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-1021042
PROJECT
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APOLLO
LUNAR EXCURSION MODULE
PRIMARY GUIDANCE, NAVIGATION AND CONTROL SYSTEM MANUA
VOLUME I
ELECTR0NICS
DIVISION OF GENERAL MOTORS
'J.
INITIAL TDRR 26432 TYPE I
APPROVED BY NASA
APOLLO .
LUNAR EXCURSION MODULE
4
PRIMARY
GUIDANCE, NAVIGATION, AND CONTROL SYSTEM « MANUAL
VOLUME I OF II
PREPARED FOR
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION * MANNED SPACECRAFT CENTER
BY
AC ELECTRONICS DIVISION OF GENERAL MOTORS Ml LWAUKEE,WISCONSI N 53201
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NASA CONTRACT NAS 9-497
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LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
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LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
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MANUAL
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LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
CONTENTS
Chapter Page
Volume I
1 SYSTEM TIE-IN . 1-1
1-1 Scope . 1_1
1-2 LEM Mission . 1-1
1-2.1 Separation and Transfer Orbit Insertion . 1-1
1-2.2 Descent Coast . 1-1
1-2.3 Powered Descent and Landing . 1-2
1-2.4 Lunar Stay . 1-3
1-2.5 Launch and Powered Ascent . 1-3
1-2.6 Rendezvous and Docking . 1-3
1-3 LEM Structure . 1-4
1-3.1 Ascent Stage . 1-4
1-3.2 Descent Stage . 1-7
1-4 LEM Systems . 1-7
1-4.1 Primary Guidance, Navigation, and Control System. . 1-7
1-4.2 Stabilization and Control System . 1-8
1-4.3 Propulsion System . 1-9
1-4.4 Reaction Control System . 1-9
1-4.5 Electrical Power System . 1-10
1-4.6 Environmental Control System . 1-10
1-4.7 Communications and Instrumentation System . 1-10
1- 5 PGNCS Interface . 1-10
1-5.1 Systems . 1-12
1-5.2 Displays and Controls . 1-12
1- 5.3 Landing Radar . 1-12
2 SYSTEM AND SUBSYSTEM FUNCTIONAL ANALYSIS . 2-1
2- 1 Scope . 2-1
2-2 Primary Guidance, Navigation, and Control System . 2-1
2-3 LEM and PGNCS Axes . 2-2
2- 3.1 LEM Spacecraft Axes . 2-2
2-3.2 Navigation Base Axes . 2-2
2-3.3 Inertial Axes . 2-2
I-xi
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ND-1021042
MANUAL
CONTENTS (cont)
Chapter Page
2-4 Inertial Subsystem . 2-4
2-4.1 Stabilization Loop . 2-5
2-4.2 Fine Align Electronics . 2-9
2-4.3 Accelerometer Loop . 2-14
2-4.4 IMU Temperature Control System . 2-21
2-4.5 ISS Modes of Operation . 2-25
2-4.6 ISS Power Supplies . 2-37
2-5 LEM Optical Rendezvous Subsystem . 2-41
2- 6 Computer Subsystem . 2-42
2-6.1 Programs . 2-47
2-6.2 Machine Instructions . 2-48
2-6.3 Timer . 2-50
2-6.4 Sequence Generator . 2-51
2-6.5 Central Processor . 2-52
2-6.6 Priority Control . 2-54
2-6.7 Input-Output . 2-55
2-6.8 Memory . 2-56
2-6.9 Power Supplies . 2-58
2- 6.10 Display and Keyboard . 2-59
3 PHYSICAL DESCRIPTION . 3-1
3- 1 Scope . 3-1
3-2 PGNCS Interconnect Harness . 3-1
3-3 Navigation Base Assembly . 3-5
3-4 Inertial Measuring Unit . 3-5
3- 4.1 Stable Member . 3-6
3-4.2 Middle Gimbal . 3-7
3-4.3 Outer Gimbal . 3-7
3-4.4 Supporting Gimbal . 3-7
3-4.5 Inter-Gimbal Assemblies . 3-10
3-5 Optical Tracker . 3-10
3-6 Luminous Beacon . 3-12
3-7 Pulse Torque Assembly . 3-13
3-8 Power and Servo Assembly . 3-17
I-xii
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
CONTENTS (cont)
Chapter Page
3-9 LEM Guidance Computer . 3-20
3-9.1 Logic Tray A . 3-21
3- 9.2 Tray B . 3-21
3-10 Coupling Data Unit . 3-22
3-11 Signal Conditioner . 3-24
3- 12 Display and Keyboard . 3-24
4 COMPONENT THEORY OF OPERATION . 4-1
4- 1 Scope . 4-1
4-2 Apollo II Inertial Reference Integrating Gyro . 4-1
4- 2. 1 Gyro Wheel Assembly . 4-3
4-2.2 Float Assembly . 4-3
4-2.3 Case . 4-4
4-2.4 Normalizing Network . 4-4
4-2.5 Apollo II IRIG Ducosyns . 4-4
4-3 16 Pulsed Integrating Pendulum . 4-10
4-3. 1 Float Assembly . 4-13
4-3.2 Housing Assembly . 4-13
4-3.3 Outer Case Assembly . 4-13
4-3.4 Normalizing Network . 4-13
4-3.5 PIP Ducosyns . 4-13
4-4 Coupling Data Unit . 4-15
4-4. 1 Coarse System Module . 4-15
4-4. 2 Quadrant Selector Module . 4-23
4-4.3 Main Summing Amplifier and Quadrature
Rejection Module . 4-29
4-4.4 Read Counter Module . 4-32
4-4. 5 Error Angle Counter and Logic Module . 4-33
4-4. 6 Digital Mode Module . 4-34
4-4.7 Interrogate Module . 4-35
4-4.8 Digital to Analog Converter . 4-37
4-4. 9 Mode Module . 4-41
4-4. 10 4 VDC Power Supply . 4-43
4-5 LEM Guidance Computer . 4-44
4-5. 1 Programs . 4-44
I-xiii
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ND-1021042
MANUAL
CONTENTS (cont)
Chapter Page
4-5.2 Machine Instructions . 4-49
4-5.3 Timer . 4-204
4-5.4 Sequence Generator . 4-229
Volume II
4-5. 5 Central Processor . 4-365
4-5.6 Priority Control . 4-428
4-5.7 Input -Output . 4-435
4-5.8 Memory . 4-439
4-5.9 Power Supply . 4-460
4- 5. 10 Display and Keyboard . 4-491
4-6 Signal Conditioner . 4-492
4- 7 LEM Optical Rendezvous Subsystem . 4-492
5 MISSION OPERATIONS . 5-1
5- 1 Scope . 5-1
5-2 IMU Coarse Alignment . 5-1
5-3 IMU Fine Alignment . 5-1
5-4 Transfer Orbit . 5-2
5-5 Powered Descent . 5-2
5- 5.1 Phase I - Braking . 5-2
5-5.2 Phase II - Final Approach . 5-2
5-5.3 Phase III - Landing . 5-7
5-6 Lunar Stay . 5-7
5-7 Ascent . 5-7
5- 8 Rendezvous and Docking . 5-7
6 CHECKOUT AND MAINTENANCE EQUIPMENT . 6-1
6- 1 Scope . 6-1
7 CHECKOUT . 7-1
7- 1 Scope . 7-1
7-2 Primary Guidance, Navigation, and Control System . 7-1
7-2.1 Preparation . 7-1
7-2.2 Checkout . 7-1
7-2.3 Test Descriptions . 7-1
I-xiv
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
CONTENTS (cont)
Chapter Page
7-3 Inertial Subsystem . 7-1
7-3. 1 Preparation . 7-1
7-3.2 Checkout . 7-2
7-4 Computer Subsystem . 7-2
7-4. 1 Preparation . 7-2
7- 4.2 Checkout . 7-2
7- 5 LEM Optical Rendezvous Subsystem . 7-2
8 MAINTENANCE . 8-1
8- 1 Scope . 8-1
8-2 Maintenance Concept . 8-1
8-3 Malfunction Isolation . 8-2
8-4 Double Verification . 8-2
8- 4. 1 Malfunction Verification . 8-2
8-4. 2 Repair Verification . 8-6
8-5 Pre-Installation Acceptance Test . 8-6
8-6 Removal and Replacement . 8-6
8-7 Maintenance Schedule . 8-6
8-8 Optical Cleaning . . . 8-6
APPENDIX A LIST OF TECHNICAL TERMS AND ABBREVIATIONS . A-l
APPENDIX B RELATED DOCUMENTATION . B-l/B-2
APPENDIX C LOGIC SYMBOLS . C-l
