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Full text of "Command service module system handbook"

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FC007 
1/16/67 



NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 



COMMAND SERVICE MODULE 
SYSTEM HANDBOOK 



AS-501 



m 

1 iSSS? JANUARY 16, 1967 

^•xy:-:':-: prepared by "%sizSS^ 

FLIGHT CONTROL DIVmON 

^ -'*'■-•'■ ^^!P^ I 

' "^ ~ MANNED SPACECRAFT CENTER 

1 V ^BHB^ \i^^^if^M^ HOUSTON, TEXAS 

(NASA-TM-X-68878) COMMAND SERVICE MODULE N73-70019 
SYSTEM HANDBOOK (NASA) 16 Jan. 1967 

'.•'.■'.•'.'.'. Unclas 

'■<<:■<: 00/99 37764 

I I I I I ! ! f f f T f f T T i I T T F T 




APOLLO 
COMMAND SERVICE MODULE SYSTEMS HMDBOOK 
AS- 5 01 

PREFACE 

This handbook has "been prepared by the Flight Control Division, 
Manned Spacecraft Center, Houston, Texas, with technical support by 
North American Aviation. Information contained within this section 
represents the Command Service Module Systems for the AS-501 Mission 
as of January l6, 196? . 

Information as shown reflects spacecraft systems with major 
emphasis on material for use by Flight Controllers in real time; 
however, caution should be exercised in using these systems drawings 
for any purpose other than flight control. 

This handbook is not to be reproduced without written approval of 
the Chief, Flight Control Division, Manned Spacecraft Center, Houston, 
Texas . 



Approved 




Arnold D. Aldrich 

Chief, CSM/Gemini Systems Branch 




fodge 
Chief, Flight Control Division 



\ 



11 



I I ill III I I I II I i i i I i 



TABLE OF CONTENTS 



Section 



Page 



1 INTRODUCTION 1_1 

2 SEQUENTIAL 2-1 

2.1 Pyrotechnics, General Notes 2-1 
2.1.1 Initiator, Standard Apollo 2-1 

2.2 Power Distribution 2-2 

2.3 Laiinch to Insertion Notes 2-U 

2.3.1 General 2-1+ 

2.3.2 Launch Escape System Aborts 2-5 
2.k Normal Preentry and Descent 2-11 

2.1+.1 General 2-11 

2.U.2 Earth Landing System 2-11 

2.5 Landing 2-li| 

3 ELECTRICAL 3-1 

3.1 Fuel Cell Notes 3-1 

3.2 DC System Notes 3-5 

3.3 AC System Notes 3-8 
3»k Circuit Breaker Summary 3-10 
3.5 AC Equipment Summary 3-20 

k ENVIRONMENTAL U-1 

U.l CM Pressurization k-1 

i+.l.l Cabin Pressure Control U-1 

1|.1.2 Water Supply k-k 

k»2 Coolant Notes 1+-7 

i+.2.1 General k-J 

U.2.2 Coolant Circuit Notes U-T 

1|.2.3 Glycol Evaporator Temp Control Subsystem k-Q 

k.2.k Cabin Temperature Control Valve U-9 

5 CRYONGENICS 5-1 

5.1 Notes, Cryogenics, General 5-1 

iii 



II ill I I I I I I I i i i i i I i 



* TABLE OF CONTENTS (Cont'd) 

Section Page 

6 COMMUNICATIONS 6-1 

6.1 Notes, Communications System Block Diagram 6-1 

6.2 Commiini cat ions System Equipment 6-U 

6.2.1 C-Band Transponder and Antenna Equipment 6-U 

6.2.2 VHF FM Transmitter 6-U 

6.2.3 VHP AM Transceiver 6-U 
6.2,^4 Updata Link Equipment 6-9 

6.2.5 Updata Real-Time Commands 6-9 

6.2.6 The Unified S-Band Equipment (USBE) 6-9 

6.2.7 S-Band Power Amplifier 6-l6 

6.2.8 HF Transceiver /Beacon 6-l6 

6.2.9 VHF Recovery Beacon 6-l6 

6.2.10 VHF Survival Transceiver /Beacon . 6-l6 

6.2.11 Flashing Light Beacon 6-21 

6.2.12 VHF Recovery Antennas 6-21 

6.2.13 HF Recovery Antenna 6-21 
6.2.1^ Scimitar Antennas 6-21 

6.2.15 S-Band OMNI Antennas 6-21 

6.2.16 Audio Center Equipment 6-21 

6.3 Notes, Communication System MCP, CUB, and UDL 
Interface Diagram 6-29 

6.k Notes, Commionication System Detailed Block Diagram 6-32 

6.5 Commxinication System Circuit Margins 6-35 

6.5.1 Communication Links 6-35 

6.5.2 Ranges of Communication Links 6-35 
6.5-3 Calculated Ranges 6-35 

6.6 Communication System Final Closeout Switch and 

Circuit Breaker Configuration 6-38 

6.6.1 MDC - 13, Audio Control Panel 6-38 

6.6.2 MDC 23 and 26 Audio Control Panel 6-38 

6.6.3 MDC-25 Circuit Breaker Panel 6-38 



IV 



I I I I 1 £ EI I I I I I I i i i i I i 



* TABLE OF CONTENTS (Cont'd) 

Section Page 

6.6.1+ MDC-19 Control Panel 6-38 

6.6.5 MDC-20 Commimications Control Panel 6-38 

6.6.6 MDC-14 GScN Computer Control Panel 6-UO 

6.6.7 VHP Survival Beacon (Switch ON Beacon) 6-UO 

6.6.8 MDC-22 Circuit Breaker Panel 6-i+O 

6.6.9 RHEB Circuit Breaker Panel 6-1+0 

T INSTRUMENTATION 7-1 

7.1 General System Notes 7-1 

7.1.1 Operational Measurements 7-1 

7.1.2 Flight Qualification Measvirements 7-1 

7.2 Instrumentation Control System Notes 7-2 

7.2.1 Pulse Code Modulation Telemetry (PCM TM) 
Equipment 7-1+ 

7.2.2 Central Timing Equipment 7-U 

7.2.3 Signal Conditioning Equipment (SCE) 7-5 
7.2.1+ Flight Qualification Recorder (FQR) 7-9 
7.2.5 Bata Storage Equipment (DSE) 7-10 

7.3 PCM Signal Flow Notes 7-l6 

7.3.1 Programer 7-17 

7.3.2 Analog Data 7-17 

7.3.3 Parallel Digital Data 7-17 
7.3.1+ PCM Telemetry Parameters 7-l8 
7.3.5 PCM Telemetiy Fonnats 7-21 

8 GUIDANCE 8-1 

8.1 Notes, Systems, General 8-1 

8.2 Inert ial Subsystem 8-3 

8.2.1 G&N Modes 8-1+ 

8.2.2 Attitude Control - IX IMU and IX CDU 

Resolvers Utilized (IG, MG, OG) 8-6 

8.2.3 Fine Align - IX & l6X IMU and the IX & l6X 

CDU Resolvers Utilized (IG, MG, OG) 8-6 



II I I if II I I I I 1 I Li II I 



* TABLE OF CONTENTS (Cont'd) 

Section Page 

8.2.1+ Coarse Align - IX IMU and IX CDU Resolvers 

Utilized (IG, MG, OG) 8-T 

8.2.5 CDU Manual - No Resolvers Utilized (DACS 
Inhibited) 8-8 

8.2.6 Zero Encode - 1/2X & l6x CDU Resolvers 

Utilized (IG, MG, OG) 8-9 

8.2.7 Entry - l6X CDU & IX IMU Re solver Combination 

in OG (IG & MG,^ IMU = IX CDU) 8-10 
1a 

8.3 Optical Subsystem 8-12 

8.3.1 Notes General - Optical Subsystem 8-12 

8.3.2 IMU Alignment 8-13 

8.3.3 Optics Nav Sightings 8-lU 
8,3.1+ Orbit Determination 8-l6 

8.U Notes, Electrical Power Diagram 8-l8 

8.5 IMU Temp Mode Selector 8-20 

8.5.1 Notes General - IMU Temp Mode Selector 8-20 

8.5.2 ZERO Button (SU) 8-22 

8.5.3 IRIG GAIN Button (Sl) 8-22 
8.5.1^ PIPA GAIN Button (S2) 8-22 

8.6 Apollo Guidance Computer (AGC) 8-2li 

8.6.1 AGC Characteristics 8-26 

8.6.2 AGC Updates 8-2T 

8.6.3 Command Data Processor (CDP) 8-27 
8.6.1+ S/C Receiver 8-28 

8.6.5 KEYTEMP 1 8-28 

8.6.6 ASSEMBLY REGISTER 8-28 

8.6.7 OUT REGISTER k 8-29 

8.6.8 Program No. and Routines For AS-501 8-29 

8.6.9 Verb Definitions AS-501 8-31 

8.6.10 Noun Definitions AS-501 8-33 

8.6.11 Computer Worksheet 8-39 

8.7 Display and Keyboard 8-75 

vi 



I I I I I if 1 i II I 11 I I ill 1 



* TABLE OF CONTENTS (Cont'd) 

Section p^ge 

8.8 Lower Equipment Bay 8-77 

9 CONTROL 9-1 

9.1 Notes, Stabilization and Control Subsystem 9-1 
9 • 1 • 1 Intr oduct ion 9-1 

9.1.2 Rate Gyro Package 9-1 

9.1.3 Attitude Gyro Accelerometer Package 9-2 

9.1.^ Electronic Control Assemblies: Pitch, Roll 

and Yaw 9_i+ 

9.1.5 Electronic Control Assembly, Auxiliary 9-i+ 

9.1.6 Display and Attitude Gyro Accelerometer 
Assembly Electronic Control Assembly 

(D/AGAA EGA) 9-5 

9.1.7 Displays and Controls 9-6 

9.1.8 Flight Director Attitude Indicator (Not 
Implement ed ) 9-8 

9.1.9 Attitude Set/Gimbal Position Display 

(AS/GPD) - MDC6 9-8 

9.1.10 Delta V Display (AV) - MDC 7 (Not Implemented) 9-9 

9.1.11 Rotation Controller (Not Implemented) 9-9 

9.1.12 Translation Controller (Not Implemented) 9-9 

9.1.13 Attitude Impulse Switch LEB-i+ (Not 
Implemented) 9-9 

9.2 SCS Power Distribution Subsystem 9-10 

9.2.1 Introduct ion 9-10 

9.2.2 28 Vdc Non-Switched Power 9-10 

9.2.3 Reaction Jet Circuit Breakers 9-10 
9.2.U Group 1 Circuit Breakers 9-11 

9.2.5 Group 2 Circuit Breakers 9-12 

9.2.6 Direct Control Circuit Breakers 9-12 

9.2.7 Group 1 Power Switches 9-12 

9.2.8 Group 2 Power Switches 9-13 

9.2.9 Additional Circuit Breakers and Switches 9-lU 

vii 



I I I I IE I I I 1 I I I I i i i I ! 



* TABLE OF CONTENTS (Cont'd) 

Section Page 

9,3 Attitude Reference Subsystem (AES) 9-l8 

9.3.1 Operation 9-l8 

9.U Attitude and Thrust Vector Control Electronics 9-21 

9.i+.l Monitor 9-21 

9.U.2 Automatic Control 9-21 

9.i+.3 Manual Control 9-21 

9.U.U Modes of Operation 9-22 

9.5 Notes: SCS Signal List 9-30 

9.6 Notes: SCS Pover List 9-33 

9.7 Relay Identification 9-35 

9.7.1 Attitude Control and Thrust Vector Control 9-35 

9.7.2 Attitude Reference System Logic 9-35 

9.8 SCS Telemetry Measurement Points 9-37 

10 PROPULSION 10-1 

10.1 Notes, Service Propulsion 10-1 

10.1.1 SPS Engine 10-1 

10.1.2 Propellant Quantity 10-1 

10.1.3 Propellant Utilization 10-3 
lO.l.U Gimhal Actuators 10-3 

10.1.5 Helium Isolation Valves 10-5 

10.1.6 Gaseous Nitrogen (GN ) Tanks 10-5 

10.1.7 Normal, Off, Direct-On Switch 10-5 

10.2 Notes, Service Module, Reaction Control System 10-12 

10.2.1 General 10-12 

10.2.2 RCS Engines 10-12 

10.2.3 Propellant Quantity 10-12 
10.2.ii SM/RCS Propellant Quantity Computation 10-12 

10.2.5 Helium Isolation Valves 10-12 

10.2.6 Propellant Isolation Valves 10-13 

10.2.7 RCS Quad Heaters 10-13 

viii 



1 I I I I If I I I I I 1 i i i i i I 1 



* TABLE OF CONTENTS (Conr.lud^c^) 

Section Page 

10.2.8 Service Module RCS Jettison Controller 

(SMJC) 10-13 

10.2.9 Propellant Solenoid Injector Control 

Valves (Fuel and Oxidizer) 10-13 

10.2.10 Attitude Control Maneuvers 10-l4 
10.3 Notes, Command Module, Reaction Control System 10-16 

10.3.1 General 10-16 

10.3.2 RCS Engines 10-16 

10.3.3 Propellant Quantity Gaging 10-16 

10.3.^ Propellant Isolation Valves (Fuel and 

Oxidizer) 10-16 

10.3.5 CM/RCS Pressurization 10-16 

10.3.6 CM/RCS Dimp/Bum and Purge 10-17 

11 MISSION CONTROL PROGRAMER 11-1 

11.1 Mission Control Programer (MCP) , General 11-1 
11.1.1 MCP Components 11-1 

11.2 Ground Conmiand Controller (GCC) 11-3 

11.3 Spacecraft Command Controller (SCC) 11-9 
11. U Attitude and Deceleration Sensor (ADS) 11-15 
11.5 MCP Circuit Logic 11-16 

12 MISCELLANEOUS 12-1 



IX 



I I I I I i ill I MI t i i i I i 



* LIST OF FIGURES 

Figure Page 

ii.l Water Glycol Temperature Control Diagram ii-10 

6.1 Antenna Locations 6-3 

6.2 Communication Mode Range Capabilities 6-37 

10.1 Propulsion - SPS Total in Pulse Vs AV 10-T 

10.2 Propulsion - Settling time Vs Propellant Remaining 10-8 
11.5.1A MCP/S-IVB lU Logic 11-19 
11. 5. IB MCP/Seq (Recovery) Logic 11-20 
11.5.1C MCP/MESC Logic 11-21 
11.5.2A MCP/EPS Logic 11-22 
11.5.2B MCP/EPS Logic 11-23 
11.5.3 MCP/ECS Logic 11-2U 
11. 5. ^A MCP/TC Logic 11-25 
11. 5. ^B MCP/TC Logic 11-26 
11.5-^C MCP/TC Logic 11-2? 
11.5. 5A MCP/G&C Logic 11-28 
11.5. 5B MCP/G&C Logic 11-29 
11.5.5c MCP/G&N-SCS Logic 11-30 

11.5.6 MCP/SPS Logic 11-31 

11.5.7 MCP/RCS Logic 11-32 



I I III E E I I I I I II i I ill I 



^ LIST OF TABLES 

Table Page 

2-1 Tower Aborts 2-7 

6-1 C-Band Transponder Data 6-5 

6-2 C-Band Antenna Data g.g 

6-3 VHF - m Transmitter Data 6-7 

6-U VHF - AM Transceiver Data 6-8 

6-5 Updata Link Data 6-11 

6-6 UDL Real-Time Commands For Communication System 6-13 

6-7 Unified S-Band Data 6-15 

6-8 S-Band Power Ampl Data 6-17 

6-^9 HF Transceiver Equipment Data 6-l8 

6-10 VHF Recovery Beacon Data 6-19 

6-11 VHF Survival Transceiver /Beacon Data 6-20 

6-12 Flashitig Light Beacon Data 6-23 

6-13 VHF Recovery Antenna 6-2U 

6-l4 HF Recovery Antenna Data 6-25 

6-15 VHF Scimitar Antennas 6-26 

6-l6 S-Band Antennas 6-27 

6-17 Mission Control Programer 6-28 

7-1 Instriimentation System Equipment Power Requirements 7-3 

7-2 PCM Telemetry - Structure 7_23 

7-3 PCM Telemetry - Electrical Power System 7-2U 
7-^ PCM Telemetry - Launch Escape And Earth Landing Systems 7-27 

7-5 PCM Telemetry - Environmental Control System 7-29 

7-6 PCM Telemetry - Guidance And Navigation System 7-31 

7-7 PCM Telemetry - Stabilization And Control System 7-33 

7-8 PCM Telemetry - Service Propulsion System 7-35 

7-9 PCM Telemetry - Reaction Control System 7-36 

7-10 PCM Telemetry - L/V Emergency Detection System 7-37 

7-11 PCM Telemetry - Communications And Instrumentation 7-38 

8-1 Inertial Subsystems Modes And Associated SCS Modes 8-5 



XI 



11 III I I I ! E I I I I i i i I i 



LIST OF TABLES (Concluded) 



Table 

9-1 SCS Signal List 

9-2 SCS Power List 

10-1 SPS Point Sensor Location 

11-1 MCP Relay Logic For A Programer Reset 



Page 
9-31 
9-3it 
10-2 
11-17 



XI 1 



I I I I I I I 1 I II I II I i i i [ I 



LIST OF DRAWINGS 



Draving Page 

1.3.1 Interface - Electrical Saturn Launch Vehicle And 

Spacecraft (Sheet 2 of 2) 1-13 

1.3.1 Interface - Electrical Saturn Laianch Vehicle And 

Spacecraft (Sheet, 1 of 2) I-II+ 

2.2.1 Sequential Power Distribution 2-3 

2.3.1 Sequential Abort Initiate 2-9 

2.3.2 Sequential Les And SPS Abort System 2-10 

2.U.1 Earth Landing System 2-13 

2. 5.1 Sequential C/M Uprighting System 2-15 

3.1.1 Electrical Fuel Cell 3,ii 

3.2.1 Electrical - Direct Current System & Distr. 3-7 

3.3.1 Electrical Alternating Current System And Distribution 3-9 

i^-.l.l CM Pressurization And H2O Systems U-6 

U.2.1 Clycol Cooling l^.H 

5.1.1 Cryogenics Gas Storage System 5-3 

6.1.1 Coimnunications System Block Diagram 6-2 

6.3.1 Communications System CUB, MCP, + UDL Interface Diagram 6-30 
(Sheet 2 of 2) 

6.3.1 Communications System CUB, MCP, + UDL Interface Diagram 6-31 
(Sheet 1 of 2) 

6.U.1 Communications System Detailed Diagram 6-33 
(Sheet 2 of 2) 

6.U.1 Communications System Detailed Diagram 6-3U 
(Sheet 1 of 2) 

7-2.1 Instrumentation Control 7-15 

7.3.1 Instrumentation PCM Signal Flow 7-22 

8.1.1 Guidance And Navigation Block Diagram 8-2 

8.2.1 Guidance - Inertial Measurement Unit, Diagram 8-11 

8.3.1 Optical Subsystem 8-17 

8.i|.l Guidance And Navigation Power Supplies 8-19 

8.5-1 IMU Temp Controller 8-23 

8.6.1 G&N/MCD Controller 8-7I+ 



Xlll 



1 11 III III I III ill i L 1 



LIST OF DRAWINGS (Concluded) 



Drawing 

9.2.1 Stabilization & Control Power Distribution Subsystem 

9.3.1 Stabilization & Control Attitude Thrust Vector Reference 
Subsystem 

9.^.1 Stabilization & Control Attitude Thrust Vector Control 
Attitude 

9.^.2 Stabilization Reaction Jet ON/OFF Control 

9.^.3 AV Remaining & SPS Thrust ON /OFF Electron 

9.^.^ Stabilization And Control Mode Logic 

10.1.1 Service Propulsion System Detailed Diagram 

10.1.2 SPS Propellent Gaging System 

10.1.3 SPS Propellant Utilization Valve Control 

10.2.1 Propulsion - Service Module Reaction Control System, 
Diagram 

10.3.1 Propulsion Command Module, Reaction Control System, 
Detailed Diagram 

11.1.1 MCP Pwr Distribution 

11.2.1 Mission Control Program GCC (F/C & Cryo) 

11.2.2 Mission Control Program GCC (Battery Switching) 

11.2.3 Mission Control Program GCC (Comm & Instr) 
11. 2. U Mission Control Program GCC (RCS) 
11.2.5 Mission Control Program GCC (Prop, and G&C) 

11.3.1 Mission Control Program SCC (G&C) 

11.3.2 Mission Control Program SCC (G&C/SPS) 

11.3.3 Mission Control Program SCC (Seq'l) 
11.3-ii Mission Control Prograjin SCC (Seq'l) 
11.3.5 Mission Control Program SCC (EPS, ECS) 

12.1.1 Controls & Displays Main Display Console Panels 

12.1.2 Controls & Displays Lower Equipment Bay Panels 



9-17 

9-20 

9-26 

9-27 

9-28 

9-29 

10-9 

10-10 

10-11 

10-15 

10-19 

11-2 

11-lt 

11-5 

11-6 

11-7 

11-8 

11-10 

11-11 

11-12 

11-13 

ll-lU 

12-1 

12-2 



XIV 



I 1 1 I I IE I I I I I I i 1 i i i I i 



CSM 
SECTION 1 AS-501 



INTRODUCTION 
1.1 DRAWING SYMBOL STANDARDS 



1-1 



I 11 I I EI I I I I I i I i i i I i 



DRAWING SYMBOLS STANDARDS 



1 . ZONE 



1.2.1 



2 . LEADS 

A. CROSSOVER 



B. CONNECTED 



-X- 



3 . BATTERY 



NOMENCLATURE 
AMP HOURS 



^. BUS 



DESIGNATION 



\ 



-' SEE BUS DESIGNATION 
PAGE 1-12 



5. GROUNDS 

A. SYSTEM 



B. FRAME 



"TT 



6. CIRCUIT BREAKERS 



A. TOGGLE 



-^P- PNL 



NUMBER 



AMPS 



B. PUSHBUTTON 



J NUMBER 

H— o o 

■■ NAME ^W 

f PNL ^^ AM 



7. FUSES 

A. GENERAL 



-<5\^ 



B. FUSISTOR 



■<5XP^ 



1-2 



i 



I 1 i i I I I 1 I 1 I I I 1 i i i i I i 



DRAWING SYMBOLS STANDARDS 



8. DIODES 

A. GENERAL 

B . ZENER 

C. TUNNEL 



>h 



Wr 



* 



D. CONTROL RECTIFIER 



^ 



9. POTENTIOMETER 



vvv 



10. HEATER 



-nilr 



11. THERMOSTAT 



-nL^ 



12. ANTENNA 



NAME 
(TYPE OR FUNCTION) 



w 



13. EXPLOSIVE 



NOMENCLATURE 




PWR 
SOURCE 



li+. AMPLIFIER 



INVERTED INPUT 
INPUT 

DC SERVO WITH 
ISOLATED RETURN 






PWR INPUT 



OUTPUT 




DC, PRE, OR BUFFER 
AS INDICATED 



ER |\^ 



1-3 



I I I I I I I I I I I Mil ill 1 



DRAWING SYMBOLS STANDARDS 



13. GATES 

A. AND 



B. NAND 



C. OR 



D. NOR 







t. EXCLUSIVE OR 




15. TIME DELAY 



TD 



17. RELAY DRIVER OR SOLENOID DRIVER 



OR 



RD 
SD 




18. C0.^1MAND 




CONTENTS 

VEHICLE 

OPTION 



19. SENSOR 



A. SINGLE SOURCE 

PWR 
SOURCE 



TTTab 

1X1 RE F 



IN 



LETTER INDICATES TYPE -▼ 

PA - PRESS PS I A W = WETNESS 

PD = PRESS PSID POS = POSITION 

T = TEMP 

1 = QUANTITY, ETC. 

A - DELTA 



B. DCUBLE SOURCE 



PWR 
SOURCE, 



TM 



ON/BOARD 

J 



© 



VELOCITY GENERATOR 
FEEDBACK TRANSDUCER 



1-^ 




POS) POSITION FEEDBACK TRANSDUCER 



( 



I I 11 I IE I i I I I I 1 1 i i i I ! 



DRAWING SYMBOLS STANDARDS 



20. TELEMETRY 

A. MEASUREMENTS TELEMETERED 



TAB 
(OPTIONAL) 




MEASUREMENT NO, 
MEASUREMENT NAME 



DATA AVAILABILITY 

(OPTIONAL) 

PCM CODE 



B. ONBOARD METERS 




21. RELAYS 

A. MOMENTARY CONTACTS 



RANGE 
METER NO.""^"* ^ PNL 



C. B I -LEVEL CMD 



COMMAND TITLE 





COMMAND NUMBER 
£ BINARY INDICA- 
TION 

'instrumentation 
system reference 



"^^ 



CLOSED 



OPEN 



B. SOLID CONTACTS 



C. NON- LATCHING 
RELAYS SHOWN 
IN DE-ENERGIZED 
POSITION 




J 



NOMENCLATURE 
LOGIC NO. 



D. LATCHING 




NOMENCLATURE 
LOGIC NO. 



1-5 



I I I I I EI I I I I I I i i i M i 



DRAWING SYMBOLS STANDARD 



21. RELAYS (CONT'D) 



LATCHING RELAYS SHOWN 
IN RELEASED POSITION 



RELEASED 




LATCHED 



NOMENCLATURE 
LOGIC NO. 



22. SWITCHES 

A. IVIO POSITION 




SWITCH 
[NUMBER 



B. THREE POSITION 

r 1 

MOMENTARY 




SWITCH 
I NUMBER 



PNL 



C. BARO 




SHOWN AT 
SEA LEVEL 



22. SWITCHES (CONT'D) 
C. PUSH BUTTON 
1. SOLID 



Jl 



NAME 
PNL 



2 . MOMENTARY 

(A) T 



(B) 



NAME 
PNL 

A A 

NAME 
PNL 



D. COAX 



RF 




1-6 



1 I ill I EI I II I II I i i i I I 



DRAWING SYMBOLS STANDARDS 



22. SWITCHES 

E. PRESSURE 



1. CLOSED DECREASE PRESSURE 



A 



2. CLOSED INCREASED PRESSURE 



A 



F. MOTOR 

1. BREAK BEFORE MAKE 




2. MAKE BEFORE BREAK 




23. LIGHTS 

A. TELELIGHTS 



COLOR 


NAME 

J 


COLOR 

V 



^OLORj 



NAME 



NAME 
COLOR 
• PNL 



] 



r 

COLOR 


NAME 


COLOR 

V 








2i+. ANNUNCIATOR FLAG 



^^ 



N>flME 



(GENERALLY, THIS IS ASSOCIATED WITH A 
SWITH POSITION AND SHOULD BE LOCATED 
IN THE SAME BLOCK AS THE SWITCH,) 



1-7 



I I I I I E i I I I I I 1 I i i i I 1 



DRAWING SYMBOLS STANDARDS 



25. LINES 

A. GLYCOL 

B. HELIUM 



yiHp K. He ^ 



C. NITROGEN ^ N2 N2 N2 ^ 



D. HYDROGEN V H2 H2 H2 {j 



E . OXYG EN 9 02 02 02 h 



F. WATER yH2C) H20 ^2^;^ 



G. STEAM p STEAM STEAM ^ 



H. 


FUEL 


I . 


FUEL 




RETURN 


J. 


OXIDIZER 


K. 


OXIDIZER 




RETURN 


L. 


HYD HIP 



(IT T. X X X /! 



nz 



X- ' -x-* 



2 



M, HYD LOP J[ 



3 



26. PIPE CROSSOVER 



D 



£ 



3 



□ 



27. FILTER 



£ 



3 



28. BURST DIAPHRAGM 



'^=[C 



BURST VALUE 



29. VENTURI 




30. VENT 



31. FILL AND DRAIN 



< 



^] 



\ 



I I I I IE I i I 1 I I I I L i i i [ I 



DRAWING SYMBOLS STANDARDS 



32. 



A. MOTOR CONTROL 




B. MANUAL 
CONTROL 



£=D>^Ct3 



C. PYRO ISOLATIOI 

SZ 




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CM DEADFACE CHARGE NO. 2 

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SM DEADFACE CHARGE NO. 2 



35. ^ WIRE RESOLVER 




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CSM 


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AS-501 


kk. BUS DESIGNATIONS 




DC MAIN BUS A 


Vma 


DC MAIN BUS B 


Vmb 


BATi'ERY BUS A 


Vba 


BATTERY BUS B 


Vbb 


FLIGHT AND POST LANDING BUS 


Vfp 


BA'i'i'ERY RELAY BUS 


Vbr 


AC BUS NO. 1 


Vaci 


AC BUS NO. 2 


VaC2 


NON ESSENTIAL BUS 1 


Vnei 


NON ESSENTIAL BUS 2 


%E2 


SERVICE MODULE BUS A 


VSMA 


SERVICE MODULE BUS B 


Vsmb 


MESC PYRO BUS A 


%QA 


MESC PYRO BUS B 


Vmqb 


MESC LOGIC BUS A 


Vmla 


MESC LOGIC BUS B 


Vmlb 


RCSC LOGIC BUS A 


Vrla 


RCSC LOGIC BUS B 


VrLA 


RCSC PYRO BUS A 


Vrqa 


RCSC PYRO BUS B 


Vrqb 


ELSC LOGIC BUS A 


Vela 


ELSC LOGIC BUS B 


Velb 


ELSC PYRO BUS A 


Veqa 


ELSC PYRO BUS B 


Veqb 


EMERGENCY DETECTION SYSTEM BUS NO. 1 


Vedi 


EMERGENCY DETECTION SYSTEM BUS NO. 2 


VeD2 


EMERGENCY DETECTION SYSTEM BUS NO. 3 


VeD3 


SECS SWITCH LOGIC BUS 


VSL 


AOX BATTERY BUS A 


Vaa 


AUX BATTERY BUS B 


Vab 


SIGNAL CONDITIONING (5VDC) 


vsc 



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ELECTRICAL. 5ATURN 
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-INTERFACE - 
ELECTRICAL, SATURN 
LAUNCH VEHICLE 
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SECTION 2 AS-501 

SEQUENTIAL 



I ■ ^ 2.1 PYROTECHNICS, GENERAL NOTES 

2.1.1 Initiator, Standard Apollo 

Fire Current: The initiator will ignite when subject to a 

current of 3.5 amperes on the bridge wires. 
Operating Voltage: The operating voltage of the initiator is 

35 volts open circuit. 
Operating Time: The operating time is approximately 2 to i| 

milliseconds at operating voltage-. 



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2-1 



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CSM 
AS.5OI 



2,2 POWER DISTRIBUTION 

A. Switch Configuration 

1. The EDS POWER Switch should be placed in the ON position 
to supply power to the EDS "buses. 

2. The MESC LOGIC and PYRO Switches should be placed to the 
ARM (on) position; however, these functions are controlled 
by the MCP. 

B. The following circuit breakers are slugged: 

1. Pyro A Seq A RHEB-203 CBlU 

2. Pyro B Seq B RHEB-203 CB17 

C. Battery Buses A and B and the Entiy Battery .0 furnish power 
for the EDS displays and logic circuitry. 

D. Battery Bus A and B furnish power to the Master Events 
Sequence Controller A and B respectively. 

E. Pyro Batteries A and B furnish power for the detonation of 
pyrotechnic devices of System A and B respectively. 

F. The Master Event Sequence Control Pyro and Logic buses can be 
armed in one of the following manners: 

1. By GSE prior to launch: 

a. Logic Armed at T - 60 seconds 

b. Pyro Armed at T - Uo seconds 

2, The MCP will arm the respective bus as a result of RTC 61 
or by the G&N Computer commanding a SM/CM separation. 

G. The Master Event Sequence Control Pyro and Logic buses can be 
disarmed in one of the following manners: 

1. By GSE 

2. The MCP will "SAFE" the respective bus as a result of a 
LV-SC + 6 second discrete signal. 

3. The MCP will SAFE the logic bus at impact + 11 seconds 
and the Pyro at an impact + 12 seconds providing the CSM 
is in a Stable I configuration. 



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AS-501 



2.3 LAUNCH TO INSERTION NOTES 

2.3.1 General 

A. Switch Configuration 

1. Since the EDS is to he launched in an opened-loop con- 
figuration, the following switches should be placed in 
the respective positions: 



SWITCH 


POSITION 


MDC 16 




2 ENG OUT (AUTO/OFF) S8 


AUTO 


L/V RATES (AUTO /OFF) S9 


AUTO 


EDS (AUTO/OFF) S2 


OFF 


MDC 2k 





EDS POWER (ON/OFF) SI ON 

2. . The LES MODE/TVR JETT SPS MODE (S15-MDCl6) should be in 
the TWR JETT SPS MODE position. The functions of this 
switch are controlled by the MCP. 

B. The normally' closed first-motion relays (K3 System A; K2 
System B) remain energized by GSE until actual liftoff occurs. 
The liftoff enable relays (K65 - System A; k66 - System B) 
are latching types, which are latched closed by GSE prior to 
liftoff and latched open by the switch selector of the lU 
after the first 5 seconds of flight. 

C. The excessive Launch Vehicle Rate Parameter (BS0020) will be 
sent via telemetry when the launch vehicle rates are in excess 
of 20 + 0.5 degrees/second in roll and 5+0.5 degrees /second 
in pitch and yaw. 

D. The Guidance Fail Parameter (BSOOI6) will be sent via telem- 
etry to indicate failure of the guidance unit of the lU. 

E. The Launch Vehicle Engine Out partlmeters will be sent via 
telemetry when a respective launch vehicle engine develops 
less than 90 percent of total thrust output. 

2-1^ 



I I II I I I I I I I M I i i i I i 



( 



CSM 
AS-501 



F. The Q-Ball provides an electrical signal input to the L/V 
AOA/SPS PC indicator on MDC 3, and an electrical signal input 
to ground control via telemetry LSOOOl. The indicator is 
graduated to 150 percent which corresponds to a differential 
pressure of to i+ psid when selected for the L/V AOA function. 
The position of the red line (approximately 100 percent) on 
the indicator is based on vehicle structural limits and launch 
vehicle capahilities. 

G. At T+42 seconds when the launch vehicle has attained an 
altitude of approximately 10,000 feet, the automatic oxidizer 
d\jmp relay (Kl) times out. 

H. Max Q will occur at approximately 1 minute l8 seconds. If 

the launch vehicle rates become excessive throiigh this region 
an automatic abort will not be triggered by the EDS. 

I. The S-IC inboard engines will stop thrusting at approximately 
2 minutes, 23 seconds and k seconds later the outboard engines 
will cut off. 

J. Staging of the S-II will occur at approximately 2 minutes, 
31 seconds. 

K. The launch escape tower will be jettisoned by the S-IVB/IU 
at approximately 3 minutes h seconds. 

2.3.2 Launch Escape System Aborts 

A. The opening of the high and low altitude baroswitches defines 
the abort mode. The high altitude baroswitch controls auto- 
matic LET jettison, apex cover jettison, and drogue motor 
fire. This baroswitch is designed to close at 25,000 + 1000 
feet and open at 36,000 + 2000 feet; however; because of a 
venting lag it will not open until approximately U0,500 feet. 
The low altitude baroswitch is designed to close at 10,750 
+750 feet and open at 15 > 900 + 1900 feet. 

B. The MESC will attempt to shut down the booster on all launch 
aborts; however; the Instrumentation Unit of the launch vehicle 

2-5 



I 1 11 I I I I ! ! I I I 1 I i i i I i 



CSM 
AS-501 



will delay BECO for the first k2 seconds from liftoff. 

C. The launch escape motor fires for approximately 8 seconds. 

D. The canards will be deployed on all LES aborts following a 
time delay of 11 seconds after the initiation of the abort. 

E. When the Command Module descends to an altitude of approximately 
2U,000 feet, the SCS/RCS activate relays will be reset and the, 
tower jettison motor and leg bolts relay will be activated. 

2.3.2.1 T 0.0 to 1^2.0 Seconds (10,000 Feet) 

A. The capability exist for a low altitude LES abort from 
lambilical drop away through time T-U2 seconds. Upon receipt 
of the abort command, the Launch Escape Motor, the Pitch 
Control Motor, and the CSM tension tie pyrotechnics are 
initiated simultaneously separating the Command Module from 
the Service Module. 

B. The pitch control motor will burn in this abort mode only 
for a duration of approximately 0.6 seconds. 

2.3.2.2 T Greater than U2.0 Sec. 

A. The Command Module Reaction Control System/Stabilization 
Control System will be automatically sequenced to start 
functioning at 1.0 seconds following the abort initiation. 

B. Entry into more dense atmosphere will be accomplished with 
the LES attached to the Command Module with the Canards ex- 
tended. 

C. After approximately 1+2 seconds of flight the U2 second TD 
relays^ AUTO OX DUMP DISABLE (K1),< times out. 

2.3.2.3 SPS Aborts 

A. The capabilities of an SPS abort will exist after jettison 
of the LET. 



2-6 



I ill III! Ill I I i i i i I i 



CSM 
AS-501 



TABLE 2-1 TOWER ABORTS 



Time 


A Time 


Initiated 


Event 


Output 


Hr:Min:Sec 


(Sec) 


By 




To* 




0.00 


RTC 


Abort 


MCP 






MCP 


Separation /Abort Command On 


MESC 






MCP 


DSE Recorder ON 


TC 






MCP 


Inhibit k2 Second Timer in MCP 


MCP 






MCP 


Inhibit G&N Control Relays 


G&N 






MCP 


Heat Shield Instrumentation ON 


EPS 






MCP 


Inhibit Removal of Entry Batteries 
from Main Buses 


EPS 






MESC 


CM RCS Pressurization 


RCS 






MESC 


RCS Transfer (SM to CM) 


RCS 






MESC 


Initiate RCS Oxidizer Diamp 


RCS 






MESC 


CSM Deadfacing 


CSM 




0.10 


MESC 


CSM Separation Squibs Fired 


CSM 






MESC 


Ng Purge Valve Cutoff 


CSM 






MESC 


CSM Umbilical Destruct 


CSM 




1.00 


MESC 


SCS/RCS Enable ON 


SCS/RCS 




2.80 


MESC 


CSM Separation Pyros Cutout 


MESC 




11.00 


MESC 


Canard Deploy 


LES 




lit. 00 


MESC 


ELS Activate 


ELS 






MESC 


2UK Baroswitch Armed 


ELS 




0.00 


ELS 


2UK Baro Signal 


MESC 






MESC 


LES Jettison Motor Ignited 


LES 






MESC 


Fire Tower Explosive Bolts 


LES 






MESC 


Fire Tower Separation Linear Charges 


LES 






MESC 


SCS/RCS Enable OFF 


SCS/RCS 




o.Uo 


MESC 


Apex Cover Jettisoned 


CM 




2.00 


ELS 


Drogue Chutes Deployed 


CM 




10.50 


ELS 


Drogue Chutes Disreefed (U cutters) 


CM 




lU.OO 


ELS 


12K Baroswitch Armed 


ELS 



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TABLE 2-1 TOWilK ABORTS (Cont'd) 


CSM 
AS-501 




Time 
Hr:Min:Sec 


A Time 
(Sec) 

0.00 


Initiated Event 
By 




Output 
To 




ELS * 12K Baro Signal 




MOP 



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2.k NORMAL PREENTRY AND DESCENT 

1 ^ 

I 2.U.1 General 

A. Switch Configuration 

1. The ELS LOGIC Switch (Sl6, MDC 8) should he in the DOWN 
or OFF position. 

2. The ELS Switch (S15, MDC l6) should be in the AUTO position. 

B. A CM/SM separation signal will be generated by the Guidance 
and Navigation System sending a separation command to the MCP 
or by ground control sending RTC 6l. The Master Event Sequence 
Controller receives the separation command 1,0 seconds after 
initiation of the separation signal. 

C. After initiation of CM/SM separation, a -X Translation will 
' I ^e started immediately. This -X Translation will continue 

Tintil depletion of the SM/RCS propellants or until the SM 
batteries fail. At the same time the RCS control is transformed 
from the SM to the CM. After a time delay of 2 seconds, a 5.5 
seconds positive roll is initiated. 

D. The pyro power to the tension tie plates and the guillotine 
] is cut off 1.8 seconds after the initiation of separation. 

2.U.2 Earth Landing System 

A. With the sensing of 0.05 g's the MCP (2K20) will send an 

ELSC activate signal to the two and gates of the Earth Landing 
Sequence Controller; however, the and gate circuit is not 
.' completed until the closing of the 2i+,000 foot baroswitches. 

■ B. Closing of the 2ii,000 foot baroswitches in the ELSC applies 

voltage to the disarming coils of the SCS RCS activate relay 
since the ELS Switch (MDC l6) will be in the AUTO position. 
This signal also energizes redundant O.k second time delays 
(apex cover Jettison 2-second time delays (drogues lockout) 
— » and Ik second time delays (mains lockout). 

