equa-tions obtained with the calculus of var iati ons . Th iseconomy is obtained even under s eve re perturbati onssuck as an engine fail ure in the fi rs t stag e of a 2stag multi-engine vehicle.2.3.3 G U I D A N C E F U N C T I O N S I N F L I G H T

The following is a brief summarizing descrip-tion of guidance functions during the var ious phases ofthe Saturn Launch Vehicle Mission in the ApolloProgram.

Before launch, the platform is erected withthe y~ axis vertical and the XI axis pointing in thedirecti on of the launch azimuth. Since the launchazimuth is varying with time , the platform is torquedto maintain this orientation. Just pri or to lift-off, the

Figure 2. 3-5 Linkage Between Vehicle and Ground2. 3-10

Ground Computer Accelerometerfor Launch Functions of Launch Veh icle MeasurementsOpera tion Launch Time Dig ital Computer

Attitude

II II

Mission From PO

I I1

II Note:

Integrated MissionControl Center

IBM ~ 7 0A

MSFNRemoteSites

To update for alternate missionfrom parking o rbit requires up to27 constants.(Automatic ign iti on sequence i s used)


40/42

Astrionics SystemI Section 2. 3

platform is released and becomes space-fixed oriented.Th e XI axi s now deter mine s the flight azimuth.FIRST STAGE FLIGHT

The vehicle lifts off ve rtic ally fr om the launchpad and maintains its lift-off orientat ion long enough tocl ea r the ground equipment. It then per for ms a rollmaneuver to align the vehic le with the flight azimuthdirection. This maneuver gives the vehicle controlaxes the cor rec t alignment to the flight plane thussimplifying the computations in the attitude c ontrolloop. On the launch pad, th e vehicle always has aro ll or ientation fixed to the launching s ite.

During firs t-st age propulsion, a time tilt pro-gra m, sto red in the LVDC, is applied simultaneouslywith the describ ed roll maneuver. The pitch angle ofthe vehicle i s commanded according to the tilt pro-gram which is a function of ti me only and is independ-ent of navigation measu rem ent s. However, navigationmeasurem ents and computations ar e performedthroughout the flight, beginning at the tim e the plat-form is rele ased (i. e. , 5 seconds before lift-off).Cutoff of the f ir s t st age engines oc cur s when the fuellevel in the tanks reaches a predetermin ed level.Thereafter, the firs t stage is separated from thelaunch vehicle.

SECOND AND THIRD STAGE FLIGHTAfte r ignit ion of the S-I1 Stage, adap tive guid-

ance (i. e., the iter ative guidance mode) is used d

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