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ND-1021042
MANUAL
ILLUSTRATIONS Volume I
Figure Page
1-1 LEM Primary Guidance, Navigation, and Control System . .I-xxxiii/I-xxxiv
1-2 LEM Mission Phases . 1-2
1-3 LEM . 1-5
1-4 LEM External Dimensions . 1-6
1- 5 LEM PGNCS Functional Interface, Block Diagram . 1-11
2- 1 PGNCS Subsystems Interface, Block Diagram . 2-3
2-2 LEM and PGNCS Axes . 2-4
2-3 ISS, Block Diagram . 2-6
2-4 Stabilization Loop, Block Diagram . 2-7
2-5 Fine Align Electronics -Computer Inputs . 2-9
2-6 Fine Align Electronics -Gyro Selection . 2-10
2-7 Binary Current Switch . 2-12
2-8 DC Differential Amplifier and Precision Voltage Reference . 2-13
2-9 Accelerometer Loop . 2-14
2-10 AC Differential Amplifier and Interrogator Module . 2-16
2-11 Accelerometer Timing . 2-19
2-12 PIPA Calibration Module . 2-20
2-13 IMU Temperature Control System . 2-23/2-24
2-14 ISS-CDU Moding . 2-27/2-28
2-15 IMU Cage Mode . 2-31/2-32
2-16 Display Inertial Data Mode . 2-36
2-17 Pulse Torque Power Supply . 2-38
2-18 -28 VDC Power Supply . 2-40
2-19 800 CPS Power Supply . 2-40
2-20 3, 200 CPS Power Supply . 2-42
2-21 Computer Subsystem, Block Diagram . 2-43/2-44
2-22 Program Organization . 2-47
2-23 Timer, Block Diagram . 2-51
2-24 Sequence Generator, Block Diagram . 2-52
2-25 Central Processor, Block Diagram . 2-53
2-26 Priority Control, Block Diagram . 2-54
2-27 Input -Output, Block Diagram . 2-55
2-28 Memory, Block Diagram . 2-57
2-29 Power Supplies, Block Diagram . 2-58
2-30 Display and Keyboard (DSKY), Block Diagram . 2-59
I-xvii
ND-1021042
ILLUSTRATIONS (cont)
Figure Page
3-1 Location of LEM PGNCS Components . 3-3
3-2 Navigation Base Assembly . 3-5
3-3 Inertial Measuring Unit . 3-6
3-4 IMU Stable Member . 3-8
3-5 Optical Tracker . 3-11
3-6 Luminous Beacon . 3-12
3-7 Pulse Torque Assembly . 3-13
3-8 Power and Servo Assembly . 3-17
3-9 LEM Guidance Computer . 3-20
3-10 Logic Tray A . 3-21
3-11 Tray B . 3-22
3-12 Coupling Data Unit . 3-23
3-13 CDU Module Locations . 3-25
3- 14 Display and Keyboard . 3-27
4- 1 Apollo II IRIG, Simplified Cutaway View . 4-2
4-2 Apollo II IRIG Normalizing Network . 4-5
4-3 IRIG Signal Generator and Suspension Microsyn . 4-7
4-4 IRIG Torque Generator and Suspension Microsyn . 4-8
4-5 Ducosyn RLC Equivalent Circuit . 4-9
4-6 Definition of 16 PIP Axes . 4-11
4-7 Result of Acceleration Along Input Axis . 4-12
4-8 PIP Torque Generator . 4-14
4-9 Read Counter Relationship to Coarse and Fine System Switching. 4-16
4-10 Coarse System Module, Block Diagram . 4-17
4-11 Resolver Sine and Cosine Phase Relationships . 4-18
4-12 Coarse Switch Circuit and Logic Equations . 4-19
4-13 Coarse Switching Diagram . 4-20
4-14 Quadrant Selector Module, Block Diagram . 4-25
4-15 Fine Switching Diagram . 4-26
4-16 Main Summing Amplifier and Quadrature Rejection Module,
Block Diagram . 4-30
4-17 Simplified 3 Bit Converter and Switch Configurations . 4-39
4-18 4 VDC Power Supply, Block Diagram . 4-44
4-19 Basic Instruction Word Format . 4-50
4-20 Subinstruction TC0, Data Transfer Diagram . 4-107
4-21 Subinstruction TC0, with Implied Address Code EXTEND,
Data Transfer Diagram . 4-108
4-22 Subinstruction CCS0, Branch on Quantity Greater Than
Plus Zero, Data Transfer Diagram . 4-109
4-23 Subinstruction CCS0, Branch on Minus Zero, Data
Transfer Diagram . 4-110
4-24 Subinstruction CCS0, Branch on Quantity Less Than
Minus Zero, Data Transfer Diagram . 4-111
4-25 Subinstruction CCS0, Branch on Plus 0, Data Transfer Diagram 4-112
4-26 Subinstruction STD2, Data Transfer Diagram . 4-113
4-27 Subinstruction STD2, with Implied Address Code INHINT,
Data Transfer Diagram . 4-114
I-xviii
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ND-1021042
MANUAL
ILLUSTRATIONS (cont)
Figure Page
4-28 Subinstruction STD2, with Implied Address Code RE LINT,
Data Transfer Diagram . 4-115
4-29 Subinstruction STD2, with Implied Address Code EXTEND,
Data Transfer Diagram . 4-116
4-30 Subinstruction TCFO, Data Transfer Diagram . 4-117
4-31 Subinstruction TCFO, with Implied Address Code EXTEND,
Data Transfer Diagram . 4-118
4-32 Subinstruction DASO, without Overflow or Underflow, Data
Transfer Diagram . 4-119
4-33 Subinstruction DAS1, without Overflow or Underflow, Data
Transfer Diagram . 4-120
4-34 Subinstruction DASO, with Overflow and Implied Address
Code DDOUBL, Data Transfer Diagram . 4-121
4-35 Subinstruction DAS1, with Overflow and Implied Address
Code DDOUBL, Data Transfer Diagram . 4-122
4-36 Subinstruction DASO, with Underflow, Data Transfer Diagram. . 4-123
4-37 Subinstruction DAS1, with Underflow, Data Transfer Diagram. . 4-124
4-38 Subinstruction LXCHO, Data Transfer Diagram . 4-125
4-39 Subinstruction INCRO, Data Transfer Diagram . 4-126
4-40 Subinstruction ADSO, Data Transfer Diagram . 4-127
4-41 Subinstruction CAO, Data Transfer Diagram . 4-128
4-42 Subinstruction CSO, Data Transfer Diagram . 4-129
4-43 Subinstruction NDXO, Data Transfer Diagram . 4-130
4-44 Subinstruction NDX1, Data Transfer Diagram . 4-131
4-45 Subinstruction NDXO with Implied Address Code RESUME,
Data Transfer Diagram . 4-132
4-46 Subinstruction RSM3, Data Transfer Diagram . 4-133
4-47 Subinstruction RSM3 with Implied Address Code EXTEND,
Data Transfer Diagram . 4-134
4-48 Subinstruction DXCHO, Data Transfer Diagram . 4-135
4-49 Subinstruction DXCH1, Data Transfer Diagram . 4-136
4-50 Subinstruction TS0 without Overflow or Underflow, Data
Transfer Diagram . 4-137
4-51 Subinstruction TS0 with Overflow, Data Transfer Diagram . . . 4-138
4-52 Subinstruction TS0 with Underflow, Data Transfer Diagram . . . 4-139
4-53 Subinstruction XCH0, Data Transfer Diagram . 4-140
4-54 Subinstruction ADO, Data Transfer Diagram . 4-141
4-55 Subinstruction MSK0, Data Transfer Diagram . 4-142
4-56 Subinstruction RE ADO, Data Transfer Diagram . 4-143
4-57 Subinstruction WRITE 0, Data Transfer Diagram . 4-144
4-58 Subinstruction RAND0, Data Transfer Diagram . 4-145
4-59 Subinstruction WAND0, Data Transfer Diagram . 4-146
4-60 Subinstruction RORO, Data Transfer Diagram . 4-147
4-61 Subinstruction WORO, Data Transfer Diagram . 4-148
4-62 Subinstruction RXORO, Data Transfer Diagram . 4-149
4-63 Subinstruction RUPT0, Data Transfer Diagram . 4-150
I-xix
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ND-1021042