2-11 



I I I ! I ! I ! f f I f F II I [Mil 



CSM 
AS-501 



second time delays (apex cover jettison 2 second time delays 
(drogues lockout) and lU second time delays (mains lockout). 

C. The apex drogue parachute will be deployed approximately 
10 milliseconds before the apex cover is jettisoned. 

D. The two drogues are deployed in a reefed condition 1.6 
seconds after apex cover jettison. Within 9 seconds after 
deployment, the two drogues are disreefed. 

E. When the 10,000 foot baroswitches in the ELSC's close, the 
drogue parachutes are released and the three pilot parachutes 
are deployed, which in turn deploys the three main landing 
parachutes in a reefed condition. The main chutes are 
disreefed approximately 10 seconds following deployment of the 
pilot chute. 

F. The VHF recovery antennas and flashing beacon are deployed 
8 seconds after main chute deployment. 

G. Touchdown velocity on all three parachutes is approximately 
28 feet /second. If one chute fails to open the remaining 
two parachutes will' allow the CM to impact at approximately 
33 feet/second. 



2-12 



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AS-501 



2.5 LANDING 
I ! A. An impact signal is generated within the MCP in one of the 

following manners: 

1. The 12K foot baroswitches activate a ik minute time delay 
which in turn generates an impact signal. 

2. The 12K foot baroswitches activate a 20 second time delay 
which in turn supplies 28 Vdc to the impact switches. The 

I ! output of this impact, switch provides the same function 

as the 1^4 minute time delay. 

B. This impact signal supplies the necessary logic to disconnect 
the main chute and set up the Stable I, Stable II control logic, 

C. If the CM becomes inverted (Stable II), after landing, the 
MCP will turn on both pumps (by energizing 2K66) and start a 
sequence to fill the three float bag at five minute intervals, 
however, because of the internal wiring within the upright ing 
system all three float bags will fill upon receiving any of the 
commands from the MCP to fill. In the event that the CM 
uprights (stable l), prior to the completion of the fill cycle, 
the pumps will be shut off. The float bags will vent when 
the MCP is reset. 

D. The HF RECOVERY ANT will be deployed only in a Stable I 
c onf igurat ion . 

E. Each of the three float bags are k3 inches in diameter and has 
a capacity of approximately 2k cubic feet of air when inflated. 

F. A U.25 + psi relief valve is located in the inlet of each 
bag. Backup relief valves set at 13.5 psi are located in the 
outlet of each compressor. 

G. For safing of Logic and Pyro buses see 2.2-6. 



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SECTION 3 CSM 

AS-501 
ELECTRICAL 



3.1 FUEL CELL NOTES 

A. The fuel cells will be activated prior to launch and will 
remain on thoughout the mission until jettisoned at CM/SM 
separation. 

B. Any two of the three fuel cells can provide sufficient power 

to complete the mission. All three fuel cells will be on during 
normal operation. In case two fuel cells should malfunction, 
the remaining fuel cell will provide enough power, with all 
nonessential loads removed from the buses, to assure a safe 
return during an early mission termination from any point. 

C. The minimiim regulator inlet pressures for proper fuel cell 
operation are as follows: 

1, Oxygen Regulator 

power (with purge) - 75 psia 
IU20 watts (with purge) - 95 psia 

power (without purge) - 63 psia 
IU20 watts (without piirge) - 85 psia 

2, Hydrogen Regulator 

power (with purge) - 90 psia 
IU20 watts (with purge) - 98 psia 

power (without purge) - 60 psia 
li+20 watts (without purge) - 65 psia 

D. In case of a leak in the fuel cell's nitrogen system, the regu- 
lated pressure could drop as low as 23 psia and still maintain 
a high performance level. A 1 volt drop in performance can be 
expected at 23 psia. The performance drops sharply as the 
regulated N2 pressure drops below 23 psia. 

E. There are no scheduled inflight purges for the AS-501 mission, 
however, the ground stations have the command capabiltiy to 
perform purges should fuel cell performance degrade during 
the mission. Each purge command will purge both O2 and H2 
simultaneously for the particular cell. The nominal purge 
duration is 2 minutes. 



I I I I IE I I II I II 1 I i i I 1 



CSM 
AS-501 



F. Purge Rates: In addition to normal flow 

1. Oxygen - 0.6 Ib/hr/fuel cell 

2. Hydrogen - 0.75 lb /hr /fuel cell 

G. H2 Regulator - Regulates the H2 pressure to 8,5 psi above 
sensed N2 pressure. The regulated H2 pressure is 6O.O + 
2.0 psia. The regulator vents overboard at 0.2 to 0.i| 
psi back pressure. 

H. O2 Regulator - Regulates the O2 pressure to 11.0 psi above 
sensed N2 pressure. The regulated O2 pressure is 62.5 ± 
2.0 psia. The regulator vents at 0.2 to O.U psi back pressure. 

I. N2 Regulator - Regulates the N2 pressure to 51-5 ± 2.0 psia. 
The regulator vents into the SM at 2 to 3 psi above the 
regulator setting. 

J. Coolant Bypass Valve - When the temperature sensor detects 
low condensor exhaust temperature, the valve passes more 
glycol returning from the radiators through the coolant 
regenerator, thus raising glycol temperature and effecting 
a mean temperature- increase in the system. When the sensor 
detects 155°F or less, the valve gives full flow through 
the collant regenerator. When the sensor detects 170°F or 
above, the valve diverts full flow bypassing the coolant 
regenerator. 

K. H2 Bypass Valve - When the temperature sensor detects low 
fuel cell exhaust temperature, the valve passes more H2 
and water through the H2 regenerator, increasing the 
temperature of H2 and water going to the fuel cell. When 
the sensor detects U25°F or below, the valve gives full flow 
through the regenerator. When the sensor detects H95°F 
or above, the valve diverts full flow bypassing the regenerator. 

L. Fuel Cell Temperatures: 

Low - The critical temperature of KOH is 300°F. The KOH 

becomes a liquid above 300°F and the F/C electrochemical 
reaction can begin. The in-line heater should keep the 
3-2 



I I I I I I EI I I I I I I I i i i [ f 



CSM 
AS-501 



module temperature at or atove 385 + 5°F while the fuel 
cell is in a standby condition (open circuited) or at a 
lov pover level. Temperatures below 360*^F are abnormal 
for skin temperature. The fuel cell performance will 
degrade below 360*^F and the cell will be lost between 
360^ and 300°F. High - Temperatures above 500°F are 
abnormal for skin temperature. The teflon seals will break 
down between 500 and .550°F and cause a loss of the cell. 
The cell should be open circuited and allowed to cool to 
prevent its loss. Purging hydrogen gas is another method 
of cooling. 
M. Flow Rates 

1. Oxygen Consumption - 2.0U X lO'^ lb/amp-hr(lOO percent 
efficient cells) 

2. Hydrogen Consmption - 2. 57 X lO"*^ Ib/amp-hrdOO percent 
efficient cells) 

3. Water Production - 2.297 X 10 lb/amp-hr(lOO percent 
efficient cells) ' 

h. Glycol Flow - 35 to 80 Ib/hr (viscosity and temperature 
dependent) 
N. Purge Line Heater - The H2 purge line heaters will be turned 

ON prior to launch and left on for the mission duration. 
0. High pH - A high pH is normally caused by KOH entering into 

the H2/H2O loop. A pH of 9 or greater is abnormal. Any 

KOH entering the H2/H2O loop will normally cause a failure 

of the hydrogen pump. 



3-3 



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CSM 
AS-501 



3.2 DC SYSTEM NOTES 

A. The total spacecraft power source consists of three fuel cells, 
three entry batteries, three auxiliary batteries, two pyrotechnic 
batteries, and two SM jettison controller batteries. 

1. Fuel Cells (3) - The fuel cells supply the bulk of space- 
craft power from liftoff through CM/SM separation. One 
fuel cell will be connected to each main bus and the third 
fuel cell will be connected to both main buses. A malfunction- 
ing cell can be disconnected only through automatic circuit- 
ry (overload >T5 amps or reverse current >k amps for 
specific times)^ 

2. Entry Batteries (3) - The entry batteries supply backup 
power during SPS burns , primary power after CM/SM separa- 
tion, and help the axoxiliary supply the flight and post- 
landing bus after impact. The MCP perfoims the battery bus 
switching operations. At impact the entry batteries are 
switched onto the flight and postlanding bus. At impact 
plus 11 seconds the entry batteries are removed from the 
main buses. 

3. Auxiliary Batteries (3) - The auxiliary batteries supply 
power to the MCP inflight and to the flight and postlanding 
bus after impact. 

U. Pyrotechnic Batteries (2) - The pyro batteries supply power 
to initiate ordinance devices in the spacecraft. 

5. Jettison Controller Batteries (2) - The jettison controller 
batteries supply power to two jettison sequencers to sus- 
tain the SM RCS retrofire, as well as firing the SM posi- 
tive roll RCS engines 2 seconds after CM/SM separation. 

B. Battery Performance 

1. Entry and Auxiliary Batteries - These batteries will be 

charged at liftoff with 50 amp-hr. The battery open circuit 
voltage is 37.2 volts (inflight). 

3-5 



IIIIIMIIIII IIIHTT f I 



CSM 
AS-501 



2. Pyrotechnic and Jettison Controller Batteries - These 

batteries are rated at 0.T5 amp-hr. and 23 Vdc. The battery 
open circuit voltage is 37.2 volts (inflight). The 
pyrotechnic batteries are considered lost at 35 Vdc (open 
circuit) . 

C, Battery Relief Valves - Each cell of each entry battery contains 
a relief valve set to relieve at U5 ± 5 psid (amb. ref . ) . These 
valves vent into the battery case which in turn vents to the waste 
water and urine disposal line. The pyrotechnic batteries have 
relief valves in each cell set to relieve at 35 psid (amb. ref,) 
into the battery case. The case has a valve set to relieve at 

30 ± 5 psid (amb. ref.) into the surrounding environment and 
reseat at 20 psid (amb. ref.). 

D. Overload and Reverse Current Sensing Units - The fuel cells 
are connected to the main dc buses through the FUEL CELL 1 (2 
or 3) MAIN BUS A (or B) switches. The sensing circuit provides 
an automatic disconnect for an overload of 75 amps for 15 
minutes or 112 amps for 30 to 300 seconds and for a reverse 
current of h to 11 amps for 10 seconds or 11 to 20 amps for 1 
to 10 seconds. 



3-6 



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MAIN BUS B 
BAT 2 
CCCP 



MAIN BUS B 
BAT 3 
CCCP 

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BAT. 2 
CCCP 



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SYSTEM AND DISTRIBUTION 



AS -501 



3.2.1 



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CSM 
AS-501 



3.3 AC SYSTEM NOTES 

^ " 1 ^ ^ 

j A. The spacecraft ac power is generated "by three solid state 

inverter xinits . Each unit produces 115/200 Vac, 3;^, i+00 cycle 

power and is rated at 1250 volt amps with a life span of 1200 

hours . 

B. Operational Configuration - At liftoff and during normal flight. 
Inverter No. 1 will supply AC Bus 1 and Inverter No. 2 will 
supply AC Bus 2. Invertor No. 3 will be OFF unless some problem 
develops . 

C. AC Over-Undervoltage Sensing Unit - Six of these units, one for 
each phase, will sense an undervoltage (below 95 + 3 Vac) or 

an overvoltage (above 130 + 2 Vac) and will automatically 
. 1 disconnect the inverter from the ac bus after 9 seconds. The 

MCP will then replace the inverter (inverter No. 1 or Inverter 
No, 2) with the spare inverter (inverter No. 3). If the affected 
bus is AC Bus 1 then the essential ac loads will transfer to AC 
Bus 2. The transfer time is 0.5 seconds. For an undervoltage 
condition the MCP will attempt a reset before disconnecting the 
1 inverter. 

D. AC Overload Sensing Unit - This circuit monitors the rectified 
dc signal representing current output. The trip circuit will 
be activated when total 3^ inverter output exceeds 250 percent 
of rated current (approximately 27. T amps) or when current 
output on any one phase exceeds 300 percent of rated current 

— j (approximately 11 amps). Any overload condition (15 + 5 seconds 

for 3^ or 5 + 1 seconds for 1^) will cause the inverter to be 
removed from the bus. In this case the inverter would not ' 
be replaced and the affected bus would be lost. The MCP will 
not attempt a reset for an overload condition. For an overload 
condition on AC Bus 1 the essential ac loads will be switched 
I to AC Bus 2. 

3-8 



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CSM 
AS-501 



3.h CIRCUIT BREAKER SUMMARY 
MAIN BUS A 



EPS 



CRYO 



ECS 



CB NOMENCLATURE 



PANEL LOCATION CB NO. RATING 



POST LDG MAIN A 


RHEB-203 


cb6 


10 AMP 


INVERTER PWR NO. 1 MN A 


RHEB-203 


cbU 


70 AMP 


INVERTER PWR NO. 3 MN A 


RHEB-203 


CB2 


70 AMP 


FUEL CELL 1 PURGE 


MDC-22 


CB66 


5 AMP 


FUEL CELL 2 PURGE 


MDC-22 


CB85 


5 AMP 


FUEL CELL 3 PURGE 


MDC-22 


CB89 


5 AMP 


DC SNSR SIG MN A 


MDC-22 


CB15 


5 AMP 


BA'i"i'EHY CHARGER MN A 


MDC-22 


CB80 


5 AMP 


COUCH ATTEN MN A 
FLOODLIGHTS 


MDC-25 


CB38 


15 AMP 


CRYOGENIC SYSTEM TANK 
HEATERS 02 1 MN A 


MDC-22 


CB5 


15 AMP 


CRYOGENIC SYSTEM TANK 
HEATERS H2 1 MN A 


MDC-22 


CBT 


5 AMP 


ECS H2O ACCUM MN A . 


MDC-22 


CB9lt 


5 AMP 


ECS STEAM DUCT HEATERS 
MN A 


RHEB-206 


CB23 


5 AMP 



ECS TRANSDUCER WASTE 
& POT H2O MN A 



MDC-22 



CB92 



5 AMP 



ECS POT H2O HTK MN A 


MDC-22 


CB3 


5 AMP 


ECS TRANSDUCER PRESS. 


MDC-22 


CBU3 


5 AMP 


GROUPS 2 MN A 








ECS TRANSDUCER PRESS. 


MDC-22 


CB9 


5 AMP 


GROUPS 1 MN A 








BIOMED COMM MN A 


MDC-25 


CB59 


5 AMP 


EC TRANSDUCER TEMP 


MDC-22 


CBll 


5 AMP 


MN A 








INSTRUMENTATION 








CENTRAL TIMING SYS MN A 


MDC-22 


CB53 


5 AMP 


ESS INST MN A 


LEB 


CBjlt 


10 AMP 


BIOMED COMM MN A 


MDC-25 


CB59 


5AMP 



3-10 



1 I I III I i II I II 111 II 1 



\ 









CSM 




* 






AS-501 






CB NOMENCLAUTE 


PANEL LOCATION 


GB NO. 


RATING 


SEQUENTIAL 












EVENT TIMER MN A 


MDC-25 


CB5T 


5 AMP 


G&N 


G&N OPTICS MN A 


MDC-22 


CB55 


10 AMP 




G&N IMU MN A 


MDC-22 


CB59 


25 AMP 




G&N IIW HTH MN A 


MDC-22 


cb6i 


7.5 AMP 




G&N COMPUTEE MN A 


MDC-22 


CB5T 


10 AMP 


SCS 


SCS DIRECT CONT MN A 


MDC-25 


CBU2 


30 AMP 




SCS A&C ROLL MN A 


MDC-25 


CB36 


20 AMP 




SCS B&D ROLL MN A 


MDC-25 


CB3U 


20 AMP 




SCS PITCH 1 MN A 


MDC-25 


cbUg 


20 AMP 




SCS YAW 1 MN A 


MDC-25 


CB32 


20 AMP 




SCS GROUP : MN A 


MDC-25 


CBltU 


15 AMP 




SCS GROUP 2 MN A 


MDC-25 


CB65 


15 AMP 


SPS 


SPS GAGING MN A 


MDC-25 


CB22 


5 AMP 




SPS HE. VALVE MN A 


MDC-25 


CB28 


T.5 AMP 


RCS 


RCS GAGING MN A 


MDC-25 


CB18 


5 AMP 




RCS PROP ISOL 


MDC-25 


CBl6 


15 AMP 




MN A 










RCS C/M-5/M TRANSFER 


MDC-25 


CBIO 


15 AMP 




MN A 










RCS HEATERS B MN A 


MDC-21 


CB2 


5 AMP 




RCS HEATERS D MN A 


MDC-21 


CBU 


5 AMP 



3-11 



I III I i E i I I I I I I i i i I i i 



CSM 
AS-501 



MAIN 


BUS B 










CB NOMENCLATURE 


PANEL LOCATION 


CB NO. 


RATING 


EPS 


POSTLDG MAIN R 


RHEB-203 


CB5 


10 AMP 




INVERTER PWR NO. 2 
MN B 


RHEB-203 


CB3 


70 AMP 




INVERTER PWR NO. 3 
MN B 


RHEB-203 


CBl 


70 AMP 




FUEL CELL 1 PURGE 


MDC-22 


CB66 


5 AMP 




■ FUEL CELL 2 PURGE 


MDC-22 


CB85 


5 AMP 




FUEL CELL 3 PURGE 


MDC-22 


CB89 


5 AMP 




DC SNSR SIG MN B 


MDC-22 


CBll» 


5 AMP 




BAT'i'ilKY CHARGER MN B 


MDC-22 


CB79 


5 AMP 




COUCH ATTEN MN B 
FLOODLIGHTS 


MDC-25 


CB37 


15 AMP 


CRYO 


CRYOGENIC SYSTEM 
TANK HEATiilKS O22 MN B 


MDC-22 


CB6 


5 AMP 




CRYOGENIC SYSTEM 
TANK HEATiilRS H22 MN B 


MDC-22 


CBlt 


15 AMP 


ECS 


ECS H2O ACCUM MN B 


MDC-22 


CB95 


5 AMP 




ECS STEAM DUCT 
HEA'i'KHS MN B 


RHEB-206 


CB2lt 


5 AMP 




ECS TRANSDUCER 
WASTE & POT H2O MN B 


MDC-22 


CB91 


5 AMP 




ECS POT H2O HTR 
MN B 


MDC-22 


CB2 


5 AMP 




ECS TRANSDUCER 
PRESS . GROUPS 1 MN B 


MDC-22 


CB8 


5 AMP 




ECS TRANSDUCER 
PRESS. GROUPS 2 MN B 


MDC-22 


CB3U 


5 AMP 




ECS TRANSDUCER 
TEMP MN B 


MDC-22 


CBIO 


5 AMP 




BIOMKI) COMM MN B 


MDC-25 


CB60 


5 AMP 


INST 


& COMM 










COMM FT QUAL RCDR 


MDC-22 


CB96 


5 AMP 




CENTRAL TIMING SYS MN B 


MDC-22 


CB52 


5 AMP 



3-12 



I I i I E i i I I i I 1 I i I i I I i 



« 




CSM 
AS -501 




CB NOMENCLATURE 


PANEL LOCATION 


CB NO. 


RATING 


ECS ESS INST MN B 


LEB 


CB15 


10 AMP 


NGN ESS BUS MN B 


LEB 


CB13 


35 AMP 


TELECOMMUNICATIONS 
GROUP 3 


MDC-22 


CBl+7 


2 AMP 


COMM C-BAND 


MDC-22 


CB9T 


7.5 AMP 


SEQUENTIAL 









G&N 



SCS 



SPS 



RCS 



EVENT TIMER MN B 


MDC-25 


G&N OPTICS MN B 


MDC-22 


G&N IMU MN B 


MDC-22 


G&N IMU HTR MN B 


MDC-22 


G&N COMPUTER 


MDC-22 


SCS DIRECT CONT MN B 


MDC-25 


SCS A&C ROLL MN B 


MDC-25 


SCS B&D ROLL MN B 


MDC-25 


SCS PITCH 2 MN B 


MDC-25 


SCS YAW 2 MN B 


MDC-25 


SCS GROUP 1 MN B 


MDC-25 


SCS GROUP 2 MN B 


MDC-25 


SPS GAGING MN B 


MDC-25 


SPS HE. VALVE MN B 


MDC-25 


RCS GAGING MN B 


MDC-25 


RCS PROP ISOL MN B 


MDC-25 


RCS C/M-S/M TRANSFER 
MN B 


MDC-25 


RCS HEATKHS A MN B 


MDC-21 


RCS HEATERS C MN B 


MDC-21 



CB58 


5 AMP 


CB5U 


10 AMP 


CB58 


25 AMP 


cb6o 


7.5 AMP 


CB56 


10 AMP 


CBl+1 


30 AMP 


CB35 


20 AMP 


CB33 


20 AMP 


CB39 


20 AMP 


CB31 


20 AMP 


CBU3 


15 AMP 


CB6U 


15 AMP 


CB21 


5 AMP 


CB2T 


7.5 AMP 


CB17 


5 AMP 


CB15 


15 AMP 


CB9 


15 AMP 


CBl 


5 AMP 


CB3 


5 AMP 



3-13 



I III I i I II I I I I I L i i i I 1 



CSM 
AS-501 



BATTERY BUS A 





CB NOMENCLATURE 




PANEL LOCATION 


CB NO. 


RATING 


EPS 


BAT A PWR ENTRY 




T,EB-150 


CB21 


100 AMP 




AMIN A BAT BUS A 




RHEB-203 


CB12 


80 AMP 




POSTLDG BAT BUS A 




RHEB-203 


CB8 


25 AMP 




BAT RLY BUS BAT A 




MDC-22 


CB13 


20 AMP 




BATTERY CHARGER 




MDC-22 


CB78 


10 AMP 




BAT A CHGE 










SEQUENTIAL 












EDS 1 BAT A 




MDC-25 


CB52 


5 AMP 




MERC LOGIC A BAT A 


MDC-22 


CB21 


15 AMP 




ELS A BAT A FLOAT 


BAG 3 


MDC-25 


CB8 


5 AMP 




UPRIGHTING SYSTEM 


COMPR 


RHEB-205 


CBlt 


25 AMP 




NO.l 












MESC ARM A BAT A 




MDC-22 


CB19 


5 AMP 


SPS 


SPS GIMBAL MOTOR ■ CONTROL 


MDC-25 


CB26 


15 AMP 




PITCH 1 BAT A 












SPS GIMBAL MOTOR CONTROL 


MDC-25 


CB2U 


15 AMP 




YAW 1 BAT A 










BATTERY 


BUS B 











EPS 



SEQUENTIAL 



BAT B PWR ENTRY LEB-150 

MAIN B BAT BUS B RHEB-203 

POSTLDG BAT BUS B RHEB-203 

BAT RLY BUS BAT B MDC-22 

BATTERY CHARGER BAT B MDC-22 
CHGE 

EDS 3 BAT B MDC-25 

MESC LOGIC B BAT B MDC-22 

MESC ARM B BAT B MDC-22 

ELS B BAT B MDC-22 

UPRIGHTING SYSTEM COMP RHEB-205 
N0.2 



CB22 


100 AMP 


CB9 


80 AMP 


CB7 


25 AMP 


CB12 


20 AMP 


CB77 


10 AMP 


CB53 


5 AMP 


CB20 


15 AMP 


CB18 


5 AMP 


CB7 


5 AMP 


CB5 


25 AMP 



3-lU 



I 1 i I i i 1 i I I I I I III ill 



SPS 



CB NOMENCLATURE 

SPS GIMBAL MOTOR CONTROL 
PITCH 2 BAT B 

SPS GIMBAL MOTOR CONTROL 
YAW 2 BAT B 



CSM 
AS-501 



PANEL LOCATION CB NO. RATING 



MDC-25 



MDC-25 



BATTERY C 



CB25 



CB23 



15 AMP 



15 AMP 



EPS BAT CHGR BAT C 


LEB-150 


CB20 


5 AMP 


BAT C. PWR POSTLANDING 


LEB-150 


CB19 


100 AMP 


ENTRY 








EDS 2 BAT C 


MDC-25 


CB5lt 


5 AMP 


CB19 (BAT C PWR POSTLANDING ENTRY) 


RHEB-203 


CBll 




EPS MAIN A BAT C 


80 AMP 


MAIN B BAT C 


RHEB-203 


CBIO 


80 AMP 


POST LDG BAT C 


RHEB-203 


CB13 


25 AMP 


FLIGHT AND POSTLANDING BUS 








ECS POSTLANDING VENT 


MDC-25 


CB63 


5 AMP 


FAN PL BUS FLOAT BAG 2 








POSTLANDING FLOAT 


MDC-25 


CB68 


5 AMP 


BAG 1 PL BUS 








INSTR & COMM 








TELECOMMUNICATIONS 


MDC-22 


CBl+6 


7-5 AMP 


GROUP 1* 








TELECOMMUNICATIONS 


MDC-22 


CBl+5 


7-5 AMP 


GROUP 5 








PYRO BUS A 








SEQUENTIAL 








PYRO A SEQ A 


LEB-150 


CBllt 


20 AMP 


RCS PYRO A RCS FUEL DUMP 


LEB-150 


CB15 


20 AMP 


PYRO BUS B 








SEQUENTIAL 








PYRO B SEQ B 


LEB-150 


CBIT 


20 AMP 


RCS PYRO B RCS FUEL 


LEB-150 


CBl6 


20 AMP 


DUMP 









3-15 



I I 1 I i i I i I I I I 1 I I k i i I i 



CSM 
AS-501 



BATTERY RELAY BUS 





CB NOMENCLATURE 




PANEL LOCATION 


CB NO. 


•RATING 


EPS 


FUEL CELL 1 H2 & 
VALVE 


02 


MDC-22 


CB63 


10 AMP 




FUEL CELL 2 H2 & 
VALVE 


O2 


MDC-22 


CB82 


10 AMP 




FUEL CELL 3 H2 & 
VALVE 


O2 


MDC-22 


CB86 


10 AMP 




SNSR UNIT DC BUS 


A 


MDC-21 


CB5 


5 AMP 




SNSR UNIT DC BUS 


B 


MDC-21 


CB6 


5 AMP 




INVERTER CONTROL 


1 


MDC-22 


CB39 


10 AMP 




INVERTER CONTROL 


2 


MDC-22 


CB38 


10 AMP 




INVERTER CONTROL 


3 


MDC-22 


CB37 


10 AMP 




SNSR UNIT AC BUS 


1 


MDC-21 


CBT 


5 AMP 




SNSR UNIT AC BUS 


2 


MDC-21 


CB8 


5 AMP 




FUEL CELL 1 BUS CONT 


MDC-22 


CB65 


10 AMP 




FUEL CELL 2 BUS CONT 


MDC-22 


CB81* 


10 AMP 




FUEL CETiT, 3 BUS CONT 


MDC-22 


CB88 


10 AMP 


AC BUS NO. 


1 










EPS 


AC SNSR SIG AC 1 




MDC-25 


CBU8 


2 AMP 




FUEL CELL 1 CIR & 
MOTORS 


: SEP 


MDC-22 


CB6U 


3 AMP 




FUEL CELL 2 CIR & 
MOTORS 


SEP 


MDC-22 


CB83 


3 AMP 




FUEL CELL 3 CIR & 
MOTORS 


SEP 


MDC-22 


CB87 


3 AMP 


ECS 


GAS ANAL AC 1 




MDC-22 


CBII6 


2 AMP 




ECS CABIN AIR FAN 
1 AC Uk 




MDC-22^ 


CB76 


2 AMP 




ECS CABIN AIR FAN 
1 AC 1«5B 




MDC-22 


CBT5 


2 AMP 




ECS CABIN AIR FAN 
1 AC l^C 




MnC-22 


QSlk 


2 AMP 




ECS SUIT COMPRESSORS 
AC lik 


MDC-22 


CB33 


2 AMP 



3-16 



I ill I I 1 I I I I I 1 I ill I 1 



ECS 



CRYO 



CB NOMENCLATURE 

ECS SUIT COMPRESSORS 
AC 1 (6B 

ECS SUIT COMPRESSORS 
AC 1 «5C 

ECS RAD VALVE 
AC 1 lA 

ECS RAD VALVE 
AC 1 2B 

ECS GLYCOL PUMPS 
AC 1 (|)A 

ECS GLYCOL PUMPS 
AC 1 *B 

ECS GLYCOL PUMPS 
AC 1 $C 

CRYOGENIC TANK FAN 
MOTORS 1 AC 1 «I>A 

CRYOGENIC TANK FAN 
MOTORS 1 AC 1 *B 

CRYOGENIC TANK FAN 
MOTORS 1 AC 1 *C 

CRYOGENIC SYSTEM 
QTY AMPL 1 AC 1 *C 



INSTRUMENTATION & CCMM 



CSM 
AS-501 



PANEL LOCATION CB NO. RATING 
MDC-22 

MDC-22 

MDC-22 

MDC-22 

MDC-22 

MDC-22 

MDC-22 

MDC-22 

MDC-22 

MDC-22 

MDC-22 



CB32 


2 AMP 


CB31 


2 AMP 


CBTO 


2 AMP 


CB69 


2 AMP 


CB27 


2 AMP 


CB26 


2 AMP 


CB25 


2 AMP 


CBIOO 


2 AMP 


CB99 


2 AMP 


CB98 


2 AMP 


CBUlt 


2 AMP 





TKLECOMMUNI CATIONS 


MDC-22 


CB51 


2 AMP 




S-BAND 










TELECOMMUNI CATIONS 


MDC-22 


CB50 


2 AMP 




GROUP 6 










TELECOMMUNICATIONS 


MDC-22 


CBU9 


2 AMP 




GROUP 1 










TELECOMMUNICATIONS 


MDC-22 


CBU8 


2 AMP 




GROUP 2 










COMM C-BAND 


MDC-22 


CB120 


2 AMP 




XPNDR 








G&N 


G&H VIEWER AC 1 


MDC-22 


CB93 


2 AMP 


SCS 


SCS GROUP 1 AC 1 


MDC-25 


CBlt6 


2 AMP 




SCS GROUP 2 AC 1 


MDC-25 


CB67 


2 AMP 


SPS 


SPS GAGING AC 1 


MDC-25 


CB20 


2 AMP 



3-lT 



I 1 ill IE I I I I I M I i i i [ i 











CSM 




» 








AS-501 






CB NOMENCLATURE 




PANEL LOCATION 


CB.NO. 


RATING 


AC BUS 


NO. 2 










EPS 


FUEL CELL 1 CIR & 
MOTORS 


SEP 


MDC-22 


CB6U 


3 AMP 




FUEL CELL 2 CIR & 
MOTORS 


SEP 


MDC-22 


CB83 


3 AMP 




FUEL CELL 3 CIR & 
MOTORS 


SEP 


MDC-22 


CB87 


3 AMP 




AC SNSR SIG AC 2 




MDC-25 


cbUt 


2 AMP 


ECS 


ECS CABIN AIR FAN 
2 AC 2 «$A 




MDC-22 


CB73 


2 AMP 




ECS CABIN AIR FAN 
2 AC 2 {5B 




MDC-22 


CB72 


2 AMP 




ECS CABIN AIR FAN 
2 AC 2 ?$C 




MDC-22 


CB71 


2 AMP 




ECS SUIT COMPRESSORS 
AC 2 (z5A 


MDC-22 


CB30 


2 AMP 




ECS SUIT COMPRESSORS 
AC 2 (^B 


MDC-22 


CB29 


2 AMP 




ECS SUIT COMPRESSORS 
AC 2 iC 


MDC-22 


CB28 


2 AMP 




ECS RAD VALVE 
AC 2 IB 




MDC-22 


CB68 


2 AMP 




ECS RAD VALVE 
AC 2 2A 




MDC-22 


CB67 


2 AMP 




ECS GLYCOL PUMPS. 
AC 2 (/>k 




MDC-22 


CB2lt 


2 AMP 




ECS GLYCOL PUMPS 
AC 2 ?5B 




MDC-22 


CB23 


2 AMP 




ECS GLYCOL PUMPS 
AC 2 ^Q, 




MDC-22 


CB22 


2 AMP 


:ryo 


CRYOGENIC TANK FAN 
MOTORS 2 AC 2 ?5A 




MDC-22 


CB103 


2 AMP 




CRYOGENIC TANK FAN 
MOTORS 2 AC 2 «5B 




MDC-22 


CB102 


2 AMP 




CRYOGENIC TANK FAN 
MOTORS 2 AC 2 ^C 




MDC-22 


CBlOl 


2 AMP 




CRYOGENIC SYSTEM 
QTT AMPL 2 AC 2 ^C 




MDC-22 
3-18 


CB90 


2 AMP 



1 III M i I I I I I I 1 i i i [ 1 









CSM 




* 






AS-501 






CB NOMENCLATURE 


PANEL LOCATION 


CB NO. 


RATING 


INST 


& COMM 










TELECOMMUNICATIONS 


MDC-22 


CB51 


2 AMP 




S-BAND 










TELECOMMUNI CATIONS 


MDC-22 


CB50 


2 AMP 




GROUP 6 










TELPICOMMUNI CATIONS 


MDC-22 


CB49 


2 AMP 




GROUP 1 










TELECOMMUNI CATIONS 


MDC-22 


CBU8 


2 AMP 




GROUP 2 










COMM C-BAND 


MDC-22 


CB120 


2 AMP 




XPNDR 








G&N 


G&N VIEWER AC 2 


MDC-22 


CB62 


2 AMP 


SCS 


SCS GROUP 1 AC 2 


MDC-25 


CBl+5 


2 AMP 




SCS GROUP 2 AC 2 


MDC-25 


GB66 


2 AMP 


SPS 


SCS GAGING AC 2 


MDC-25 


CB19 


2 AMP 



3-19 



I I 1 i I I I i I 1 I I I 1 I i i i I i 



3.5 AC EQUIPMENT SUMMARY 



CSM 
AS-501 



COMPONENT 

F/C 1 COOLANT PUMP 

F/C 1 H2 PUMP 

F/C 1 pH SENSOR 

F/C 2 COOLANT PUMP 

F/C 2 Hg PUMP 

F/C 2 pH SENSOR 

F/C 3 COOLANT PUMP 

F/C 3 H2 PUMP 

F/C 3 pH SENSOR 

SUIT COMPRESSOR NO.l 

SUIT COMPRESSOR NO. 2 

CABIN AIR FAN NO.l 

CABIN AIR FAN NO. 2 

WASTE MGT BLOWER 

CABIN TEMP CNTL BOX 

GLYCOL PUMP NO.l 

GLYCOL PUMP NO. 2 

GLYCOL EVAP BACK PRESS. CNTL 

GLYCOL TEMP CNTL 

GLYCOL EVAL H2O FLOW CNTL 

AC SENSING UNIT 

AC VOLTMETER SELECT SW 

ECS RAD lA ISOL VLV 

ECS RAD IB ISOL VLV 

ECS RAD 2A ISOL VLV 

ECS RAD 2B ISOL VLV 

CRYO Og TANK NO.l FANS 

CRYO O2 TANK NO. 2 FANS 



PHASES USED 
(NORMAL OPERATION) 


BUS USED 


3?5 


AC1/AC2 


-hi 


AC1/AC2 


ik 


AC1/AC2 


U 


AC1/AC2 


3?5 


AC1/AC2 


<hk 


AC1/AC2 


U 


AC1/AC2 


3^ 


AC1/AC2 


M 


AC1/AC2 


3(z5 


AC1/AC2 


395 


AC1/AC2 


3^ 


ACl 


3i> 


AC2 


3«5 


ACl 


«$C 


AC2 


3«5 


AC1/AC2 


3i> 


AC1/AC2 


TL iZ 


ACl 


M 


ACl 


M 


AC2 


3i 


AC1/AC2 


3i 


AC1/AC2 


M 


ACl 


M 


AC2 


iz 


AC2 


$& 


ACl 


34, 


ACl 


34> 


AC2 



3-20 



I III I ill I I I I i 111 M 1 



* 








CSM 
AS-501 


COMPONENT 




PHASES USED 
(NORMAL OPERATION) 


BUS USED 


CRYO H TANK 


NO. 1 FANS 




34. 


ACl 


CRYO H TANK 


NO. 2 FANS 




B* 


AC 2 


CRYO TAUK 
QTY UNIT 


NO. 1 FANS 




<|)C 


ACl 


CRYO TANK 
QTY UNIT 


NO. 2 FUEL 




<}>C 


AC2 


CRYO H TANK 
QTY UNIT 


NO. 1 FUEL 




(frc 


ACl 


CRYO TANK 
QTY UNIT 


NO. 1 FUEL 




<t>c 


AC2 


TELECOM PCM 


& DATA STORAGE EQUIP 


s* 


AC1/AC2 


TELECOM SIG COND EQUIP 
& S-BAOT) PWR AMPL 




3(t) 


AC1/AC2 


PREMOD PROCESSOR & UNIFIED 
S-BAND (GR. 2) 


3^ 


AC1/AC2 


FLIGHT QUAL RCDR & C-BAND 
XPONDER & VHF/FM XMITTEH 


(GR. 1) 


3* 


AC1/AC2 


G&N VIEWER ] 


LIGHTING 




<|>B 


AC1/AC2 


SCS GROUP 1 






3* 


AC1/AC2 


SCS GROUP 2 






3<0 


AC1/AC2 


SCS GAGING UNIT 




(t'C 


AC1/AC2 


GAS ANALYZER 




<}>A 


ACl 



3-21 



I I I I 1 i i I I I I I I I I i i i [ i 



^ SECTION 1+ ?!^.^. 

* AS-501 

ENVIORNMENTAL 

I i+.l CM PRESSURIZATION 

i+.l.l Cabin Pressure Control 

A. Cabin Pressure Relief Valve 

1 . Conf igurat ion 

a. Two independent pressure relief valve sections with 
manual overrides. 

b. Either section will outflow to limit cabin-to-ambient 
AP (positive pressure relief). 

c. Either section will inflow to limit ambient to cabin 
AP (negative pressure relief). 

2. Positions (LHEB 30?) 

a. CLOSED - Will not open automatically. 

b. NORMAL - Automatic with, full flow. 

c. BOOST and ENTRY - Automatic with full flow. 

d. DUMP - Manually open with full flow (one relief valve 
does not have dump position). 

3. Positive Pressure Relief 

+0 2 
a. 6.0 _^* ^ psi cabin-to-ambient AP. 

U. Negative Pressure Relief 

a. 25 inch H^O ambient-to-cabin AP maximum. 
B. Cabin Pressure Regulator 

1 . Conf igurat ion 

a. Two flow-limited, absolute pressure, demand regiilators 
operating in parallel and inflowing to cabin. 

b. Regulators maintain cabin at 5 +, 0.2 psia and shutoff 
during depressurized cabin (pressure below 3.5 psia). 

2. (LHEB 3lU) Cabin Repress 

a. Manual control bypasses regulators to repressurize 
cabin after emergency. 

3. Flows 

a. Pressure regulators: l.i; Ib/hr 0^ maximum (flow 
limited) through both regulators at 3.5 to 5.2 psia 
cabin (0.7 Ib/hr O2 flow through each regulator). 
U-1 



I 1 1 1 1 11 1 III 1 1 1 1 i II 1 



CSM 
AS-501 



b. Repressurization Valve: 0.1 lb /minute. ^ 

C. ::^irergency O2 Inflow Control Valve 

1. Valve Position (Emergency Cabin Press. )(LHEB 31^) 

a. "1" - Flow to regulator 1 only. 

b. "2" - Flow to regulator 2 only. 

c. NORMAL - Flow to regulator 1 and 2. 

d. OFF - Shuts OFF flow to both regulators. 

2. Flow Capacity 

a. Limiter in Oxygen Pressure Regulator limits emergency O2 
flow to 0.6T lb/minute 0^ with U.2 to U.6 psia cabin. 
(Either valve can flow 0.6T lb/minute.) 

D. Oxygen Pressure Regulator 

1. Valve Positions (Main Regulator) LHEB 31^ 

a. "1" - Flow to 'regulator 1 only. 

b. ''2" - Flow to regulator 2 only. 

c. NORMAL - Flow to regulator 1 and 2. 

d. OFF - Shuts OFF SM and CM supply to both regulators 

2. Flow Capacity 

a. Limiter at outlet of O2 Pressure Regulator limits flow 
to G.6T lb /minute Oo- (either regulator can supply 0.67 
lb /minute. ) 

3. Regulated Pressure 
a. 90 to 110 psig. 

U. Relief Valve 

a. With failed open regulators, limits downstream pressvire 
to 130 ± 10 psig. 
E. Water and Glycol Tank Pressure Regulator Valves (LHEB 3lU) 

1. Valve Position 

a. "1" - Flow to regulator 1 only. 

b. "2" - Flow to regulator 2 only. 

c. NORMAL - Flow to regulator 1 and 2. 

d. REGULATOR OFF - Shuts OFF flow to both regulators. 

2. Regulated Pressure 
a. 20+2 psig. 

k-2 

I I I I I i II I I I I I i L i i i [ I 



CSM 
^ AS-501 

I F. Water and Glycol Tank Pressure Relief Valves (LHEB 3lii) 

1. Valve Position 

a. "1" - Relieves from 1 only. 

b. "2" - Relieves from 2 only. 

c. NORMAL - Relieves from 1 and 2. 

d. RELIEF OFF - Will not relieve. 
F 2. Relief Pressure 

a. 25+2 psig. 



U.3 



I I I ill i I 1 I I I i i i i i I i 



CSM 
AS-501 



1+.1.2 Water Supply 

A. Potable Water Tank 

1. Configuration 

a. Cylindrical tank with pressurized bladder for water 
expulsion at zero "g". 

b. Transducer for water quantity consists of potentiometer 
driven by a string attached to bladder. 

2. Storage Capacity 
36 ^Q lb H2O 

3. Expulsion Gas Pressure 
18 to 27 psig 

k. Water Delivery Rate 

5 lb /minute minimum at I8 psig expulsion pressure, 

B. Waste Water Tank 

1. Configuration 

Same as potable water tank. 

2. Storage Capacity 
56 ^Q lb H2O 

3. Expulsion Gas Pressure 
18 to 27 psig 

h. Water Delivery Rate 

11 lb /minute minimum at I8 psig expulsion rate. 

C. Waste Tank Inlet Valve (LHEB 315) 

1. Valve Position 

CLOSE - Prevents flow of fuel cell or SM water to Waste 
Water Tank. 

AUTO - Allows fuel cell or SM water to flow into Waste 
Water Tank. 

2. Crack and Reseat 

Crack, full flow and reseat with U.5 to 6.5 psi AP range. 

3. Flow Capacity 



12 Ib/hr H2O minimum. 



I I I I I I i i I I 1 I I I I k i i I 1 



CSM 
AS-501 



Water Pressure Relief Valve (LHEB 315) 
1, Valve Position 

a. "1" - Allows overboard venting or dumping of H2O 
through relief valve 1 only. 

b. "2" - Allows overboard venting or dumping of H2O 
through relief valve 2 only. 

c. BOTH - Allows overboard venting or dumping of HpO 
through relief valves 1 and 2. 

d. OFF - Prevents overboard venting or dumping of HpO. 

2. Flow Capacity 

0.3 gpm H2O minimum at hQ psi AP. 

3. Reseat 

i+0 psi AP minimum 



h-3 



I III I I I I 1 I I 1 1 i i ill I 



I ! 



■ 1 



ff I 



I 1 



I I 



11 



II 



1 1 1 1 1 1 1 I 

SSaSSstsa 

A & £. i. ^. i ^ A 



i! 



■ ■111 



'it 






1 




3 11!' 



Ill 



f Z t 1 



I I ¥ 



Y^Y^T 



i ]L -1 



w m 



) . 



CSM 
AS- 5 01 



k.2 COOLANT NOTES 

U.2.1 General 

The ECS radiator is not connected into the water-glycol coolant 
loop for Mission AS-501. GSE will supply the cooling during ground 
operation, and evaporative cooling will "be employed during flight. Both 
glycol piomps will he turned on prior to hatch closeout for redundancy. 
PiJinp 1 will be turned on by closing Circuit Breakers 25 through 27 on 
MDC 22. P\amp 2 will be turned on by closing Circuit Breakers 22 through 
2k and setting ECS GLYCOL PUMP Switch on MDC 21 to PUMP 2 - ac 2. 

The following events will be accomplished by the MCP at the 
specified times indicated. 

A. Start ECS Operation (T-T5 seconds) 

1. Steam Pressure Control to AUTO by closure of 2K56 in MCP. 

2. Glycol Shutoff Valve closed by closure of 2K8 in MCP. 

3. If there is a hold after T-T5 seconds, the Glycol Shutoff 
valve will be opened by a COUNTDOWN RESET command through 
GSE. 

B. LET JETT (Approx l85 seconds) 

1. Glycol Wetness Control START by closiire of 2K6 C and D 
in MCP. 

C. CM/SM SEP 

1. 0^ Isolation Valve CLOSED by 2KT in MCP. 

2. Glycol Shutoff Valve CLOSED by 2K8 in MCP. 

U.2.2 Coolant Circuit Notes 

A. Glycol Evaporator 

1. Heat Transfer Rate 
7620 Btu/hr maximum 

2. Water Evaporation Rate 
7«5 Ib/hr maximiJm 

B. Glycol Reservoir 

Capacity - 9.76 lb water-glycol. 



h-l 



I I I I ill i I I I I I I I i i I 1 



CSM 
AS- 501 



C. Glycol Pump 

Flow capacity - 200 Ib/hr water-glycol at lOO^F inlet and 29-5 

psi minimum pressure rise. Inlet pump pressure 7-5 j; 1*5 psig. 
D- Accumulator 

Quantity - 1.36 lb water-glycol maximum • 
E. Cabin Heat Exchanger 

Cooling Heat Transfer AUTO 1250 Btu/hr. 

NOTE 

Total water-glycol quantity in 
system approximately 20 pounds. 

1+.2.3 Glycol Evaporator Temp Control Subsystem 

The glycol evaporator is designed to maintain the temperature 
output of the evaporator between UO to 50.5°F. 

Activation of the steam pressure control section takes place when 
the input to the evaporator from the glycol pumps rises to the U8°F to 
50.5°F range. The evaporator outlet temperature sensor supplies the 
controlling signal to the steam pressure control section for positioning 
the steam pressure control valve. Upon command from the control lonit , the 
valve is opened a specific amount required to regulate steam duct pressure. 



U-8 



I I I I I I I I i I I I I 1 i i i i I i 



CSM 
AS- 5 01 



Initial movement of the valve out of the CLOSED position, positions a 
switch within the valve assembly that activates the water control section. 
After LET JETT, this unit automatically controls the amount of water 
admitted to the evaporator hy the solenoid-operated water inlet control 
valve. The water control unit activates the valve in response to signals 
received from the evaporator wick temperature sensor and the reference 
glycol inlet temperature sensor. ON-OFF cycling of the water inlet control 
valve and steam pressure control valve continues as long as evaporative 
cooling is required. 

Steam pressure indications should read between 0.125 and 0.135 
for low heat loads, and between 0.097 and 0.106 for high heat loads. 

U.2.i+ Cabin Temperature Control Valve 

The cabin temperature control valve is a motor driven modulating 
type, consisting of two mechanically-linked, three-way valves, and a 
manual override. The motor produces rotary mechanical motion for posi- 
tioning the three-way valves, which function as inflow and outflow 
valves for the cabin heat exchanger. 



CONDITION 


FULL FLOW 


ZERO FLOW 


MAX COOLING 


A&C 


B&D 


MAX HEATING 


B&D 


A&C 


NEITHER COOLING 
OR HEATING 


B&C 


A&D 



1+-9 



I I I 1 I I I I I I I I I ( LI i i i 



VAC2 



ECS-GLYCOL 
PUMP-AC 2 
♦ A 

(MDC-22) 
CB24 



. CB3 

VmA 11 (J^^b— I 

JP. ECS 

' DOT U_rt 



S22 



POT H2O 

HTR-MN A 

(MDC-22) 



<H 



GLYCOL 

EVAP-H2O 

FLOW 

(MDC-13) 



• MAN 
^•OFF 



• AUTO 



ECS- 
POT H2O 
HTR-MN B 
(MDC-22) 
CB 2 



ecs-glycol 

PUMPS-AC 

l-<t>C 

(MDC-22) 



HI 



^ 



2K6 



E 






GLYCOL EVAP. 
STEAM PRESS 

tMDC-13) 
S24 

TO DC 
RETURN 



MANi 



INCR 

▼ — 



ECRl 



DECR 



OUTNESS) 
IC WATER CONTROL 
— SECTION 



n a 



EVAP TEMP 
CONTROL UNIT 



2K36. 




0825 



S23 



HI 



Vac I 



GLYCOL EVAP. 
STEAM PRESS 
(MDC-i3 



-^ 



STEAM PRESSURE 
CONTROL 
SECTION 



J 



GLYCOL TEMP 
SENSOR ffy 



WICK TEM 
SENSOR Q_ 



FROM: 

..ASTE 5 TTTT 

.vATER '^ 
NETWORK 



J 

I 



WATER INLET 
CONTROL VALVE 



t^ 



TK 



i 



WATER- 

GLYCOL 

EVAPORATOR 



A 



GLYCOL 

TEMP SENSOR ^ 

■■■■■■■■ 



WATER 
CONTROL 
ACTIVATION 
SWITCH 



*»>? 



ZJ 






■ ■ ■ k GLfivL 



JEL 



GLYCOL TEMP 
SENSOR 



jTEAM 




lOVER OARD 



STEAM PRESS 
CONTROL VALVE 



■1} 

OUTLET 



FIGURE 4.1 WATER-GLYCOL TEMPERATURE CONTROL-DIAGRAM 



I 1 I I I i E I I II III I I i i [ I 



n 



SM I I CM 



nipjf^^liJ^Rj firiTi i iiii i ■■ ■■■■■■■■■ 



4 PASS 
SECTOR n 



b PASS 
SECTOR n 



* PASS 
SECTOR V 



b PASS 
SECTOR V 



VENT 



GLYCOL 
PRESSURE 
RELIEF 
VALVE 







-OSAO- 
POTH2OHTR 5 
MDC 22 I 



082 



I* 



Tjl 



I 



I S34 MDC-13 J 



I CB6S 1 

— I— 2A0 — 
I ECS RAD 
I MDC-22 



2KBB 2K8A 



I 



GLYCOL 

Si-;uroFF 



r 



I I I I I I I ! I I I T I m [ T T F I 



f 




STEAM 

VENT 

OVERBOARD 



III 1 1 nil imiii iTTf I 



rm 



Ilt« | ;ome|~ 



"J^l 



I 




I * GLYCOL PUMP I 
' ! MOC22 I 



1 




. ECS GLYCOL 

I PUMP 

I SI4 MDC 21 J 



■ ■■■■■■■■■■■■■■■■■■rm 



CFo.;o:^ 

CAIN PRESSURE 
TRANSDUCER 



cfooo:t 

CA m TEMP 
TRANSDLCER 



CFOOI-T I 

GLYEVAPOUT TEMP 1 
TRAr.Si/bCER 



CF00016P 

CLV PUMP OUT PRESS 

TRAMSDOCER 



CF0019Q 

CLY ACCUM QTY 
TRAHSDUCtR 



[J> SEE HOTES: 
^ SEE NOTES: 
^ SEE NOTES: 
^ SEE NOTES; 
^ SEE MOTES: 
^ SEE NOTES: 
[^ SEE NOTES: 



4.Z.ZA 

4.2.2B 

4.2,20 

4. 2 .2D 

4.2 .2E 

4.2,3 

4.2.4 



ECS 



J CF0034P 

GLYEVAPBACK PRESS 
H TRANSDUCER 



-^^. 






*^ t^u'-^Jt/ ii 



'^'"^^ 



NATIONAL AERONAUTICS i SPACE ADMINISTRATION 
MANNED SPACECRAFT CCNT6R • HOUSTON. TIEXAS 



GLYCOL COOLING 



rsMi 



vS - 5 I 



42.1 



1 1 1 1 1 1 1 1 f ! I r 



[ TTl riT 



f I 



n 



T - ^ 



H 



n 



CSM 
* SECTION 5 AS-501 

CRYOGENICS 

5.1 NOTES, CRYOGENICS i GENERAL 

A. Hydrogen tanks will be loaded to \5 percent of maximum, 
(approx 13.0 lbs) 

B. Oxygen tanks will be loaded to 33 percent of maximum, 
(approx 105 lbs) 

C. Real-Time commands can be sent to turn on the heaters and fans 
in the 0^ ^^^ ^2 "^Q^s. The commands bypass the heater and fan 
switches ( MAN-OFF-AUTO ) and apply power directly to the heaters 
and fans until a reset command is sent. The following is a list- 
ing of the Real-Time Commands for cryogenics: 

RTC 6U H2 No. 2 Heater-Fans 

RTC 65 02 No. 2 Heater-Fans 

RTC SS H2 No.l Heater-Fans 

RTC 67 02 No.l Heater-Fans 

RTC 70 .Reset RTC 6i+-67 

D. For AS-501 the O2 and H2 heater switches on MDC 13 will be turned 
OFF, and commanded ON ^.s required. The O2 and H2 Fan switches on 
MDC 13 will be turned to AUTO. 

E. Cryogenic heaters thermo-switches in the Hydrogen Tanks, open to 
shut off heaters when the sensed temperature is 90°F or above. 
The thermo-switches close to turn heaters on when the sensed 
temperature is 70°F or below. Sensed temperature is that of the 
inner wall of the Hydrogen Tank. 

F. Cryogenic heaters thermo-switches in the Oxygen Tanks, open to 
shut off heaters when the sensed temperature is 90°F or above. 
The thermo-switches close to turn on heaters when the sensed 
temperature is 70°F or below. Sensed temperature is that of 
the inner wall of the Oxygen Tank. 

G. The Hydrogen Tanks pressure switches close when the sensed pressure 
is 230 psia or below and open when the sensed pressure is 260 
psia or above. Both pressure switches must be closed to operate 
Hydrogen Tanks fans. 

5-1 



mil I ! I ! ! f r f IT i ITT IT 



H. The Oxygen Tanks pressure switches close when the sensed pressure 
is 865 psia or "below and open when the sensed pressure is 935 
psia or above. Both pressure switches must be closed to operate 
Oxygen Tanks fans. 
I. H2 Flow Rate for Fuel Cells. Hg Ib/hr = 2.57 x 10"3 x I (ampere) 
J. Og Flow Rate for Fuel cells. O2 Ib/hr = 20.56 x 10-3 x I (ampere) 



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SECTION 6 ?5^.^n 

AS-501 

COMBJUNICATIONS 



6.1 NOTES, COMMUNICATIONS SYSTEM BLOCK DIAGRAM 

A. Unified S-Band. System Equipment 
I W •^' ^^^ Ranging 

1 2. Downlink Voice (1|00 HZ Tone) 

3. PCM HBR and LBR Data (Real-Time Only) 
k. UDL Updata 

B. VHF AM Transceiver Equipment 

1. Downlink Voice (iiOO HZ Tone) 
, C. VHF FM Transmitter Equipment 

1. PCM HBR Data (Real-Time Only) 

2. PCM LBR Data (Real-Time Only) 

D. C-Band Transponder Equipment 

1. Tracking 

E. Updata Link Equipment 
[ ^ I 1. UHF FM Updata 

2. S-Band Updata 

F. HF Transceiver Equipment 
1. Recovery Beacon (CW) 

G. VHF Recovery Beacon Equipment 
1. Recovery Beacon 

^ H. VHF Survival Beacon Equipment 

1. Recovery Beacon 
I. See Table 6-6 for Updata Link Controlled Functions and Table 
6-17 for MCP Automated Control Fxinctions which operate 
Commiinication Equipment. 
J. See Drawings 11.2.1 thru 11.2.5 for all real-time commands to 
I the MCP. 



6-1 



11 1 III ! I ! ! f f I nr IT T f I 



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■ ■ ■ t I 



ill*. 



Z X 7 



Iff 



I I 113 



r T T 



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D M :i 

& A. A. 



VHF RECOVERY 
ANTENNAS C2)' 
S-BAND ANTENNA 
NO. 2 (LOWER) 

HF RECOVERY 
ANTENNA- 
C-BAND 
ANTENNA NO. 3 

-Y 
= 180' 

VHF ANTENNA 
(UPPER) 

S-BAND ANTENNA 
N0,3 (UPPER) 

C-BAND ANTENNA NO. 2 



VHF RECOVERY 
ANTENNAS 




S-BAND ANTENNA NO.l 
(UPPER) 

VHF ANTENNA (UPPER) 
FLASHING LIGHT 
C-.BA^D 
ANTENNA NO.l 



RECOVERY INTERCOM 
SWIMMER CONNECTOR 

S-BAND ANTENNA NO. 4 

(LOWER) 



= 270** 



HF RECOVERY 
ANTENNA (14F) 





VIEW (7) 
VHF ANTENNAS (2) 



Xs = 200 




Xs = 355 



XS = 307 



VHF ANTENNAS (2) 



VIEW (C) 
S-BAND ANTENNAS (4) 



C-BAND ANTENNAS (4) 
FIGURE 6.1 ANTENNA LOCATIONS 



I I I I ill I I I I I 1 I ill I I 



CSM 
AS-501 



6.2 COMMUNICATIONS SYSTEM EQUIPMENT 

6.2.1. C-Band Transponder and Antenna Equipment 

The G-Band Transponder is used in conjunction with ground instrumen- 
tation radars to provide range, rate, bearing and elevation tracking data 
during laiinch, boost, orbital and entry phases of the mission. 

The Transponder receives either single or double pulse interrogations 
from the ground radar. When the two-pxilse mode is chosen, the code interro- 
gation pulse spacing must be shown in Table 6-1 for the transponder to reply. 
The transponder replies in a single pulse for each pulse or double pulse 
received with a fixed internal time delay. The time delay is to allow 
sufficient time for the antenna switch to select the proper antenna and 
to allow the ground operator to observe the skin track reply pulse for 
comparison with the transponded pulse. 

The transponder will respond to interlaced interrogations from more 
than one radar with the limitations specified in Table 6-1. Nominal 
interrogation rate is l60 pps. 

The transponder operates. in conjxmction with four circularly 
polarized resonant cavity Helix antennas, flush mounted and spaced 90 
degrees apart about the spacecraft. The transponder decoder determines 
which antenna is receiving the strongest signal and switches the trans- 
mitter reply pulse to this antenna. 

6.2.2 VHF FM Transmitter 

The VHF FM Transmitter's sole function is to transmit NRZ PCM 
data to MSFN during launch, boost, earth orbit, and entry phases of the 
mission. 

6.2.3 VHF AM Transceiver 

The VHF AM Transceiver provides voice communications with MSFN 
during launch, boost, orbit, and entry phases of the mission. The 
transmitter portion of the transceiver caji be used as a recovery beacon. 

Dxiring Mission AS-501, a UOO HZ Tone from the audio center will 
be transmitted via the VHF AM Transceiver to the MSFN for voice simulation 
tests. 



6-U 



I 1 I I 1 1 I 1 I I I I I 1 I i i i i I 



TABLE 6-1 C-BMD TRANSPONDER DATA 



CSM 
AS-501 



ITEM 



POWER 



RECEIVER 
FREQUENCY 



RECEIVER 
SENSITIVITY 

RECEIVER 
BANDWIDTH 

DECODER CODE 
SPACING 



COMPARATOR 
AND ANTENNA 
SELECTION SWITCH 



TRANSMITTER 
FREQUENCY 



TRANSMITTER 
REPLY PULSE 



TRANSMITTER 
REPLY POWER 



C-BAND TRANSPONDER DATA 



UOO HZ, 3<(>, 115/200 Vac 

75 Watts maximum at 500 PRF 

+28 Vdc Control Power 



Tuneable : 
Assigned: 
Stability: 



56UO to 57^0 MHZ 
5690 MHZ 
+ 2 MHZ 



to -70 DBM with 95 percent reply minimum 
measured at RCVR input 

12 MHZ minimum at 30 DB points 
60 MHZ maximimi at 60 DB points 

Accept: Double Pulse spaced within 6+0.3 

microseconds 
Reject: Double Pulse spaced outside 6+0.8 

microseconds 
Measured: Leading edge to leading edge 

-70 to -55 DBM range requires at least k DB 

differential for Antenna Selection. 

-55 to -25 DBM range requires at least 6 DB 

Differential. 

-25 to range not defined. 

Tuneable: 5715 to 5815 MHZ 

Assigned: 5765 MHZ 

Stability: + k MHZ 

Drift: Less than 1 MHZ per minute 

Pulse Width: O.75 + 0.2 microseconds 
Rise Time: Less than Q.l microseconds between 
10 and 90 percent amplitude points 
Fall Time: Less than 0.25 microseconds 

Peak Power greater than 2.5 kW averaged over 
pulse width at 70 percent amplitude. 



6-5 



I I I I I I I ! II I II I i i I I i 



CSM 

AS-501 
TABLE 6-1 C-BAND TRANSPONDER DATA (Concluded) 



ITEM 


C-BAND TRANSPONDER DATA (Concluded) 


INTERROGATION 
RATE 


1 to 1300 Interrogations per second 
Countdown less than 5 percent 


NOISE 
FIRING 


Less than 5 pulses per minute 


TRANSMITTER 
RECOVKKY TIME 


Less than TO microseconds after last reply- 


TRANSMITTER 
LOCKOUT 


Yes - See TRANSMl'i'TKK RECOVERY TIME 


FREQUENCY 
SEPARATION 


Greater than kO MHz 


REPLY DELAY 


A. Differential: 0.05 microseconds between re- 
ceiver-transmitter combinations . 

B. Artifical: 3.0 plus or minus 0.2 microseconds 

C. Variations: 0.1 microseconds including antenna 
paths . 


WARMUP 
TIME 


2 minutes to reach frequency stability spec. 
32 seconds to become operational 



ITEM 


TABLE 6-2 C-BAND ANTENNA DATA 


C-BAND 

ANTENNA 
PATTERN 

ANTENNA 
NULL 

ANTENNA 
GAIN 


Helix J Four, Spaced at 90 degrees about the 
spacecraft, (See Figure 6-1) 

Omnidirectional in S/C YZ Plane. 
(RCP Polarization) 

36. U DB with Respect to Isotropic 

DB on Axis (Referenced to Isotropic Source) 



6-6 



1 I i i i i I 1 i I I I I I I i i i I i 



CSM 
AS-501 



TABLE 6-3 VHF - FM TRANSMITTER DATA 



ITEM 



POWER 



FREQUENCY 



RF POWER 

MODULATION 
DEVIATION 

WARMUP 

TIME 

ANTENNAS 



VHF - FM TRANSMITTER DATA 



+28 Vdc Control Power 

1.5 Watts dc 
UOO HZ, 3<J>, 115/200 Vac 

50 Watts ac 

237.8 MHz 

Stability ± 0.713^ MHz 

10 Watts (UO DBM) Minimun 

250 ± 20 kc for 3V P-P Input from PCM m 

2 minutes to meet specifications* 

h3 seconds for minimum operation stability, 

(See Table 6-15) 



6-7 



III I I i i I I I I 1 I I i i i I 1 



TABLE 6-1* VHF-AM TRANSCEIVER DATA 



CSM 
AS-501 



ITEM 


VHF AM TRANSCEIVER DATA 


POWER 


+28 Vdc 




1.5 Watts Receive only 




15.5 Watts Standby T/R 




61.5 Watts Trajismit 


TRANSMIT 


296.8 MHz 


FREQUENCY 


Stability + 0.890^+ MHz 


RECEIVER NO. 1 


296.8 MHz (Simplex) 


FREQUENCY 




RECEIVER NO. 2 


259.7 MHz (Duplex) 


FREQUENCY 




RECEIVER 


70 kc at 6 DB Points 


BAND PASS 




RECEIVER 


0.5 Micro Volt at signal-to-noise ratio of 


SENSITIVITY 


10 DB or better. 


RECEIVER 


Less than 10 percent 


DISTORTION 




RECEIVER 


Less than 0.1 seconds 


WARMUP 




TRANSMIT 


5 Watts Minimum 


POWER 




TRANSMITTER 


Less than 120 seconds to meet specifications 


WARMUP 





6-8 



I I I I I I i i I I I I I I I i i i I i 



CSM 
* AS-501 

6.2.U Updata Link Equipment 

The HDL provides the means for MSFN to update the CTE Time 
Accumulator, update the G&N Computer, and select certain vehicle 
functions during a mission. The equipment basically contains a UHF 
FM Receiver and a Command Decoder Unit. 

The UDL has two operational modes, UHF and S-Band. These modes 
allow selection of either UHF FM received commands or S-Band received 
commands. The UHF FM command link is used for near-earth operation 
and is the primary mode. The S-Band link can he used in near-earth or 
deep space. 

6.2.5 Updata Real-Time Commands 

Real-Time Commands effecting the commiani cations systems are shown 
in Table 6-G. 

6-2.6 The Unified S-Band Equipment (USBE) 

The USBE provides near-earth backup for the C-Band Transponder 
(tracking), VHF AM Transceiver (voice), VHF FM Transmitter (PCM Data) 
and UHF Updata Receiver (Updata Commands) on Block I spacecrafts. 
The following S-Band information modes will be used during Mission AS-501, 
S-Band uplink information (PM only) 

A. Voice (Not Used) 

B. Updata Link Commands 

C . Ranging 

S-Band downlink information: 

A. PM Mode 

1. Voice (1^00 cps Tone) 

2. PCM Data (Real-Time only) 
3* Ranging 

^. Emergency Key (Not Used) 
5. Emergency Voice (Not Used) 

B. FM Mode (NotUs?d) 
1. Voice 



6-9 



1 I II I i III I I 1 1 I 1 ill I 



2. PCM Data 

3. TV 

1*. Recorded PCM Data 



6-10 



CSM 
AS-501 



I 1 11 II i 1 I ill I 1 I i i i i f 



TABLE 6-5 UPDATA LINK DATA 



CSM 
AS-501 



ITEM 



\ 



POWER 



RCVR 
FREQUENCY 

FM CARRIER 
DEVIATION 

MODULATION 

SIGNAL 



RECEIVER 
SENSITIVITY 

MESSAGE 
STRUCTURE 

SUB-BIT 
CODE 

INFORMATION 
BITS 



REAL-TIME 

COMMAND 

MESSAGE 

G&N COMPUTER 
MESSAGE 

CTE TIME 

ACCUMULATOR 

UPDATA 



UPDATA LINK DATA 



+28 Vdc 

9.6 Watts (UHF Mode) 

^•S Watts (S-Band Mode) 

i+50 MHz Stability + 2.25 MHz 



+ 50 MHz (See Modiaation Signal) 



2 KHz phase shift keyed l80 degrees at 1000 sub- 
bits per second composite with A 1 KHz Sync Tone 
(See subbit characteristics.) 

-112 DBM for not more than 1 message rejection in 
1000 messages sent. 



5 sub-bits per information bit 
(See sub-bit characteristics.) 

30 bits total 

a. Bits 1, 2, and 3 Vehicle Address 

b. Bits i+, 5, euid 6 System Address 

c. Bits T thru 30 Message. 

Bits 7 to 12, Bit T is MSB. 

Actuation time is normally 10 milliseconds after 

receipt of message. 

Bits 7 thru 22, Bit 7 is MSB. 



N 



Bits 7 thru 30 

a. 7 thru 12 for seconds 

Bit 7 is LSB 

13 thru l8 for minutes 

Bit 13 is LSB 

19 thru 2k for hours 

Bit 19 is LSB 

25 thru 30 for days 

Bit 25 is LSB 



b. 



c. 



d. 



6-11 



I 1 I I I ill II I 11 I i i II 1 



TABLE 6-5 UPDATA LINK DATA (Concluded) 



CSM 
AS.5OI 



ITEM 



UDL TEST 

MESSAGE 

A&B 

SIGNAL 

ACCEPTANCE 

CRITERIA 



VALIDITY 
SIGNAL 



NON-VALIDITY 
CRITERIA 



SUB-BIT 
CHARACTERISTICS 



SUB-BIT ONE 



SUB-BIT ZERO 



SUB-BIT PERIOD 



UPDATA LINK DATA 



Bits k thru 30 

a. ^, 5» and 6 Test Message ID 

b. 7 thru 30 Data Word 



a. 
h. 
c, 
d. 
e. 



a. 



d. 



e. 



f. 



Correct Vehicle Address 

Valid System Address 

Correct Message Length 

Correct Sub-bit Coding 

Correct PSK Timing. (See sub-bit 

characteristics ) 

8-bit word to PCM TM present for, 50 milli- 
seconds on PCM Word 51D (Items cj d & e) 
Standby word 
Bit No. 1 2 3 U 5 6 T 8 
Binary 10100100 
System Validity Word 
Binary 01011011 
Test Message "A" Word 
Binary 10111000 
Test Message "B" Word 
Binary 01000100 
UDL OFF 
Binary 00000000 



No change from standby word to test message or 
system validity word following transmission of a 
command. 



One cycle of a one KHz reference signal and two 
cycles of a two KHz information signal which are 
both in-phase and positive going at the begin- 
ning of the sub-bit period. 

One cycle of a one KHz reference signal and two 
cycles of a two kH.z information signal which are 
180 degrees out of phase and where one KHz signal 
is positive going and two KHz is negative going 
at the beginning of the sub-bit period. 

1 millisecond. 
6-12 



1 III 1 I 1 i I I I I I I 1 i i i I 1 



CSM 
AS-501 

. TABLE S'S UDL REAL-TIME COMMANDS FOR COMMUNICATION SYSTEM 



NO. 


COMMAND 


UDL RELAY 
BOX 


MCP 


FUNCTION 




00 


ABORT LIGHT OFF 
(SYSTEM A) 


Kl 
(reset) 


NA 


Turns ABORT LIGHT A OFF. 
Binary "O" on PCM. 
(CS0080X/11E25-07) 




01 


ABORT LIGHT ON 
(SYSTEM A) 


Kl 
(Set) 


NA 


Tiirns ABORT LIGHT A ON. 
Binary "1" on PCM. 
(CS0080X/11E25-07) 




06 


ABORT LIGHT OFF 
(SYSTEM B) 


K2 
(reset) 


NA 


Turns ABORT LIGHT B OFF. 
(no PCM) 




07 


ABORT LIGHT ON 
(SYSTEM B) 


K2 
(set) 


NA 


Tiorns ABORT LIGHT B ON. 
(no PCM) 




51 


-Z ANTENNA ON 
(VHF SCIMITAR) 


NA 


1K33R 
lK3itS 


Switches VHF AM, VHF FM, and 
UHF UDL to -Z ANTENNA on S/M. 




52 


+Z ANTENNA ON 
(VHF SCIMITAR) 


NA 


1K33S 
1K3UR 


Switches VHF AM, VHF FM, and 
UDL to +Z ANTENNA on S/M. 




53 


G&N ANTENNA 
SWITCHING 


NA 


NA 


Not implemented 




61 


CM/SM SEPARATION 


NA 


1K38S 
lKlt2S 
lKlt3R 


Turns OFF the following equip- 
ment: 

a. VHF AM TRANSCEIVER 

b. VHF FM TRANSMITTER 

NOTE: RTC 6l also initiates CM/SI 
separation and is not used as a 
direct command fimction to turn 
off the above equipment for this 
reason. 




62 


UDL S-BAND RCVR 
SELECT 


NA 


1K3TS 


Routes S-Band received updata 
from S-Band RCVR via PMP 70 kc 
discriminator to UDL Decoder. 
MDC 20, 8, UDL Switch may be 
in UHF or S-Band position. 





6-13 



I I I I I I I I II I II I I i I I i 



CSM 
AS-501 



TABLE 6-6 UDL REAL-TIME COMMANDS FOR COMMUNICATION SYSTEM (Cont'd) 



NO. 


COMMAND 


1 
UDL RELAY 


FUNCTION 






BOX 


MCP 




63 


UDL UHF RCVR 
SRT.ECT 


NA 


1K3TR 


Routes UHF received updata from 
UHF RCVR in UDL to UDL Decoder. 
MDC 20, S8, UDL Switch must be 
in UHF position for command 
operation. 


Ik 


C-BAND OFF 


K3 
(reset) 


NA 


Effect is as follows depending 
upon MDC 20, S13, C-Band Switch 
position: 

(a) 1 Pulse: No effect 

(b) OFF : Turns C-BAND OFF 

(c) 2 Pulse: No effect 


75 


C-BMD ON 


K3 


NA 


Effect is as follows depending 




{2-PULSE) 


(set) 




upon MDC 20, S13, C-BAND Switch 

position: 

(a) 1 Pulse: No effect 

(h) OFF : C-Band ON in 

2-Pulse Mode 
(c) 2 Pulse: C-Band ON in 

2-Pulse Mode 


76 


PCM LBR VHF FM 


Kk 


NA 


Effect is as follows depending 




XMTR OFF 


(reset) 




upon MDC 20, S12, and S20 (VHF 
FM XMTR POWER Switch and PCM 
MODE Switch) positions: 
(a) VHF FM ON: No effect 
(h) VHF FM OFF: Turns VHF FM 
XMTR OFF 

(c) PCM HIGH : No effect 

(d) PCM LOW : PCM operation 

to LBR. 


77 


PCM HBR VHF FM 


Kk 


NA 


Effect is as follows depending 




XMTR ON 


(set) 




upon MDC 20, S12 and S20 (VHF 
FM POWER Switch and PCM MODE 
Switch positions: 

(a) VHF FM ON: No effect 

(b) VHF FM OFF: Turns VHF FM 

XMTH ON 

(c) PCM HIGH: No effect 

(d) PCM LOW: PCM operation 






1 


to HBR. 


NOTE 


See Paragraph ( 


5.5 for switch -Dositio 


ns at liftoff. 



6-iU 



I I i I I I i I I I I I I 1 ill i I [ 



TABLE 6-7 UNIFIED S-BAND DATA 



CSM 
AS-501 



ITEM 



POWER 



RECEIVER 

RECEIVER 
FREQUENCY 

TRACKING 
THRESHOLD 

TRACKING 
RANGE 

TRANSMITTER 
FREQUENCY 

TURN-AROUND 
RATIO 

TRANSMITTER 
POWER 

ANTENNA 
SWITCHING 



UNIFIED S-BAND DATA 



115/200 Vac 3 0, 400 HZ 
13.5 Watts kc 

Double Superheterodyne Phase Lock Loop 

2106.U0625 MHz 



-132 DBM to maintain phase lock with no 
Modulation 

Center Frequency 
+^ 60 KHz minimum 

2287.5 MHz, stability + O.OO5 percent 



2I+O/221 



250 to UOO milliwatts 



a. Level: -122 DBM or below switches ANT 

-104 to -122 Undefined, may switch • 
-lOl* DBM or above holds switching 

b. Delay: U.5 to 10 seconds before switching 

back to original antenna, switch 
will continue to switch from one 
ANT to the other until signal 
level is increased. 



6-15 



I ill ill I I I I I I I till 1 



CSM 
AS-501 



6.2.7 S-Band Pover Amplifier 

The S-Band power amplifier is ^ used to amplify the USBE RF signal 
for increase range and data capabilities. The lonit contains a diplexer 
for uplink/dovnlink signal isolation. 

During Mission AS-501 the S-Band Power Amplifier will be operated 
in the HIGH POWER Mode. 

6.2.8 HF Transceiver/Beacon 

The HF transceiver equipment is used to provide beyond line-of- 
sight direction finding and voice communications during postlanding 
phase of the mission. The HF transceiver has three modes of operation: 
(a) Single-Sideband transmit-receive, lower sideband only^ (B) Amplitude 
modulation transmit-receive, with full carrier and lower sideband, 
(C) Beacon on CW with full carrier no modulation. The SSB Mode is used 
for increased voice range capability. 

During Mission AS-501 the HF transceiver will be operated in the 
BEACON Mode. 

6.2.9 VHF Recoveiy Beacon 

The VHF recovery beacon is used to provide line- of- sight direction 
finding operation during the recovery phase of the mission. The beacon 
transmits an interrupted AM modulated CW carrier. 

6.2.10 VHF Survival Transceiver/Beacon 

The survival transceiver is part of the crew survival equipment 
used as a backup to the VHF/AM transceiver and VHF recovery beacon. The 
\init will operate from its own internal battery and provide beacon, 
voice transmission, and voice reception. In the voice mode, it will 
provide simplex two-way voice between vehicle, astronaut, and the recovery 
force after main parachute deployment for kQ hours. 

The survival beacon which is part of the unit consists of a beacon 
and battery pack. The unit will operate for 2k hours. 

The survival beacon will be connected to VHF Recovery Antenna No. 2. 

During Mission AS-501, the VHF Survival Transceiver /Beacon will be 
operated in the BEACON Mode. 

6-16 



I II i i i I i I I I I II Mi i [ i 



TABLE 6-8 S-BAND POWER AMPL DATA 



CSM 
AS.5OI 



ITEM 



POWER 



RF OUTPUT 
POWER 



INSERTION 
LOSS 

WARMUP 
TIME 



S-BAND POWER AMPL DATA 



115/200 Vac. 3 0, UOO Hz 

27.1* Watts Low pwr, 80 Vatts high pwr 

+28 Vdc control 

3 Watts dc 

a. High Power: Ik.k Watts minimum 

b. Low Power: 3.3 + 0.6, -O.U Watts 
with. 250 MW drive 



a. Bypass Receive: 

b. bypass Transmit: 



-1 DB maximum 
-1.8 DB maximimi 



90 seconds (Time Delay Relay Delays TWT 
operation during warmup). Power amplifier 
will i-emain in the bypass Mode during warmup, 



6-17 



I I I ill I I III II I Li 11 1 



TABLE 


- ' CSM 

AS.5OI 
6-9 HF TRANSCEIVER EQUIPMENT DATA 


ITEM 


HF TRANSCEIVER EQUIPMENT DATA 


POWKH 


+28 Vdc . : 

Receive AM and SSB 0.6 Watts 
Transmit AM or BCN 29 Watts 
Transmit SSB kO Watts 
Keying Relay** 3.5 Watts 


TRANSMIT 
FREQUENCY 


10.006 MGz 
Stability ± 50 Hz 


RECEIVER 
SENSITIVITY 


a. SSB: 1.0 microvolts 

b. AM: 3.0 microvolts with 
both sidebands 


RECEIVER 
MUTING 


Applied when transmitting 


TRANSMIT RF 
POWER 


a. SSB: 10 Watts average 

20 Watts PEP 

b. AM: 5 Watts average 

c. Beacon: 5 Watts average 


ACTIVATION 

TIME 


With' the equipment energized the transmitter 
will achieve full transmit operation within 
100 milliseconds of PPT key closure. 



6-18 



1 I III I i i I I I I I 1 i i i i I i 



CSM 
AS-501 



TABLE 6-10 VHT RECOVERY BEACON DATA 



ITEM 



VHP RECOVERY BEACON DATA 



POWER 
FREQUENCY 
RF POWER 
MODULATION 



INTERRUPTION 
RATE OF IKHz 
MODULATION 



+28 Vdc, 8 Watts 

21+3.0 MHz stability + 0.01265 MHz 

3 Watts minimvun peak 

AM, 20 to ilO percent, 1000 ± 50 Hz 
squarewave 

a. ON; 1.8 to 2.2 seconds 

b. OFF: 2.7 to 3.3 seconds 



6-19 



I I I I II Ml I I I 1 I i ill I 



TABLE 6-11 VHP SURVIVAL TRANSCEIVER/BEACON DATA 



CSM 
AS-501 



ITEM 


VHF SURVIVAL TRANSCEIVER/BEACON DATA 


POWER 


Internal Battery Pack 


OPERATION 


2k Hours in CW Mode 


LIFE 


U8 Hours in voice mode 


RF POV/EK 


2 Watts average 


OUTPUT 




FREQUENCY 


2l+3 MHZ 


BEACON MODE DUTY 


a. Modulation: 1 KHz sine Wave 


CYCLE 


b. On Time: 2 sec (Carrier + IKHz) 




c. Off Time: 3 sec (Carrier + IKHz) 


MICROPHONE/ 


' Integral to unit 


SPEAKER 




MODE SWITCH 


BCN/Voice/OFF 


PTT BUTTON 


Integral 



< 



6-20 



1 I i I I E E I I I 1 I I i I k i i I 1 



CSM 
AS-501 

6.2.11 Flashing Light Beacon 

The Flashing light is used as a visual beacon during recovery 
phases of the mission. 

■6.2.12 VHF Recovery Antennas 

There are two VHF Recovery Antennas, No. 1 and No. 2, stowed in 
the forward compartment of the C/M. They are deployed by the main para- 
chute riser lines which actuate redundant pyrotechnic line cutters (delayed 
8. seconds) releasing the spring-loaded antennas to the upright position. 
6.2.13 HF Recovery Antenna 

The HF recovery antenna is coiled in a small container stored in 
the forward equipment compartment. After the C/M has reached Stable I 
and impact plus 11 seconds, the antenna control circuits will be actuated 
by the mission control programer. The antenna is pyrotechnically released 
and imfurls into a tubiQar pole. 