MANUAL
ILLUSTRATIONS (cont)
Figure Page
4-64 Subinstruction RUPT1, Data Transfer Diagram . 4-151
4-65 Subinstruction DVO, Data Transfer Diagram . 4-152
4-66 Subinstruction DV1, Data Transfer Diagram . 4-153
4-67 Subinstruction DV3, Data Transfer Diagram . 4-154
4-68 Subinstruction DV7, Data Transfer Diagram . 4-155
4-69 Subinstruction DV6, Data Transfer Diagram . 4-156
4-70 Subinstruction DV4, Data Transfer Diagram . 4-157
4-71 Subinstruction BZFO with Branch on Non-Zero Quantity,
Data Transfer Diagram . 4-158
4-72 Subinstruction BZFO with Branch on Plus Zero, Data
Transfer Diagram . 4-159
4-73 Subinstruction BZFO with Implied Address Code EXTEND,
Data Transfer Diagram . 4-160
4-74 Subinstruction MSUO with Positive Resultant, Data
Transfer Diagram . 4-161
4-75 Subinstruction MSUO with Negative Resultant, Data
Transfer Diagram . 4-162
4-76 Subinstruction QXCHO, Data Transfer Diagram . 4-163
4-77 Subinstruction AUGO with Positive Quantity, Data
Transfer Diagram . 4-164
4-78 Subinstruction AUGO with Negative Quantity, Data
Transfer Diagram . 4-165
4-79 Subinstruction DIMO with Positive Quantity, Data
Transfer Diagram . 4-166
4-80 Subinstruction DIMO with Negative Quantity, Data
Transfer Diagram . 4-167
4-81 Subinstruction DCAO, Data Transfer Diagram . 4-168
4-82 Subinstruction DCA1, Data Transfer Diagram . 4-169
4-83 Subinstruction DSCO, Data Transfer Diagram . 4-170
4-84 Subinstruction DCS1, Data Transfer Diagram . 4-171
4-85 Subinstruction NDXXO, Data Transfer Diagram . 4-172
4-86 Subinstruction NDXX1, Data Transfer Diagram . 4-173
4-87 Subinstruction SUO, Data Transfer Diagram . 4-174
4-88 Subinstruction BZMFO with Quantity Greater Than
Plus Zero, Data Transfer Diagram . 4-175
4-89 Subinstruction BZMFO with Plus Zero, Data Transfer Diagram . 4-176
4-90 Subinstruction BZMFO with Negative Quantity, Data
Transfer Diagram . 4-177
4-91 Subinstruction BZMFO with Implied Address Code EXTEND,
Data Transfer Diagram . 4-178
4-92 Subinstruction MP0 with Two Positive Numbers, Data
Transfer Diagram . 4-179
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ILLUSTRATIONS (cont)
Figure Page
4-93 Subinstruction MPO with Positive Number in A and
Negative Number in E, Data Transfer Diagram . 4-180
4-94 Subinstruction MPO with Negative Number in A and
Positive Number in E, Data Transfer Diagram . 4-181
4-95 Subinstruction MPO with Two Negative Numbers, Data
Transfer Diagram . 4-182
4-96 Subinstruction MP1, Data Transfer Diagram . 4-183
4-97 Subinstruction MP3, Data Transfer Diagram . 4-184
4-98 Subinstruction MP3 with Implied Address Code EXTEND,
Data Transfer Diagram . 4-185
4-99 Subinstruction GOJ1, Data Transfer Diagram . 4-186
4-100 Subinstruction PINC, Data Transfer Diagram . 4-187
4-101 Subinstruction MINC, Data Transfer Diagram . 4-188
4-102 Subinstruction DINC with Positive Quantity, Data
Transfer Diagram . 4-189
4-103 Subinstruction DINC with Plus Zero, Data Transfer Diagram . . 4-190
4-104 Subinstruction DINC with Negative Quantity, Data
Transfer Diagram . 4-191
4-105 Subinstruction DINC with Minus Zero, Data Transfer Diagram . 4-192
4-106 Subinstruction PCDU, Data Transfer Diagram . 4-193
4-107 Subinstruction MCDU, Data Transfer Diagram . 4-194
4-108 Subinstruction SHINC, Data Transfer Diagram . 4-195
4-109 Subinstruction SHANC, Data Transfer Diagram . 4-196
4-110 Subinstruction TCSAJ3, Data Transfer Diagram . 4-197
4-111 Subinstruction FETCH0, Data Transfer Diagram . 4-198
4-112 Subinstruction FETCH1, Data Transfer Diagram . 4-199
4-113 Subinstruction STOREO, Data Transfer Diagram . 4-200
4-114 Subinstruction STORE 1, Data Transfer Diagram . 4-201
4-115 Subinstruction INOTRD, Data Transfer Diagram . 4-202
4-116 Subinstruction INOTLD, Data Transfer Diagram . 4-203
4-117 Timer, Functional Diagram . 4-205/4-206
4-118 LGC Oscillator, Schematic Diagram . 4-209/4-210
4-119 Clock Divider Logic . 4-213/4-214
4-120 Scaler . 4-219/4-220
4-121 Scaler Waveforms . 4-223
4-122 Time Pulse Generator Logic . 4-227/4-228
4-123 Time Pulse Generator Waveforms . 4-230
4-124 Sync and Timing Logic . 4-231/4-232
Volume II
4-125 Order Code Processor, Block Diagram . 4-233
4-126 Command Generator, Block Diagram . 4-235
4-127 Control Pulse Generator, Block Diagram . 4-236
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ILLUSTRATIONS (cont)
Figure Page
4-128 Register SQ Control, Logic Diagram . 4-239/4-240
4-129 Register SW and Decoder, Logic Diagram . 4-243/4-244
4-130 Stage Counter and Decoder, Logic Diagram . 4-247/4-248
4-131 Subinstruction Decoder, Logic Diagram . 4-257/4-258
4-132 Instruction Decoder, Logic Diagram . 4-269/4-270
4-133 Counter and Peripheral Instruction Control Logic . 4-273/4-274
4-134 Crosspoint Generator, Logic Diagram . 4-281/4-282
4-135 Control Pulse Gates, Logic Diagram . 4-351
4-136 Branch Control, Logic Diagram . 4-359/4-360
4-137 Word Formats . 4-366
4-138 Central Processor, Functional Diagram . 4-369/4-370
4-139 Flip-Flop Register, Single Bit Positions . 4-371
4-140 Write, Clear, and Read Timing . 4-372
4-141 Addressable Registers Service . 4-373/4-374
4-142 Flip-Flop Registers . 4-375/4-376
4-143 Register A Service . 4-391/4-392
4-144 Register L Service . 4-395
4-145 Register Q Service . 4-396
4-146 Register Z Service . 4-397
4-147 Z15 and Z16 Set (Sign Test During DV1) . 4-398
4-148 Register B Service . 4-399
4-149 Register G Service . 4-401/4-402
4-150 Editing Control . 4-403
4-151 Editing Transformations . 4-404
4-152 Adder Service (Registers X and Y) . 4-409/4-410
4-153 Carry Logic . 4-412
4-154 Memory Address Register (S) . 4-417/4-418
4-155 Address Decoder . 4-421/4-422
4-156 Counter Address Signals . 4-427
4-157 Parity Logic . 4-429/4-430
4-158 Priority Control, Functional Block Diagram . 4-433/4-434
4-159 Input-Output Channels, Functional Diagram . 4-437/4-438
4-160 Inlink Functional Diagram . 4-440
4-161 Outlink, Functional Diagram . 4-441/4-442
4-162 Erasable Memory, Functional Diagram . 4-445/4-446
4-163 Erasable Memory Timing Diagram . 4-448
4-164 X and Y Selection, Simplified Diagram . 4-451/4-452
4-165 Fixed Memory, Functional Diagram . 4-453/4-454
4-166 Fixed Memory, Timing Diagram . 4-459
4-167 Power Supply, Functional Diagram . 4-461/4-462