6.2.lit Scimitar Antennas • 

The Scimitar Antennas , so called because of their shape are 
located on the service module.- This type of antenna has a broad band- 
width to cover VHF and UHF frequencies. 

6.2.15 S-Band OMNI Antennas 

Four S-Band OMNI Antennas are mounted in the APT heat shield of 
the command module at 90 degree intervals about the X-Axis. The antennas 
operate in pairs 180 degrees apart. Each antenna of a pair is fed from 
a common power divider. Only one pair is selectable at a time. 

6.2.16 Audio Center Equipment 

The Audio Center Equipment consists of three audio modules, one 
each for the command pilot, senior pilot, and pilot positions. Each 
audio module is used to process and control audio signals between the 
VHF AM, HF, and S-Band equipment and the crew headsets. The intercomm 
bus between each module is used for intra crew communications and 



6-21 



I I I li I I I III I I i Li ill 



CSM 
AS-501 



hardline voice prior to liftoff. A swimmer umbilical disconnect is tied 
to the intercom bus for voice communications during recovery operation 
after landing. 

Since Mission AS-501 is unmanned, a itOO Hz tone derived ft:om the 
UOO Hz.ac power, will be used to assimulate downlink voice communications. 
The senior pilot's audio module will process this tone to the VHF AM and 
S-Band equipment. 



6-22 



I I III ill I II I I 1 M i i I 1 



TABLE 6-12 FLASHING LIGHT BEACON DATA 



CSM 
AS-501 



ITEM 



FLASH RATE 
INTENSITY 
FLASH DURATION 
SERVICE LIFE 



FLASHING LIGHT DATA 



30 + flashes per minute 

2 million lumens 

3 milliseconds 
2U hours minimum 



6-23 



I I I I ill I I I I I III ill 1 



TABLE 6-13 VHF RECOVERY MTEMA 



CSM 
AS-501 



ITEM 



LOCATION 

TYPE 

FREQUENCY 

GAIN 

MAXIMUM NULL 

BEAMWIDTH 
(MAIN LOBE) 

POLARIZATION 



VHF RECOVERY ANTENNA 



See antenna location drawing (Figure 6.1) 

Stub, quarter wave length, omnidirectional 

2i+3 MHz tuned 

+6 DB with respect to isotropic 

-l8 DB with respect to isotropic 

+_ TO degrees elevation, 
360 degree azimuth 

Vertical 



6.2U 



I 1 I I I I i i I I 1 I 1 I i i i I I i 



TABLE S-lk HP RECOVERY ANTENNA DATA 



CSM 
AS-501 



ITEM 



LOCATION 

TYPE 

PATTERN 



. FREQUENCY 
RANGE 



GAIN 



HF RECOVERY ANTENNA DATA 



See antenna location drawing (Figure 6.1) 
Monopole 

Omnidirectional in a plane perpendicular 
to C/M X-Axis 

9.95 to 10.05 MHz 

DB minimum vith respect to isotropic 



6-25 



I I I 1 I I I I I III I i i i i I 1 



CSM 
TABLE 6-15 VHF SCIMITAR ANTENNAS AS-501 



ITEM 


VHP SCIMITAR ANTENNAS 


BANDWIDTH 

RADIATION 
PATTERN 

GAIN 
LOCATION 


225 to 1+50 MHz 
Hemispherical 

-3 DB with respect to isotropic 
(See DWG 6.1) 



6-26 



I I II 1 i I ill I I I i LI i i I i 



TABLE 6-16 S-BMD ANTENNAS 



CSM 
AS-501 



ITEM 



S-BAND ANTENNA DATA 



BANDWIDTH 

POLARIZATION 

RADIATION PATTERN 
AND GAIN* 



LOCATION 



2100 to 2300 MHz 

Right-Hand circular 

Nulls: Less than 20 DB with respect to isotropic 
level external to a solid angle of * 15 . 
degrees about the S/C ± X-Axis. 

Gain: Greater than -3 DB with respect to iso- 
tropic level for a minimum of 10 steradians 
of spherical coverage, {kit steradians in 
a sphere) 

(See DWG 6,1) 



± 



*ALL FOUR ANTENNAS RADIATING SIMULTANEOUSLY. 



6-27 



I I I 1 ill III I I 1 Hi i I 1 




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I 1 111 IE I I II I II I i i i I I 



CSM 
AS-501 



6.3 NOTES, COMMUNICATION SYSTEM MCP, CUB, AND UDL INTERFACE DIAGRAM 

A. This drawing is provided to show automated and remote control 
functions which operate the COMM system equipment. 

B. See communication system detailed diagram. Drawing ;6.U.l, for 
complete system schematic. 

C. See Table G-6 for communication system control functions by 
the Updata Link (UDL), 

D. See Table 6-17 for communication system control functions^ by 
the MCP. 

E. The Communication System Circuit Utilization Box (CUB) provides 
control of ac and dc power to the following: 

1. VHF FM Transmitter (ac) 

2. VHF AM Transceiver (dc) 

3. C-Band Transponder (ac) 
h. Unified S-Band (ac) 

5. S-Band Power Amplifier (ac) 

6. Updata Link (dc) . 

7. Premodulation Processor (dc only) 

Hardline commands to the CUB by Ground Support Equipment (GSE) can 
remove power from the above equipment after closeout and before liftoff 
if reqiiired. 



6-29 



I I I I I I 1 I I I III I I I I I 1 



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* AS-501 



6.1+ NOTES, COMMUNICATION SYSTEM DETAILED BLOCK DIAGRAM 

A. MAIN DISPLAY Switches are shown in the final closeout position 
according to Mission AS-501 flight plan. 

B. See communication system MCP, CUB, and UDL Diagram, Drawing 
6.3.1, for simplified diagram of automated and remote control 
fTjnctions effecting the COMM system. 

C. See Drawing 6.3.1 and Table 6-17 for further detail of control 
wires which interface with' the MCP. 

D. See Drawing 6.3.1 and Table 6-6 for detail of real-time commands 
which effect the COMM system. 

E. The command pilot's audio module of the audio center equipment 
will be used to process a 1+00 Hz tone to the VHF AM and S-Band 
equipment for downlink voice simulation test. 



1 



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AS-501 



6.5 COMMUNICATION SYSTEM CIRCUIT MARGINS 

6.5.1 Communication Links 

Communication links with the MSFN are basically sensitive to 
two mission parameters: vehicle-to-MSFN slant range and antenna gain 
in the direction of the vehicle-to-MSFN line of sight. S-Band communi- 
cations are additionally sensitive to combinations of uplink and downlink, 
modulation modes. 

I I 6.5.2 Ranges of Communication Links 

Figure 6.k shows the ranges of the communication links where the 
circuit margins are positive. Negative circuit margins res\ilt in ex- 
cessive bit error rates, voice distortion, command rejects, and ranging 
errors . 

J - I 6.5.3 Calculated Ranges 

Ranges shown in Figure 6,k were calculated using the following ground 
rules and data obtained from CSM circuit margin interface control documents « 

A. S-Band system 

1. MSFN 30 foot dish {k3 DB Gain Uplink) 

2. CSM antenna gain of -3 DB (OMNi) 

] 3. MSFN power of 10 kw 

J 

k. CSM power of 10 watts (P.A. HIGH) 

5. Worse case margin: 

a. Voice -T.99 DB (Downlink) 
\ b. PRN ranging 1+.38 DB 

c. HBR PCM -13.i*2 DB 
I d. Updata IT.87 DB 

6. Reference range of 15,000 n.mi. 

B. C-Band system 

1. MSFN AN/FPQ6 or AN/TPQ - I8 RADARS 

2. CSM antenna gain DB 
._, 3. CSM power of 2.5 kW peak ' 

1 

6-35 



I I I I I M I ! ! I [ I n i 1 Iff ! 



CSM 
AS-501 



k. Worst case margin of 9-56 DB (Downlink) 

5. Reference range of UOOO n.mi. ^ » ^ 

C. VHF FM PCM system I | 

1. MSFN antenna gain of l8 DB 

2. CSM antenna gain of -3 DB 

3. CSM power of 10 watts 

U. Worst case margin of -9-12 DB 

5. Reference range of 2,000 n.mi. 

D. VHF AM voice and UHF FM updata. systems K " 1 

1. MSFN antenna gain of l8 .DB 

2. CSM antenna gain of -3 DB 

3. Worst case margin -3.09 DB (VHF AM) and 12.12 DB (UHF FM) 
h. Reference range 2,000 n.mi. 

NOTE: Data shown in Figure S.k represents minimum design ■ ■ 

requirements and worse case conditions with, the exception 
of optimum antenna look angles. Actual circuit margin 
parameters are expected to exceed design requirements. 



[i 



1 1 

6-36 



1 1 1 1 1 1 1 1 ! 1 1 1 1 r 1 1 1 1 1! I 



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CSM 
AS- 501 



6.6 COMMUNICATION SYSTEM FINAL CLOSEOUT SWITCH AND CIRCUIT BREAKER 

CONFIGURATION. (REFERENCE: AS-501 FLIGHT PLAN, SEPTEMBER 12, I966) 



6.6.1 MDC - 13^ Audio Control Panel 
Power (PTT/OFF/VOX) (S5) 
S-Band (T/R / 0FF/REC'(S1) 

HF (T/R / OFF/REC) (S2) 
VHF/AM (T/R /OFF/REC) (S3) 
Intercom (T/R /OFF/REC) {Sk) 
VOX Sens (Thmntwheel) (R3) 
Intercom Balance (Thumbwheel) (R2) 
Volume (Thumbwheel) (Rl) 

6.6.2 MDC 23 and 26 Audio Control Panel 

All audio center control switches will be OFF. 

6.6.3 MDC-25 Circuit Breaker Panel 
ANTENNA Deploy (HF) 

A. (A/OFF) (s8) 

B. (B/OFF) (Sl6) 
BEACON LIGHTS/OFF (S9) 
BIOMED COMM 

A. MNA (CB59) 

B. MNB (CB6o) 

6.6.U MDC--19 Control Panel 

UP TM CMD (RESET/OFF) (S12) 

6.6.5 MDC-20 Communications Control Panel 
S-Band 

Xponder (Xponder/Off /Xponder Pwr Amp) (Sl) 
Pwr Amp (High/Low) (S2) 
OSC (Prim/Sec) (S3) 



VOX 
T/R 
OFF 
T/R 
OFF 

MAX (9) 
NOT USED 
NOT USED 



OFF 
OFF 
BEACON LIGHTS (UP) 

CLOSED 
CLOSED 

OFF 



I I 



II 



[] 



XPONDER PWR AMP 

HIGH 

PRIM 



6-38 



[ ] 



I I I I I If I I f [ I I I [ i i I [ I 



CSM 
AS-501 



I 1 ' 



I 1 



I I 



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\ 



Voice 

Rng (Rng/Off/Rng Only) (S5) 
Tape/Analog (Tape/Off /Analog) (S6) 
TV/PLSS ( TV/Of f/PLSS) (S?) 
Emerg (Voice/Off/Key) {Sk) 
Updata (S-Band/Off/UHF) (S8) 
S-Band Antenna (Auto/Upper /Lower) (S32) 

VHF/AM 

Squelch (Thumbwheel) (Rl) 

T/R/Off/Rec (Sll+) 

Rcvr (1/2) (S15) 
C-Band (l Pulse/Of f/2 Pulse) (S13) 
VHF/FM (On/Off) (S12) 

Recovery 

VHF BCN (ON/OFF) (Sll) 

HF 

ON/OFF (SIO) 
SSB/BCN/AM (S9) 

Tape Recorder 

PLAY (PCM/NORM/ ANALOG) (S19) 
SPEED (HIGH/NORM/LOW) (Sl8) 
RECORD/OFF /PLAY (Sl6) 
FWD/STOP/REV (S17) 

Power 

SCE (ON/OFF) (S22) 
PMP (ON/OFF) (S23) 

TM Inputs 

PCM (HIGH/LOW) (S20) 
ANALOG (1/2) (S21) 

6-39 



RNG 
OFF 
OFF 
OFF 
UHF 
UPPER 

TBD 

T/R 

1 

OFF 

ON 

ON 

ON 
BCN 

NORM 
NORM 
RECORD 
FWD 

ON 
ON 

HIGH 
NOT USED 



I I II I M II fill IF I 1 [ f I 1 



CSM 
AS-501 



BIOMED (1/2/3) (S2U) 
VHF Antenna (Recovery /Upper/Lower) (S35) 

6.G.6 MDC-ltt G&N Computer Control Panel 
UPTEL (ACCEPT/BLOCK) (Sl) ,. 

6.6.7 VHF Survival Beacon (Switch ON Beacon) 
VHF Survival Beacon (BCN/ VOICE /OFF) 

6.6.8 MDC-22 Circuit Breaker Panel 
Central Timing System 

MN A (CB53) 
MN B (CB52) 
COMM C-BAND XPNDR (CB120) 

Communications 

ESS A (AC 1/OFF/AC 2) (S13) 

NONESS B (AC 1/OFF/AC 2) (SlS) 
Telecommunications 

GROUP 1 (CBlt9) 

GROUP 2 (CBltS) 

GROUP 3 (CBltT) 

GROUP 1* (CBlt6) 

GROUP 5 (CBlt5) 

GROUP 6 (CB50) 

S-BAND (CB51) 

Comm 

FLT QUAL RCDR (CB96) 
C-BAND (CB97) 

6.6.9 RHEB Circuit Breaker Panel 
POSTLDG 

MAIN A (CB6) 
MAIN B (CB5) 



NOT USED 
UPPER 

ACCEPT 

BCN 



CLOSED 
CLOSED 
CLOSED 

AC 1 (AUTO) 



CLOSED 
CLOSED 
CLOSED 
CLOSED 
CLOSED 
CLOSED 
CLOSED 

CLOSED 
CLOSED 



CLOSED 
CLOSED 



6-I1O 



I I I I I I I I I f I T I f [ I i [ 1 1 I 



n 



! 1 ' 



SECTION 7 ^^ 

' AS-501 

INSTRUMENTATION 



7.1 GENERAL SYSTM NOTES 

All Apollo spacecraft instrumentation measrirements are divided into 
two classifications: Operational and Flight Qualification. 

7.1. 1 Operational Measurements 

Operational measurements are those measurements which remain fixed 
for a block of vehicles fulfilling a similar type of mission. They are 
1 divided into three categories as follows: 

A. Preflight Checkout of the Spacecraft 

These preflight checkout measurements are used to insure proper 
system operation and flight readiness of the spacecraft. 

B. Inflight Management of the Spacecraft 

These are onboard display measurements used by the astronaut 
I during real time for inflight spacecraft management. The 

measurements give status of consumable items , provide spacecraft 
performance information and indicate proper sequencing of 
operations . 

C. Mission Evaluation and Subsystem Performance 

These telemetered measurements are 'used by groimd personnel to 
j monitor spacecraft performance for real-time flight control. 

The measurements are recorded on the ground and tised in post- 
flight mission evaluation. These measurements along with tele- 
metered flight qualification measurements are listed in 
\ Section 7.3. 

7-1.2 Flight Qualification Measurements 

Flight qualification measurements are those measurements relating 
to qualification and verification of spacecraft, and subsystem engineering 
design. Some flight qualification measurements are displayed onboard the 
spacecraft and/or telemetered for ground observation. Flight qualification 
measurements available for analysis during reaijl time are listed with the 
I operational measurements in Section 7.3. 

7-1 



I I I I I M ! I I M I I I n IT f f 



" 1 1 



CSM 
AS-501 



7.2 INSTURMENTATION CONTROL SYSTEM NOTES 

The CSM Instrumentation System has five major components: 

A. The Pulse Code Modulation Telemetry Eqxiipment (PCM) 

B. The Central Timing Equipment (CTE) 

C. The Signal Conditioning Equipment (SCE) 

D. The Flight Qualification Recorder (FQR) 

E. The Data Storage Equipment (DSE). 

Other components include transducers for sensing CSM system parameters, — > -: ^ 

high and low level commutators for flight qualification measurements and 
instrumentation power distribution circuits. 

Table 7-1 presents the power requirements of the five major components. 



[ ] 



E ] 

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! 1 



' 1 



TABLE 7-1 INSTRUMENTATION SYSTEM EQUIPMENT POWER REQUIREMENTS 



EQUIPMENT 


MODE/ (STATUS) 


DC WATTS 


AC WA'i'i'b 


PCM TM 


High or Low Bit Rate 





10.5 


CTE 


(ON) 


21.0 





DSE 


Record and Playback 


0.5 


32.5 


FOR 


Record 


3U.0 


15.0 


SCE 


(ON) 





1*5.0 



\ 



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cm 

AS-501 



7.2.1 Pulse Code Modulation Telemetry (PCM TM) Equipment 

The Piilse Code Modulation (PCM) telemetry equipment, receives data 
inputs from various sampling points throughout the spacecraft and converts 
them into a single Non-Return to Zero (NRZ) digital, output to the DSE for 
recording and to the PMP for transmission to the grovmd. 

Two modes of operation are possible: the High Bit Rate (HBR) mode . 
of 51.2 kilobits per second (kbps) ,and the Low Bit Rate (LBR) mode of 
1.6 kbps. The desired mode is selected by placing the TM INPUTS PCM 
switch on MDC 20 to HIGH or LOW, or, when this switch is in the LOW position, 
HBR mode can be commanded by transmitting RTC 05 via the UDL. The DSE 
recording and playback operations will be effected by changes in PCM bit 
rate. In HBR mode, all telemetry inputs to the PCM equipment are processed 
and appear in the output signal. In LBR mode, some of the parameters will 
not appear in the PCM output signal. See Section T.3 for LBR parameters. 

The PCM timing is normally derived from the CTE 512 KHz clock input 
signal and the CTE 1 PPS signal is used for subframe synchronization. If 
the CTE 512 KHz clock is lost for more than 10 microseconds, PCM timing is 
switched to an internal clock and the 1 PPS subframe sync is disabled. 
If the CTE clock recovers after a switchover, timing is automatically 
switched to the CTE clock and a new subframe is started with the next 
1 PPS sync input . 

,7.2.2 Central Timing Equipment 

The CTE provides basic clock signals to time correlate all CSM 
time sensitive functions. 

The CTE supplies a 512 KHz signal to the PCM telemetry equipment, 
to the PMP, and to groirnd checkout equipment. Three 6.U KHz signals are supplied 
to the three ac power inverters for synchronization of the i+00 Hz 
frequency of the ac power. A h KHz signal is supplied to the Delta V 
section of the SCS for change in velocity calculations. 

Synchronization of the subframe of the PCM is accomplished by a 
1 pps signal from the CTE. 

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A 1 piase per 10 minutes signal is supplied to the ECS for use in 
I "" ' ^ "the water separator accumulator. The CTE time accvmiTolator supplies a 

32 hit time word (CT0llt2F) to the PCM which indicates the day of the month, 
the hour, the minute, and the second. The time accumiaator may be set to 
any predetermined time up to 31 days, 23 hours, 59 minutes, and 59 seconds 
via the UDL. Any interruption of power to the CTE will cause the time 
accumulator to reset to 01 day, 00 hour, 00 minutes and 00 seconds. 

^^ "tlie primary or normal mode of operation, the Apollo giii dance 
computer (AGC) provides a 1024 KHz signal to the CTE. This automatically 
synchronizes (phase locks) the CTE with the AGC and provides a stability 
of ±2 X 10" parts in lU days. In the event of sync signal failure, the 
CTE automatically switches to its own crystal oscillator with a stability 
of +2.2 X 10-6 parts in 5 days. 

' 1 

T.2.3 Signal Conditioning Eqviipment (SCE) 

The signal conditioning equipment is designed to process inputs 
from spacecraft sensors and produce analog outputs ranging from to 
+5.0 volts dc. The signal conditioning equipment can consist of up to 
U5 modules of the following kind: frequency demodulators , bi-phase 
demodulators, dc differential amplifiers, dc differential bridge amplifiers, 
active attenuators, active inverter attenuators, ac to dc converters and 
dc low gain amplifiers. 

7.2.3.1 Frequency Demodulator Module (FD) 
. The frequency demodulator modtile contains one circuit capable 

of transforming an input signal frequency change from 380 to U20 Hz to 

, an output signal from to 5 Vdc. Input signals are used to trigger a 

one shot multivibrator feeding an active filter. The filter along with 
a dc differential amplifier and output driver transforms the multivibrator 
frequency to a to 5 Vdc signal. The dc differential amplifier and output 
driver circuits are adjustable for zero output volts and gain. 
A. Requires SCE +20 and -20 Vdc power. 

i B. Input Signal - 105 to 150 VRMS, UOO Hz ac voltage. 



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C. Input Impedance - > 500 K ohms 

D. Output Impedance - < 500 ohms. 

E. Output signal - to 5Vdc, varies linearly with input from 
380 to i+20 Hz. Output is zener diode clamped for voltages 
less than -0.8 Vdc or greater than +6.2 Vdc. 

F. Overall response is flat for input frequency changes up to 
15 Hz/sec. 

7.2. 3. 2 Bi -Phase Demodulator Module (PP) 

The biphase demodiilator module contains one circuit capable of 
measuring and I80 degrees phase relation between input and reference 
ac voltages. The module can be adjusted to handle input signal levels 
between 0.5 and 50 VRMS and UOO to 16OO Hz. 

Zero and gain adjustments are provided to adjust Vdc (18O 
degrees out of phase) and +5 Vdc (O degrees in phase) outputs for a specific 
input voltage level. The input signal is gated by a reference signal so 
that a half wave signal is fed to an integrator circuit. With volts 
input or with a ±90 degrees phase relation the output signal will be 
+2.5 Vdc. 

A. Requires SCE +20 and -20 Vdc power 

B. Input signals are dc isolated 

C. Input Impedance - > 500 K ohms 

D. Output Impedance - < 500 ohms 

E. Output signal varies linearly, 2.5 to or +5 Vdc, with input 
voltage. Output voltage is zener diode clamped for voltages 
less than -0.8 Vdc or greater than +6.2 Vdc. 

7.2.3.3 DC Differential Amplifier Modules (DA) 

The dc differential amplifier mod\ile consists of two circuits, 
each capable of amplifing a differential 20 to 250 MVdc input signal to 
5 Vdc output or a single ended 0.25 to 7.0 Vdc input signal to 5 Vdc 
output. After making connections for proper range, potentiometer adjustments 
are made for gain, balance and zero. 



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A. Requires SCE +20 and -20 Vdc power 

B. Input Ranges: 

! 1 ^ 1. Differential - 20 to 200 MVdc , 

^ i and 

200 to 250 MVdc 
2, Single ended -^250 to 320 MVdc, 

320 to 1^10 MVdc, 
O.lil to k.k6 Vdc 
i+.l+f^'to 7.0 Vdc 

C. Output voltage, to 5 Vdc, varies linearly with input 
voltage. Output is zener diode clamped for voltages less 
than -0.8 Vdc or greater than +6.2 Vdc. 

D. Frequency Response - flat to 200 Hz. 

E. Input Impedance - > 500 K ohms 

F. Output Impedance - < 500 ohms 

T.2.3.U DC Differential Bridge Amplifier Modiile (DBA) 

The dc differential bridge amplifier module consists of two 
circuits, each capable of providing a to 5 Vdc output for a change in 
resistance of an external sensor. Bridge and ajr5)lifier can be adjusted 
to provide linear output voltage with sensor resistance changes over the 

} following ranges: Ii55 to 65O ohms, 328 to 355 ohms, 270 to 525 ohms, 

2800 to 5700 ohms, or 315 to 369 ohms. For the above ranges, Vdc output 
corresponds with minimum resistance. The output is zener diode claii5)ed 
for voltages less than -0.8 Vdc or greater than +6.2 Vdc. A short circuited 
input will cause -0.8 Vdc to appear at the output. An open circuited 
input will cause +6.2 Vdc to appear at the output. The bridge circuit 

j reqxiires SCE +10 Vdc power and the amplifier circuits requires SCE +20 

and -20 power. 

7.2.3.5 Active Attenuator Modiile (AA) 

The dc active attenuator module contains seven separate circuits, 
each capable of attenuating a 7.0 to 50 Vdc input signal to 5 Vdc output. 
—1 After connecting for the proper attenuation range, a potentiometer is 

7-7 



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adj-usted for the desired gain, 

A. Requires SCE +20 and -20 Vdc power. 

B. Input Ranges - 7.0 to l8.T Vdc 

18. T to 50.0 Vdc 

C. Output voltage varies from to +5*0 Vdc linearly with the 
input voltage. Output is zener diode clamped for voltages 
less than -0.8 Vdc or greater than +6.2 Vdc 

D. Frequency Response - flat to 200 Hz 

E. Input Impedance - 20 K ohms. 

F. Output Impedance - < 500 ohms. 

7.2.3.6 Active Inverter Attenuator Module (AIA) 

The active inverter attenuator module contains seven circuits , 
each capahle of attenuating and inverting a -7.0 to -50 Vdc input signal 
to a 5 Vdc output. After connecting for the proper attenuation range, a 
potentiometer is adjusted for the desired gain. 

A. Requires SCE +20 and -20 Vdc power. 

B. Input Ranges - 7.0 to -l8.7 Vdc, 

- 18.7 to -50.0 Vdc 

C. Output voltage, to 5 Vdc, varies linearly with input 
voltage. Output is zener diode clamped for voltages less 
than -0.8 Vdc or greater than +6.2- Vdc 

D. Input Impedance - 20 K ohms 

E. Output Impedance - < 500 ohms. 

7.2.3-7 AC to DC Converter Module (AC) 

The ac to dc converter Modiile contains two circuits, each capable 
of converting a 1 to 150 VRMS ac voltage to a 5 Vdc output. After an 
input ac voltage range is selected, the gain is adjiisted for the maximum 
desired ac voltage input. The ac signal is rectified to produce a dc 
output proportional to the amplitude of the imput signal. 

A. Requires SCE +20 and -20 Vdc power 



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B. Input Ranges - 1.0 to 10.0 VRMS and 

10.0 to 150 VRMS for 
] 1+00 to 3200 Hz signals. 

C. Input Impedance - > 500 K ohms 

D. Output Impedance - < 500 ohms 

E. Output is zener diode clipped for voltages less than -0.8 Vdc 
or greater than +6.2 Vdc. 

, T.2.3.8 Lov Gain DC Amplifier Module (LGA) 

The low gain dc amplifier module contains two circuits, each 
capable of providing +2.5 Vdc output for zero signal input, Vdc output 
for -2.5 to -35 Vdc signal input and 5 Vdc output for +2.5 to +35 Vdc 
signal input. The gain of the conditioner is adjusted for the maximum 
desired input signal for the ±2.5 to ±35 Vdc range i 

j A. Requires SCE +20 and -20 Vdc 

B. Input Impedance - > 20 K ohms 

C. Output Impedance - < 500 ohms 

D. Output is zener diode clipped for voltages less than -0.8 Vdc 
or greater than +6.2 Vdc 

7.2. U Flight Qualification Recorder (FQR) 
J The FQiR is a li+-track magnetic tape recorder used to 

record certain flight qualification measurements during critical phases 
of the mission. This data will be used for postf light analysis only; 
the FQR has no inflight playback or transmission capability. It will be 
\ activated to record the ascent and entry phases of the mission. 

The FQR has four external commands: 
' A. REWIND 

B. CALIBRATE 

C. RECORD 

D. STOP 

The REWIND command moves the tape transport in the reverse direction 
at 120 ips and locks out RECORD. The CALIBRATE command removes the data 



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AS-501 



input signal and applies a calibration signal to each VCO. The RECORD 
command energizes the electronics and moves the tape transport in the 
forward direction at 15 ips. Thirty one minutes of RECORD time is avail- 
able. The entire tape width is erased prior to recording in the RECORD 
mode. The STOP command de-energizes the RECORD electronics and de-activates 
the transport. End of tape sensing with automatic tape transport and 
electronics shutoff is provided in both directions. 

7.2.5 Data Storage Equipment (DSE) 

T.2.5.1 DSE Record Capability 

The DSE provides the capability to record and reproduce one 
track of PCM data clock, four tracks of PCM data and nine tracks of 
analog data in parallel. The DSE uses Main Bus A power for the control 
circuits and ac Bus 1 or ac Bus 2 ac power for the electronics and tape 
drive motors . 

The DSE is a bi-directional machine with a tape transport 
mechanism capable of forward or reverse operation at any of three speeds: 
3-75 ips (low speed), 15 ips (normal speed), and 120 ips (high speed). 
The low speed of 3.75 ips is used only to record low bit rate PCM TM data. 
The normal speed of 15 ips is used to record and playback high bit rate 
PCM TM data. The high speed of 120 ips is used for fast dump of the low 
bit rate PCM TM data which was recorded at 3.75 ips. The high speed is 
also used for fast forward or reverse rewinding. Maximum operating times, 
to record or playback the entire 2250 feet of tape, are as follows: 
2 hours at the low speed of 3.75 ips, 30 minutes at the normal speed of 
15 ips, and 3.75 minutes at the high speed of 120 ips. 

7.2.5.2 DSE Control Switches 

DSE control switches located on MDC 20 along with relays 1K39 
and IKUo in the MCP provide the following DSE control signals. 

A. RECORD - Energizes the record, erase (the ERASE/OFF switch 
on the DSE must be in ERASE) and tape motion electronics 



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provided: 
f 1 ^ 1. A HIGH signal and a NORMAL SPEED signal are both present » or 

a LOW signal and a LOW SPEED signal are both present. 
If conditions 1 or 2 are not met, tape movement will be in- 
hibited and all electronics de-energized. 

B. PLAYBACK - Energizes the tape motion electronics. The PCM 
playback electronics will be energized provided there is no 

^ 1 I'OW SPEED signal preseht. The analog playback electronics 

will be energized provided there is no LOW SPEED or HIGH 
SPEED signal present. Inhibits erase circuits. 

C. FORWARD - Tape will move in the forward direction provided: 

1. End of tape has not been reacted, or 

2. A record inhibit signal is not present. 

I D. REVERSE - Tape will move in the reverse direction provided: 

1. Start of tape has not been reached, or 

2. A record inhibit signal is not present. 

E. HIGH SPEED - Tape will move at 120 ips when tape movement 
is commanded. Inhibits analog playback electronics. 

F. NORM SPEED - Tape will move at 15 ips when tape movement 

j is commanded. Inhibits tape movement if RECORD signal is 

present and HIGH signal not present. 

G. LOW SPEED - Tape will move at 3.75 ips when tape movement 
is commanded. Inhibits tape movement if RECORD signal is 

\ present and LOW signal is not present. Inhibits playback 

electronics. 
H. HIGH - Inhibits tape movement if RECORD signal is present 

and NORM SPEED signal is not present. 
I. LOW - Inhibits tape movement if RECORD signal is present 

and LOW SPEED signal is not present. 



I 



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7.2.5.3 Internal DSE Controls 

The DSE provides the following internal controls. 

A. An ERASE/OFF switch on the DSE allows recording without 
erasure. 

B. The DSE is completely tiorned off if no RECORD, PLAYBACK, 
FORWARD, or REVERSE control signal is present. 

C. End of tape sensing and automatic tape transport shutoff 
without loss of tape end is provided in "both directions. 

D. Tape motion indication (CT0012) will not be valid unless a 
RECORD or PLAYBACK control signal is present. 

E. There will he no tape movement unless a FORWARD or REVERSE 
control signal is present. 

F. LBR recorded PCM data must be played back at 120 ips, HBR 
recorded PCM data must be played back at 15 ips and analog 
recorded data must be played back at 15 ips or the output data 
will be inhibited or garbled. 

7. 2. 5.^ DSE - PMP Interface 

The DSE - PMP interface is controlled by relays in the PMP and 
switches on MDC 20 (shown on Dwg. 6.U.I). When the TAPE RECORDER - 
RECORD/PLAY switch is placed to the RECORD position, relays in the PMP 
are energized connecting nine channels of analog signals to the DSE 
analog record circuits. During AS-501 flight, four of these channels 
will be used to record the output from the flight qualification HL and LL 
communtators. Placing the TAPE RECORDER - RECORD/PLAY switch to OFF 
(center) or to PLAY resets these relays disconnecting the analog signals 
from the DSE. 

In order to transmit DSE playback data to the groimd via S-Band 
or VHF-FM, the S-BAND - VOICE - TAPE switch must be in the TAPE position. 
This places the S-Band (if powered) in the FM mode of operation and arms 
the TAPE RECORDER - PLAY switch enabling one of the following operational 
modes . 



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A. PCM - Energizes circuitry in PMP to transmit PCM playback 
data and analog real-time data. 

B. NORMAL (center) - Energizes circixitry in the PMP to transmit 
PCM and analog playback data, 

C. ANALOG - Energizes circuitry in the PMP to transmit analog 
playback data along with real-time PCM data. 

There will not be any playback of data dvtring the AS-501 flight. 
Analog data can only be transmitted via S-Band whereas PCM data can be 
transmitted via VHF-FM or S-Band. 

7.2.5.5 DSE Time Code Signal 

The DSE receives a time code modulated 25 KHz signal from the 
FQR which is used as a timing reference for the analog parameters. This 
time reference will not be available when the FQR is stopped. 

7.2. 5. 6 DSE PCM Data 

The DSE PCM data is obtained from the PCM equipment as a single 
NRZ serial pulse train at a high bit rate of 51.2 kbps or a low bit rate 
of 1.6 kbps. In the DSE, a serial-to-parallel converter converts this signal 
into four parallel digital channels, each of which has a resiilting pulse 
repetition rate (PRR) of only 12.8 kbps or O.U kbps (one fourth of the 
original PRR). The PCM TM equipment furnishes a 51.2 KHz or 1.6 KHz square 
wave data clock signal to the DSE, which is also divided by four. 

During playback, the four parallel channels of recorded digital 
data and the data clock signal are routed to a parallel-to-serial converter 
where the four digital signals are converted back to a single NRZ serial 
pulse train output to the PMP. The bit rate of this output will always 
be 51.2 kbps, even if the recorded data was originally 1.6 kbps. The 
51.2 kbps high bit rate PCM signal is recorded at 15 inches per second 
(ips) and played back at the same speed. The low bit rate signal of 
1.6 kbps is recorded at 3.75 ips, and played back at 120 ips, an increase 
of 32 times. This increases the 1.6 kbps to 51.2 kbps. 



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T.2.5.T DSE Characteristics 

A. PCM Input Signal Levels - Binary "l" = 6.0 (±0,6) Vdc 

Binary "0" = 0.0 (+0.5, -0.0) Vdc 

B. PCM Output Signal Levels - Binary "l" = 6.0 (±0.5) Vdc 

Binary "O" = 0.0 (+0.5, -O.O) Vdc 

C. Audio circuitry and patching capability is available for 
recording and playback of two wire audio signals; however, 
this capability will not be used. 

D. Analog channel inputs and outputs are single wire unbalanced 
signals (O to 5 volts peak to peak). 

E. The tape transport and electronics will reach operational 
stability in less than 3 seconds after application of power. 

F. Tape speed will be controlled within ±0,5 percent of selected 
speed. The tape drive motors are i+00 cps synchronous motors. 

G. Rewind time from end to end of tape in either direction at 
120 ips is 3-75 minutes. 

H. Tape motion monitor is bi-directional. 

Tape in motion output - 7 (+3, -2) Vdc 

Tape stopped output - (+0.5, -O.O) Vdc 
I. One inch wide Mylar tape 2250 feet long is used. 
J. DSE Operational Mode Characteristics 



: 1 1 



11 



PARAMETER 


DIGITAL CHANNELS 


ANALOG CHANNELS 




HI BIT 


LOW BIT 


HI BIT 


LOW BIT 


Input Data Rate 


51.2 kbps 


1.6 kbps 


50-10,000 cps 


12.5-2500 cps 


Record Speed 


15 ips 


3.75 ips 


15 ips 


3-75. ips 


Playback Speed 


15 ips 


120 ips 


15 ips 


15 ips 


Output Data Rate 


51-2 kbps 


51.2 kbps 


50-10,000 cps 


50,10,000 cps 


Total Record Time 


30 Minutes 


2 Hours 


30 Minutes 


2 Hours 


Total Playback Time 


30 Minutes 


3.75 Minutes 


30 Minutes 


30 Minutes 



[ ] 



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MISSION CONTROL PROGAAMEfi (WCP) 



A, 



1K32 - SET - (NOT RTC 71 ANO 2ND GPS) + f3SE TM C*U +(d/dl LET JETT) 

1K32 -RESET -(PH)+CTAPE RCDR REWIND) + {1M1 -SET) 

IMl -SET - (C/M - S/M SEfi) + (TAPE RCDR START RCD) 

1K41 - RESET - (PR) + (1K32 - SET) 

IKSS - SET - (TAPE RCDR REWIND) 

1K55 - RESET - (PR) + (CSE TM CAU 

1K39, 1K40 - SET - CSE COMM CONTL TRANSFER PRELAONCH 

1K39, 1K40 - RESET - CPfl) + (ABORT) + (C/M - S/M SEP) 




I I III I If ! ! 11 I T [ I I f T f [ 



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ANALOG 
INPUTS 

PRE MOOULATfON 
PROCESSOR 

(PMP) 


C26-1AI 








ERREF 









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I ! 



7.3 PCM SIGNAL FLOW NOTES 

Each analog and digital input to the PCM equipment is assigned to a 

I I channel and given a channel code number. A fixed program is provided for 

HBR and for LBR sampling. In HBR, all inputs will be sampled at least once 
during a subframe at sampling rates from 1 s/s to 200 s/s. In LBR, only 
selected inputs will be sampled during a subframe at a sample rate of 
1 s/s (except channel code 51DS, AGC parameters, which will be sampled 
at 10 s/s). The sampling sequence for HBR and for LBR are shown in the 

1 1 HBR format and LER format respectively. 

In the HBR mode the PCM equipment is capable of handling the following 
input signals: 

A. Low Level Analog (O-lfO MVdc) - 50 channels at 1 s/s 

B. High Level Analog (0-5 Vdc) - 270 channels {k at 200 s/s, l6 at 
100 s/s, 25 at 50 s/s, 125 at 10 s/s, and 100 at 1 s/s). 

C. Parallel Digital - 1 - l6 bit channel at 200 s/s 

1 - .16 bit channel at 50 s/s 
1 --32 bit channel at 10 s/s 
1 - 2l+ bit channel at 10 s/s 

2B - 8 bit channels at 10 s/^s 
1 D. Serial Digital - 1 - Uo bit channel at 50 s/s 

In the LBR mode the PCM equipment is capable of handling the following 
input signals: 

A. Low Level Analog (O-Uo MVdc) - 10 channels at 1 s/s 

B. High Level Analog (0-5 Vdc) - 100 channels at 1 s/s (these channels 
are not necessarily the same as the HBR 1 s/s channels). 

- C. Parallel Digital - 1-32 bit channel at 1 s/s 

28-8 bit channels at 1 s/s 
D. Serial Digital - 1 - 1|0 bit channel at 10 s/s 

In the HBR mode, the PCM output data is shifted out serially bit by 
bit (one bit period equals approximately 19.5 ysec - 51.2 kbps); eight bits 
make a word (one word period equals approximately 156 ysec - 61iOO words/sec); 



T-16 



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AS-501 



128 words make a prime frame (one prime frame period equals 20 milliseconds - 
50 prime frames/sec); and 50 prime frames make a subframe (one subffame 
period equals 1 second - 1 subframe/sec) . In the LBR mode, the PCM output 
data is shifted out serially bit by bit (one bit period equals 625 milli- 
seconds - 1.6 kbps); 8 bits make a word (one word period equals 5 milli- 
seconds - 200 words/sec); 200 words make up a prime frame and a subframe 
(one prime frame or subframe period equals 1 second). 

7. 3.1 Prosramer 

The programer maintains primary control of all functioning units of 
the PCM equipment. It uses CTE or internal clock pulses to generate gating 
and control signals for sampling and processing data. 

T.3.2 Analog Data 

Each analog data input is connected to an analog gate. The analog 
gates are grouped in matrix combinations, basically arranged such that 
five channel inputs (columns) are provided per row with multiple rows used 
to form the desired matrices. The five analog gates in each row utilize a 
common Sequencer Gate (SG). Drawing T.3.1 shows the matrix combinations 
used for low level and high level analog signals. A row command and a 
column command must both be present for an analog gate to transfer the 
data signal via the sequencer gate and high-speed gate or low-level 
amplifier to the coder. The coder uses a successive approximation tech- 
nique to code the analog to 5 Vdc signal into an 8 bit binary coded 
decimal word. This 8 bit word is fed to the digital mutliplexer where it 
is transferred to the output shift register when the coder gate is inter- 
rogated. 

7.3. 3 Parallel Digital Data 

Each parallel digital (or event) signal is connected to a gate in 
the digital multiplexer. When the digital multiplexer gate is interrogated, 
the output shift register will be loaded with a "l" if the input signal is 
greater than 3.5 Vdc or if the input is open circuited. The output shift 
register will remain a zero provided the input signal is less than 0.5 Vdc 

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and the input impedance is less than 5 K ohms. 

' W T.3.U PCM Telemetry Parameters 

Tables 7-2 through 7-11 list the PCM telemetry parameters that 
will be available for real-time analysis. Analog parameters recorded 
on the flight qualification recorder and the data storage equipment are 
not listed. Also, onboard display and GSE parameters are not listed. 
Information on parameters not listed is available in the Apollo Baseline 
Master Measurements list and/or the Apollo Instrumentation Equipment List 
for Spacecrafts 17 and 20. 

T. 3.^.1 Column Definitions for Tables 7-2 Through 7-11 
A. Column 1 - Parameter No. 

The Parameter Niamber consists of seven characters: 
Two letters follow by four niambers and one letter. 

1. Module Code 

The first letter designates the measxarement location by 
module. 

a. A - ADAPTER 

b. B - BOOSTER 

c. C - COMMAND MODULE 

d. L - LAUNCH ESCAPE TOWER 

e. S - SERVICE MODULE 

2. Fimctional Subsystem Code 

The second letter denotes the subsystem wherein the < 
\ measurement originates. 

a. A - STRUCTURES 

b. C - ELECTRICAL POWER 

c. D - MASTER EVENTS SEQUENCE CONTROLLER 

d. E - EARTH LANDING SEQUENCE CONTROLLER 

e. F - ENVIRONMENTAL CONTROL 

f . G - GUIDANCE & NAVIGATION 

g. H - STABILIZATION & CONTROL 



7-18 



I I I I I M I I ! I T f r I i [ f T f f 



I 



CSM 
AS-501 



B. 



C. 



D. 



h. J - CREW EQUIPMENT 

i. K " FLIGHT TECHNOLOGY 

J . L - SCIENTIFIC EQUIPMENT 

k. P - SERVICE PROPULSION ' 

1. R - REACTION CONTROL 

m. S - L/V EMERGENCY DETECTION 

n, T - COMMUNICATIONS & INSTRUMENTATION 
3- Discrete N-umber 

Characters three through six are discrete ntanbers listed 

sequentially vithin each subsystem. 
Column 2 - Nomenclature 

The nomenclature is a brief, definitive title given to each 
measurement. Standard abbreviations are used, where applicable. 
Column 3 - TM 

The telemetry col-umn contains two characters. The first 
character indicates the type of signal measured as follows: 
H - High level analog 
L - Low level analog 
E - Event (one bit) 

D - Digital word (more than one bit) 
S - Serial digital word (more than one bit) 
The second character indicates the telemetry format as 
follows: 

1 - HBR format only 

2 - HBR and LBR fomats 
Column k - Channel Code 

The channel code contains information giving the high bit 
rate (HBR) sample rate, the type of parameter measurement 
(analog, digital, etc.), and the channel code number associated 
with that type of measurement. 



T-19 



1 1 1 1 1 1 1 1 1 f ! 1 1 r 1 1 1 ! [ f I 



n 



CSM 
AS-501 



EXAMPLE: 



\ 



L 



■CHANNEL CODE NUMBER 

A-ANALOG, HIGH LEVEL (0-5 Vdc) 

AL-ANALOG, LOW LEVEL (O-UO MVdc) 

D-DIGITAL, PARALLEL 

DS-DIGITAL, SERIAL 

E-EVENT 

-NUMBER OF ZEROS FOLLOWING FIRST 
DIGIT IN HER SAMPLE RATE 



n 



— ^FIRST DIGIT IN HBR SAMPLE RATE 

E. Column 5 - INST BUS 

The Instrtnnentation Bus column will indicate the power 
source for that parameter. Some measurements are generated 
and/or signal conditioned within a particular system, using 
that systems' power and will be indicated by G&N, S&C, ELECT, 
and C&I. 

The instrumentation systjsm distributes +28 Vdc power directly 
to certain sensors and signal conditioners. These instrumen- 
tation power buses are shown on Dwg. 7.2.1 and designated 
Vli. Vj2. Vj3g, Vjj^g. Vi8> VjQg, Vjp, VjiQ. Vg^^, and Vsc5s- 
VjQg and Vg^^g are created from Vjg and Vg^ respectively 
and feed sensors located in the service module. 
F. Column 6 - SCE Cond 

The SCE Conditioner coliomn uses the following code to indicate 
the type of SCE module that is used to condition the signal. 
See Section 7.2.3 notes for module description. 
AA - dc Active Attenuator module 
AIA - dc Active Inverter Attenuator module 
DA - dc Differential Amplifier module 
DBA - do Differential Bridge Amplifier module 



7-20 



1 1 1 1 1 1 1 1 1 ! II 1 1 r 1 1 If 1 1 



CSM 
AS-501 



LGA - dc Low Gain Amplifier modtile 
AC - ac to dc Converter modiile 
PD - Bi-phase Demodulator modiile 
FD - Frequency Demodxilator module 

7. 3.5 PCM Telemetry Formats 

The HBR format, shown on drawing 7.3.1, gives the bit stream 
word and prime frame location for each channel. The 128 word blocks 
make up a prime frame. A 200 s/s channel will appear in k word blocks, 
a 100 s/s channel will appear in 2 word blocks and a 50 s/s channel will 
appear in 1 word block in each prime frame. The 10 s/s channels will 
appear in every fifth prime frame. The 1 s/s channels will appear in 
only one prime frame. Prime frame locations are given in the word 
blocks by the following code: 

A. A - All prime frames 

B. B - Prime frames 1, 6, 11, I6....U6 
CO- Prime frames 2, 7, 12, 17 ^7 

D. D - Prime frames 3, 8, 13, I8 k8 

E. E - Prime frames h^ 9, 1^, 19 ^+9 

F. F - Prime frames 5, 10, 15, 20 50 

G. 1, 2, 3, 50 - Prime frames 1, 2, 3 50. 

The LBR foimat gives the word location (l thru 200) of each 
channel appearing in the LBR bit stream. 



7-21 



1 1 1 II M 1 1 1 1 1 1 in ill ! I 



HBR FOPMAT 





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79 

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12A6 
74 






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110 Id 
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51A22 




12A9 
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103 






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12A13 


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47 
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117|l22 




22A3 


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19- C 

23 

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12A14 
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74 

99 

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50 
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61 


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64 




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a 


81 


20 


87 


21 


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103 


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110 Id 


104 






9 


131 


10 


122 


23 


123 


23 


134 






11 


141 


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25 


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144 






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161 


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162 


26 


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164 










27 


182 


28 


18 




104 






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25 


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21 


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26 


46 


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31 


65 


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66 


33; 67 




61 






41 


85 


51 


86 


43 87 


52 


St 






51 


109 


63 


106 


53 


107 










6II125 


76 


126 


63 


137 


77 


13( 






71 i 145 


88 


146 


73 


147 


89 


14f 






81.165 


101 


166 


•3 


161 


102 


16f 






91 


185 


113 


186 


93 


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188 








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29 




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5DS 


69 
89 

109 
129 
149 
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5DS 


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110 
13t 
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190 


MS 


71 
91 
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13: 

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92 
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153 
173 
193 






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5: 


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7: 


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11; 


66 


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13: 


79 


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15; 


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76 


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17: 


104 


17' 


86 


17: 


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193 


116 19 


96 


19: 


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77 157 


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POWER SERVO 
ASSEM&LY 

(PSA) 



li 



,, Y, Z 
CYBOTWQ^INfi~X,Y,Z 



^ a VELOCITIES 



OPTICAL 
CONTROLLER 



INERTIAL 
MEASURE- 
MENT 
UNIT 

tIMU) 



SCANNING 

TELESCOPE 

(SCT) 



OPTICS 
COUPLING 
DISPLAY 

UNITS 
(ODCO'S) 



JTT 



T*mm< > 



OCDUj A 



OCDU 
MANUAL 

CONTROLS. 



MANUAL 

MINIMUM 
IMPULSE 
CONTROL 



scs 



QtH SYNC SWITCH 01 



MANUAL 
DCADBANO 

SELECT 



'' ^1 ; 



TX 




ORBITAL RATE 



B-jfa 



r. 



THRUST ON/ OFF 



I 1 



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L PROPORTIONAL CONTROL 



MANOAL 
ROTATIONAL 
CONTROL 



DIRECT ROTATION 



TTTTTJ 



ENTRY 
ROLL 
YAW 
COUPLMC 



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DRECT 
MODE 
SW ON 



MANUAL 
CHANNEL 
DISABLE 



MANUAL 
TRANSLATIQNAL 

CONTROL 



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" t T t T " " 



,^_" p^.vw, 



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SP5 THRUST 
OIL/ OFF 
LOGIC 



PROtnOES GND 
TO SPS FUEL, 
AND OX SOLS 



INHIBIT P«Y HCS 



DWECT ULLAGE 



P 



NOT APPLICABU TO AS-Hl 



SA ^ SHAFT ANGLE 

TA = TRUNNION ANGLE 

OG, MC, tG, * OUTER, MIDDLE, MMER GWBAL 

• =ROLL 

t= PITCH 

f = YAW 

Ets, E*^ Efs' ROLL, PITCH AND YAW STABH.ITY ERROI 

AOG, AMC, AIG s OUTER, MIDDLE AND INNER GIMBAL AM 

FDAI - FLIGHT DIRECTOR ATTITUDE tNDKATOR 

AGCU " ATTITUDE GYRO COUPLING UNIT 

RGP = RATE GYRO PACKAGE 

itC, tK"^ ENGINE GIMALLINC PITCH AND YAW 

Eta E«a EfB, » ROLL, PITCH AND YAW BODY ERRORS 

BMAG = BODY MOUNTED ATTITUDE GYRO 

i == ROLL RATE 

K == ROLL ACCELERATION 

KDU = tNERTIAL COUPLING DISPLAY UHtT 



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BLOCK DIAGRAM 



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MISSIQM 

AS-501 



8.1.1 



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8.2 INERTIAL SUBSYSTEM 

The inertial subsystem, one of the three subsystems of the G&N 
system, consists of an IMU (which contains the inertial sensors) a. 
Coupling Display Unit (CDU), power supplies in the Power Servo Assembly, 
and the required controls for the operation of said equipment, all found 
in the lower equipment bay. 

The inertial system is aligned before flight by gyrocompassing 
and since the system will be turned off for approximately 80 percent 
of the time, it will be aligned in flight by the flight crew using 
the optical subsystem. In addition to turning the system off completely, 
the gyros could drift over a prolonged period of use and thus require 
realignment. 

The inertial subsystem is used in spacecraft guidance to determine 
the direction and magnitude of the required velocity corrections to 
be applied to the spacecraft. The Apollo stable member (IMU) is a 
three gimbaled system containing three integrating gyros for sta- 
bility and three orthogonally mounted pendulous accelerometers for 
velocity and position determination. Once the inertial subsystem is 
energized and aligned, any rotational motion of the spacecraft will be 
about the gimbaled stable member, which will remain fixed in space. 
Resolvers mounted on the gimbal axis act as angular sensing devices 
and measure the attitude of the spacecraft with respect to the stable 
member. These angular mesaurements are displayed to the flight crew and 
are sent to the AGC via the CDU's. 
J The CDU's couple the IMU to the AGC and generate steering and 

alignment signals. The inertial subsystem contains three CDU's, one 
for each gimbal. Each CDU contains resolvers, which are positioned 
either manually by the astronaut or automatically by the AGC. One 
of the major signals generated by the CDU's is the A angle which is 
proportional to the angular difference between the IMU gimbal and 
I CDU resolver positions. This signal can be used to control the 

8-3 



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position of the IMU or the CDU's, to display spacecraft attitude error, 
or to develop attitude error signals for steering during thrusting man- 
euvers or for spacecraft attitude control. 

If the attitude of the spacecraft is to be changed by the AGC, the 
AGC drives the CDU's to the desired P, R, Y settings. The error (IMU vs 
CDU) is transmitted to the SCS and the appropriate RCS jets fire to null 
the error. In this manner the attitude of the spacecraft is under 
control of the AGC. 

8.2.1 G&N Modes 

There are six inertial modes of operation that are used in con- 
junction with the eight Stabilization and Control modes of operation. 
The eight SCS modes are: 



A. G&N Attitude Control 
G&N AV 
G&N ENTRY 
Monitor 



B. SCS Attitude Control 
SCS AV 
SCS ENTRY 
Local Vertical 



When any of the modes in Group A is selected, the IMU is the reference 
as displayed on the Flight Director Attitude Indicator (FDAI). When 
Group B is being utilized, the FDAI reflects the Body Mounted Attitude 
Gyro (BMAG) information as the prime attitude data. 

Group A is utilized when the primary guidance system is in control 
of the spacecraft. That is, the inertial subsystem will interface with 
the above mentioned modes in the SCS. The IMU however, is the prime 
reference system. Group B is utilized when the guidance system is proven 
to be inoperative or has been turned OFF by the flight crew. These 
modes of operation will not interface with the prime system and will 
utilize the BMAG's as their reference system as opposed to the IMU. 

The exception to this rule is the utilization of the SCS Attitude 
Control Mode (for spacecraft stability) for IMU alignment. 

The following Table 8-1 shows the Inertial Subsystem modes and their 
associated SCS modes. A brief description of the inertial modes is included. 

8-U 



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TABLE 8-1 INERTIAL SUBSYSTEMS MODES AND ASSOCIATED SOS MODES 



I ■ 



INERTIAL SUBSYSTEMS 

1. Attitude Control 

2. Fine Align 

3. Coarse Align 
k. CDU Manual 

5. Zero Encode 

6. Entry 



SCS MODES OF OPERATION 

G&N Attitude Control 
G&N AV 

SCS Attitude Control 

SCS Attitude Control 

SCS Attitude Control 

SCS Attitude Control 

G&N Entry 



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8.2.2 Attitude Control - IX IMU and IX CPU Resolvers Utilized (IG, MG, OG) 
In the Attitude Control Mode of operation, velocity steering or 

attitude control signals are generated so that the IMU gimbal angles 
and the CDU shaft angles remain the same. Any movement of the spacecraft 
causing a change of gimbal position about the stable member will result 
in an error output from the CDU IX resolver which will be applied (SCS 
in the Attitude Control, AV, or Entry mode) to the SCS for spacecraft 
corrective rotational thrusting. 

The error signal which is used for attitude control and velocity 
steering is generated by the IX CDU resolver and is sent to the Stabi- 
lization and Control system through the pitch-yaw resolver and fixed 
resolution transformer. The error signal results from the difference 
between the angular position of the IMU and CDU IX resolvers. While 
in the attitude control mode, the CDU resolver can be positioned 
automatically by the AGC. If velocity steering is being performed, the 
AGC drives the CDU's to control the direction of the thrust vector and 
the CDU error signal (P&Y) is used to position the service propulsion 
engine. 

8.2.3 Fine Align - IX & l6X IMU and the IX & l6X CDU Resolvers Utilized 
(IG, MG, QG) 

Fine alignment will be accomplished by AGC pulsing of the stabili- 
zation gyro torque ducosyn to provide a stable reference from which the 
computer can track spacecraft attitude movements. The CDU's at this time 
are slaved to the IMU and will follow the IMU gimbal angles. The niimber 
of pulses required to fine align the stable member are calculated by the 
AGC based upon optical measurements. This mode may be selected by manual 
operation, computer control, and also by the G&N sync switch. 

The fine align mode performs two tasks; (a) the gyro floats are 
torqued so the stabilization loops reposition the stable member (b) 
the CDU's are driven to repeat the IMU gimbal angles. The torquing 
is accomplished by the AGC plus and minus torquing pulses at a rate of 
3,200 PPS, applied to the ternary current switch. 

8-6 



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The ternary current switch output pulses displace the stabilization 
I ; Syro float from a null, and generate error signals in the stabilization 

loop. The error signal in the stabilization loop causes the torque motors 
to drive the gimbals to position the stable member. The CDU's follow the 
gimbal angles. The angular difference between the IMU and CDU IX and l6x 
resolvers generates a CDU resolver error signal proportional to the 
sine of the angular difference. This signal is sent to the motor drive 
amplifier. The motor drive amplifier output excites the control winding 
of the CDU drive motor and drives the CDU shaft until the IMU and CDU 
angles are equal. Upon completion of the mode, the stable member is fine 
aligned to a reference coordinate system and the CDU readouts indicate 
the gimbal angles. 

When the G&N sync switch is ON and the SCS relay is energized, the 
system will also be in fine align. Two conditions are required to 
energize the SCS relay. First, the SCS must be in the G&N attitude 
control mode. Second, the astronaut must be using the rotational hand 
controller to control the spacecraft. Thus, when a new attitude is 
established, the CDU's having been slaved to the gimbals will have the 
capability of providing the reference for holding the spacecraft at its 
new orientation. 

8.2.1+ Coarse Align - IX IMU and IX CDU Resolvers Utilized (IG, MG, OG) 
The purpose of the Coarse Align Mode is to slave the IMU gimbals 
to an angular position as determined by the CDU. The mode is entered 
by one of four means: (a) pressing the coarse align pushbutton with 

I the inertial subsystem in manual control, (b) AGC command of coarse 

align mode with the system in computer control, (c) IMU tvirned ON for 
less than 100 seconds (d) system in manual CDU mode and the manual align 
pushbuttom pressed. The coarse align light will come on except in the 
case of condition (d) when the basic mode is CDU manual. 

The gimbal IX resolver is excited with an 800 CPS reference and 

j the motor sine and cosine windings are connected to the CDU IX resolver 

8-T 



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motor. When the two motor angles differ, an error signal proportional 
to the sine of the difference is generated by the CDU resolver. The 
signal is sent to the IMU-CDU difference meter and also to the coarse 
align amplifier where it is amplified and demodulated. The demodulated 
signal is sent to the gimbal servo amplifier which drives the IMU 
gimbal until the error between the IMU gimbal resolver and CDU shaft 
resolver is driven out. 

Positioning of the CDU's is, performed by the AGC's pulsing the 
digital-to-analog converter. When the AGC is used to position the IMU 
gimbals, a 3,200 PPS signal is applied to the digital-to-analog converter 
for a 1/12 second interval. The three digital-to-analog converters 
receive pulses sequentially, requiring l/k second to supply all the 
dig it al-to- analog converters with an input. The cycle is repeated 
until the required nimiber of pulses has been issued by the AGC. For 
a small amount of shaft rotation, the digital-to-analog converter provides 
an 800 cps output proportional in amplitude to the number of pulses 
received. The output goes to the motor drive amplifier and drives the 
CDU shaft. The CDU encoder electronics provides pulses to the AGC and 
digital-to-analog converter. The pulses are generated by shaft rotation 
(approx 2,200 pps max). Each pulse is equal to approximately Uo arc 
seconds of rotation of the one-speed resolver. Theoretically, one 
pulse out of the AGC to the digital-to-analog converter should cause 
the encoder to generate one pulse of feedback to the AGC and digital- 
to-analog converter. The pulse input from the AGC to the digital-to- 
analog converter and the feedback from the encoder should cancel each 
other when CDU shaft angle equals the AGC commanded angle. 

8.2.5 CDU Manual - No Resolvers Utilized (DACS Inhibited) 

The manual CDU mode is used as a backup mode to provide the 
coarse align capability in the event of an AGC malfunction, since the 
CDU is positioned only by the AGC in the coarse align mode. For this 
mode, three methods of entry are possible: 



8-8 



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A. Pressing the CDU MANUAL pushbutton with the TRANSFER switch 
in MANUAL. 

B. By AGC command with the TRANSFER switch in COMPUTER 

C. By turning on the G&N SYNC switch on the main D&C panel 
during the attitude control mode. The G&N SYNC switch 
initiated manual CDU mode is used to lock the CDU's and 
hold the spacecraft in the previous attitude reference 
position obtained by pilot manipulation of the rotational 
hand controller. The CDU's are slewed to the desired shaft 
angle by the MANUAL SLEW switch located on the CDU. When 
the manual align pushbutton is depressed, the IMU gimbals 
are driven to the CDU shaft angles. At this time, the DAC 
is disabled. 

8.2.6 Zero Encode - 1/2X & l6x CDU Resolvers Utilized (IG, MG, OG) 

The purpose of the Zero Encoder mode of operation is to drive the 
CDU shaft angle to a zero position utilizing the CDU l6x and 1/2X 
resolvers. AGC CDU registers are cleared and encoder outputs indicative 
of CDU shaft angle changes are thereafter referenced to the zero position 
for use by the AGC, essentially referencing the computer and CDU to 
each other. The zero encoder mode is manually Selected by pressing 
ZERO ENCODER on the IMU control panel. If the system is in manual 
control, has been on for 100 seconds minimum, and zero encoder is 
\ initiated, the following occurs: (a) the ZERO ENCODER lamp lights 

and (b) a signal is provided to disable the digital-to-analog converter 
(DAC) since it will be receiving piilses and could provide an input to 
the CDU motor drive amplifier. 

In the CDU ZERO ENCODER loop, an 800 cps reference voltage is 
applied to the stators of the l6x and 1/2X CDU resolvers. The sine 
windings of the resolvers are applied through a limiter circuit or two 
speed switch to the motor drive amplifier inputs. The motor drive 
amplifier output signal is applied to the CDU servo motor which repositions 

8-9 



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the CDU shaft until the shaft is at the zero angle. At this time the 
sine of the angle is zero and the resolver outputs are zero. The 
maximum time for CDU zeroing is approximately 20 seconds. 

8.2.7 Entry - l6X CDU & IX IMU Resolver Combination In OG (IG & MG^x ^^ 

IX CDU) 

The purpose of the Entry mode is to reduce the response time 
required for generating an attitude error signal in the outer CDU which 
controls the roll of the command module during the entry phase. It 
is also necessary to provide ISS error output to the SCS in terras of 
the spacecraft entry axes. 

The entry mode may be entered in manual control or in computer 
control. Mechanization of the entry mode is similar to mechanization 
of the attitude control mode, except that the output of the IMU outer 
gimbal IX resolver is switched to the outer gimbal CDU l6X resolver. 
The roll commands to the Stabilization and Control System during entry 
are supplied by the outer gimbal CDU l6X resolver. By connecting the 
IX IMU resolver to the l6X CDU resolver, the amount of angular rotation 
of the CDU shaft for a given IMU angle is reduced by a factor of l6. 
Therefore, when the AGC commands a roll maneuver of X degrees, the CDU 
is positioned at X°/l6. 

In addition, the fixed resolution transformer is not being 
used at this time since the Navigation Base axes coincide with the 
spacecraft entry axes. 



8-10 



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AS-501 



8.3 OPTICAL SUBSYSTEM 

8.3.1 Notes General - Optical Subsystem 

The optical subsystem for the Apollo spacecraft consists of a 
scanning telescope (SCT) with a 60 degree field of view and a sextant 
(SXT) with a 1.8 degree field of view. The SCT has a single line of sight 
used for acquisition and orbit determination,- while the SXT utilizes two 
lines of sight for angular measurement . A single line of sight of the SXT 
is used to align the IMU. 

The optical instruments are physically attached to the navigation base 
to insure alignment; fiinctionally they perform two services. First, they 
provide the AGC with data obtained by measuring angles between lines of 
sight to celestial objects and, secondly, they provide measurement for 
establishing an inertial reference. 

The optics can be positioned by the flight crew utilizing the optics 
control stick or the AGC can. position the optics. If the optics are to 
be positioned by the AGC, the Optical Coupling Display Unit (OCDU) is 
driven by the AGC to the desired shaft and trunnion settings. The SXT 
will follow as will the SCT depending upon the switch settings on the 
Optics control panel in the Lower Equipment Bay (LEB). The OCDU's 
function exactly as the ICDU's in that they both receive and transmit 
angular data to the AGC. In this manner the computer is cognizant 
of sighting data as performed by the flight crew. Selection as to 
which mode of operation is to be utilized is done manually from the 
optics control panel in the LEB. In addition, the LEB is equipped to 
receive a rotational hand controller on the cable should spacecraft 
orientation be required while performing navigation sightings. 

At the instant the optical sighting is taken, the time of sighting, 
the angles of the optical instruments and the IMU gimbal angles if, the ISS 
is enabled, are recorded by the AGC. Data pertaining to the location of 
the celestial objects and programs for navigational calcTilations have 
been stored in the AGC. The navigational measurements will consist of 
earth orbit determination, midcourse navigation, and IMU alignment. The 

8-12 



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IMU is aligned during flight by measuring the angle between the navigation 
base and each of tvo stars with the sextant, 

8.3.2 IMU Alignment 

Since the Guidance and Navigation System will be turned OFF 
during approximately 80 percent of the mission, inflight alignment is 
required and is accomplished by utilizing the optics in conjiinction with 
the AGC. The navigator will mark; using the SCT, on two coarse align 
stars which will in turn enable the AGC to determine the spacecraft 
orientation. With this data known, the AGC can now complete the coarse 
alignment process by driving the CDU's to the desired settings and the 
gimbal angles will in t\arn follow. After coarse alignment is completed, 
fine alignment is initiated and this time the navigator, utilizing the SXT, 
marks singularly on two fine align stars. The AGC, through the optical 
CDU's will record the angle the star line of sight makes with the stable 
member, and will in turn provide the necessary number of torquing pulses 
required to position the stable member to its exact orientation. After 
torquing, the stable member is inertially fixed, or fine aligned. 



8-13 



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8.3.3 Optics Nay Sightings 
A. R 




Shaft ^ = 0, Trunnion ^ = 



B. 




C. 



Shaft ^ Si li5° Trunnion s 12° 



The navigator has found the desired star 
in the SCT 60 degree FOV and will now, 
utilizing the optics hand controller, 
center the star. The SLAVE TELESCOPF 
switch is in the star LOS position. 



The star is now centered and the navigator 
will press the MARK button. The AGC will 
record the IMU gimbal angles, the shaft 
and trunnion settings of the OCDU's and the 
AGC time. 



The above process is repeated for a second 
star and after identifying the two stars, 
via the DSKY to the AGC, the computer now 
determines spacecraft orientation and 
completes coarse alignment. 



\ 



SXT 




Shaft ^ s 1+5° Trunnion = 12° 



Fine alignment is accomplished using the 
SXT. The SCT is slaved to the SXT and 
will function as an acqiiisition aid for 
the SXT. While centering the star in 
the SCT FOV, it will not necessarily be 
centered in the SXT due to the smaller, 
thus more accurate, FOV. 



8-ll| 



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The SXT now has the star centered in its 
FOV, using the optics hand controller, 
and after identifying the star to the AGC, 
the MARK button is depressed and the 
shaft and trunnion angles, time, and IMU 
angles are recorded by the AGC. 



Shaft < ^ -90° Trunnion < - 10*= 



The above process is repeated, for a 
second star, and the computer initiates 
the required gyro torquing to complete fine 
alignment . 



8-15 



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8.3.^ Orbit Determination 

The scanning telescope (SCT) vith its 6o degree field of view 
is used for orbit determination since the 1.8 degree field of view of 
the sextant is too small. The IMU is aligned and the +X axis of t^e 
spacecraft is along the flight path. The object to be MAEKED comes 
into view and if it is not traveling along the R line of the spacecraft 
(shaft ♦ = 0), the S/C will be rolled until this condition exists. When 
the landmark has traveled down the R line until it is centered in the 
crosshairs, the operator will press the MARK button, at which time the 
SCT shaft and trunnion angles, time, and the IMU angles will all be 
read by the AGC. A series of three or four sightings will determine the 
spacecraft orbit. 



Direction of Motion 




Landmark 



After Roll Correction 




. MARK At This Time 



Trunnion Angle Tracking 



8-16 



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8-18 



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8.5 IMU TEMP MODE SELECTOR 

8.5.1 Notes General - IMU TEMP MODE Selector 

The IMU TEMP MODE Selector is a four-position rotary switch. 
The four positions and their functions are as follows: 



IMU TEMP MODE 



PROPORTIONAL 



AUTO 



fcACKUP 



GAIN PIPA^ 
lyC=^vl 




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A. Auto/OVerride 

The AUTO/OVERRIDE switch position is the normal operating 
position. In this mode, both the temperature 'controJ circuit 
and temperature indicating circuit are utilized. Temperature 
control is furnished by the temperatiire control amplifier 
using the IRIG temperature control sensors. Temperatiire 
indication, utilizing the IRIG and PIPA temperature indicating 
sensors, may be obtained from the temperature indicating bridge 
(test point) or the temperature indicating bridge amplifier 
(telemetry) . 

Normally, heating is accomplished by the control heaters and 
. emergency heaters operating in parallel. If the IMU exceeds 
normal temperature tolerances (+ 1+°F), the system will auto- 
matically switch to the emergency mode and the IMU TEMP light 
will illuminate. When this occurs, the temperature alarm 
relays turn off the control heaters. The emergency heaters 



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continue to operate under the control of an emergency mercury 
thermostat. When the temperature returns to within +1°F of 
normal, the IMU TEMP light will be extinguished and the system 
will switch back to the normal control mode. The system will 
continue to cycle between auto override and emergency modes 
with an attendant cycling of the IMU TEMP light as long as the 
malfiinction exists and the selector switch remains in the AUTO/ 
OVERRIDE position. In this event, the navigator should select 
another mode. The mode selected will depend on whether the 
malfunction exists in the control circuit or indicating circuit. 

B . Pr oport i onal 

The PROPORTIONAL mode is used if a malfunction occurs in the 
temperature indicating circuit causing undesirable cycling to 
the EMERGENCY mode. In the proportional mode, temperature 
control is furnished by the same heater control circuitry 
used in the AUTO/OVERRIDE mode. The temperature indicating 
circuitry is not used. Abnormal temperature will cause the 
IMU TEMP light to illuminate, but will not switch the system 
to the EMERGENCY mode. 

C. Backup 

The BACKUP mode is used if a malfunction occurs in the normal 
control circuit. Temperature control is then furnished by the 
temperature indicating circuits. In this mode, the IMU TEMP 
light will illuminate when the heaters are OFF and will extin- 
guish when the heaters are ON. 

D. Emergency 

In the EMERGENCY switch position, IRIG temperature is con- 
trolled by means of an emergency safety mercury thermostat and 
the emergency heaters. The safety thermostat provides overheat 
protection for the IMU by opening the emergency heater circuit 
when the stable member temperature exceeds 13U°F. When this temp- 
eratiire drops below 132°F the thermostat closes the circuit to re- 
activate the heaters. 

8-21 



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8.5.2 ZERO Button (Sk) 

The ZERO pushbutton is a monentary contact pushbutton used for 
checking the calibration of the temperature monitoring devices. Depres- 
sing this button replaces the TRIG and PIPA temperature indicating sensors 
in the bridges with standard resistances equal to the resistances of the 
indicating sensors at their normal operating temperature thus providing 
a calibration point for the condition of normal temperature. 

8.5.3 IRIG GAIN Button (Sl) 

The IRIG GAIN pushbutton is similar in function to the ZERO button. 
Depressing the IRIG button replaces the IRIG temperature indicating sensor 
in the bridge with a standard resistance equal to the resistance of the 
temperature indication sensor when ,it is operating at 5^F below normal. 
Depressing the IRIG button also tests the temperature alarm circuits and 
will light the IMU TEMP light. 

8.5.i+ PIPA GAIN Button (S2) 

The PIPA GAIN pushbutton performs the same function as the IRIG 
GAIN button with the exception that the resistance used is equivalent 
to PIPA indicating sensors operating at 5°F above normal. 



8-22 



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AS-501 



8.6 APOLLO GUIDANCE COMPUTER (AGC) 

The computer subsystem (CSS) consists of the AGC, the navigation 
panel display, and keyboard (DSKY) , and the main panel DSKY. The AGC 
is a digital computer and is the major lonit of the CSS. The DSKY's 
provide two-vay communications between the flight crew and the AGC. 

The AGC is the processing center of the G&N system. It is both 
a general purpose and control computer. The AGC performs calculations 
and control functions according to a series of instructions which are 
a result of conditions both internal and external to the AGC. 

The AGC provides the following capabilities for the Apollo missions: 

A. It solves the Guidance and Navigation problems required for 
all phases of the Apollo Mission. This falls under the 
general purpose category. 

1. Orbital Integration - Given pertinent data, the AGC can 
calculate, at time T^^, the state vector of the spacecraft 
at time Tx» where X > N. 

2. Thrust Vector Control Calculations - During a thrust maneu- 
ver, the AGC can calculate Vq (velocity to be gained), 

Aq (acceleration to be gained), and from their cross 
product. The cross product is then used by the G&N System 
and other spacecraft systems to steer the spacecraft to 
a given point in space at a given time. 

B. It controls the G&N system and other spacecraft systems. 
Under this control category would be items such as positioning 
the stable member, positioning the optical unit, turning on 
and off the SPS engine, and the generation of timing pulses 
which are used throughout the spacecraft for synchronization 
purposes. 

C. It maintains cognizance of its own operation, that is, it can detect 
many of its own malfunctions. 

The AGC consists of a logic tray assembly and a memory tray assembly. 
A tray assembly consists of a tray, the modules, and the connectors mounted 
on it. The modules contain the AGC's circuitry. 

8-24 



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\ 



CSM 
AS-501 



The AGC is functionally divided into seven blocks: 

Timer 

Sequence Generator 

Central Processor 

Memory 

Priority Control 

Input-Output 

Power 

A. TIMER - The Timer generates all the necessary synchronization 
pulses to ensure a logical data flow from one area to another 
within the AGC. It also generates timing waveforms which 
are used "by the AGC's alarm circuitry and other areas of 

the spacecraft for control and synchronization purposes. 

B. SEQUENCE GENERATOR - The Sequence generator directs the 
execution of machine instructions. It does this "by genera- 
ting control pulses which logically sequence data throughout 
the AGC. The control pulses are formed by combining the 
order code of an instruction word with synchronization 
pulses from the timer. 

C. CENTRAL PROCESSOR - The central processor per f onus all 
arithmetic operations required of the AGC, buffers all 
information coming from and going to memory, checks for 
correct parity on all words coming from memory, and generates 
a parity bit for all words written into memory. 

D. MEMORY - Memory provides the storage for the AGC and is 
divided into two sections : erasable memory and fixed 
memory. Erasable memory can be written into or read from 
and its readout is destructive. Fixed memory cannot be 
written into and its readout is non-destructive, 

E. PRIORITY CONTROL - Priority control establishes a processing 
priority of operations which must be performed by the AGC. 
These operations are a result of conditions which occur both 
internally and externally to the AGC. Priority control 
consists of counter priority control and interrupt priority 
control. Counter priority control initiates actions which 



8-25 



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CSM 
AS-501 



\ 



update cotmters in erasable memory. Interrupt priority- 
control transfers control of the AGC to' one of six interrupt 
subroutines (priority programs) stored in fixed memory, 

F. INPUT-OUTPUT - The input-output section routes and condi- 
tions signals between the AGC and other areas of the 
spacecraft. 

G. POWER - This section provides voltage levels necessary for 
the proper operation of the AGC. 

8.6.1 AGC Characteristics 



Computer Type 

Internal Transfer 

Memory 

Erasable 

FIXED 

Word length 

Number System 

Hardware registers 

Machine Instruction 

Memory Cycle Time (MCT) 

Basic Clock oscillator 

AGC power supplies 

COUNTERS 

Logic 

Parity 



Automatic, electronic, digital, general 
purpose and control 

Parallel 

Random Access 

Core, Capacity = 102^+ words 

Core, rope, capacity = 2^,576 

16 bits 

Binary one's complement 

22 total 

21 total 

11.7 y seconds = 12 ACTION times 

2.0U8 mc 

+3 VDC, +13 VDC 



20 total 

Positive logic 1 = +3 VDC, logic 0=0 VDC 
Odd 

The following is a brief discussion of several AGC characteristics 
lister above. 

A. INTERNAL TRANSFER - Parallel internal transfer implies that 
all bits of a word are transferred simultaneously when a word 
is routed from one area to another within the AGC. 

B. HARDWARE REGISTERS - A hardware register is a flip-flop register. 
Most of the AGC*s hardware registers consist of l6 flip-flops 
and their associated read and write gates. 

8-26 



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CSM 
AS-501 



C. MACHINE INSTRUCTIONS - A machine instruction is one the 
AGC can directly interpret and execute. Machine instruc- 
tions are divided into three groups; regular, invol\m- 
tary, and miscellaneous. Regular instructions consist of 
an order code and a relevant address. They are divided 

into two groups: basic groups; basic and extra code. Involuntary 
instructions are associated with priority control. 
Miscellaneous instructions are associated with testing 
the AGC on the ground. All machine instructions consist 
of one or more subinstructions. A subinstruction takes one 
MCT to perform. 

D. COUNTER - A counter is a l6-bit (word) location in erasable 
memory which is updated when certain conditions (counter 
interrupts) occur. A counter is not a flip-flop register. 

E. LOGIC - The AGC employs positive logic. Positive logic 
implies that the voltage level associated with a Logic 1 
is algebraically higher than the voltage level associated 
with a Logic 0. 

F. PARITY - The AGC uses odd parity, that is all words stored 
in memory containing an odd number of Logic I's. This 
condition makes it possible for the AGC to detect a mal- 
fimction when a word is incorrectly transferred from memory. 

8.6.2 AGC Updates 

The following is a description of AGC updating utilizing the 
portions of the loop from the UPLINK XMITTER to the AGC downlink. 
For this description, we shall consider three bits of data (an octal five) 
starting at the remote site. The AGC however, receives data in terms of 
a keycode, which is a five bit code, recognizable by the computer. 

8.6.3 Command Data Processor (CDP ) 

The CDP will take the three bits and convert it to a keycode 
(5 bits). To this is added the vehicle address, system address, and 



8-27 



1 1 1 1 1 11 1 1 1 ! 1 1 r [ 1 1 IT 1 1 



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CSM 
* AS-501 

a "1" is added to bit 7. The three "bits now have the following format: 

' F 12 3 k ^ 6 7 8 9 10 11 12 13 lit 15 16 17 l8 19 20 21 22 

XXX XXX 100101 11010 00101 

VEHICLE SYSTEMS K R K 

ADDRESS ADDRESS 

The three bits have now grown to 22 bits. The 22 bits are now 
sub-bit encoded and now becomes 110 bits. The sub-bit encoding for the 
vehicle address is different than the sub-bit encoding for the systems 
address and data. A total of 110, bits will be transmitted to the space- 
craft (via the uplink transmitter UHF or USB) to get three bits of 
data onboard. 

8.6.4 S/C Receiver 

The sub-bit encoding is stripped here and the vehicle address 
and systems address is checked. If all check satisfactorily, the data 
is transmitted serially to the uplink counter in the AGC, Bit 7 first. 
When Bit 7 overflows, the uplink counter is left with KKK where the K 
represents the keycode, its' complement , and the keycode again. The 
overflow causes an interrupt alarm in the AGC and then the data (15 bits) 
is transferred to an area of AGC memory called KEYTMP #1. 

8.6.5 KEYTEMP 1 

In the area of KEYTEMP 1, a check is made to see if the KKK 
conditions exist. If so. "K" (5 bits) of data is transferred to an 
area where keycode processing is performed which consists of interpre- 
^ ting keycodes. This keycode process is also used to interpret DSKY inputs 

since the DSKY utilizes the 5-bit Keycode also. 

8.6.6 ASSEMBLY REGISTER 

The keycode ^ now reduced to the original three bits and 
these three bits are transferred to the assembly register. When five 
words are transmitted, (keycodes) the 15-bit assembly register is full 
and an ENTER command will transfer all 15 bits to an area called 



TEMP STORAGE. 



This is an area of core (l4 regs) memory which is set 
8-28 



II III ! I I I! II I II I I [f f I 



CSM 
AS- 5 01 



aside for updating. When all the words associated with this type up- 
date, are into TEMP STORAGE, a FINAL ENTER command will enable the 
AGC to utilize this data in its computations. If no FINAL ENTER is 
received hy the AGC, the data will remain in TEMP STORAGE and not be 
utilized, 

8.6.7 OUT REGISTER k 

The OUT REGISTER k is the method whereby the data transmitted 
to the AGC is returned to the ground for verification purposes. The 
data is transmitted as a UO-bit word in the following manner: 
1 Word 1 / P / Word 2 / P / WORD ORDER CODE 1 
1 15 16 IT 31 32 33 ko 

Words 1 and 2 are identical and the word order code identifies 
the Words 1 and 2 as either marker, identification, or data words. 
The i+O-bit word arrives via the TM system as 5, 8-bit words. On the 
ground, the data received is checked against the data transmitted. 
If it checks a FINAL ENTER it is transmitted to enable the AGC to 
utilize the update. If the data does not compare, corrective action 
is automatically initiated. 