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ILLUSTRATIONS (cont)
Figure Page
4-168 +4 VDC Power Supply, Schematic Diagram . 4-465/4-466
4-169 +14 VDC Power Supply, Schematic Diagram . 4-469/4-470
4-170 Alarm Detection Circuits, Schematic Diagram . 4-487/4-488
4- 171 DSKY, Functional Diagram . 4-493/4-494
5- 1 LEM Mission . 5-3/5-4
5-2 LEM IMU Coarse Alignment . 5-3
5-3 LEM IMU Fine Alignment . 5-3
5-4 Powered Descent . 5-6
5- 5 Powered Ascent . 5-8
6- 1 Typical Universal Test Station Layout . 6-11/6-12
7- 1 Primary Guidance, Navigation, and Control System Master
Checkout Flowgram . 7-17/7-18
7-2 Primary Guidance, Navigation, and Control System
Checkout Preparation Flowgram . 7-19/7-20
7-3 Primary Guidance, Navigation, and Control System
Checkout Flowgram . 7-21/7-22
7-4 Inertial Subsystem Master Checkout Flowgram . 7-23/7-24
7-5 Inertial Subsystem Checkout Preparation Flowgram . 7-25/7-26
7- 6 Inertial Subsystem Checkout Flowgram . 7-27/7-28
8- 1 Maintenance Flowgram . 8-3
C-l NOR Gate Symbols . C-2
C-2 NOR Gate Schematic . C-4
C-3 NOR Gate Flip-Flop . C-5
C-4 Logic Diagram Symbols . C-6
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TABLES
Number
Volume I
Page
1-1 SCS Interface Signals . 1-13
1— II Displays and Controls . 1-14
1 -ITT Description of Landing Radar Interface Signals . . 1-16
2-1
Instruction Classes
2-49
3-1 LEM PGNCS Components . 3-1
3-II PGNCS Harness Interconnections . 3-4
3-ID Locations and Functions of IMU Electronics . 3-9
3-IV Locations and Functions of PTA Modules . 3-14
3-V PTA Test Points . 3-16
3- VI Locations and Functions of PSA Modules . 3-18
3-VD Functions of CDU Modules . 3-26
3-VHI DSKY Controls and Indicators . 3-28
4-1 Program Storage Allocation . 4-45/4-46
4-H Functional Organization of Machine Instructions . 4-53
4-HI Counter Instructions . 4-59
4-IV Machine Instructions, Alphabetical Listing . 4-60
4-V Subinstructions . 4-68
4-VI Control Pulses . 4-73
4-VH Subinstruction Codes and Control Pulses . 4-81/4-82
4-VHI Scaler Outputs (Stages 1-17) . 4-225
Volume II
4-IX Commands Per Sub instruction . 4-251
4-X Subinstructions Per Command . 4-264
4-XI Counter Cell Signals . 4-278
4-XU Subinstruction CCSO . 4-280
4-XIH Subinstruction DVO . 4-303
4-XIV Subinstruction DV1, Part 1 . 4-304
4-XV Subinstructions DV3, DV7, and DV6, Part 1 4-305
4-XVI Subinstructions DV1, DV3, DV7, and DV6, Part 2 . 4-306
4-XVH Subinstruction DV4 . 4-307
4-XVIH Subinstruction MP0 . 4-309
4-XIX Subinstruction MP1 . 4-310
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TABLES (cont)
Number Page
4-XX Subinstruction MP3 . 4-311
4-XXI Crosspoint Pulse ZIP . 4-312
4-XXH Subinstruction STD2 . 4-314
4-XXHI Subinstruction TCO . 4-314
4-XXIV Subinstruction TCFO . 4-315
4-XXV Subinstruction TCSAJ3 . 4-315
4-XXVI Subinstruction GOJ1 . 4-315
4-XXVII Subinstruction DASO . 4-316
4-XXVm Subinstruction DAS1 . 4-317
4-XXIX Subinstruction LXCHO . 4-318
4-XXX Subinstruction INCRO . 4-318
4-XXXI Subinstruction ADSO . 4-319
4-XXXn Subinstructions CAO and DCAl . 4-320
4-XXXHI Subinstructions CSO and DCS1 . 4-320
4-XXXTV Subinstruction NDXO . 4-321
4-XXXV Subinstruction RSM3 . 4-321
4-XXXVI Subinstruction NDX1 . 4-322
4-XXXVD Subinstruction XCHO . 4-323
4-XXX VIII Subinstruction DXCHO . 4-324
4-XXXIX Subinstruction DXCH1 . 4-324
4-XL Subinstruction TSO . 4-325
4-XLI Subinstruction ADO . 4-326
4-XLII Subinstruction MASKO . 4-327
4-XLIH Subinstruction BZFO . 4-328
4-XLIV Subinstruction MSUO . 4-329
4-XLV Subinstruction QXCHO . 4-330
4-XLVI Subinstruction AUGO . 4-330
4-XLVII Subinstruction DIMO . 4-331
4-XLVHI Subinstruction DCAO . 4-332
4-XLIX Subinstruction DC SO . 4-333
4-L Subinstruction SUO . 4-334
4-LI Subinstruction NDXXO . 4-334
4-LII Subinstruction NDXX1 . 4-335
4- LIH Subinstruction BZMFO . 4-336
4-LIV Subinstruction READO . 4-337
4-LV Subinstruction WRITE 0 . 4-338
4-LVI Subinstruction RANDO . 4-339
4-LVII Subinstruction WANDO . 4-340
4-LVHI Subinstruction RORO . 4-341
4-UX Subinstruction WORO . 4-341
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TABLES (cont)
Number Page
4- LX Subinstruction RXORO . 4-342
4-LXI Subinstruction RUPTO . 4-343
4-LXII Subinstruction RUPT1 . 4-343
4-LXIII Subinstruction PINC . 4-344
4-LXTV Subinstruction MINC . 4-344
4-LXV Subinstruction PC DU . 4-345
4-LXVI Subinstruction MCDU . 4-345
4-LXVH Subinstruction DINC . 4-346
4-LXVEI Subinstruction SHINC . 4-347
4-LXIX Subinstruction SHANC . 4-347
4-LXX Subinstruction INOTRD . 4-348
4-LXXI Subinstruction INOTLD . 4-348
4-LXXH Subinstructions FETCHO and STOREO . 4-349
4-LXXHI Subinstruction FETCH1 . 4-349
4-LXXIV Subinstruction STORE 1 . 4-350
4-LXXV Control Pulse Orgin . 4-357
4-LXXVI Register A and L Write Line Inputs . 4-393
4-LXXVH Write Amplifiers External Inputs . 4-413/4-414
4-LXXVIII Erasable Memory Address Selection . 4-425/4-426
4-LXXIX E Addressing . 4-447
4-LXXX F Addressing . 4-455
4-LXXXI Power Distribution . 4-472
6-1 Checkout and Maintenance Test Equipment . 6-1
6-U Checkout and Maintenance Tools . 6-5
6- III List of Operating Procedure JDC's for GSE . 6-6
7- 1 Equipment Required for Checkout . 7-2
7-n PGNCS Interconnect Cables . 7-4
7— III Inertial Subsystem Interconnect Cables . 7-9
7- IV Computer Subsystem Interconnect Cables . 7-14
8- 1 PGNCS and ISS Loop Diagrams and Schematics . 8-4
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MANUAL
LIST OF RELATED MANUALS
-1021038 -1021039 -1021040 ND-1021043
Packing, Shipping and Handling Manual Auxiliary Ground Support Equipment Manual Bench Maintenance Ground Support Equipment Manual Block II Primary Guidance, Navigation, and Control System Manual
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MANUAL
INTRODUCTION
This manual provides information necessary for checkout, maintenance, and re¬ pair of the lunar excursion module (LEM) primary guidance, navigation, and control system (PGNCS) (figure 1-1). Included in the manual are functional analysis, detailed theory of operation, component description, system tie-in, and description of flight operations. The manual also provides for an introduction and complete familiarization with the PGNCS.