8.6.8 Progr a m No. and Routines For AS-$01 

8.6.8.1 Program No. For AS-$Q1 
OX PRELAUNCH 

00 IDLING 

01 G&N STARTUP AND CHECKOUT 

02 INITIALIZATION 

03 GYROCOMPASSING 

OU OPTICAL VERIFICATION OF AZIMUTH 

05 INERTIAL REFERENCE 

07 SYSTEM TEST 

IX BOOSTER MONITOR 

11 PRE- LET (LAUNCH ESCAPE TOWER) JETTISON 

8-29 



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CSM 
AS-501 



12 POST-LET JETTISON 

IT LET ABORT 

2X COASTING PHASE 

22 LANDMARK TRACKING 

23 STAR/LANDMARK SIGHTING 
27 AGC UPDATE 

3X PRE-THRUSTING PHASES 

I ] 31 ORBIT CHANGE 

32 RETUEN TO EARTH 

33 SPS MINIMUM IMPULSE 
hX THRUSTING PHASES 

Ul ORBIT CHANGE 

1+2 RETURN TO EARTH 

I ' 1 k3 SPS MINIMUM IMPULSE 

5X ALIGNMENT 

51 IMU ORIENTATION DETERMINATION 

52 S-IVB/IMU ALIGNMENT 

53 CSM/mu ALIGNMENT 

5h IMU REALIGNMENT PROGRAM 

^ 6X ENTRY 

61 MANEUVER TO CM/SM SEPARATION ATTITUDE 

62 CM/SM SEPARATION AND PREENTRY MANEUVER 

63 INITIATE ENTRY STEERING 
6k ,05 G INDICATION 
67 FINAL PHASE 
7X ABORT PHASES 



71 FIRST ABORT BURN 

8.6.8.2 Mission AS-501 AGC 

ROUTINE NO. ROUTINE TITLE 

1 ATTITUDE CONTROL MODE CHECK 

2 THRUST CONTROL MODE CHECK 

3 ENTRY CONTROL MODE CHECK 
h FINE ALIGNMENT 



8-30 



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CSM 
AS-501 



ROUTINE NO. 
21 
22 
2U 
25 
27 
28 
29 
30 
31 
33 
3U 
35 
36 
37 
38 

8.6.9 Verb Definitions AS-501 
REGULAR VERBS 
01 
02 



ROUTINE TITLE 

ATTITUDE MANEUVER 

SOS DISCRETE MONITOR 

DELTA V MONITOR 

COARSE ALIGNMENT 

SIGHTING MARK 

AUTO OPTICS POSITIONING 

STAR DATA TEST 

GYRO TORQUING 

BACKUP DELTA V COUNTER 

PRE-THRUST SPS MINIMUM IMPULSE DATA LOAD 

ORBIT PARAMETER DISPLAY 

PRE-THRUST ORBIT CHANGE DATA LOAD 

PRE-THRUST RETURN TO EARTH DATA LOAD 

SPS ENGINE IGNITION 

SPS ENGINE THRUST FAIL 



I I 



I I 



03 
Ok 
05 
06 
07 
10 
11 
12 
13 
Ik 
15 



DISPLAY OCTAL COMP 1 (Rl) 

DISPLAY OCTAL COMP 2 (Rl) 

DISPLAY OCTAL COMP 3 (Rl) 

DISPLAY OCTAL COMP 1, 2 (Rl, R2) 

DISPLAY OCTAL COMP 1, 2, 3 (Rl, R2, R3) 

DECIMAL DISPLAY 

DP DECIMAL DISPLAY (Rl, R2) 

REQUEST WAITLIST 

MONITOR OCT COMP 1 (Rl) 

MONITOR OCT COMP 2 (Rl) 

MONITOR OCT COMP 3 (Rl) 

MONITOR OCT COMP 1, 2 (RiTj^) 

MONITOR OCT COMP 1, 2, 3 (R1,.R2, R3) 



[ ] 



[ ] 



8-31 



I I I I I I I I ! f ! ! I r I I i IT [ I 



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CSM 
AS-501 



] ] ' W 16 MONITOR DECIMAL 

17 MONITOR DP DECIMAL (R1,R2) 

20 REQUEST EXECUTIVE QJON-FLIGHT USE ONLY) 

21 LOAD COMP 1 (Rl) 

22 LOAD COMP 2 tR2) 

23 LOAD COMP 3 (R3) 

' ' 2lt LOAD COMP 1, 2 (Rl, R2) 

25 LOAD COMP 1, 2, 3 (Rl, R2, R3) 

26 SPARE 

27 SPARE 

30 SPARE 

31 BANK DISPLAY (FIXED MEMORY) 
' I 32 SPARE 

33 PROCEED WITHOUT DATA 

3^ TERMINATE CURRENT TEST OR LOAD REQUEST 

35 RELEASE DISPLAY SYSTEM 

36 FRESH START . 

37 CHANGE MAJOR MODE 
END OF REGULAR VERBS 

i 

EXTENDED VERBS 

^^ ZERO (USED WITH WOUN ICDU OR OCDU) 

^^ COARSE ALIGN (USED WITH NOUJM ICDU OR OCDU) 

^2 FINE ALIW IMU 

^3 LOCK IMU 

^^ SET IMU TO ATTITUDE CONTROL 

^5 SET IMU TO REENTRY CONTROL 

^6 RETURN IMU TO COARSE ALIGN 

'*7 OPTICAL TRACKER ON (NOT Bl USE YET) 

50 PLEASE. PERFORM 

51 PLEASE MARK 

52 MARK REJECT (UNTIL BUTTON AVAILABLE) 

8-32 . 



I I II I I 1 I I ! I f I [ [ t I FT f I 



CSM 
AS-501 



53 

55 
56 
57 
60 
61 
62 
63 
6U 
65 
66 
67 
70 
71 
72 
73 
7i* 
75 
76 
77 



FREE (USED WITH NOUN ICDU OR OCDU) 

PULSE TORQUE GYROS 

ALIGN TIME 

PERFORM BANKSUM 

DO SYSTEM TEST (NON-FLIGHT USE ONLY) 

PREPARE FOR STANDBY 

RECOVER FROM STANDBY 

ILLEGAL VERB 

ILLEGAL VERB 

DISPLAY ORBITAL PARAMETERS 

CALCULATE TIME OF ARRIVAL AT LONGITUDE 

CALCULATE LATITUDE AND LONGITUDE AT SPECIFIED TIME 

CALCULATE TIME OF ARRIVAL AT MAXIMUM DECLINATIOW 

PERFORM MANUAL ATTITUDE MANEUVER 

WTVC TAKEOVER 

UPDATE MINIMUM IMPULSE TARGETTING 

UPDATE RETURN TO EARTH TARGETTING 

UPDATE ORBITAL CHANGE TARGETTING 

PERFORM BACKUP LIFTOFF 

PERFORM STATE VECTOR (RVT) UPDATE 

UPDATE LIFTOFF TIME 



I 1 



II 



n 



8.6,10 Noun Definitions AS-501 
NORMAL NOUNS 



COMPONENTS 



00 NOT IN USE 

01 SPECIFY MACHINE ADDRESS (FRACTIONAL) 

02 SPECIFY MACHINE ADDRESS (WHOLE) ICOMP 

03 SPARE 
Ok SPAKE 

05 ANGULAR ERROR ICOMP 

Ob SPARE 

07 CHANGE OF PROGRAM (MAJOR MODE) ICOMP 

8-33 



SCALE & DECIMAL POINT 

.xxxxx 
xxxxx. 

XXX. XX DEG 
OCTAL ONLY 



[ 1 



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CSM 




AS-501 


NORMAL NOUNS 


COMPONENTS 


(USED WITH PTpRASE PERFORM ONLY) 




10 SPARE 




11 ENGINE ON ENABLE 


_^ 



(USED WITH PLEASE PERFORM ONLY) 

12 GIMBAL AMGLES 3C0MP 

13 DELTA VELOCITY MEASURED (VECTOR MAG) ICOMP 
Ik DELTA VELOCITY COUNTER SETTING ICOMP 

15 INCREMENT MACHINE ADDRESS ICOMP 

16 AGC CLOCK TIME (HRS, MIN, SEC) 3C0MP 



IT FINAL ICDU ANGLES 

20 ICDU 

21 PIPAS 

22 NEW ANGLES I 

23 DELTA ANGLES I 

2k DELTA TIME FOR AGC CLOCK 
(HRS, MIN, SEC) 

25 CHECKLIST 

(USED WITH PLEASE PERFORM ONLY) 

26 PRIO /DELAY 

27 SELF TEST ON /OFF SWITCH 

30 STAR NUMBERS 

31 FAILURE INFORMATION 

(FAILREG, SFAIL, ERCOUNT) 

32 SPARE 

33 TIME OF IGNITION 



3C0MP 
3C0MP 
3C0MP 
3C0MP 
3C0MP 
3C0MP 

ICOMP 



SCALE & DECIMAL POINT 



INFORMATION CONVEYED 
BY NOUN CODE ONLY 



XXX. XX DEG FOR EACH 

XXXXX. FT/SEC 

XXXXX. FT/SEC 

OCTAL ONLY 

OOXXX. HRS 
OOOXX. MIN 
OXX.XX SEC 

XXX. XX DEG FOR EACH 

XXX. XX DEG FOR EACH ' 

XXXXX. PULSES FOR EACH 

XXX. XX DEG FOR EACH 

XXX. XX DEG FOR EACH 

OOXXX. HRS 
OOOXX. MIN 
OXX.XX SEC 
XXXXX. 



ICOMP 


XXXXX. 


ICOMP 


XXXXX. 


3C0MP 


XXXXX. FOR EACH 


3C0MP 


OCTAL ONLY FOR EACH 


3e0MP 


OOXXX. HRS 




OOOXX. MIN 




OXX.XX SEC 



1 



8-31* 



1 1 1 1 1 11 ! I ! 1 1 1 r [ \ 1 Iff I 



CSM 
AS- 5 01 



l~l 



NORMAL NQIMS 

3U EVENT TIME (HRS, MIN, SEC) 

35 DELTA EVENT TIME (HRS, MIN , SEC) 

36 DELTA EVENT TIME (MIN/SEC) 

37 SIGHTING IDENTIFICATION 

MIXED NOUNS 
kO GAMMA, 

INERTIAL VELOCITY (Vl) , 

ALTITUDE ABOVE LAUNCH PAD (HPAD) 
kl MAX ACCELERATION (GMAX), 

PERIGEE ALTITUDE (HP), 

FREE-FALL TIME (Tjrp) 
k2 MISS DISTANCE (DELTA Rj , 

PERIGEE ALTITUDE (HP), 

FREE-FALL TIME (Tpp) 
k3 APOGEE ALTITUDE (HA), 

PERIGEE ALTITUDE (HP), 

FREE-FALL TIME (Tpp) 
hk LATITUDE , 

LONGITUDE, 
ALTITUDE (ABOVE FISCHER ELLIPSOID) 
k5 APOGEK ALTITUDE (HA) , 

PERIGEE ALTITUDE (HP)% 

DELTA VELOCITY REQUIRED (DELTA VREQ) 
kb TIME TO EVENT, 

VELOCITY TO BE GAINED (VG) , 

PERIGEE ALTITUDE (HP) 

8-35 



COMPONENTS 


SCALE & DECIMAL POINT 


3C0MP 


OOXXX. HRS 




OOOXX. MIN 




OXX.XX SEC 


3C0MP 


OOXXX. HRS 




OOOXX. MIN 




OXX.XX SEC 


ICOMP 


XXBXX MIW/SEC 


ICOMP 


OCTAL ONLY 


3C0MP 


XXX. XX DEC 




XXXXX. FT/SEC 




XXXX.XN.M. 


3C0MP 


XXX. XX G 




XXXX.X N.M. 




XXBXX MIN/SEC 


3C0MP 


XXXX.X N.M. 




XXXX.X N.M. 




XXBXX MIN/SEC 


3C0MP 


XXXX.X N.M. 




XXXX.X N.M. 




XXBXX MIN/SEC 


3C0MP 


XXX. XX DEC 




XXX. XX DEC 




XXXX.X N.M. 


3C0MP 


XXXX.X N.M. 




XXXX.X N.M. 




XXXXX. FT /SEC 


3C0MP 


XXBXX MIN/SEC 




XXXXX. FT/SEC 




XXXX.X N.M. 



II 



11 



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MIXED NOUNS 

U7 GAMMA AT EI, 

MISS DISTANCE (DELTA R), 
INERTIAL VELOCITY (VI AT El) 

50 TIME TO EVENT, 
DELTA TIME BURN, 
DELTA VELOCITY MEASURED 

51 TIME TO EVENT, 

VELOCITY TO BE GAINED (VG), 
MEASURED VELOCITY CHANGE ALONG XSC 

52 TIME TO EVENT, 

VELOCITY TO BE GAINED (VG), 
FREE-FALL TIME (Tff) 

53 MAX ACCELERATION (GMAX), 
GAMMA AT EI, 
FREE-FALL TIME (Tff) 

5lt COMMAND ROLL ANGLE (BETA) , 
PRESENT ACCELERATION (G), 
RANGE TO TARGET 

55 OCDU X (SHAFT), 

Y (TRUNNION) 

56 MARK DATA X (SHAFT), 

Y (TRUNNION) 

57 NEW ANGLES OCDU X (SHAFT), 

Y (TRUNNION) 

60 IMU MODE STATUS (IN3, WASKSET, OLDERR) 

61 TARGET AZIMUTH, 
ELEVATION 

62 IMPACT LATITUDE, 
IMPACT LONGITUDE, 
HEADS UP/DOWN 





CSM 
AS-501 


COMPONENTS 


SCALE & DECIMAL POINT 


3C0MP 


XXX. XX DEG 




XXXX.X N.MI. 




XXXXX. FT/SEC 


3C0MP 


XXBXX MIN/SEC 




XXBXX MIN/SEC 




XXXXX, FT/SEC 


3C0MP 


XXBXX MIN/SEC 




XXXXX. FT/SEC 




XXXXX. FT/SEC 


3C0MP 


XXBXX MiN/SEC 




XXXXX. FT/SEC 




XXBXX MIN/SEC 


3C0MP 


XXX. XX G 




XXX. XX DEG 




XXBXX MIN/SEC 


3C0MP 


XXX. XX DEG 




XXX. XX G 




XXXX.X N.MI. 


2C0MP 


XXX. XX DEG 




XXX. XX DEG OR XX. XXX DEG 


2C0MP 


XXX. XX DEG 




XXX.XX DEG OR XX. XXX DEG 


2C0MP 


XXX. XX DEG 




XXX.XX DEG OR XX. XXX DEG 


3C0MP 


OCTAL ONLY FOR EACH 


2C0MP 


XXX.XX DEG 




XX, XXX DEG 


3C0MP 


XXX.XX DEG 




XXX.XX DEG 




+/-00001 



8-36 



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CSM 
AS-501 



MIXED NOUNS 


COMPONENTS 


SCALE & DECIMAL POINT 


63 


LATITUDE , 


3C0MP 


XX. XXX DEG 




LONGITUDE /2, 




XX. XXX DEG 




ALTITUDE (ABOVE FISCHER ELLIPSOID ) 




XXX. XX N.MI. 


6k 


PRESENT LATITUDE 


3C0MP 


XXX. XX DEG 




PRESENT LONGITUDE 




XXX. XX DEG 




PRESENT ALTITUDE 




XXXX.X N.MI. 


65 


SAMPLED AGO CLOCK TIME (HRS.MIN ,SEC) 


3C0MP 


OOXXX, HRS 




(FETCHED IN INTERRUPT) 




OOOXX. MIN 
OXX.XX SEC 


66 


SYSTEM TEST RESULTS 


3C0MP 


XXXXX., .XXXXX, 
XXXXX. 


67 


DELTA GYRO ANGLES 


3C0MP 


XX. XXX FOR EACH 


70 


PITCH TRIM 


3C0MP 


XXX. XX DEG 




YAW TRIM, 




XXX. XX DEG 




DELTA TIME TAILOFF 




XXX. XX SEC 


71 


COMMAND ROLL ANGLE (BETA), 


3C0MP 


XXX. XX DEG 




PRESENT ACCELERATION (G) 




XXX. XX G 




PREDICTED RANGE-RANGE TO TARGET 




XXXX.X N.MI. 


72 


DELTA POSITION 


3C0MP 


XXXX.X KILOMETERS 
FOR EACH 


73 


DELTA VELOCITY 


3C0MP 


XXXX.X METERS/ 
SEC FOR EACH 


7U 


DELTA VELOCITY ALLOWABLE, 


2C0MP 


XXXXX. FT/SEC 




DELTA TIME TAILOFF 




XXX. XX SEC 


75 


DELTA POSITION MAGNITUDE, 


3C0MP 


XXXX.X N.MI. 




DELTA VELOCITY MAGNITUDE, 




XXXXX. FT/SEC 




MULTIPLE MARK COUNTER 




XXXXX. 



76 SPAEE 

77 SPARE 



8-3? 



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ri 



CSM 
AS-501 



! 1 



' 1 



00001 
00002 
00003 
OOOOU 

00007 
00011 
00.012 
00013 
OOOlU 
00015 
00016 
00031 
00035 
00036 
OOOUl 
00051 
00052 
00053 
OOO5U 
00060 
00061 
00062 



CHECKLIST FOR USE WITH NOUN 25 

SCS MODE - GJsN ATT CONT 

SCS MODE - G&N DELTA V 

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CSM 
AS-501 



8.7 DISPLAY AND KEYBOARD 

The computer has two DSKY's associated with it, a navigation panel 
DSKY and a main panel DSKY. The navigation panel DSKY is mounted to the 
right and. slightly above the G&N indicator control panel. The main 
panel DSKY is mounted on the Main Display and Control panel. Except 
for a few differences, the DSKY's are functionally identical. 

The DSKY's provide the interface between the flight crew and the 
AGC-. They enable the flight crew to load information into the AGC, 
request information from the AGC, initiate various programs stored in 
the AGC's memory, and perform tests on the AGC and other portions of 
the G&N system. The DSKY's also provide indications of failures and 
operational changes which may occur within the AGC and G&N system. The 
AGC can, under its own control, initiate a display of information or 
request the flight crew to perform an action. 

The navigation panel DSKY keyboard consists of ten digit keys (0-9), 
nine operation keys (+, -, VERB, NOUN, ENTER, CLEAR, KEY RELEASE, ERROR 
RESET, TEST ALARM), and a brightness control. 

Whenever a key, with the exception of the TEST ALARM key is pressed, 
+13 vdc is applied to a diode matrix which generates a unique five bit 
code (keycode) associated with that key. The keycode is inserted into 
bits five through one of the AGC's in register. A key must be released 
before another key is pressed if information is to be properly processed 
by the AGC. The TEST ALARM key does not have a keycode associated with 
it. It is hard-wired into the AGC's circuitry. ERROR RESET has a key- 
code associated with it; however, it is also hard-wJ.red directly to the 
AGC's alarm circuitry. 

The main panel DSKY keyboard consists of ten digit keys, eight 
operation keys, a brightness control, and an UPTL switch. The main 
panel DSKY does not contain a TEST ALARM key. 

The UPTL switch provides the flight crew with the capability of 
blocking or accepting uplink telemetry information. If this switch is 
placed in the BLOCK POSITION, a Logic 1 is inserted into Bit 6 of the 
IN register. 

.' 8-75 ' , 



1 1 1 1 1 M I n 1 1 1 [ [ 1 1 [ T f I 



CSM 
AS-501 



The following is a brief functional description of the DSKY 



keys, 

KEY 

0-9 

+and - 
VERB 

NOUN 

ENTER 

CLEAR 

TEST ALARM 

ERROR RESET 
KEY RLSE 



Enters digital data into the DATA, 
VERB, NOUN, and PROGRAM display. 

Enters the proper sign into the DATA 
displays for decimal data. 

Blanks the VERB display and conditions 
the AGO to interpret the following 
two digital inputs as a verb code. 

Blanks the NOUN display and conditions 
the AGC to interpret the following two 
digital inputs as a noun code. 

Informs the AGC that the assembled 
data is complete and to execute 
the requested function. 

Clears data contained in DATA display 
currently being used. Successive 
pressings clear the upper DATA displays 

Lights the PARITY FAIL, COUNTER 
FAIL, RUPT LOCK, TC TRAP, and 
COMP FAIL alarm indicators to 
insure that they are properly working. 

Tests for the presence of a transient 
or continuous alarm in the AGC. 

Transfers control of the DSKY displays 
from the keyboard to the AGC program. 



! I I 



II 



[ ] 



8-76 



[ 1 



1 1 1 1 1 1 1 1 ! f I ! I r [ I ii [ [ I 



n 



CSM 
AS-501 



8-8 LOWER EQUIPMENT BAY 

The following is a description of the various controls and displays 
foiind in the Lower Equipment Bay (LEB) . 

The LEB is utilized by the flight crew for navigational sightings; 
however, it also functions as the physical housing for the guidance and 
navigation subsystems. The panel members listed herein are referenced 
to the LEB drawing included in this handbook. 

Switch/Button Function 

Provides flooding intensity 



Panel 
100 



Primary Rheostat 
Sec Switch 

Clocks Switch 



IMU/CDU DIFFERENCE 
INDICATOR 



Provides floodlighting intensity for 
secondary floodlights. 

Provides integral lighting for the 
mechanical clocks in the LH LEB (GMT, 
TO EVENT, FROM EVENT) 

Displays the difference between the IMU 
gimbal angles and the CDU shaft angles 
in degrees. The signals applied to the 
meter are the demodulated outputs of the 
CDU single speed resolvers. 



\ 



8-77 



1 1 1 1 1 1! 1 1 ! 1 1 1 1 n [ f f f I 



# 






CSM 
AS-501 


Panel 


Switch/Button 


Function 




100 
(Cont'd) 


Computer Transfer 
Switch 


Controls 
modes . 


selection of I 



Manual 



Computer 



MODE Switches 



ZERO ENCODE 



COARSE ALIGN 



FINE ALIGN 



CDU MANUAL 



ATT CONT 



Permits normal selection of ISS 
operating modes through the use 
of MODE switches. 

Enables AGC selection of ISS 
operating modes by setting' mode 
relays directly. 

Allow selection and display of ISS 
operating modes. Computer controlled 
when TRANSFER switch is in COMPUTER 
position. Manually controlled when 
TRANSFER switch is in MANUAL position. 

Selects zero encoder mode of ISS 
operation. This mode sets the 
shafts and encoders of the CDU's 
and CDU registers in the computer 
to zero. 

Selects coarse align mode of IMU 
operation. This mode positions the 
stable member of the IMU. This mode 
positions the stable member to within 
1.5 degrees of desired inertial re- 
ference attitude. 

Selects fine align mode of ISS 
operation. This mode completes 
stable member alignment to desired 
inertial reference attitude. 

Selects manual CDU mode of ISS 
operation. This mode provides 
for backup-manual alignment of 
the stable member. Stable member 
drives to CDU angles when MANUAL 
ALIGN switch is depressed. 

Selects attitude control mode of 
IMU operation. This mode provides 
attitude and velocity change sensing 
with respect to the space stabilized 
stable member. 



I 1 



I 1 



[i 



1 1 



8-78 



1 1 i 1 1 1 1 1 ! f I! I T n n [ [ I 



n 



CSM 
AS-501 



Panel 

100 
(Cont'd) 



102 



103 



\ 



Svitch/Button 
ENTRY 



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readouts 2X 
TRUNNION, SHAFT 
ANGLE, OUTER GIMBAL 
(ROLL), INNER 
GIMBAL (PITCH), 
MIDDLE GIMBAL 
(YAW) 

G&N condition 
lights 

PONS 

AGO PWR FAIL 
IMU FAIL 
CDU FAIL 
ACCEL FAIL 
GIMBAL LOCK 



IMU TEMP 



Function 

Selects entry mode of IMU operation. 
This mode is similar to attitude 
control mode; however, the gain of 
the roll control loop is increased 
by a factor of l6 to increase roll 
rate and decrease response time. 

Provide visual representation of 
ICDU and OCDU angles. These are 
dnam type readouts. The first 
cylinder represents 360^/rev, the 
second 10*^ /rev and the 3rd .2°/rev. 



Warning lights denote detected 
malfunction. 

Indicates an IMU, CDU, PIPA and/or 
AGO failure. 

Indicates a failure in the Apollo 
guidance computer power supply. 

Indicates a failure in the inertial 
measurement unit. 

Indicates a failure in one or more 
of the coupling display units. 

Indicates a failure in one or more 
of the PIPA's. 

Indicates a potential gimbal lock 
condition in the IMU (middle gimbal 
angle is greater than ± 60° with 
respect to outer gimbal). 

Indicates that the IMU tempera- 
ture is out of tolerance, i.e., 
exceeds normal temperature by 
±4°F (if IMU TEMP MODE switch is 
in AUTO OVERRIDE or PROPORTIONAL 
positions) . 



8-79 



I I I I I If I I f M T II I i If I I 



Panel 

103 
(Cont'd) 



Switch/Button 
ZERO ENCODER 



IMU DELAY 



MASTER ALARM 



10 U 



Sextant 



CSM 
AS-501 



Fimction 

Indicates that the CDU encoders 
are being zeroed; lamp is 
extinguished after all encoders 
have been zeroed. (Remains 
illuminated for approximately 
30 seconds. If AGO commanded 
zero encode mode, indicator may 
be illuminated for approximately 
60 seconds) . 

IMU coarse alignment to CDU 
angles taking place. Stable 
member is caged. Remains 
illuminated for approximately 
100 seconds. 

Red light illuminates to alert 
crev members at lower equipment 
bay of malfunction or out-of- 
tolerance condition. This is 
indicated by illiamination of 
applicable system status lights on 
MDC 10 or 11. 

Upon illumination of MASTER ALARM 
light, the MASTER ALARM switch- 
lights on MDC 3 and lb are simulta- 
neously illuminated and an audio tone 
is sent to each headset. 
The MASTER ALARM light does not 
contain an integral switch. Light may 
be extinguished only by pressing the 
MASTER ALARM switch-light on 
MDC 3 or 18. 

Optical instrument for measuring 
the angle between two objects. 
The SXT is a dual line-of-sight 
instrument used to determine the 
following : 

a. The angle between a landmark 
and a star. 

b. The angle between a star line 

of sight and the navigation base. 

c. Tracking an unknown landmark. 
This information is used by the AGO 
for the following: 

a. Determine spacecraft position. 

b. Calculate required AV corrections 

c. To fine align the IMU 

The SXT has a 1.8 degree field of view 
with a magnification of 28. 



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AS-501 



1 1 " 



Panel 

lOU 
(Cont'd) 



Svfitch/Button 
Scanning Telescope 



Function 



Tmnnion and 
Shaft control 



Trunnion and Shaft 
Angle Display 



105 



Optics hand 
controller 



\ 



CHECK COOLANT 
switch 



Optical instrument used for the 
following: 

a. Tracking a landmark in earth 
orbit. 

b. Identifying and centering a 
star within the sextant field 
of view. 

The scanning telescope has a 60 
degree field of view and a IX 
magnification. 

These controls are manually operated 
by means of a universal tool. They 
are used in the event of optics elec- 
tronic failure. 

Enables manual positioning of the 
shaft and trunnion angles. 

Provides a mechanical readout and 
shaft of the SCT trunnion angle 
that is commanded manually 
or by the computer. 
Drum- type readouts in degrees, 
mechanically connected to the 
scanning telescope (SCT) trunnion 
and shaft drives, respectively. 
(SXT angles are identical). 

Provides electrical commands to 
the optics shaft and trunnion 
drive motors. 

Five-position switch spring- 
loaded to center OFF position. 
Direction of target movement with 
respect to controller movement 
depends upon mode selected by 
CONTROLLER MODE switch. Rate 
of image movement is proportional 
to amount of hand controller 
movement . 

When pressed, applies power to 
floodlamps behind the display and 
control panel. Enables the crew 
to view the IMU quick-disconnect 
couplings through CHECK 
COOLANT windows. 



8-81 



1 1 1 1 1 ! 1 1 T f [ I r [I I [ [fi I 



Panel 

105 
CCont'd) 



Switch /Button 

CHECK COOLANT 
windovs (two) 

MARK switch 



PANEL BRIGHTNESS 
Control 



CHECK MODE LAMPS 
switch 

CHECK CONDITION 
LAMPS switch 

Condition Lamps 
Switch 

ATTITUDE IMPULSE 
group 

Attitude impulse 

control 



ENABLE switch 
ON /OFF 



CSM 
AS-501 



Function 

Permit observation of IMU coolant 
supply system quick-disconnect 
couplings for detecting leaks. 

Supplies an interrupt signal to the 
AGO which commands it to read 
the optics CDU angles, the time,. 
and the IMU gimbal angles (if the 
IMU is operating). 

Provides adjustment of the 
illumination level of all integrally 
lighted G&N system controls and 
displays. In addition, the control 
provides power to the lamp in 
the THRUST ON switch. 

Applies power to all MODE indi- 
cators on the IMU control panel 
LEB-101. 

Applies power to all condition 
(caution and warning) lamps on 
panel LEB-103. 

Provides dc power to G&N condition 
lamps. 

A control stick used to apply 
small rotational thrust impulses 
to the spacecraft by means of 
the service module reaction jets. 
This seven-position switch is spring- 
loaded to center off. The control is 
used to apply one or any combination 
of pitch, roll, or yaw minimum thrust 
impulses to the S/C providing rate 
damping impulses of 2.U arc-minutes/ 
second/pulse or less. One pulse is 
produced each time the control is 
moved from the center position. 

Supplies a signal to the G&N and 
SCS systems which disables the 
active S/C attitude control mode, 
allowing the S/C to drift freely, 
and enables the attitude impiilse 
control. 



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n 



1 " 



Panel 

105 ^ ' \ 
(Cont'd) 



Switch /Button 

ENABLE switch 
(Cont'd) 



' 1 



IMU TEMP MODE 

group 

Mode switch 



AUTO OVERRIDE 



\ 



PROPORTIONAL 



BACKUP 



CSM 
AS-501 



Function 

Toggle-type, solenoid-held-to-on 
microsvitch. The SCS must be in 
the G&N or SCS attitude control 
modes to enable solenoid holding. 
Attitude impulse is not enabled 
in the SCS local vertical mode. 
To enable the attitude impulses ^ 
control circuit, the LIMIT CYCLE 
switch must also be ON. 

Rotary, four-position, three-wafer 

switch. 

Enables the crew to select any 

one of the four modes of IMU 

temperature controls. 

Normal mode of temperature 
control. If IMU temperature 
exceeds normal tolerance by 
-i*°F, system automatically 
switches to emergency mode and 
IMU TEMP light illuminates. 
When a malfunction occurs, the 
temperature alarm relays turn off 
the control heaters. Emergency 
heaters continue to operate under the 
control of an emergency mercury 
thermostat. When the temperature 
returns to within 1°F of normal, the 
IMU TEMP light extinguishes and 
the system switches back to 
the normal control mode. 

Temperature control is furnished 
by same control circuit used in 
AUTO /OVERRIDE mode. 
PROPORTIONAL mode is used if a 
malfunction occurs in the temper- 
ature indicating circuit, causing 
cycling to the EMERGENCY mode. 

Temperature control is furnished 
by temperature sensing circuits. 
BACKUP mode is used if a malfunc- 
tion occurs in the nomal control 
circuit. In this mode, the IMU TEMP 
light will illuminate when the 
heaters are off and will extinguish 



8-83 



I I I I I ! I I I ! I T r T [ I I f i f T 



in 



CSM 
AS-501 



Panel 

105 
(Cont'd) 



Svitch/Button 
BACKUP (Cont'd) 



EMERGENCY 



ZERO svitch 



GAIN svitch 



IRIG 



PIPA 



MAP AND DATA 
VIEWER group 

OPTICS group 
SLAVE 
TELESCOPE switch 

STAR LOS 



LANDMARK 
LOS 0° 



Function 

when the heaters are on. Manual 
switching to emergency mode is nece- 
ssary if the temperature becomes 
excessive. 

A Mercury thermostat provides overheat 
protection for the IMU by opening 
the emergency heater circuit 
when the temperature exceeds 
ISH^F. When this temperature 
drops below 132°F, the thermostat 
closes the circuit to reactivate 
the heaters. 

Used to check calibration of 
IRIG and PIPA temperature 
monitoring devices. 

Pushbutton, moment aary- contact 
switch. When it is pressed, the 
IMU TEMP indicator (LEB-103) 
should be illuminated. If not, a 
malfunction exists in the system. 
Used to check temperature 
alarm circuit. 

Simulates low temperatxire error 
(-5°F) to check IRIG temperature 
sensors and also tests alarm 
circuit . 

Simulates high temperatiire error 
(+5°F) to check PIPA temper- 
atiire sensors and also test 
alarm circuits. 

Function deleted as a result of map 
and data viewer deactivation. 

Single pole, three-position toggle 
switch. 

Slaves SCT trunnion axis to SXT 
trunnion. 

Drives SCT trunnion to zero 
independently of CDU trunnion. 
Zero position is parallel to SXT 
shaft axis. 



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n 



CSM 
AS-501 



Panel 

105 
(Cont»d) 



Switch/Button 
OFFSET 25® 



OPTICS HOLD 

MODE SWITCH 
ZERO OPTICS 



MANUAL 



\ 



COMPUTER 



CONTROLLER 
SPEED switch 



HI 

MED 

LOW 



HI 

MED 

LOW 

CONTROLLER MODS 
switch 



Function 

Drives telescope trunnion to 25® 

offset from shaft axis. 

This will bring the star and landmark 

within the 60° field-of-view of the 

SCT, while rotating ahout the shaft 

axis. 

Opens input to motors to prevent 
CDU creep. 

Selects optics mode of operation. 

Optics and CDU resolvers are 
driven to zero. AGC register is 
set to all zeros. 
Completed in approximately 
60 seconds. 

Normal operating position 
enabling crew to position optics 
by means of optics hand controller. 
In this position, the crew may select 
either the DIRECT or RESOLVED 
controller mode of operation. 

Optics are automatically posi- 
tioned by the AGC. Panel mounted 
controls are disabled. 

Provides attenuation of shaft and 
trunnion slew commands from optics 
hand controller. 

Direct Mode Maximvmi Drive Rates: 
Trunnion Shaft 

8.8°/sec 17.3°/sec 

1.05°/sec 2.06°/sec 
0.105°/sec 0.206°/sec 

Resolved Mode Meiximiaii Drive Rates : 
Trunnion Shaft 

8.8°/sec 8.8®/sec 

1.05°/sec 1.05*^/sec 

0.l05''/sec 0.105°/sec 

Controls control stick output 
configuration. 



8-85 



1 1 1 1 If 1 1 !! II I r r 1 1 f F f F 



CSM 
AS-501 



Panel 

105 
(Cont'd) 



Switch/Button 

DIRECT 



106 



RESOLVED 



Alarm condition 
indicators 

PROG ALM 

COUNTER FAIL 



RUPT LOCK 

TC TRAP 

SCALER FAIL 
PARITY FAIL 
TM FAIL 
CHECK FAIL 
KEY RLSE 



Function 

Applies control movements 
directly to CDU's. Right/left 
movement commands shaft 
rotation. Up /down movement 
commands trunnion angle 
increase/decrease . 
Target moves about arcs with speed 
of target movement varying with 
magnitude of angle. 

Applies control stick movements 

to angular resolving circuits 

before applying commands to the 

CDU's. 

Target moves horizontally and 

vertically at a linear rate. 

Indicate abnormal conditions of 
computer operation. 

AGC program error detected. 

Counter increment instruction not 
executed or not completed within 
10 msec of initiation. 

No interrupt within 80 msec or 
interrupt not completed within 
10 msec. 

Transfer control not executed 
within 10 msec or not completed 
within 10 msec. 

100-pps signal from scaler A of 
computer timing section failed. 

Parity error exists in data word 
from memory. 

Telemetry data rate incorrect or 
transmission incorrect. 

Incorrect DSKY operation 
attempted. 

Computer program cannot pro- 
ceed until operator releases 
DSKY control. 



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AS-501 



I I " 



Panel 

106 
(Cont'd) 



Svitch/Button 
ACTIVITY lights 



I 1 



I ' I 



\ 



UPTL 



COMP 



PROGRAM indicator 



VERB indicator 

NOUN indicator 

REGISTER 1 
indicator 



REGISTER 2 in- 
dicator* 



REGISTER 3 
indicator 



BRIGHTNESS control 



KEY RELEASE 
pushbutton 

TEST ALARM push- 
button 

ERROR RESET 
pushbutton 



Function 

Indicates activity the computer 
is presently engaged in. 

Computer is receiving information 
from telemetry updata link. 

Computer is engaged in 
computation. 

A tvo-digit display, indicating 
the number of the program 
(major mode) presently in 
progress. 

A two-digit display, indicating 
verb code selected. 

A two-digit display, indicating 
notm code selected. 

Displays contents of selected 
register or memory location. 
First component of extended 
data word, if applicable. 

Displays contents of selected 
register or memory location. 
Second component of extended 
data word, if applicable. 

Displays contents of selected 
register or memory location. 
Third component of extended 
data word, if applicable. 

Varies brightness of electro- 
luminescent data displays: 
REGISTER 1, REGISTER 2, 
and REGISTER 3. 

Enables program control of 
DSKY. Releases operator control 
of DSKY circuits. 

Illuminates the alarm displays 
for b;ilb test. 

Resets alarm light relays. AGC 

recycles to start of current 

operation. 

Verifies alarms. Alarms triggered 

by transients should not repeat. 



8-8T 



1 1 III 1 1 1 1 ! M I r r I I f FT f 



CSM 
AS-501 



Panel 

106 

(Cont'd) 



Switch/Button . 
KeylDoard switches 



CLEAR 
VERB 
NOUN 
ENTER 



107 



to 9 

AGC MODE 
switch 

ON 
STANDBY 



Function 

Provides for entering data into 

or commanding operations of the 

AGC. 

Pushhutton-type switches (selectors) 

Each key generates a specific 5-'bit 

key code denoting the instruction or 

number being selected. 

Places all zeros (logic O's) in 
register being loaded. 

Prepares Computer to accept 
next two digits as verb code. 

Prepares computer to accept 
next two digits as noun code. 

Transfers contents of input 
register to central processor 
and initiates execution of 
instructions. 

Denotes data to follow has 
positive decimal value. 

Denotes data to follow has 
negative decimal value. 

Enters binary equivalent of key 
pressed. 

Applies normal power to AGC 

Applies power timing section of 
AGC 

In STANDBY, AGC maintains 
timing signals and other circuits 
necessary to restart. STANDBY 
operation is used to conserve 
electrical power. 



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8-88 



E 1 



1 1 1 1 1 1 E 1 1 f I T I r 1 1 1 [ 1 1 1 



SECTION 9 
CONTROL 



CSM 
AS-501 



1 ' ^ 9.1 NOTES, STABILIZATION AND CONTROL SUBSYSTEM 

9.1.1 Introduction 

The Stabilization and Control System (SCS) is mechanized vithin 
foair subsystems. These subsystems are as follows: 

A. Attitude Reference Subsystem 

B. Altitude Control Subsystem 

] ■ C. Thrust Vector Control Subsystem 

D. Power Distribution Subsystem 

These subsystems and their associated controls and displays in 
conjunction with the other spacecraft systems (Propulsion, Guidance and 
Navigation, Mission Control Programer, and Electrical Power) provide 
the capability for the following: 

E. Automatic control of spacecraft attitude and inert ial velocities 

F. Manual/semiautomatic control of spacecraft attitude and 
inertial velocities 

G. Manual/direct attitude control (ground initiated) 

The intraconnections of the stabilization and control end items, 
^ or black boxes, and their interconnection with the other spacecraft 

i systems are controlled by relays in the Mission Control Programer. The 

following narrative describes the functions provided by each end item. 

9.1.2 Rate Gyro Package 

V The rate gyro package contains three identical rate gyros, mounted 

orthogonally along the spacecraft body axes, and associated gyro elec- 

I tronics. No provision is made for heaters or temperature control of the 

gyros. Each gyro is a single-axis unit, with the input axis determined 
by the gyro mounting fixture. Self -test capabilities are provided by 
torquing coils, which enable the gyro to be displaced at a known rate 
and by spin motor rotation detection circuits, which allow monitoring 
of the gyro spin motor speed. All self -test circuits are completely 

j isolated from operational circuits to prevent a failure in the former 

from affecting gyro opieration. The function of the rate gyros is to 

9-1 



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CSM 
AS-501 



provide an indication on the FDAI of the rate of attitude change in 
pitch, roll, and yaw axes. An attitude change about any of the ajces 
results in an output signal which is representative of the rate of 
displacement. The gyro outputs are used by the SCS as primary damping 
or stabilization signals and, in addition, as negative feedback to null 
rotational control commands and provide a proportional maneuver rate 
capability. 

Each miniature rate gyro assembly consists of a spin motor, 
damping system, gimbal assembly, quadrilever spring, and self -check cir- 
cuitry. The gyro spin motor is a UoO-cps three-phase synchronous hysteresis 
motor powered by 26 volts ac. The maximum time allowed for the gyro to 
come up to operating speed is IT seconds. Damping is accomplished by 
positive displacement of the damping fluid throiogh temperature-controlled 
orifices. The quadrilever spring provides the torsional restraint re- 
quired by the gyro, together with radial support for the gimbal assembly. 

9-1.3 Attitude Gyro Accelerometer Package 

9.1.3.1 General 

The attitude gyro accelerometer package contains three body- 
mounted attitude gyros (BMAG's) and an accelerometer. Electronic control 
circuits for the gyros and accelerometer are contained in the display 
and attitude gyro accelerometer package electronic control assembly. 

9.1.3.2 Body-Mounted Attitude Gyros 

The three BMG's are identical units, mounted orthogonally along 
the spacecraft body axes to sense attitude displacement along the pitch, 
roll, and yaw axes. Each gyro is a single-axis unit, with the input 
axis determined by the physical mounting in the spacecraft. The BMAG's 
have heaters and temperature control circuits which maintain the gyro 
temperature at 170° + 2®F. If a single gyro exceeds the temperature limits, 
a light will be illuminated on the Caution Warning Panel MDC-10. A spin 
motor detection circuit is included in each gyro to allow monitoring 
of gyro spin motor speed. This parameter is on PCM. 

9-2 



1 1 II i 1 1 II f 1 1 1 rn nn 



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I 1 



n 

^ CSM 

AS-501 

The BMAG's provide information denoting the angular displace- 
ment of the spacecraft from a preset attitude. They are initially set 
to a specific space-stabilized orientation; thereafter, any displacement 
from this initial setting results in output signals which are represen- 
tative of the amount of angular displacement. The output signals are 
used to produce attitude error signals for an attitude-hold mode or for, 
display on the flight director attitude indicator. The outputs may also 
be applied to the attitude gyro coupling unit (AGCU) for conversion to 
inertial measurement unit (IMU) axes. (IMU axes differ from the space- 
craft body axes.) The AGCU makes the transformation from body axes to 
IMU axes to allow the BMAG*s to be used as a substitute or backup inertial 
guidance unit for the IMU during the periods when the IMU is turned off 
or malfunctioning. The BMAG's can also be used to provide spacecraft 
f ' I rotational velocity signals to back up any one or all of the rate gyros. 

The BMAG's are single-degree-of- freedom, miniature integrating 
gyros contained in electrically heated individual packages. The heater 
will maintain individual BMAG temperature at 170° + 2°F. Degraded gyro 
operation will result if this temperature is not maintained. The gyro 
spin motors are three-phase 2U,000-rmp synchronous devices, powered by 
I 13.6 volts 1+00 cps from a supply in the attitude gyro accelerometer 

package electronic control assembly. With C/M temperature at 80°F and 
the mounting plate temperature at 55*^F, the maximum time allowed for 
the BMAG to reach operating limits is kO minutes. If the BMAG Power 
V Switch should be cycled and if the BMAG warm-up logic should fail or a 

' single gyro should be out of the + 2° temperature band, none of the gyros 

I will operate until the out of tolerance gyro reaches operating temperature; 

in the case of a logic circuit failure, none of the gyros can be used. 

9.1.3.3 Accelerometer 

The accelerometer is mounted along the spacecraft X-Axis and 
senses velocity changes along this axis. It is a pendulous type accel- 
— I erometer with electronic null and balaiice . The temperatiire is maintained 

at ITO^F + 2°-F under normal operating conditions; however, it will operate 

9-3 



I I III M ! ! ! II I I r i I ITi F 



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AS-501 



with predictable, slightly reduced accuracy vhen the temperature is out 
of tolerance. A signal output is generated by an acceleration along the 
X-Axis. This acceleration causes the pendulous mass to move, resulting 
in a change of coupling between the primary and secondary windings of a 
signal generator. This results in an output signal which is demodulated 
and amplified to provide acceleration information in the form of digital 
'signals to a counter in the Delta V display. The pendulous mass is re- 
turned to null by the balancing action of the electronic caging signal 
when the velocity change ceases. 

9.1.U Electronic Control Assemblies: Pitch, Roll, and Yaw 

The pitch, yaw, and roll electronic control assemblies (ECA's) 
are nearly identical, with slight differences due to different require- 
ments for each ajcis. The ECA's provide the circuitry for input control 
signal processing and SCS mode control and configuration. Input control 
signals consist of attitude -error signals from the SCS BMAG's and from 
the G&N system, rate gyro angular velocity signals +X translational and 
direct rotation commands from the G&N and MCP respectively. Service 
Propulsion System (SPS) engine gimbal position, and rate feedback. Mode 
control inputs are received from relays in the MCP. These input signals 
are applied to logic-controlled relays which enable circuit configurations 
corresponding to the desired mode. ECA output signals consist of re- 
action jet firing commands and SPS gimbal position commands and SPS gimbal 
position commands. The reaction Jet commands are generated in the Jet 
selection logic portions of the ECA's. Preignition SPS engine gimbal 
position commands are generated manually at the AS/GPD by thumbwheels 
which provide input signals to the gimbal control circuits. Postignition 
gimbal control is provided by automatic thrust vector control (TVC) 
circuitry in the pitch and, yaw ECA's. 

9.1.5 Electronic Control Assembly, Auxiliary 

The auxiliary electronic control assembly contains the Attitude 
Gyro Coupling .Unit (AGCU) and service propulsion system thrust on/off 



9"U 



1 1 II 1 1 1 1 1 f II I r [ 1 i [ [ f ] 



! 1 



* CSM 

AS-501 

command circuitry. The AGCU portion of the auxiliary EGA receives BMAG 
signals from the DISP/AGAA EGA and transforms them from spacecraft hody 
axes to IMU axes for display on the FDAI. It also transforms attitude 
set dial signals to hody axes. The SPS engine on/off control circuitry 
in the AUX EGA receives engine on/off commands from the G&N system (G&N 
.AV mode), and off commands from the Delta V display electronics. These 
commands are conditioned and applied through appropriate logic to the SPS 
j en'gine solenoid valves. 

9.1.6 Display and Attitude Gyro Accelerometer Assembly Electronic 
Gontrol Assembly (D/AGAA EGA) 

The display and attitude gyro accelerometer package electronic 

control assembly provides the electronic circuitry required to control 

and power the displays, BMAG's and accelerometer. The DISP EGA portion 

consists of the circuitry necessary to receive and condition the following: 

A. Attitude error signals from the G&N system or the BMAG's to 
the FDAI attitude error indicators 

B. Attitude rate-of-change signals from the rate gyros or BMAGS 
to the FDAI attitude rate indicators 

1 ^ C. Feedback signals from the SPS engine gimbal position transducers 

^ to the gimbal position indicators 

D. Accelerometer signals from the AGAA to the Delta V display 
integrator 

E. Total attitude signals from the G&N subsystem or AGGU to the 
^ FDAI. 

The AGAA EGA portion consists of circuitry necessary to accom- 
plish the following: 

F. Accept and condition BMAG inputs for the AGGU. 

G. Accept and condition AGGU torquing commands to the BMAG^s. 
H. Gontrol BMAG and accelerometer temperature controls. 

I. Gontrol and condition the accelerometer rebalance loop and 
I inputs to the integrator. 

9-5 



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CSM 
AS- 501 



J. Supply reference voltages to the BMG's and accelerometer. 

9.1.7 Displays and Controls 

The Display and Control System provides optimiam capability for 
the crew to monitor, evaluate automatic control maneuvers, and/or manually 
fly the spacecraft. 

.9.1.7.1 MCP RTC kk^ li5, and k6 Backup Rate Commands jg.k.k) 

A. Backup Rate Relays IKU, 1K5, and 1K6 in the MCP 

The Backup Rate Relays provide redundant contacts (two for 
each channel pitch, yaw, and roll) which are energized by 
ground initiated Real-Time Commands (RTC's) through the 
uplink telemetry receiver. When the command is sent and 
a relay energized, the selected BMAG provides rate information 
to the display and control electronics; otherwise the rate 
gyro performs this function. 

B. Attitude Deadband Relay 2K69 (controlled by the G&N system) 
The Attitude Deadband Relay selects minimum or maximxam 
attitude deadband by energizing or de-energizing relay 2K69. 

C. Limit Cycle Switch Relay 2K68 MCP 

The Limit Cycle Relay is energized by a G&N command to 
simulate closure of the Limit Cycle switch and apply pseudo 
rate in all three channels. 

D. 0.05G Switch (2K30 and 2K29) 

During a G&N entry both 2K30 and 2K29 will be energized by 
a G&N O.O5G command. The programer (MCP) provides a backup 
to the G&N at the O.IG to 0.5G level. If G&N FAIL is 
transmitted by RTC i+1, the flight controller has the addi- 
tional option of selecting lifting entry RTC-10 which holds 
spacecraft roll attitude during entry using roll attitude 
gyro inputs and rate damping in the pitch and yaw axes. If 
RTC-10 is not transmitted, the SCS will perform a rolling 
entry upon receipt of the 0.05G backup from the MCP and a 
roll rate signal from the MCP, applied through the input 
normally used by the roll Rotation Controller Transducer. 

9-6 



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CSM 
AS-501 

E. Direct Mode Switch 
I 1 ^ The Direct Mode Switch is a two-position switch which pro- 

■ vides power for direct rotation commands (RTC-13 through 
21 inclusive) when in the ON position. Two Direct Mode 
Switch contacts are provided; one set uses power fron direct 
Bus A and the other from direct Bus B; however, only Bus A 
inputs are used for direct control. 
I 1 F. G&N/SCS Switch (Not Used) 

The G&N/SCS Switch selects whether G&N system or the SCS 
system controls the displays and control electronics. The 
monitor configuration is an exception since this configuration 
can be selected from either the G&N or SCS switch position. 
G. Attitude/Monitor /Entry Switch (in Monitor Position at Liftoff) 
^ ' "^is three-position switch selects whether the system is in 

an attitude, monitor, or entry configuration. The fimctions 
obtainable when the switch is in the ATTITUDE position are 
explained in the switch sections, Delta V switch, and Local 
Vertical Switch. When the Attitude/Monitor /Entry Switch 
is in the ENTRY position, either G&N entry or SCS entry is 
J selected, depending on the position of the G&N/SCS switch. 

When the Attitude/Monitor/Entry Switch is in the MONITOR 
position, monitor is selected. 
H. Delta V Switch (Not Used) • 
\ - I. Local Vertical Switch (Not Used) 

J. Channel Enable Switches (Left in the ON or "UP" Position 
1 at Liftoff) 

Four Channel Enable Switches are provided; one for the 
pitch channel, one for the yaw channel, and two for the roll 
channel. The A&C Roll Channel Enable Switch will be dead- 
faced after CM-SM separation. When the Channel Enable 
_ Switch is in the ON position, power is supplied to the 

1 reaction Jet coils. The mission sequencer supplies power 

9-7 



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CSM 
AS-501 



to enable the driver amplifiers throiigh the Channel 
Enable Switches at latmch vehicle separation or for 
abort conditions at altitudes greater than 25,000 
feet. The mission sequencer removes this pover diiring 
entry when the command module reaches an altitude of 
less than 25,000. - The OFF function of the channel 
disable switches is provided by RTC 5^5 55, 56, and 
57 which grounds the 28U solenoid driver enable signal 
in the respective channel. 

9.1.8 Flight Director Attitude Indicator (Not Implemented) 

9.1.9 Attitude Set/Gimbal Position Display (AS/GPD) - MDC6 

The Gimbal I Position Display portion provides for control and 
display of the SPS gimbal angles. The attitude set portion provides 
for FDAI/AGCU alignment and provides a reference attitude for the space- 
craft. 

9.1.9.1 Gimbal Position Display 

Provides indication of SPS engine gimbal position and allows 
for positioning of the gimbals. 

A. Gimbal Position Thumbwheels (Set to +3.6 degrees yaw and 
+0.U degrees pitch at liftoff.) 

Two thumbwheels, one each for pitch and yaw gimbal. Move- 
ment of the thumbwheels results in the generation of control 
and control logic, which drives the AGCU resolver shaft to 
the commanded position resulting in an output from an angle 
generator to the FDAI ball. The switch must be held until 
alignment is completed. 



9-8 



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9.1.10 Delta V Display (AV) - MDC 7 (Not Implemented) 
' w 

9.1.10.1 AV Remaining (Not Implemented) 

9.1.10.2 AV Set Switch (Not Implemented) 

9.1.10.3 Thrust Control Svitch (Not Implemented) 

This is a three-position toggle svitch (NORMAL-OFF-DIRECT ON). 
The NORMAL position permits the SPS engine to he ignited from the AGC 
through the SCS electronics or from the Thrust-On Momentary Switch (anns 
SPS). The OFF position provides a positive off command to the SPS. The 
DIRECT ON position provides a positive on command to the SPS. 

9.1.10.1+ Thrust-On Pushl3utton Svitch (Not Implemented) 

This is a pushbutton momentary contact svitch and is illvuni- 
nated during engine firing sequences. After the SPS is armed and direct 
ullage has been commanded, this svitch is used to ignite the engine in 
either SCS or G&N Delta V modes. After ignition, the engine continues 
to fire until the AV Remaining counter counts to zero or an engine off 
command is received from the. AGC or a positive off command is generated 
by the manual OFF svitch. 

9.1.10.5 Direct Ulla^^e Relay 1K31 and 1K72 

The Direct Ullage Relay is commanded by RTC 22. Contacts of 
1K31 applies 28 Vdc to the four +X reaction jet direct solenoids (l, 2, 
5, and 6) for a ullage maneuver. Contacts of 1KT2 energize the 
V SPS gimbal motors. 

9.1.11 Rotation Controller (Not Implemented) 

9.1.12 Translation Controller (Not Implemented) 

9.1.13 Attitude Impulse Svitch LEB-U (Not Implemented) 



9-9 



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AS- 501 



9.2 SCS POWER DISTRIBUTION SUBSYSTEM 

9.2.1 Introduction 

The Electrical Pover System supplies the SCS with 28 Vdc from 
two separate buses, V and V ^. In addition, it supplies 115 volt, 
three-phase, UOO cps power from a Y-Connected, four-wire source which 
•also uses two separate buses, Vac and Vac^. The two buses for ac and 
dc power distribution provide redundancy for critical SCS functions in 
case of a power failure on one bus relay. It will be energized to power 
the SCS electronics from the other bus. This power is supplied to the 
SCS through l8 circuit breakers located on the left-hand circuit breaker 
panel. From these circuit breakers, the power is distributed to the SCS 
components through six power switches on the sequence controller and SCS 
power panel. These six switches supply power to those SCS components 
which provide the backup (manual TVC) capability for controlling the 
spacecraft during the critical periods when the SPS engine is firing. 
The Group 1 power switches supply power to the rest of the SCS, except 
for 28 Vdc power which is not routed throiigh the main power switches. 
This is termed the non-switched power. 

9.2.2 28 Vdc Non-Switched Power 

The 28 Vdc non-switched power is distributed to the SCS through 
li+ of the l8 circuit breakers on the left-hand circuit breaker panel. 
These lU circuit breakers are labeled; Yaw, Pitch, A&C Roll, B&D Roll, 
Group 1, Direct Cont, and Group 2. This power is called non-switched 
power since its availability to the SCS is independent of the state of 
the six SCS power switches. The power distribution from these circuit 
breakers will now be considered according to the functions provided. 

9.2.3 Reaction Jet Circuit Breakers 

The Yaw, Pitch, A&C Roll, and B&D Roll circuit breakers apply 
28 volts to the high side of the automatic coils for the indicated jets 
in the reaction control system. The jets can be activated by the SCS 



9-10 



I I I I I ! I I ! f f I I T I 1 I! [ [ 



1 " 



n 



• CSM 

AS-501 

Jet drivers, vhich supply a ground when the Jet should "be on. The power 
from the circuit "breakers goes through the Channel Enahle Switches on 
the control panel. Any of the four channels can he enabled separately 
by means of these switches. Power to the reaction Jets will enable half 
the reaction Jets with V and the other half with V ^. 

9.2.U Group 1 Circuit Breakers 

The two Group 1 dc circuit breakers supply 28 volts to the reaction 
Jet drivers, translation controllers, and control panel. They also supply 
dc power through two of the SCS power switches. The power to the reaction 
Jet drivers goes through a set of relays in the two Master Event Sequence 
Controllers (which are not part of the SCS), through the Channel Enable 
Switches on the control panel, through a set of relays in the SCS, and 
finally to the jet drivers. This power does not turn on the drivers. 
It is merely an enable signal. Thus, it can be seen that the Channel 
Enable Switches on the control panel disable a given channel both by 
removing the power from the high side of the coils, and also by disabling 
the Jet drivers for that particular channel. Notice that in any given 
channel, half the Jets get their power from Bus A and half from Bus B. 
This means that if power on one bus is lost, disabling half the Jets, 
the rest of the Jets on the remaining bus are sufficient to provide 
positive or negative rotation in all three axes but cannot provide trans- 
lation. 

The power which is used for commanding normal translation maneuvers 
by means of the translation controls is routed through the CM/SM Separation 
Switch (C19A1 and C19A2) so that translation control is deactivated after 
CM/SM separation. Prior to 'separation, however, the outputs of the two 
controllers are routed to the Jet select logic to provide three-axis 
translation capability. Two isolation diodes feed the power from both 
Group 1 dc buses to the indicated switches on the control panel. Finally, 
it can be seen that dc power from the Group 1 circuit breakers is also 
—J routed to the SCS power switches. The distribution from these switches 

9-11 



\ 



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CSM 
AS.5OI 



will be covered under the section dealing with the Group 1 power switches. 

9.2.5 Group 2 Circuit Breakers 

In reference to Figure 9.1.1, it can be seen that the two Group 2 
dc circuit breakers supply power to the BMAG power switch. The distri- 
bution from the BMAG power switch will be covered in the section dealing 
with the Group 2 power switches. 

9.2.6 Direct Control Circuit Breakers 

The direct control circuit breakers are used to provide two redun- 
dant sources of power for direct ullage, direct rotation, and the SPS 
solenoid valves. Either one of the two buses is sufficient to provide 
a ullage maneuver, three axis rotational control, and turn-on of the SPS 
engine when the proper sequence of ground commands has been sent. 

Manual direct rotation can be used any time the direct control 
circuit breakers are in and the Direct Mode Switch is ON. Direct ullage 
is available whenever the circuit breakers are in. As a backup capability, 
the SPS engine could be turned on simply by putting the Thrust Control 
Switch in the DIRECT ON position. Thus, it can be seen that the power 
from direct control circuit breakers provides a backup means for the 
critical maneuvers of ullage, rotation, and SPS engine thrust on. These 
backup capabilities are independent of the SCS power switches and would 
be available even if the rest of the SCS were to lose power. 

9. 2.7 Group 1 Power Switches 

There are three SCS Group 1 power switches which supply 28 Vdc 
and 115 Vac power to the SCS. These three switches are shown in the 
upper half of Drawing 9.2.1 (SCS Power Distribution Subsystem). 

A. Partial SCS Power 

B. TVC No. 1 

C. Rate Gyro Power 



I 



I 



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AS-501 



9.2.7.1 Partial SCS Pover 

The Partial SCS Power Switch distributes power to seven units 
in the SCS; the roll, pitch, yaw, auxiliary, and D/AGAA ECA's as well as 
the Delta V Indicator and Attitude Set/Gimbal Position Display. Each 
of the five ECA's contains rectifiers and voltage regulators for gener- 
ating dc voltages from the 115 Vac supply. Unless it is indicated other- 
wise, the various voltages are generated using all three phases of the 
source. 

The dc voltages generated in the pitch and yaw ECA's are used 
to power the control electronics within those units. The roll EGA gen- 
erates power for its control electronics as well as the power for the 
signal conditioning circuits which are used in telemetering SCS infor- 
mation to the ground stations. 

The power supply in the auxiliary EGA is used for the electronics 
in the AGGU and in the SPS engine on-off electronics. The FDAI and 
gimbal position display electronics receive their power from the D/AGAA 
EGA. 

9.2.7.2 TVC No. 1 Power 

The TVC No. 1 Power Switch feeds 115 Vac to the pitch and yaw 
TVC power supplies as well as 28 Vdc and 115 Vac to a power supply in 
the D/AGAA EGA. The pitch TVC power supply is used for operating the 
pitch SPS gimbal drive mechanism which consists of a servo amplifier, 
gimbal position and rate transducers, demodulators, and gimbal trim 
potentiometer. Similarly, the yaw power supply feeds power to the yaw 
channel. 

The power supply in the D/AGAA EGA generates the voltages neces- 
sary for operating the accelerometer and Delta V indicator electronics. 
9.2.8 Group 2 Power Switches 

The Group 2 switches supply power to those components of the SCS 
which are necessary for manual thrust vector control. The three Group 
2 switches are shown in the lower half of Drawing 9.2.1 as follows: 

9-13 



1 1 1 1 1 If I ! f II I r 1 1 1 f f i } 



CSM 
AS-501 



A. BMAG Power 

B. TVC No. 2 Power 

C. Rotation Control Power 

9.2.8.1 BMAG Power Switch 

This switch supplies 115 Vac and 28 Vdc to the toll, pitch, yaw, 
and D/AGAA ECA's. The voltages generated in the D/AGAA EGA are used for 
operating the BMAG's together with the torquing amplifiers and preamplifiers 
in the AGAA. One hundred fifteen Vac, Phase A power is also supplied to 
the yaw, pitch, and roll ECA's for demodulator references. 

9.2.8.2 TVC No. 2 Power Switch 

This switch supplied only 115 Vac power to the redundant set of 
TVC power supplies in the pitch and yaw ECA's. This second set of TVC 
power supplies in the pitch and yaw ECA's is used for operating a second 
gimbal drive mechanism which consists of servo amplifiers, summing am- 
plifiers, gimbal position and rate transducers, demodulators, and gimbal 
trim potentiometers. If an improper high or low gimbal drive clutch 
current is sensed in either servo Amp No. 1 diiring an SPS engine firing, 
there is an automatic switchover to the TVC 2 mechanism. This switchover 
is accomplished in both channels if a manual thrust vector control signal 
is givei;! tjy twisting the translation control handle clockwise. 

9.2.8.3 Rotation Control Power Switch 

From this switch, 115 Vac (Phase A only) power is distributed 
to the roll to provide an ac reference for the rotation control demod. 

9.2.9 Additional Circuit Breakers and Switches 

The following are circuit breakers and switches which are not 
considered part of the SCS but which must function properly if the SCS 
is to function properly. 



9-lU 



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* CSM 

AS-501 

9.2.9.1 Crev Safety Control Panel - MDC I6 

Those switches pertaining to SCS functions are as follows: 

A. Transfer CM/SM RCS Switch 

Two-position toggle switch, CM/SM, provides "backup capa- 
bility for transfer of power from the SM to the CM RCS after 
CM/SM separation. 

B. CM RCS Pressurization Switch 

Two-position toggle switch which functions as a backup to 
the normal automatic pressurization of the CM Reaction 
Control System after CM/SM separation. 

C. CM/SM Separation Switch 

Two-position toggle switch which functions as the command 
for CM/SM separation transferring power (RCS Transfer Switch) 
' from SM to CM. 

D. RCS Command Switch 

The three-position toggle switch (ON/NORM/OFF) which functions 
as a backup to the automatic enabling of the RCS solenoid 
drivers. The center position permits the MESC to permit 

^ its normal function of enabling the RCS solenoid drivers. 

i The ON position provides a manual backup for enabling the 

RCS solenoid drivers. The OFF position inhibits activation 
of the RCS solenoid drivers. 

V 9.2.9.2 Master Caution Warning - MDC 10 and 11 

A. SPS Gimbal Drive Failure Indicator 

I The indicator illuminates indicating failure of a gimbal 

drive motor. 

B. AGAA Temperature FailTire Indicator 

The indicator illuminates indicating failure of one of the 
three BMAG heaters, signifying gyro temperature in excess 
of the design band width of 170 + 2*^F. 

"1 

9-15 



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CSM 
AS-501 



9.2.9.3 Gimbal Drive Control Panel - MDC 3 
A. SPS Gimbal Drive Motor Switches 

The SPS Gimbal Drive Motor Switches are three-position 
toggle switches { ON/AUTO/OFF ) . The ON position is spring- 
loaded and momentarily used for engine start. The AUTO 
position is normal and serves to maintain power for actua- 
tion of the gimbal motors. The OFF position removes power 
from the gimbal motors. Number 1 denotes primary system; 
N-umber 2 denotes secondary (redundant system). 



9-16 



1 1 1 1 1 1 1 1 1 T ! 1 1 r [ I in f I 



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0*TE NATIONAL AERONAUTICS 4 SPACE AOMINISTRATtON 
':i/jk6 MANNED SP*CEC»«*FT CENTER - HOUSTON. TEXAS 



'Jm^ 



STABILIZATION AND CCNTRCL 
POWER DISTRIBUTION 
SUBSYSTEM 



MISSION 



AS-501 



9.2.1 



L8E-L. 



1 1 1 1 1 11 1 1 1 II I rr I \ [TIT 



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CSM 
* AS-501 

9.3 ATTITUDE REFERENCE SUBSYSTEM (ARS) 

The Attitude Reference Subsystem consists of electronics, controls, 
and displays which provide a backup inertial attitude reference to facili- 
tate manual spacecraft attitude control in the event of IMU failure. 

9.3.1 Operation 

The subsystem has five operational configurations: Caged, Uncaged, 
Backup Rate, Alignment, and Attitude Set (not used). These configurations 
are. established by the operation of • panel switches on the Main Display Con- 
sole, lower equipment bay, and MCP which provide inputs to the logic cir- 
cuits on the SCS Logic Diagram. 

9.3.1.1 Caged 

In the Caged configuration of the ARS attitude error outputs 
from the three body axes attitude gyros are summed and converted to a 
navigation axes reference. The signals are then applied to resolvers in 
the AGCU, where a pseudo navigation to Euler rate transformation is 
accomplished and the resolved errors are applied to level detectors in 
the AGCU electronics. Each time the Euler rate inputs rise above 1.5 
Vdc, The level detector causes a stepper motor in the affected axes to 
step one quarter turn, 0.2 degrees of attitude change, and generates a 
second pulse which is applied first to a Euler-to-navigation transformation, 
then to a navigation-to-body axes transformation and then back to the atti- 
tude gyro torquer amplifiers to null outputs of the gyros. In this mode 
(Caged), the AGCU computes spacecraft body rotations and provides three 
attitude outputs, pitch, yaw, and roll which are identical, neglecting 
drift and other errors to the outputs of the inner, middle, and outer 
gimbal ajigles of the G&N inertial subsystem. 

9.3.1.2 Uncaged 

In the \incaged configuration, the attitude gyro outputs are 
applied to the Attitude and Thrust Vector Control Electronics and the 
Attitude Error Display on the Flight Director Attitude Indicator. 



9-lB 



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CSM 
AS-501 



9.3.1.3 Backup Rate 

If any of the rate gyro switches are placed in the BMAG position, 
the output of that gyro is applied to its torquer amplifier. This output 
is now equivalent to spacecraft rotational velocity and is applied to 
the command rate display and the Attitude and Thrust Vector Control 
Electronics. If the AGCU is Caged when Backup Rate is selected, logic 
circuits in the D/AGAA EGA will configure the other two attitude gyros 
into rate gyros; however, only the gyro selected by the Rate Gyro Switches 
on MDC-8 will be applied to the control circuitry and display. 

9.3.1.H FDAI Align MDC-6 

This teim is in reality a misnomer since the alignment is strictly 
a function of the AGCU. The FDAI may or may not indicate the alignment, 
depending upon whether or not an SCS control mode has been selected. Any 
time the FDAI Align button is pressed, the AGCU will align itself to the 
Euler angles indicated on the readout windows adjacent to the Attitude 
Set thumbwheels on MDC-6. If the SCS is in a mode which connects the 
AGCU outputs to the FDAI, the angles in the readout windows will be in- 
dicated by the FDAI under the Navigation Axes Index (cross-hair symbol) 
and the roll bug of the FDAI when the alignment is complete. 

9.3.1.5 Attitude Set Switch MDC-6 

The Attitude Set Switch provides a means of manually orientating 
the vehicle from one inertial attitude to another using the FDAI Attitude 
Error Display, Attitude Display, and the Attitude Set thumbwheels. When 
the Attitude Set Switch is in the ATT SET position, the body axes equi- 
valent of the difference between the attitude set resolver shaft angle 
and the shaft angle of the AGCU Attitude Resolvers is displayed by the 
FDAI Attitude Error Needles. These needles will null when the spacecraft 
has been rotated to the proper attitude. 



9-19 



I 1 I I I 1 I I I I 11 I I I I I [11 ] 



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7^ 



NATIONAL AERONAUTICS * SPACE ADMINISTRATION 
MANNED SPACECRAFT CENTER • HOUSTON. TEXAS 



STABlLIZATtON AND CONTROL 
ATTITUDE 
RFFERENCE SUBSYSTEM 






9,3.1. 



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* CSM 

AS-501 

9.^ ATTITUDE AND THRUST VECTOR CONTROL ELECTRONICS 
^ The electronics depicted on Drawings 9.^.1 and 9.^.2 represent the 

mechanization required for control of the SPS engine gimbal angle and 
the CM/SM reaction Jets. Control of the gimhaled engine and reaction 
Jets is obtained through the utilization of a combination of any of the 
four input devices to obtain one of the three basic control configurations. 

A. Monitor 

B. Automatic Control 

C. Manual Control 

The electronics are configured by logic inputs from switches 
on the Main Display Console MDC-25 and lower equipment bay. These logic 
inputs modify the basic electronic configuration into control modes 
which are each suited to a particular mission phase, i.e.. Launch, AV, 
Attitude Control or Entry. 

9.^.1 Monitor 

The monitor configuration (also a control mode) provides the 
capability of observing outputs of the G&N Inert ial Subsystem (G&N ISS) 
while having manual control of spacecraft attitude and rate damping. 
G&N ISS outputs are monitored on the FDAI Attitude and Attitude Error 
Displays. 

9.^.2 Automatic Control 

The automatic control configuration enables automatic control of 
s^ the spacecraft attitude or thrust vector direction and velocity during 

orbital non-atmospheric flight phases and to guide the spacecraft through 
atmospheric entry to a preselected landing site. This control capability 
is exercised when G&N Attitude Control, G&N AV, G&N Entry Mode has 
been selected. 

9A.3 Manual Control 

When a Manual Control configuration is selected, direct manual inputs 
from the ground are required to control spacecretft attitude. 

9-21 



1 1 1 1 1 II 1 1 ! II I r I n Iff I 



CSM 
AS-501 



These functions may or may not be aiigmented by rate damping. This 
control configuration is present at any time ground commands are trans- 
mitted. 

9.^.^ Mode s of Operation 

9.^.^.1 Monitor Mode 

This is the normal SCS mode for the boost phase and is not normally 
used for orbit or entry. When the CSM/LV separate and the monitor mode has 
been selected, the FDAI Attitude and Attitude Error Displays are driven 
by either G&N or SCS inputs and the angular velocity display is driven 
by the rate gyros. The vehicle may be rate stabilized at ±0.2 degree/second 
in all three axes, however, normally cannot be used after using CSM/S-IVB 
separation. 

9.i+.^.2 G&N At titude Control Mode 

The G&N System inputs control the spacecraft attitude automati- 
cally with the SCS Rate Gyro. Rotational maneuver rates are fixed by 
the AGO program at 0.5 degrees/second. The attitude deadband at this time 
may be selected by the crew at a maximum of 0.5 degrees or 5-0 degrees, 
according to the position of the deadband switch MDC-8 (Min-Max, 
respectively). Manual attitude control can be exercised during this mode 
but will not be effective as the G&N will command the vehicle to the CDU 
desired attitude when the Rotation Control is returned to detent. The 
computer attitude control can be inhibited by the G&N Sync Switch MDC-25. 
As long as the G&N Sync is ON, the G&N will hold the attitude selected by 
crew command, however, when the G&N Sync Switch is turned OFF, the computer 
will command the vehicle to the CDU desired attitude. When the vehicle is 
being maneuvered by CDU imputs to the reaction jets, the vehicle can respond 
to inputs which command spacecraft rotational velocities of 10 degrees/second 
in all three axes. Normally, automatic maneuver velocity will be a maximum 
of approximately k degrees per second in all axes controlled by the AGO 
program. 



9-22 



1 1 1 1 1 1 f 1 1 f ! 1 1 [ 1 1 n 1 [ ] 



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* CSM 

AS-501 

9*h.k.3 SCS Attitude Control Mode (Not Used) 
9.h.k.k G&N Delta V Mode 

During G8bN Delta V Mode, prior to SPS engine ignition, configured 
similar to its configuration for G&N Attitude Control Mode, the one 
significant departxire from the attitude control configuration is in the 
■AGCU which is Uncaged so the attitude gyros can follov spacecraft atti- 
tude changes during the hurn. After SPS engine ignition, the pitch and 
yaw reaction Jet solenoid drivers are disabled to conserve RCS fuel 
and are nbt enabled until one second after SPS engine shutdown. At 
SPS engine ignition, the G&N Control inputs are applied to the SPS Gimbal 
Electronics where they are summed with rate signals from the RCS Rate 
Gyros. Signals to the SCS Electronics are generated by the Apollo 
Guidance Computer through the coupling display unit of the G&N Subsystem. 
During the burn, the computer generates steering commands which are 
sent to the CDU's to adjust the CDU shaft angles, resxilting in an angular 
difference between the IMU gimbal angle and CDU shaft angle. (See 
Drawing 9.3.1.) The sine of the difference angle is applied to the SCS 
Electronics as an error signal. The magnitude and phase of the error 
signal is a function of the computer program and is designed to provide 
a solution to the cross-product steering laws and accomplish multi-axis 
velocity changes. The SPS Gimbal can be driven at a maximum rate of 
0.3 radians per second. Servo Amplifier No. 1 is used at all times -to 
\ control the SPS gimbal angle unless Servo No. 2 is switched online by 

over/under current sensing or Manual Thrust Vector Control is selected. 
(See SCS Mode Logic, Drawing g.h.k.) 
A. G&N MTVC (Not Used) 

9-^«^*5 SCS Delta V Mode (Selected Automatically at G&N FAIL Discrete) 

In this mode, prior to SPS engine ignition, the SCS is configured 
similar to the SCS Attitude Control Mode and has identical control capa- 
bility. At engine ignition. Just as in G&N Delta V Mode, the pitch and 
yaw reaction J^ts solenoid drives are disabled. SPS gimbal angle is 

9-23 



I 1 I I I I I ! I ! I! I M 1 I [f f ! 



CSM 
AS-501 



controlled "by attitude gyros and the integrators in the pitch and yaw 
electronic control assemblies. These integrators compute eg location 
vith error inputs from the attitude gyros and gimbal position from the 
SPS gimbals. The gimbal position input is the form of a positive feed- 
back to the Integrator and enables the mechanization of a very tight 
-servo loop. The integrator output and attitude error amplifier output 
are summed and applied to the gimbal servo electronics. The output of 
this summing point is equivalent to SPS gimbal position. 
A. SCS MTVC Mode (Not Used) 

9.h.k.6 G&N Entry Mode 

Prior to 0.05G sensing, this mode is identical to G&N Attitude 
Control Mode except that the G&N System is outputting navigation ajces 
referenced attitude error rather than body axis referenced attitude 
errors, and the Command Service Module has separated. The systems vill 
cause the spacecraft to deadband ±2 degrees/second within the five- 
degree deadband and may be maneuvered at +10 degrees/second pitch, yaw, 
and roll. After 0.05G switching pitch and yaw axis attitude control are 
inhibited by the pitch and yaw RJC attitude relays and roll angular 
velocity is summed vith yaw angular velocity for entry roll maneuvering. 
The summing gain for roll is equivalent to the tan <21.5 degrees and 
provides a coordinated lift vector roll. The G&N and SCS will maintain 
the lift vector to within +50 degrees of the selected (by roll CDU desired) 
attitudes deadband at a maximum rate of two degrees/second and maneuver 
at a maximum rate of IT degrees/second. 

Crew displays for spacecraft attitude, attitude the IMU, CDU/IMU 
difference and rate gyro package respectively. The AGCU is caged and 
can provide an inertial attitude reference should the G&N fail, providing 
a BMAG has not been switched into Backup Rate Mode. 

9.U.U.7 SCS Entry Mode (Selected at CM/SM Separation and G&N Fail) 

SCS Entry provides an alternate to G&N Entry Mode. Prior to 

9-24 ' 



I I I I I I I I f I ! I I T I 1 11 1 [ 



\ 



CSM 
AS-501 



0.05G Switch activation, the control functions are similar to the con- 
trol fiinctions provided by SCS attitude control mode, except that there 
is no translation capability. After 0.05G Switch activation, pitch and 
yaw attitude control are inhibited by RJC attitude relays and roll rate 
is coupled into the yaw axis after passing throiogh a gain change. This 
signal (roll rate tan <21.5 degrees) insures that entry lift attitude display 
on the FDAI. All rotational maneuvers must be accomplished manually with 
proportional or direct reaction Jet control. Maneuver rates are +5 
degrees/second in pitch and yaw, + 17 degrees/second in roll before 0.05G, 
and +17 degrees /second in roll after 0,05G. The rate deadband is 
+20 degrees/second in all axes before or after 0.05G and the attitude 
deadband is +5.O degrees before G.05G and inactive after 0.05G. 



9-25 



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RELAYS Kl-A AND K2<B ARE 

ENERGIZED BY A CM/SM DEADFACE 

SIGNAL ORIGINATING IN SYSTEM B 

AND A RESPECTIVELY A RESPECTIVEL" OF THE MESC 

> AUTO CM/SM TRANSFER MOTOR SWITCH INPUT 

GREEK SYMSOLOGY 

« PfTCM RATATIOMAL ACCELERATION 
f YAW ROTATIONAL ACCELERATION 
« ROLL ROTATIONAL ACCELERATION 

ALPHABETICAL S'.'MSOLXY 

X X AXIS ACCELERATION 

Y Y AXIS ACCELERATION 

Z Z AXIS ACCELERATION 



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STABILIZATION 

REACTION JET 
on/off CONTROL 



CSM 



MISSION 

AS- 501 



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NATIONAL AERONAUTICS t, SPACE AOMINtSTRATtON 

MANNED sPACecRArr cef*TER - mouston. texas 



STABILIZATION 

AND CONTROL 
MODE LOGIC 






9.4.4 



I I I I I M ! I ! [I I r I I I III ! 



n 



CSM 
AS-501 



9.5 NOTES: SCS SIGNAL LIST 

The SCS signal list, Table 9-1, indicates all the signals, analog 
or discrete, which are generated in or used by the SCS. These signals 
are identified by a three-digit code: 

FIRST DIGIT NOMENCLATURE 

6 Pitch axis control signal 

^ Yaw axis control signal 

<1> Roll axis control signal 

X Not identified by axis 

The last two digits are a niamerical identification. 

EXAMPLE: (|)15 is explained as follows: 



is Roll 



Roll axis Roll to yav coupling 



\ 



"~1 
I 

9-30 



I I I I I II 1 I ! f I I II U IT f f 



CSM 
TABLE 9-1 SCS SIGNAL LIST AS-501 



6,i(^,(t>01 Attitude Error or Backup Rate 

6,4;,<()02 Backup Rate Display 

e,i|;03 MTVC Rate 

e,i|;,<f>Oi+ Attitude Error Display 

0,i^,(()O5 Rate Gyro Output 

B^}\)^(^06 CDU Body Error Command Output 

}p^(\>OJ CDU Entry Mode Error Command Output 

e,ij;,<(>08 Rotation Control #1 

6,i^,<(>09 Rotation Control #2 

e,ij^,(()10 Command Error 

e5ip,({)ll Command Rate 

)(12 Translation Control CCW 

)(13 Translation Control CW 

0,4',(()lU Rotation Control BO 

<()15 Roll to Yaw Coupling 

ei6 + Direct Rotation 

017 - e Direct Rotation 

}\)2_Q + yi) Direct Rotation 

}l)±g - \l} Direct Rotation 

())20 + 'i Direct Rotation 

(j)23 - i Direct Rotation 

(()22 + X Translation 

(()23 - X Translation 

62i+ + Y Translation 

025 - Y Translation 

i(;26 + 2 Translation 

i|^2T " 2 Translation 

^28 Direct Ullage 

029 Pitch Axis Engage 

}P20 Y^^ Axis Engage 

(})31 Roll A&C Engage 

((,32 Roll B&D Engage 

9-31 



I I I I I M I I f [ I I I I t I! 1 1 



i~l 



\ 







CSM 




TABLE 9-1 SCS SIGNAL LIST (CONT'D) 


AS-501 


X33 


Attitude Control Mode 




X3l+ 


G&N AV Mode 




X35 


SCS AV Mode 




X36 


G«sN Attitude Input 





9-32 



1 1 ill f 1 1 1 f II r r [ 1 1 Iff f 



CSM 
AS-501 



9.6 NOTES: SCS POWER LIST 

The SCS pover list. Table 9-2, indicates all the power which is 
generated in and used by the SCS electronics. The power is identified 
by a three-digit code. The first digit identifies the SCS end item, 
from which this power originates, that is: 

FIRST DIGIT NOMENCLATURE 

Pitch ECA 

1 Yaw ECA 

2 Roll ECA 

3 Displays ECA 

k Auxiliary. ECA 

X Not identified with an ECA 

The second two digits are simply a numerical identification. 