Job Description Cards (JDC's) containing detailed step-by-step procedures are contained in separate supplementary volumes. Listings of the JDC's required for given tests and the sequence of performing the JDC's are included in the manual.
This manual and its JDC's cover PGNCS system part number 6015000-011 and shall be used in the laboratories at Kennedy Space Center, the Manned Spacecraft Center (MSC), and at Grumman Aircraft Engineering Corporation (GAEC). Portions of this manual pertaining to the luminous beacon are also applicable for use in the laboratories at North American Aviation (NAA). Source data available as of 15 January 1966 was used in preparation of the basic issue of this manual.
This manual is prepared in accordance with E-1087 Documentation Handbook and National Aeronautics and Space Administration (NASA) contract NAS 9-497, exhibit D.
Appendix A contains a listing of technical terms and abbreviations used in the manual. Appendix B explains the function and relationship of the System Identification Data List (SIDL) to the manual. Appendix C will contain the logic symbols used in the discussion of the computer logic diagrams.
Changes to the manual are requested by sending a completed Technical Data Change Request (TDCR) form to:
Apollo Field Service Publications, Department 38-01
AC Electronics Division GMC
PLT Ml
Milwaukee, Wisconsin 53201
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LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM MANUAL
LEM GUIDANCE COMPUTER-
INERTIAL
MEASURING
UNIT.
COUPLING DATA UNIT
POWER AND SERVO ASSEMBLY
DISPLAY AND
KEYBOARD
OPTICAL
TRACKER
LUMINOUS BEACON (LOCATED ON AN ADAPTER RING BETWEEN THE COMMAND AND SERVICE MODULE)
SIGNAL CONDITIONER
PULSE
TORQUE ASSEMBLY
I5775B
|
, l |
(R\ |
|
Tf ) |
|
|
II l |
jjl |
|
\ t - 1 |
' |
t
Figure 1-1. LEM Primary Guidance, Navigation, and Control System
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Chapter 1
SYSTEM TIE-IN
1-1 SCOPE
This chapter presents the lunar excursion module (LEM) mission. The chapter also describes the functional interface between the primary guidance, navigation, and control system (PGNCS) and the other spacecraft systems.
1-2 LEM MISSION
The purpose of the LEM mission is to transfer the LEM from a circular lunar orbit into a descent orbit, land two astronauts on the lunar surface, and return them to the orbiting command and service module (CSM). The LEM mission (figure 1-2), with respect to the PGNCS, is best described by dividing it into six phases: separation and transfer orbit insertion, descent coast, powered descent and landing, lunar stay, launch and powered ascent, and rendezvous and docking.
1-2.1 SEPARATION AND TRANSFER ORBIT INSERTION. Approximately one hour before the LEM enters the descent orbit, two astronauts leave the CSM and enter the LEM through the top docking hatch. The crew then checks out the various LEM systems, establishes a voice link, and, after initial PGNCS turnon, establishes a time reference for the LEM guidance computer (LGC), and coarse aligns the inertial meas¬ uring unit (IMU) using CSM data. One astronaut then manually commands reaction control system (RCS) jet firing to separate the LEM from the CSM. The IMU is fine aligned. Near the end of the second lunar orbit, the LEM descent engine is fired by the PGNCS and the LEM begins its descent. The timing and duration of LEM descent engine firing is critical, to insure the proper elliptical Hohmann transfer orbit.
1-2.2 DESCENT COAST. During the descent coast phase, the LEM is in free fall on an elliptical flight path. During free fall, the astronauts check out the landing radar (LR). At the perilune of the Hohmann transfer orbit, the LEM is at an altitude of approximately 50,000 feet and has a velocity vector essentially parallel to the lunar surface. During this phase, the PGNCS determines the flight parameters required for powered descent.
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CSM
Figure 1-2. LEM Mission Phases
1-2.3 POWERED DESCENT AND LANDING. In preparation for powered descent, an IMU fine alignment is performed. At the perilune of the descent orbit, the PGNCS issues a descent engine start discrete. The descent engine firing slows the LEM which begins the actual descent to the lunar surface. During descent, the PGNCS con¬ trols the engine trim and thrust level, controls the LEM attitude, and provides visual displays of the guidance system status. During the final approach and landing, the PGNCS holds the LEM at a constant attitude, allowing the astronaut to view the landing site. The astronaut can select a new landing site by inserting new landing site coordinates into the LGC. The LGC will automatically control the RCS and the descent engine to guide the LEM to the new landing site. Inertially derived flight parameters are up¬ dated in the LGC by comparison with the altitude and velocity parameters determined from LR measurements.
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1-2.4 LUNAR STAY. After LEM touchdown the astronauts check out all systems for damage and insure that the systems can perform the functions required for a successful ascent. All equipment not required for lunar stay is then turned off. The astronauts survey the surrounding lunar landscape, secure the hatches, and perform a final check on the portable life support system (PLSS). After the LEM is secured, one astronaut, wearing the PLSS, leaves the LEM to explore the lunar surface. The exploring astronaut inspects the LEM and sets up communication antennas. A television system sends pictures of the lunar scene to earth. The astronaut always in direct voice contact with the LEM, explores the lunar surface, makes photographic records, and collects surface samples. After approximately three hours, the astronaut must return to replenish his PLSS. Additional surface explorations depend upon the planned stay time. Near the end of the lunar stay, the PGNCS is brought to an operate con¬ dition and the IMU is coarse and fine aligned. The IMU is fine aligned to a known reference coordinate system by making star sighting measurements. The LEM optical device tracks the orbiting CSM and sends data to the LGC which calculates applicable flight parameters in preparation for the launch and powered ascent.
1-2.5 LAUNCH AND POWERED ASCENT. After the astronauts prepare the LEM, the PGNCS determines time of launch and ascent trajectory based on a fixed rendezvous aim point. Mechanical and electrical separation of the two LEM stages takes place and the LGC issues the ascent engine start discrete at a time calculated to effect a successful rendezvous.
During powered ascent, the LEM rises vertically and then is pitched to attain a Hohmann transfer orbit for the rendezvous. Because the ascent engine is a fixed- position, fixed-thrust engine, the LEM attitude during ascent is controlled by the LGC which issues commands to the RCS jets. The LGC determines necessary RCS commands by comparing calculated values with actual flight parameters obtained from the inertial subsystem (ISS), and determines required attitude changes to correct any differ¬ ences. When the injection of the LEM into the proper elliptical orbit is accomplished, the LGC issues the ascent engine off discrete and the LEM enters the coasting portion of the ascent phase.
1-2.6 RENDEZVOUS AND DOCKING. LEM guidance during this phase is a combination of optical tracking data and inertial data. Azimuth and elevation data from the optical tracking device and velocity and attitude information from the IMU are used by the LGC to control the RCS to maintain attitude and to provide a display of position and velocity information. During rendezvous, the LEM is maintained at an orientation such that the CSM is visible through the vehicle windows.
Terminal rendezvous maneuvers begin when the LEM and CSM are approximately five nautical miles apart. The LGC computes the intercept time and with this data updates the thrust vector and velocity requirements. Three ascent engine bums during terminal rendezvous reduce the closing rate to near zero. The LGC utilizes the RCS
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to maintain vehicle attitude during these burns. The final step is docking, which is initiated when the vehicles are approximately 500 yards apart. The astronaut uses the translation controller and attitude controller in a computer-aided manual operation to guide the LEM to hard docking with the CSM. The two astronauts then leave the LEM and transfer to the CSM through the vehicle's forward tunnel to prepare for the return to earth. The LEM is jettisoned following crew transfer to the CSM.
1-3 LEM STRUCTURE
The LEM (figure 1-3) has two stages mated to form one structure: the ascent stage and the descent stage. These stages and the umbilical interconnecting cables can be separated at launch from the lunar surface or because of mission abort during descent.
The approximate LEM external dimensions are shown in figure 1-4. At earth launch, the weight of the LEM is approximately 30,000 pounds.
1-3.1 ASCENT STAGE. The ascent stage, constructed mainly of aluminum alloy, con¬ sists of the crew compartment, a midsection, aft equipment bay, tankage sections, associated hatches, and windows.
From the crew compartment, the astronauts control all phases of the LEM mission. The crew also uses this compartment as their operations center during their lunar stay.
The displays and controls associated with the PGNCS are located at the front of the crew compartment. The IMU, a portion of its electronics, and the optical tracking device are located in an enclosure above the crew compartment. The remaining PGNCS components are mounted on coldplates to the rear wall of the ascent stage midsection.
The midsection is cylindrical, smaller than the crew compartment, and directly behind it. The ascent engine and related components are in the midsection, the LEM’s center of gravity. Also contained in the LEM’s midsection are the ascent engine hatch, top hatch, environmental control system (ECS), and equipment that requires crew ac¬ cessibility.
To transfer from the CSM to the LEM while in lunar orbit, the crew uses the upper docking tunnel at the top centerline of the ascent stage. The forward tunnel, at the lower front of the crew compartment, is used for entering and leaving the LEM while on the lunar surface.
The aft equipment bay, at the rear of the vehicle, is separated from the mid¬ section by a pres sure -tight bulkhead. This area houses the glycol loop for the ECS, inverters, batteries, and equipment for the electrical power system (EPS).
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Figure 1-3. LEM
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MANUAL
Figure 1-4. LEM External Dimensions
1-6
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MANUAL
The propellant tankage sections are located on either side of the midsection outside the pressurized area. The tankage sections contain the ascent engine fuel and oxidizer tanks; RCS fuel, oxidizer, and helium tanks; and ECS water tanks. The ratio by weight of oxidizer to fuel is 1.6 to 1; therefore, to maintain the lateral center of gravity on the vehicle X axis, the ascent engine propellant tanks are offset to one side.
Two triangular windows in the front face of the crew compartment provide visi¬ bility. Each window has approximately 1.6 square feet of viewing area and are canted down and to the side to increase visibility. Each window consists of two panes.
1-3.2 DESCENT STAGE. The descent stage, constructed mainly of aluminum alloy, has equipment necessary to land on the lunar surface. It is also a platform for the launching of the ascent stage after completion of the lunar exploration. The descent engine is the center of the stage surrounded by its four main propellant tanks. In addition to the descent engine and its related components, the descent stage houses the descent control instrumentation; scientific equipment; EPS batteries; and tanks for water used by the ECS. Landing gear and the LR antenna are attached to the descent stage.
1-4 LEM SYSTEMS
Functionally, there are seven LEM systems. Four of these systems control the LEM flight. The PGNCS or the stabilization and control system (SCS) receives inputs from the crew and electrical inputs from the inertial sensors to generate commands that result in rotation and translation maneuvers. The RCS or propulsion system provides external forces and mechanical couples to maneuver the LEM under the control of the PGNCS or the SCS. The crew obtains information from the LGC (part of the PGNCS), by communications (Manned Space Flight Network), or displays that indicate the necessity to initiate one or more of the basic LEM motions. The three remaining LEM systems are indirectly related to LEM control. They provide the power (EPS), environmental control (ECS), and the communications [[communi¬ cations and instrumentation system (CIS)].
1-4.1 PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM. The PGNCS provides the measuring and data processing capabilities and control functions nec¬ essary to accomplish the LEM mission. The PGNCS utilizes inertial components for guidance, an optical device for navigation, and a digital computer for data processing and issuance of flight control signals.
The inertial guidance portion of the PGNCS, the IMU, employs accelerometers mounted on a gyroscopically stabilized gimbal-mounted platform. The IMU senses acceleration and attitude changes instantaneously and provides signals to a digital com¬ puter, the LGC, for the generation of attitude control and thrust commands.
For navigation, the PGNCS utilizes an optical tracking device to take star sight¬ ings and obtain measurements. These sightings are used by the LGC to establish proper alignment of the stable platform. The LGC contains a catalog of celestial
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bodies and is programmed to calculate alignment commands using the information obtained from the optical sightings. In addition to functioning as a data processing unit, the LGC, through its flight programs, performs the function of a digital auto¬ pilot in controlling the LEM.
1-4.2 STABILIZATION AND CONTROL SYSTEM. The SCS consists of two major sections: the control electronics section (CES) and the abort guidance section (AGS). The CES processes flight control signals during all mission phases. The AGS pro¬ vides the CES automatic steering commands, derived from explicit guidance equa¬ tions, in the event of mission abort due to a PGNCS malfunction.
The CES consists of an attitude and translation control assembly (ATCA), a descent engine control assembly (DEC A), rate gyro assembly (RGA), two translation con¬ troller assemblies (TCA), and two attitude controller assemblies (AC A). The CES processes and routes signals to fire any combination of the 16 thrusters in the RCS to control LEM attitude and translation. The attitude and translational control inputs originate from any of three sources: the PGNCS during normal automatic operation, the AC A and TCA during manual operations, or the AGS during an abort.