EXAMPLE: 009 is explained as follows: 



09 



Pitch axis 26V Phase A 



9-33 



I I I I I ! I 1 I f M I T [ 1 if [ f I 



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1 







CSM 




TABLE 9-2 SCS POWEH LIST 


AS-501 


01 


Rotation Controller Demod 




02 


Pilot Direct 




03 


Co-Pilot Direct 




OU 


Translation Control #1 




05 


Translation Control #2 




06 


5 Vac Rate Self Test 




07 


28 Vdc 




08 


115V ^A 




09 


26V (j)A 




10 


17V < 90" 




11 


±20 Vdc 




12 , 


±30 Vdc 




13 


±1*0 Vdc 




lU 


±U0 Vdc 




15 


28V UoOcps 




16 


28V 800cps 




IT 


-1+ Vdc 




18 


±15 Vdc 




19 


lOV pp sq wave 




20 


-2.5 Vdc 




21 


-10 Vdc 




22 


31.5 Vdc- 




23 


11. 1| Vac UoOcps 




2U 


±12 Vdc 




25 


±25 Vdc 




26 


26 Vac (tiC 




27 


(fiB Ret. 




28 


3.6 Vac 




29 


RCS Enable 




30 


VMB Direct Power 




31 


VMA Direct Power 




32 


Reference Power, Servo Mtr & Att 


Error Amp 


33 
36 


Reference Power, Rate Amp 
AGCU Demod Reference 





9-3U 



I I I I I I I I ! ! II I r [ 1 I f T f F 



CSM 
AS- 5 01 



9.7 RELAY IDENTIFICATION 

9.7.1 Attitude Control and Thrust Vector Control 



Kl 


Rate Transfer Pitch 


K17 


Pitch Auto Control Interrupt 


K2 


Rate Transfer Yaw 


Kl6 


Yaw Auto Control Interrupt 


K3 


Yaw Auto Roll 


K15 


Roll Auto Control Interrupt 


Kk 


G&K Attitude Input 


KlU 


Attitude Impulse Enable 


K5 


MTVC Engage 


K6 


TVC Electronic Transfer 


K7 


SCS Delta Velocity Mode 


K8 


AGAA Attitude Input 


K9 


Entry Gain 


KIO 


RJC Attitude; Pitch and Yaw 


Kll 


RJC Roll Attitude 


K12 


Minimum Deadband, Select 


K13 


Pseudo Rate Cutout 


Kl8 


Roll to Yaw Rate Coupling 


KIO 


Yaw Engine Ignition 


K20 


Pitch Engine Ignition 


Attitude 


Reference System Logic 



K32 Backup Rate Pitch 

K33 Backup Rate Yaw 

KBU Backup Rate Roll 

K35 MTVC Rate Output 

K36 AGCU Cage 

K3T AGCU Cage 

K38 Temporary Hold FDAI Power 

9-35 



I I I I I I I 1 I f I ! I I I i 11 1 ! I 



i~l 



1 

i 





AS-501 


K21 


FDAI Align 


K22 


Attitude Set 


K23 


Displays Entry Gain 


K2k 


Monitor Mode 


K25 


Roll Attitude Error Scale Factor 


K26 


Rate Scale Factor 


K2T 


G&N Control Mode Selected 


K28 


G«sN Entry Mode 


K29 


Ortit Rate 


K30 


Warmup 


K31 


Outband 



\ 



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n 



' 1 



SECTION 10 If^^^^ 

PROPULSION 



10.1 NOTES, SERVICE PROPULSION 

10.1.1 SPS Engine 

The SPS is composed of a rocket engine, propellant storage 
tanks, propellant distribution system, helium pressxarization system, 
and a propellant utilization and gaging system. 

A. Fuel — Aerozine 50, Unsymmetrical Dimethylhydrazine (UBMH) 
and Hydrazine (N^ H^) .(50:50 blend) 

B. Oxidizer — Nitrogen Tetroxide (N^ Ot ) 

C. 0/F Ratio — 2:1 + .05 

D. Engine Life — 500 seconds 

E. Nominal Thrust — 21,500 lb. 

10.1.2 Propellant Quantity 

The total propellant supply is contained within four tanks: an 
oxidizer storage tank, oxidizer sump tank, fuel storage tank, and fuel 
sump tank. The storjage and sump tanks for each propellant supply are 
connected in series by a single transfer line. Each sump tank contains 
a propellant retention reservoir. Propellant quantity is measured by 
two separate sensing systems; primary and auxiliary. The primary 
quantity sensors are cylindrical capacitance probes, mounted axially in 
each tank. In the oxidizer tanks, the probes consist of a pair of 
concentric electrodes with the oxidizer used as the dielectric. In the 
fuel tanks, a Pyrex glass probe, coated with silver on the inside, is 
used as one conductor of the capacitor. Fuel on the outside of the 
probe is the other conductor. The Pyrex glass itself forms the dielectric, 
The aiixiliary system utilizes point sensors mounted at intervals along 
the primary probes to provide a step function impendance change when 
the liquid level passes their location centerline (see Table 10-1 ). 
Comparator circuits monitor propellant quantity and ratio unbalance in- 
formation from both gaging systems. The SENSOR switch will be in the 
NORM position. The propellant quantity gaging system is operational 
only during engine firing. A U.5 second firing period is required before 

10-1 



I 1 I I I f I ! ! ! II I [ [ U [ [ f [ 



\ 

r 

I 











CSM 


« 








AS-501 


TABLE 10-1 SPS POINT SENSOR LOCATION 






HEIGHT 


WEIGHT LB 


WEIGHT LB 


SENSOR STATION 


(INCHES) 


TANK ABSOLUTE 


GAGEABLE 


■ 8 


lU.55 


OX SUMP 


80U.2 


691.2 


7 


26. 8U 


OX SUMP 


2066.6 


1858.1 


6 


33.83 


OX SUMP 


2817.0 


2551+.1 


5 


U2.20 


OX SUMP 


3715.2 


3387.3 


k 


80.00 


OX SUMP 


7TT2.5 


7327.1 


3 


102.39 


OX SUMP 


IOI7I+.5 


9729.1 


2 


116.38 


OX SUMP 


11676.0 


11230.5 


1 


138.76 


OX SUMP 


1U077.9 


13632. U 


MAX LEVEL. 


153.26 


OX SUMP 


15I+8I.O 


15035.5 


T 


22.87 


OX STORAGE 


17130.1 


1668U.6 


6 


59.12 


OX STORAGE 


21033.5 


20588.0 


5 


9U.9O 


OX STORAGE 


2U887.O 


2I1I+I+I.6 


U 


117.21 


OX STORAGE 


27289.0 


268U3.5 


3 


131.06 


OX STORAGE 


28780.8 


28335.3 


2 


133.81+ 


OX STORAGE 


29080.9 


28635.5 


1 


137.1+7 


OX STORAGE 


29l+71.i+ 


29025.9 


MAX LEVEL 


11+5.70 


OX STORAGE 


303U3.I+ 


29897.2 


8 


1I+.22 


FUEL SUMP 


U5.5 


3U5.6 


7 


26.89 


FUEL SUMP 


1061.2 


929.0 


6 


3U.28 ■ 


FUEL SUMP 


IUU5.6 


1277.0 


5 


1+3. 00 


FUEL SUMP 


I89I+.8 


1693.6 


It 


81.08 


FUEL SUMP 


3881.2 


3663.5 


3 


10lt.l6 


FUEL SUMP 


5082.1 


lt86U.5 


2 


118.59 


FUEL SUMP 


5832.9 


5615.2 


1 


1U1.67 


FUEL SUMP 


7033.9 


6816.2 


MAX LEVEL 


156.72 


FUEL SUMP 


7738. 1 


7520.3 


7 


22.22 


FUEL STORAGE 


83I+2.3 


83U2.3 


6 


59.62 


FUEL STORAGE 


10512.0 


102ltU.3 


5 


96.53 


FUEL STORAGE 


12I+38.5 


12220.8 


U 


119. 5U 


FUEL STORAGE 


13639.6 


131+21. 9 


3 


133.83 


FUEL STORAGE 


1U385.3 


1I+167.6 


2 


136.70 


FUEL STORAGE 


1H53I+.I+ 


1I+317.7 


1 


lUo.1+7 


FUEL STORAGE 


1U730.6 


1I+512.9 


MAX LEVEL 


ll+O.Ol 


FUEL STORAGE 


15171.2 


11+953. 8 



10-2 



I I I I I I I I ! f [ I I I I 1 if [ f I 



n 



' ] 



1 



CSM 
AS-501 



propellant quantity information is updated, when the SENSOR switch is 
I j in the NORM position. Ground display drops to zero. 

A. Total Fuel capacity 15,300 pounds 

B. AS-501 loading 10 333 

C. Total oxidizer capacity 30,600 pounds 

D. AS-501 loading 20,66? 



10.1.3 Propellant Utilization 

A propellant utilization valve is installed to the oxidizer 
supply line "between the heat exchanger unit and the engine oxidizer 
main orifice assembly. The fuel supply line connects directly from the 
heat exchanger unit to the engine main fuel orifice assembly. The 
propellant utilization valve is only manually controlled from the CM 

J by the OXID FLOW switch (MDC 20) and is for adjustments in oxidizer flow 

which may be necessary to insure simultaneous propellant depletion. The 
propellant utilization valve consists of a motor-operated, redundant, 
double-blade valve assembly providing increased, decreased, or normal 
oxidizer flow rates. The PU. valve primary or secondary gage servo 
operation is selected manually by the VALVE switch (MDC 20). The valve 

1 flow area is controlled by a three-position OXlD FLOW switch. Increased 

{h^.lk lb/sec). Normal (1+6. 06 lb/sec), and'Decreased {kk.kQ lb/sec). 
There is no groxmd TO on the PU valve positions. There will be no 
ground control of the PU valve. It will be manually placed in the NORM 
^ position prior to CM closeout and remain in that position for the entire 

flight. 

^ 10.1.1+ Gimbal Actuators 

Thrust vector control of the service propulsion engine is effected 
by two servo-controlled electromechanical actuators. Each actuator is 
a sealed unit and contains a primary and secondary motor, a primary and 
secondary set of extend and retract electromagnetic particle clutches, 
a bull gear and a ball nut, a jackscrew, three position transducers, and 
two rate transducers. The actuator control limits in theY - Y axis 

10-3 



I I III f I ! ! f II I r I \ I IF f I 



CSM 

AS-501 



(pitch) allow engine deflections of 6 (+1/2, -0) degrees in either 
direction from a zero degree n\ill offset and 7 {± 1/2, -0) degrees in 
the Z -Z axis (yaw), about a +U degree null offset. A snubbing device 
limits the actuator overtravel to one degree beyond the control limits. 
Actuator maximum rates are 0.13 radians/sec (8 degrees/sec). An 
over cur rent /lander current relay is connected between the actuator drive 
motor and Master Control Programer. The relay provides for drive motor 
starting and current monitoring during actuator operation. The relay 
will automatically trip (50 amps over, 5 amps londer) , removing power 
from the actuator drive motor in the event of either an overcurrent or 
undercurrent condition. The secondary system is protected by a TO amp 
circuit breaker only. During normal operations, both the primary and 
secondary gimbal drive motors are operating, but only the primary set 
of clutches is commanded (two pitch, two yaw). If a malfunction occurs 
in the primaiy drive, the overcurrent relay removes power and provides 
a signal to the SCS to switch clutch commands to the secondary channel. 
The gimbal motors are turned on normally by a signal from the G&N to 
the MCP, which in turn, activates the four motors ON at 1/2 second 
intervals. The gimbal motors may also be turned on by sending RTC 22 
(direct ullage) and RTC 11 (direct thrust ON). Prior to thrusting, the 
engine is initially aligned by the MCP gimbal position set discrete. When 
an extend or retract clutch command is received, the corresponding clutch 
shaft rotates, driving the ball gear and ball nut, which causes the 
Jacks crew to reposition the engine gimbal. The velocity generators and 
position transducers provide feedback to the SCS summing logic and 
thereby controls the Jackscrew. One position transducer in each actuator 
provides a signal to drive the applicable gimbal position indicator on 
MDC 6. 



lO-U 



I I I I I ! { I ! f I I I T I I 1 f T ! ] 



n 



CSM 

« AS-501 

I 

10.1.5 Helium Isolation Valves 

There are two helium isolation valves which are solenoid-actuated 
j poppet valves. The control of these valves is either manual or auto- 

matic. The control method is selected from two three-position switches 
(AUTO, OFF, ON) on MDC 20. Normal switch position is in AUTO. This 
allows the valves to be opened at the SPS fire command signal and closed 
at removal of the fire command. The ON position is a manual override 
position which will open the vfi^ve.at any time. The switch will be in 
] the AUTO position for this flight. 

10.1.6 ftaseous Nitrogen (GNp) Tanks 

Two GN2 tanks are mounted on the engine injector assembly. GN2 
pressure is used to overcome the actuator spring pressure upon SPS fire 
command. There are two manually operated solenoid valves (INJECT PRE- 

I VALVE, MDC 3) installed downstream on each GN2 tank. The prevalves 

will be energized open by the MCP when gimbal motors are commanded ON. 
In the ON position, the valves allow GN2 pressure to flow to two pilot 
control valves, which in turn control the pneumatic actuators. When the 
SPS fire signal is received by the pilot control valves, GN2 pressure 
overcomes the actuator spring pressure and opens the main propellant 

] ball valves.. At removal of the fire command, the GN2 is vented overboard 

through the pilot control valve and the spring pressure closes the main 
propellant ball valves. Each GN2 tank is capable of 30 SPS ENGINE starts. 

10.1.7 NORMAL. OFF, DIRECT-ON Switch 

\ This switch supplies dc power to the control solenoid valves 

thro\agh the MCP. The control solenoid valves are armed by the MCP 

I at LV/SC SEP plus 2.5 seconds by the closure of latching relays Ik^T 

and lk56. This arming supplies +28 Vdc to the control valves enabling 
the SPS engine to be fired by either the thrust vector control logic 
or RTC 11 (DIRECT THRUST ON). If RTC 11 (DIRECT THRUST ON) is transmitted, 
RTC 13 (RESET CMDS 10-12) or RTC 12 (DIRECT THRUST OFF) must be sent in 

-, order to shutdown the SPS ENGINE. When RTC 12 (DIRECT THRUST OFF^ C 11 

10-5 



1 1 1 1 1 1 1 n ! [ I f [ [ 1 1 If f I 



CSM 
AS-501 



is transmitted, the SPS Arm Relays Reset (unlatch) thus removing +28 Vdc 
to the control valves thereby shutting down the SPS ENGINE. The NORMAL, 
OFF, DIRECT-ON Switch will be placed in the NORMAL position prior to CM 
closeout and remain there throughout the flight. 



10-6 



I I I I I I I I I f ! I I T [ I ill [ I 



n 



' 1 



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A. NORMAL SETTLING 
WITH 4 RCS ENGINES 

B, 2 RCS ENGINES 




NOTE: DOES NOT INCLUDE 
1 SEC OVERLAY OF 
SPS AND RCS 

NORMAL CURVES (A) 
CORRECTED FOR 
ATTITUDE CONTROL. 



::::::::s:::::;:!::ss:::;::::::;:::s!:::::e::::::::!::s::s:sss:::!::s 




5 10. 15 20 25 

WEIGHT OF PROPELLANT REMAINING-1000 LBS. 

FIGURE 10.2 PROPULSION - SETTLING TIME VS PROPELLANT REMAINING. 

10-8 



1 1 1 1 1 II 1 1 f M I r [ 1 1 IT I ] 



n 



I ' 



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DATI IaPPWOWAL 



, ^ - vy ,, SiARl 

YAWCIMBAL: I H^2 -" 

DRIVE FAIL I I ;p 



E^;;_^7J1*«1 




V 






0IM6AL MOTOR YAW 1 RELAYS 



2K21A 

« ! 



GIMBAL MOTOR START SEQUENCE CE7n 
RELAYS uro ^ 



V 



■^■^ 



2K26B 



u 



<r- 



G1M9AL MOTOR YAW 7 



u 



SERVICE PRORJLSIOM SYSTEM 
CIMSAL MOTOR CONTROL 
3Ar A YAW 1 



— o o— 



— w'^ 

YSTEM ' 



SERVICE PROPULSION SYSTEM 
GIMBAL MOTOR CONTROL 
3AT B YAW 2 



■*■• 



<r 



^Tnioeo WHEN VALVE i: 

C^es GHAT OTviERMI^ 

AUTO 



|S2(, 



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^ 



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L 



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MDC 25 



mck;2s 

CB20 



CJ. 



CB19 



MDC25 

eg, 

05A0— 



^1^ 



ID- 



|sT«IPeO WHE^J VALVE IsT 
OPfN GRAYOTHEBwiSEj 



SERVtCE PMPULSION 
SYSTEM GUA6IN6 
MNB 
M0C25 



SPS HELIUM <B) 
S 27 HOC 20 



CJ, 



HOC 20 a 



•>,- 



C- 



INJECT PRE -VALVE A 



7.M 
SERVICE PROPULSION 
SYSTEM Hi VALVE 
MNB 



V*ci 



TO THRUST rl_ 
VECTOR CONTROL J 




1KS6A llt56B 

CA SPS ARM r- 

^ RELAYS ^3 



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*2j £I*B1 



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1 -iH ,.. 



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5 DRIVE FAIL S -* ^- - 

■ , I Y ELLOW - } ■ - 



*-<^ 



*^ 



GIM9AL MOTOR PITCH 1 RELAYS 



:tM9ALM0TOfi START SEQUENCE ({T. 

Trelays ^^-^^ 



^ — ' 

1 2K28B 2KZi5* ^^ ^ ■ 
I • ^ , J 2K2 



GIMBAL MOTOR PITCH 2 




'■ SERVICE PROPULSION SYSTEM 
{ GIMBAL MOTOR CONTROL 
BAT A PITCH 1 MDC 25 



%■ 



^ REFER TO DWG 10.1,2 
^ REFER TO DWG 10.1.3 t 

'^ .NOT APPLICABLE Ift, 
CSM50I 



SPS 



Cfl?5 



SERVICE PROPULSION SYSTEM 
GIMBAL MOTOR CONTROL 

SAT 3 PITCH 2 MDC 25 



«■' 



CH|/ ^ 



r^^-l^' 



AUTH ^iZJj 






AFRM CI2 
SERVICE PROPULSION SYSTEM 
DETAILED DIAGRAM 



C5M 



AS -50 1 



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CSM 
AS-501 



10.2 NOTES, SERVICE MODULE, REACTION CONTROL SYSTEM 

10.2.1 General 
The SM/RCS consists of four individual, functionally identical 

packages located 90 degrees apart around the SM periphery (forward +X 
Axis) and offset from the S/C Y and Z axes by 7.25 degrees (CCW) and 
.canted 10 degrees from the SM surface. Each quad incoroporateS a pressure feec 
positive-expulsion, pulse-modulated, hi-propellant system to produce 
the reaction thrust required. The reaction engines may be pulse or 
continuously fired. 

10.2.2 RCS Engines 

A. Fuel — Monomehtylhydrazine (MMH) 

B. Oxidizer — Nitrogen Tetroxide (N2O4) 

C. 0/F Ration — 2.1 + .05 

D. Thruster Life — 1000 seconds pulse mode 

500 seconds continuous burn 

E. Nominal Thrust — 100 ± 5 lb 

10.2.3 Propellant Quantity 

A. Fuel — 65 each quad usable 

B. Oxidizer — 121 each quad usable 

C. Total usable {k quads) — 790 lb 

10-2.1+ SM/RCS Prnpellant QuRn titv ComnutfttinT) 

SM/RCS propellant quantity will be computed by a PVT method 
with pressures and temperatures using the RTCC and the Offline Computer 
- for computations. 

10.2.5 Helium Isolation Valves 

The helium supply flows from the helium tank into two parallel 
helium isolation valves. The valves are a two-solenoid valve and are 
mechanically latched open and spring-loaded close. The normal valve 
position is open. No ground commands (RTC) for the helium isolation 
1 valves . 

10-12 



1 1 1 1 If 1 1 f f II r M 1 1 If f T 



CSM 
AS-501 



10,2.6 Propellant Isolation Valves 

Propellant isolation valves are two-solenoid valves vhich are 
magnetically latched open and spring-loaded closed. A fuel and oxidizer 
valve is controlled by a real-time command, both open and closed commands 
(one command for each quad). 

Real-Time Commands 



Quad 


Open 


Clos( 


A 


32 


2k 


B 


33 


25 


C 


3h 


26 


D 


35 


27 


RCS Quad Heaters 







Each SM/RCS quad engine housing contains an electrical strip 
heater. Control of these heaters is automatic through thermostatic 
switches; range 115*^F minimum and 13U*^F maximum. Power is supplied by 
Main Bus A +28 Vdc, for Quads B and D, and Main Bus B +28 Vdc for quads 
A and C. Circuit breakers are. located on MDC 21. Heater operation is 
monitored by TM. (Engine package temp.) 

10.2.8 Service Module RCS Jettison Controller (SMJC) 

At CM/SM separation, the SMJC will command continuous firing 
of the -X Jets (manual coils). This function is to provide an adequate 
separation distance between the CM and SM. After two seconds, the 
+roll jet will fire for 5«5 seconds for spin stabilization. The -X 
jets will fire until propellant depletion or electrical power loss. 
The SMJC consists of two complete redundant systems. 

10 . 2 . 9 Propellant Solenoid Injector Control Valves (Fuel and Oxidizer) 
The solenoid injector valves utilize two coaxially wo\and coils; 

one for automatic and one for direct manual operation. The automatic 
coil is utilized when the command is generated within the Jet selection 
logic (SCS). The manual coils are used for commands originating from the 



10-13 



I 1 I I I M ! ! ! I ! I I n if I f 1 



' 1 



\ 



1 



CSM 
AS-501 



real-time commands and direct ullage. The valves are spring-loaded 
closed and electrically energized open. The fuel injector valves reach 
the full OPEN position in approximately T milliseconds and the oxidizer 
in approximately 9 milliseconds, giving a 2-millisecond fuel lead. The 
automatic coils (fuel and oxidizer) are connected in parallel, while 
the manual coils are connected in series. Auto coils are connected in 
parallel to prevent mismatch between opening and closing of the valves 
due to heat soak back. The series connection of the fuel manual coils 
(positive to negative) to the oxidizer coil (negative to positive) then 
to ground, increases the arc suppression, reducing the arc at the rotation 
controller in the DIRECT RCS MODE. 



Real- 


-Time 


Commands 


Direction 




RTC 


+ Pitch 




11+ 


- Pitch. 




15 


+ Yaw 




16 


- Yaw 




IT 


+ Roll 




20 


- Roll 




21 



10.2.10 Attitude Control Maneuvers 

Attitude control of the spacecraft can be accomplished by any 
two adjacent quads, either using the automatic mode (SCS) or direct 
mode (real-time commands). The ullage maneuver (SPS propellant settling) 
is accomplished by four Jets (+X). 



10-lU 



I 1 I ! I f I f ! ! f f I [ [ I I f T f T 




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CSM 
» AS-501 

10.3 NOTES, COMMAND MODULE, REACTION CONTROL SYSTEM 

' 10.3.1 General 

The CM/RCS consists of tvo separate independent systems, with 
one system capable of complete control. In low altitude (L/0 to less 
than L/0 +ii2 seconds) abort operation, provisions have been incorporated 
to automatically dump the oxidizer supply overboard, followed by a 
helium purge of the remaining helium supply. Nonnal reentry and after 
L/0 +ii2 seconds, remaining propellants are dumped and then purged 
manually by the crew. Tanks will be pressurized to lesser value (60 
psia, normal 291 psia) until system activation. 

10.3.2 RCS Engines 

A. Fuel — Monomethylhydrazine (MMH) 

- I B. Oxidizer — Nitrogen Tetroxide (N2OU) 

C. 0/F Ratio — 2:1 + .05 

D. Thruster Life — 200 seconds 

E. Nominal Thrust ~ 93 lb 

10.3.3 Propellant Quantity Gaging 

^ No propellant quantity gaging system on the CM/RCS. 

^ 10.3.^ Propellant Isolation Valves (Fuel And Oxidizer) 

The isolation valves are two solenoid valves and are magnetically 
latched open and spring loaded closed. These valves allow fuel and 
oxidizer to flow to the injector valves and are controlled by a three- 
position (ON, HOLD, OFF) toggle switch (one switch per system) on MDC 
15. The two proipellant isolation valves will be placed in the open 
position, by MCP command, prior to CM/SM SEP to insure that the valves 
remain open upon system pressurization. No ground commands (RTC) for 
the CM/RCS propellant valves. 

10.3.5 CM/RCS Pressurization 

The CM/RCS will remain in STANDBY condition until CM/SM SEP. 
j This is accomplished by isolating the high pressure helium with pyro 

10-16 



\ 



1 1 1 1 1 IT ! I ! I ! r f [ I rr T f [ 



CSM 
AS-501 



squib valves. Each RCS system incorporates tvo squib valves which are 
blown one second prior to CM/SM SEP by the MESC. The MESC receives 
the signal for pressurization from the MCP. The RCS Tank pressures 
will be approximately 60 psia until system activation, at which time 
the pressures will increase to nominal regulated pressiires. 

10.3.6 CM/HCS Dump /Burn and Purge 

There are two sequences for propellant Jettison. One sequence 
is employed in event of an abort from laionch pad to T+U2 seconds. The 
second sequence is for a normal reentry or abort after T+i+2 seconds. 

10.3.6.1 Sequence of Events For An Abort, Pad to T+U2 Seconds: 
(PROP JETT LOGIC Switch ON at L/O) 

A. At abort signal reception (LES ABORT), the four helium 
isolation squibs are initiated open (by closure of C/SM 
deadface relay), pressurizing A and B systems. 

B. Simultaneously, the oxidizer helium interconnect valve is 
initiated open, the propellant isolation valves (both 
fuel and oxidizer) are closed, and the oxidizer propellant 
interconnect is initiated open. The oxidizer overboard 
blowout plus is opened, dumping the oxidizer through the 
aft heat shield. The entire oxidizer supply is dumped in 
13 to 15 seconds. 

C. At abort plus l8 seconds, the heliiom interconnect squib valve 
and the helium overboard dump is opened, dumping helium 

into the aft equipment compartment thus relieving the 
high-pressure helium. The oxidizer tank bypasses are also 
initiated open, allowing a complete oxidizer purge. The 
fuel remains onboard at impact, with no pressure on the 
fuel tanks. 
At abort plus 80 seconds, the RCS purge is activated, thus giving 
a redundant method of reducing the high-pressure helium. At T+H2 seconds, 
the OX DUMP AUTO is disabled which inhibits auto oxidizer dump. Any 
abort after T+U2 is the same as the normal reentry procedure. 

10-17 



1 1 1 1 1 M n ! 11 1 r 1 1 i [ [ f I 



\ 



CSM 
AS-501 



10.3.6.2 Sequence of Events For A Normal Reentry or Abort Above T+U2 
Seconds: (PROP JETT LOGIC Svitch ON Prior to L/O) 

A. After the spacecraft is on the main chutes (12k baro + 
20 seconds), RCS DUMP is activated. This simultaneously 
initiates OPEN two heliim interconnect squib valves, fuel 
and oxidizer squib valves, and the injector valves of 
ten of the twelve engines (+ pitch jets not energized) 
burning the remaining propellants through the engines. 

B. Upon completion of propellant burn (250 second time delay), 
the PROP JETT PURGE is activated. This function will 
initiate the four helium bypass squib valves allowing 
helium pressure to bypass aroTind the fuel and oxidizer 
tanks and out of the ten engines. 



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SECTION 11 

MISSION CONTROL PROGRAMER ?o^cn-r 

AS-501 



11.1 MISSION CONTROL PROGRAMER (MCP) , GENERAL 



11.1.1 MCP Components 

The MCP is comprised of three major components : 

A. Groi:ind Command Controller (GCC) 

B. Spacecraft Command Controller (SCC) 

C. Attitude and Deceleration Sensor (ADS) 

^ 1 The eqiiipment is moxmted on a platform that repl^aces the crew 

couches. Coolant is provided by the ECS glycol coolant loop. MCP 
internal -power is derived from the Auxiliary Battery Buses A and B. 
The postlanding power is provided by the Flight and Postlanding Bus. 
Upon impact, the auxiliary batteries are transferred to the F&PL Bus, 
deactivating a portion of the MCP. The SCC and GCC have independent 
I internal power distribution, with crossovers for isolated functions. 

The ADS derives its power from the SCC. 



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AS-501 



11.2 GROUND COMMAND CONTROLLER (GCC) 
11^ All RTC's are received from the UDL through the GCC. Upon receipt 

of a given RTC, the required logic is set up in the GCC to perform the 
given function. In some incidences, the RTC is converted to a discrete 
and routed directly from the GCC to the SCC where the logic is set up to 
perform the function. 

' 1 



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^12K BARO * 20 SEC % 

11.3.3 



LET JET (SCO "^ 
11.3.4 



VHP/AM RCVR - XMTR BEACON 

-^ VHF/AMXMTRONCONTL 

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MIS*^ION CONTROL PROGRAMER, 
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NATIONAL AERONAUTICS 4 SPACE AO«tNISTRATION 
MAW«0»PACeCRAFT C«MTBR . HOUSTON. TEXAS 



CSM0I7 

MISSION CONTROL 

PROGRAMER.GCC (RCS) 



TSM" 



MISSION 

AS-501 



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MANNED SPACECRAFT CENTER ■ HOUSTON, TEXAS 



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MISSION CONTROL PROGRAMER 
GCC CPROP AND GiC) 



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AS-501 



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11.3 SPACECRAFT COMMAUD CONTROLLER ( SCO) 
■ ] " F The sec interfaces primarily with the MDC to replace the manual 

fiinctions of the crew. Spacecraft discretes setup the necessary 
circuit logic to perform a given function. Discretes are also routed 
from the SCC to the GCC to perform various spacecraft fionctions, thru 
the GCC relay logic. 



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Pft: PRKKAMER RESET 
* (ASTEXICKS): POlin OF 01 
OF A CtVEN DISCRETE 



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CSK' 017 
MISSION CONTROL 
hhOGUAMER sec (SEQ'L) 



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* tASTEHISKS): PfflWT OF ORtGM OF A GIVEN DISCRETE 



MCP 



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NATIONAL AERONAUTICS & SPACE ADMINISTRATION 
MANNED SPACECRAFT CENTER . HOUSTON. TEXAS 



CSM 017 

MISSION CONTROL 
PROGRAMERjSCC (SEQL) 



CSM 



AS-50( 



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CSM 
AS-501 



11. U ATTITUDE AND DECELERATION SENSOR (ADS) 

The AM provides a Stable I-Stable II Attitude Sensor, a 0.05G backup 
sensor* and an impact switch. 



\ 



11-15 



1 I II I ! I I I ! I I I r I U [ f f F 



CSM 
AS-501 



11.5 MCP CIRCUIT LOGIC 

Figures 11.5.1 thru 11. 5-7 are the logic circuits associated with the 
MCP and Spacecraft Systems. Relays numbered IKXare in the .GCC. Relays 
nioiribered 2KX are in the SCC. ADS interfaces are noted. 

Relays 'that are reset (or set) by the programer reset discrete 
(sent at T-5 hrs by GSE) are not shown in the logic where no other 
discrete is involved in the reset (or set). Table 11-1 lists all 
relay numbers and functions involved in a programer reset discrete. 

The following terms and symbols are used in the logic circuit : 

A. S: SET OR LATCH 

R: REST OR UNLATCH 



B. 
C. 



SOME* 
WORD 



INPUT 



A flag, origin of a discrete. This flag is common 
with all other identical flags shown without asterisk. 
[l] OUTPUT: An inverter. The output is the 



logical inverse of the input, 



INPUTS 



F. INPUTS 



G. INPUTS 




OUTPUT: A standard "OR" gate. 



OUTPUT: A standard "AND" gate 



OUTPUT: A latching or gate. The output 
is logically 1 if either input 
is 1 and continues to be 1 even 
if the input becomes logically 
0. (Note: A programer reset 
discrete is required to reset 
the output to 0. ) 



11-16 



1 1 1 1 1 ! f 1 1 f 1 1 1 r n I! T ! 



n 



\ 















CSM 


« 






AS- 501 




TABLE 11-1 MCP RELAY LOGIC FOR A PROGRAMER RESET 


SYSTEM 


CONTACT 
LOCATION 


CONTACT 
NUMBER 


SET 


RESET 


FUNCTION 


sec 


GCC 


SEQ 


X 




2K100EFGH 




X 


MESC A&B LOGIC PWR 




X 




2K9 




X 


E/T JET 




X 




?KTO 




X 


LES MOTOR FIRE B 




X 




2K11 




X 


CSM SEP B 




X 




2Kll* 




X 


ELS ACTIVATE B 




X 




2K15 




X 


E/T JtT A 




X 




2Kl6 




X 


LES MOTOR FIRE A 




X 




2K1T 




X 


CSM SEP A 




X 




2K20 




X 


ELS ACTIVATE A 




X 




2K55 




X 


MN CHUTE DISCONNECT A 




X 




2K63 




X 


VENT BAGS ON, UPRIGHTING CONTL 




X 




2K66 




X 


PUMPS OFF, UPRIGHTING CONTL 




X 




2K6T 




X 


MN CHUT DISCONNECT B 


EPS 


X 




2KU6 




X 


INV 3 OFF BUS 1 




X 




2KU7 




X 


INV 3 OFF BUS 2 




X 




2Klt8 




X 


ISOLATE INV 2 




X 




2KU9 




X 


ISOLATE INV 1 




X 




2K51 




X 


MN B OFF INV 3 




X 




2K52 




X 


MN A OFF IVN 3 






X 


1K50 




X 


ENTRY BATTS OFF MAINS 






X 


1K51 




X 


C BATT ON F&PL BUS 






X 


1K52 




X 


A&B BATTS ON F&PL BUS 






X 


1K81 




X 


AUX BATTS 1 & 3 ON F&PL BUS 






X 


1K82 




X 


HEAT SHIELD INSTR ON 






X 


IK85 




X 


AUX BATTS 2 & 3 F&PL BUS 


ECS 


X 




2K6 




X 


GLYCOL WETNESS CONTROL 




X 




?KT 




X 


O2 ISOLATION VALVE OPEN 




X 




2K8 




X 


GLYCOL SHUTOFF VALVE OPEN 




X 




2K56 




X 


BACKPRESSURE CONTROL 


CRYO 












NONE 


TELECOMM 




X 


IKIOO 




X 


Z ANTENNA SWITCHING 






X 


1K32 




X 


FQ RECORDER STOP OFF 






X 


1K33 




X 


-Z ANTENNA ON 






■ X 


1K3U 




X 


+Z ANTENNA ON 






X 


1K35 




X 


CINE CAMERA CONTROL (OFF) 






X 


1K36 




X 


HF ANT DEPLOY B OFF 






X 


1K3T 




X 


UHF RCVR ON 






X 


1K38 




X 


VHF/FM XMTR ON 






X 


1K39 




X 


DSE FWD-NML ON 






X 


ikUo 




X 


DSE RECORD ON 






X 


IKltl 




X 


FQ RECORDER OFF 




X 


1K1*2 




X 


VHF/AM RCVR ON 



11-17 



I 1 II I M I f [ ! I n [ U IT f f 



TABLE 11-1 MCP RELAY LOGIC FOR A PROGRAMER RESET (Cont'd) 



CSM 
AS-501 



SYSTEM 


CONTACT 
LOCATION 


CONTACT 
NUMBER 


SET 


RESET 


FUNCTION 




sec 


GCC 


TELECOMM 




X 


IKUU 




X 


HF ANTENNA DEPLOY A OFF 






X 


lKk6 




X 


PWR TO CBU5 OFi*' 






X 


IK5U 




X 


FLD SWING LIGHT ON 






X 


1K55 




X 


FQ RECORDER REMIND OFF 






X 


IKU3 


X 




VHF XMITTER ON 






X 


IKU5 


X 




HF XCVR ON 






X 


1KU8 


■X 




VHF SURVIVAL BEACON ON 






X 


1K5T 


X 




VHF SURVIVAL BEACON ON 


G&C 


X 




2K39 




X 


GIMBAL POSITION SET 




X 




2K1+0 




X 


GIMBAL POSITION SET 




X 




2KU3 




X 


SEP /ABORT A OFF 




X 




2KUU 




X 


SEP/ABORT B On' 




X 




2K53 




X 


GIMBAL POSITION SET 




X 




2K68 




X 


PSECJDO RATE OUT 




X 




2K69 




X 


DEADBAND SELECT 






X 


IKUT 




X 


DIRECT THRUST OFF 






X 


IK56 . 




X 


DIRECT THRUST OFF 


PROP 


X 




2K21 




X 


YAW 1' START 




X 




2K22 




X 


YAW 1 ON 




X 




2K23 




X 


PROP ISOLATION VALVE (n.C 




X 




2K21+ 




X 


PROP ISOLATION VALVE (n.C 




X 




2K25 




X 


YAW 2 START 




X 




2K26 




X 


YAW 2 ON 




X 




2K2T 




X 


PITCH 2 START 




X 




2K28 




X 


PITCH 2 ON 




X 




2K1 




X 


RCS DUMP A 




X 




2K2 




X 


PCS PURGE ACTIVATE 




X 




2K3 




X 


OXID DUMP A 




X 




2KT0 




X 


RCS DUMP B 




X 




2KT1 




X 


RCS PURGE ACTIVATE 




X 




2KT2 




X 


OXID DUMP B 




X 




2K100ABCD 




X 


PITCY, YAW, & ROLL BACKUP 



11-18 



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11-23 



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11-26 



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11-27 



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11-28 



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11-29 



1 1 1 1 1 M I T f f T f r n I [ [ f f 



.G6N FAIL 



LCS ABORT > [T] 1^ 

> Q— ^ 

> 



LV-SC SEP 
^2.b SEC 



SEE MCP/SPS LCGIC ' 



28 VDC \^ 

sec ^^ 



/RTCV FDAl ALIGN 
Shi/ 



/yK CSM SEP. 



c 



CBN ABOPT ARM 
^.IMBAL MOTOR ARfl 

FDAI ALIGN ARM 
- CEN AR,'"^ 



• FDAI ALIjN- 



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C 



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AR ^ 

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APM 

G£N ATT CONT 

MODE ARM 
" CSM SEP AP.-I 
.+X TRANS APM 



CSM SEP ON 



0.05C 



IoT^^ensopI 



SN FAIL- 



CSM SEP" 



rf\ w6N FAIL INHIBIT 



-m— 






+ h-*- 



> 



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<t X l-tin- entry ^ 

> 



Hp 



FIGURE U.5.5C MC"/^-&N-SCS LO": 1 C 



G6N AV " 
'^DE 



X TP^NSLATION 



FDAI ALIGNJt 



. 2K58 



CSN ABORT" 



PSEUDC RATE ^ 2K68S 



CUTOUT 

rW DEADBAND ^ 2K695 



0.05-C" 



G&N FAIL" 



',&N E^4TPV lODE 



"EN aV -'ODE 



G£N ATT CCNTL MOOE , 



iONlTOR MODE OFF 



5CS ENTRY lODE 



SCS iV '.ODE 



■- 2K54 
^ 2K36 
». 2K32 

► 2K31 
^ 2K35 
^ 2K57 



ROLL R.ATE >J 



2K37 



ROLL 0.05 C 



^> 2K30 

PITCH S YA' 0.05 : ^ 2K29 



DIRECT ulla:e 



1K51 



DIRECT T>^F.UST ON ^ ^^^^^ 



I 1 I I I I I I I I I I I I [ 1 i! [ f ] 



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