The CES converts the applicable input commands into pulsed or constant level signals and routes them to the RCS to fire the appropriate thrusters. Rate signals from the CES are displayed on the flight director attitude indicator (FDAI).
The CES also processes '’ON-OFF'* commands for the ascent and descent engines, and routes automatic and manual throttle commands to the descent engine. Trim control of the descent engine insures that the thrust vector operates through the vehicle center of gravity.
The AGS provides abort capability from any point in powered descent or powered ascent and increases crew safety by acting as a backup system to the PGNCS. The AGS has three main assemblies: abort sensor assembly, abort electronics assembly, and data entry and display assembly.
The abort sensor assembly utilizes a strap-down technique employing three single- degree-of-freedom integrating rate gyros and three accelerometers. This backup guidance provides vehicle attitude, angular velocity, and translational acceleration indications. The outputs of the abort sensor assembly go to the abort electronics assembly, a 4,096 word capacity general purpose computer. Computations are performed using the inputs from the abort sensor assembly. When the AGS is in control of the LEM, the results are displayed and control signals are issued to the vehicle's reaction control and propulsion systems.
The abort sensor assembly measures the accelerometer triad rotation from, and resolves the acceleration into, a fixed reference frame. This reference frame is provided by an initial alignment of the AGS with the PGNCS. Initial alignment is required for attitude, velocity, time, and position. Velocity and position vectors are manually entered into the computer by a data entry device available to the astronaut.
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Attitude alignment is accomplished by transferring PGNCS IMU gimbal angles to the computer. The abort electronics assembly receives this data from the coupling data unit (CDU) in the same manner and at the same time as the LGC (i.e. incremental angles accumulated from a zero reference after "CDU ZERO").
1-4.3 PROPULSION SYSTEM. The LEM utilizes separate, complete, and independent descent and ascent propulsion systems, which consist basically of a liquid propellant rocket engine and its propellant storage, pressurization, and feed components.
The descent propulsion system is in the LEM descent stage and utilizes a throttle- controlled, gimbaled engine. The engine injects the LEM into the descent transfer orbit and is used during powered descent and landing to control the rate of descent. The descent engine, developing 10,500 pounds maximum thrust in a vacuum at full throttle and 1,050 pounds minimum thrust, can be gimbaled 6 degrees in any direction. The PGNCS issues the "ON-OFF" commands for the descent engine and also provides sig¬ nals controlling thrust magnitude and gimbal trim position.
The propellant used in both propulsion systems is a 50-50 fuel mixture of hydrazine and unsymmetrical dimethylhydrazine using nitrogen tetroxide as the oxidizer and he¬ lium as the tank pressurant.
The ascent propulsion system utilizes a fixed, constant-thrust engine installed along the centerline of the ascent stage midsection and includes the associated pro¬ pellant feed tanks and pressurization components. The engine develops 3,500 pounds thrust in a vacuum, sufficient to launch the ascent stage from the lunar surface and place it in orbit. The PGNCS issues the "ON-OFF" commands for the ascent engine.
1-4.4 REACTION CONTROL SYSTEM. The RCS provides rocket thrust impulses that stabilize the LEM during descent and ascent and control the LEM attitude and trans¬ lation about or along all axes. The RCS has 16 thrust chambers supplied by two separate and independent propellant feed and pressurization sections. The thrust chambers are mounted in clusters of four on outriggers equally spaced around the LEM ascent stage. In each cluster, two thrust chambers are mounted on a vertical axis, facing in opposite directions; the other two are spaced 90 degrees apart, parallel to the LEM's Y and Z axes. The RCS utilizes the same fuel as the ascent engine. In the event of RCS fuel depletion, the remaining ascent fuel can be used for the RCS. The RCS can be operated in any of three modes: manual, automatic, or semi-automatic. The PGNCS supplies "ON-OFF" signals through the SCS to the valves on the desired thrust chambers during the automatic or semi-automatic mode. The automatic mode is normally used to provide attitude control during all mission phases except when manual control is required. It is possible to select manual control in one or two axes and retain automatic control in the other axis during all mission phases. The semiautomatic mode combines automatic attitude hold control with manual control. The LEM attitude is changeable about each axis using the astronaut’s attitude controller. This mode is used primarily to control the LEM during the rendezvous and docking phase of the mission. In the manual mode, all control commands originate from the attitude con¬ troller, including manual control of the thrust duration.
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All automatic translational commands originate in the PGNCS and are routed to the RCS similar to the attitude control signals.
1-4.5 ELECTRICAL POWER SYSTEM. The EPS provides 28 vdc and 115 vac, 400 cps power to the PQNCS. This power originates from six batteries, four in the descent stage, and two in the ascent stage. The batteries, the silver-zinc type, are rated at 80 watts per hour per pound of weight. The 115 vac, 400 cps power is obtained by routing the 28 vdc through an inverter.
1-4.6 ENVIRONMENTAL CONTROL SYSTEM. The ECS sustains life in space by providing breathable atmosphere, acceptable temperatures, food and water, and waste disposal. In addition, the ECS circulates an ethylene glycol-water coolant about the temperature sensitive electronic equipment in the PGNCS and other LEM systems to provide thermal stability. The IMU has coolant circulated through its case while the power and servo assembly (PSA), pulse torque assembly (PTA), signal conditioner, LGC and CDU are mounted on coldplates through which the coolant is circulated to provide temperature control.
1-4.7 COMMUNICATIONS AND INSTRUMENTATION SYSTEM. The CIS links the lunar astronauts, the orbiting CSM, and earth monitoring stations.
The communications portion contains two radio frequency (RF) sections, one oper¬ ating in the VHF range and the other in the UHF range; a television section; and a signal processing section. In addition to two-way voice communication, the RF section re¬ ceives and transmits tracking and range information, biomedical information, and emergency code keying in the event of voice transmission failure. The television sec¬ tion is used by the extravehicular astronaut to televise the lunar surface within an eighty foot radius of the grounded LEM. In the signal processing section, critical signals of the PGNCS are conditioned and supplied to pulse code modulated (PCM) telemetry equipment for transmission to earth. Telemetry data can be stored when direct com¬ munication with the earth is not possible.
The instrumentation portion provides the astronauts and ground facilities with LEM performance data during the mission by sensing physical status data, monitoring the various systems, and performing inflight and lunar surface checkout. This system also contains the scientific instruments which are used by the astronauts during their lunar stay.
1-5 PGNCS INTERFACE
PGNCS operation during the LEM mission requires the interface of the PGNCS with the other LEM systems, the displays and controls on the crew display and con¬ trol panels, the landing radar, and the astronauts. The functional interface of the PGNCS is shown in figure 1-5.
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Figure 1-5. LEM PGNCS Functional Interface, Block Diagram
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1-5.1 SYSTEMS. Four LEM systems (SCS, ECS, CIS, and EPS) have direct interface with the PGNCS and two systems (propulsion system and RCS) have indirect inter¬ face with the PGNCS. The indirect interface of the propulsion system and the RCS occurs through the SCS. These two systems may thus be controlled by the PGNCS or by the backup control provided by the AGS of the SCS. Descriptions and sources of the SCS interface signals are provided in table 1-1. Descriptions of the interfaces with the other systems are provided in paragraph 1-4.
1-5.2 DISPLAYS AND CONTROLS. Several displays and controls located on the crew control panels, LGC display and keyboard (DSKY) panel, and the SCS control panel interface with the PGNCS. Two sets of hand controllers are provided for manual control of the LEM and interface with the PGNCS. Descriptions of the displays and controls are in table 1-IL
1-5.3 LANDING RADAR. The landing radar (LR) provides data to the LGC from which LEM velocity (in antenna coordinates) and LEM altitude may be determined. The data is also available for visual display, independent of the PGNCS, except that the velocity is in spacecraft coordinates.
The landing radar which operates in the X-band, consists of an antenna assembly, a solid-state electronics assembly, and a control panel. Velocity data is acquired from a three beam continuous wave Doppler radar. Altitude data is provided by a one beam FM continuous wave radar altimeter. The antenna assembly accommodates the requirements of both the Doppler and the altimeter beams.
Landing radar and PGNCS interface include digital data transfer, scaling, velocity and range sensing, status, and antenna positioning. Descriptions and sources of the interface signals are in table 1-in.
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Table 1-1. SCS Interface Signals
|
Signal Name |
Source |
Description |
|
Manual translation commands (±x, ±y, ±z) |
SCS |
Signals from translation controller which fire RCS jets by LGC control. |
|
Attitude control out of detent |
SCS |
Signal from attitude controller indicating that it is not in neutral position. |
|
Rate of descent (±) |
SCS |
Discretes commanding an increase or decrease in rate of descent. |
|
Gimbal off (pitch, roll) |
SCS |
Signal to LGC indicating that descent engine pitch or roll gimbal is off null. |
|
Trim commands (± pitch, ± roll) |
LGC |
Signals which control trim of descent engine. |
|
Engine on-off |
LGC |
Signal to turn descent or ascent engine on or off. |
|
Descent engine throttle command (decrease, increase) |
LGC |
Signal to increase or decrease thrust of descent engine. |
|
RCS jets on-off |
LGC |
Signals (16) to turn RCS jets on or off. |
|
Increments of IMU gimbal angles (±Z*0IG, ±A9mg» ±^-0OG |
LGC |
Supplies changes in IMU gimbal angles to AGS. |
|
CDU zero (initial clear) |
LGC |
Sets alignment logic of AGS to zero. |
|
800 cps ±1% |
PGNCS |
Provides reference between PGNCS and SCS. |
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LEM PRIMARY 6UIDANCE, NAVIGATION, AND CONTIOL SYSTEM |
Table 1-n. Displays and Controls
|
Display or Control |
Function |
|
GUID CONT switch |
Selects either primary guidance (PGNS) or abort guidance (AGS). Normally in the PGNS position. |
|
MODE SEL selector |
Three position switch used during landing phase to select one of three inputs to be displayed on AZ RT/ELEV RT-LAT VEL/FWD VEL indicator. Inputs are landing radar (LDG RADAR), PGNS and AGS. |
|
RNG/ALT MON switch |
Controls display of RANGE/RANGE RATE- ALT/ ALT RATE indicator. Positions are RNG/RNG RT and ALT/ALT RT. |
|
RATE/ERR MONITOR switch (2) |
Selects one of two inputs for AZ RT/ELEV RT-LAT VEL/FWD VEL indicator and attitude needles of FDAI. |
|
ATTITUDE MON switch (2) |
Selects one of two inputs to FDAI total attitude dis¬ play and attitude error needles during landing. |
|
THR CONT switch |
Selects either automatic (AUTO) or manual (MAN) control of descent engine throttle. Normally in AUTO position. |
|
MAN THROT switch |
Activates either commander’s (CDR) or system engineer's (SE) translation controller for manual throttling of descent engine. |
|
ABORT |
Pushbutton to cause mission abort at any point be¬ tween LEM/CSM separation and touchdown on lunar surface with descent stage still attached. |
|
ABORT STAGE |
Pushbutton to cause mission abort using ascent stage. |
|
AZ RT/ELEV RT- LAT VEL/FWD VEL meter (2) |
Provides visual displays of vehicle forward and lateral velocity during landing. |
|
(Sheet 1 of 2) |
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Table l-II. Displays and Controls
|
Display or Control |
Function |
|
RANGE/RANGE RATE- ALT/ALT RATE meter |
Provides visual displays of range, altitude, range rate, and altitude rate. |
|
FDAI meter (2) |
Provides three visual displays, total attitude, atti¬ tude error, and attitude change rate. PGNCS or AGS provides inputs for total attitude and attitude error. Attitude rate signals are provided by SCS rate gyros. |
|
LGC and ISS warning indicators, PGNS caution indicator. |
Controlled by instrumentation system which re¬ ceives discretes from LGC when certain PGNCS troubles exist. |
|
MODE CONTROL selector |
A three-position selector located on SCS control panel concerned with attitude control. Positions are OFF, ATT HOLD, and AUTO. In AUTO posi¬ tion, fully automatic attitude control is achieved through PGNCS or AGS control of RCS jets. ATT HOLD position allows crew to manually reposition LEM and have new position automatically maintained by LGC. |
|
IMU CAGE switch |
Switch located on DSKY mounting panel to drive IMU gimbal angles to zero. |
|
Attitude controller (2) |
Three-axis, pistol-grip, right-hand device for manual attitude control of LEM. Outputs from controller are processed by PGNCS or may be routed directly to RCS. |
|
Translation controller (2) |
Three-axis, T-handle, left-hand device for manual translation control of LEM. Using switch located next to T-handle, controller can operate RCS jets or throttle the descent engine. |
|
(Sheet 2 of 2) |
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Table 1-ID. Description of Landing Radar Interface Signals
|
Signal Name |
Source |
Description |
|
Antenna positioning command |
DSKY and LGC |
Changes antenna position. |
|
Antenna position #1 (descent) |
LR |
Indicates to LGC that antenna is in posi¬ tion #1. |
|
Antenna position #2 (hover) |
LR |
Indicates to LGC that antenna is in posi¬ tion #2. |
|
Velocity data good |
LR |
Indicates to LGC that LR velocity trackers have locked on. |
|
Range data good |
LR |
Indicates to LGC that LR range trackers have locked on. |
|
Range low scale factor |
LR |
Indicates to LGC that a change in scale factor is necessary. Issued automat¬ ically at approximately 2,500 feet. |
|
LR in "0" and LR in "l" |
LR |
Digital pulses sent to LGC which contain range and velocity data. |
|
Readout command |
LGC |
Indicates that LGC is ready to receive LR data pulses. |
|
Gate reset |
LGC |
3,200 cps continuous LGC output to reset LR transfer gates. |
|
Range strobe |
LGC |
Timing pulses to enable LR transfer gates. |
|
vxa» Vya, VZa strobe pulses |
LGC |
Timing pulses to enable LR transfer gates. |
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Chapter 2
SYSTEM AND SUBSYSTEM FUNCTIONAL ANALYSIS
2-1 SCOPE
This chapter provides functional descriptions of the PGNCS and its subsystems. This chapter describes how the PGNCS subsystems perform the PGNCS operations.
2-2 PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM.
The PGNCS is functionally divided into three major subsystems: inertial, optical, and computer. The PGNCS performs three basic functions: inertial guidance, navigation, and autopilot stabilization and control. Within these functions the subsystems, or combination of subsystems, with assistance from the astronaut, perform the following operations:
(1) Establish an inertial reference which is used for measurements and compu¬ tations.
(2) Aligns the inertial reference by optical measurements and, through inter¬ face, aligns the inertial reference with the CSM PGNCS.
(3) Calculates the position and velocity of the LEM by inertial navigation.
(4) Accomplishes a LEM and CSM rendezvous by optical navigation and inertial guidance.
(5) Generate attitude control and thrust commands to maintain the LEM on a satis¬ factory trajectory.
(6) Control throttling of descent engine during lunar landing.
(7) Display pertinent data related to guidance status.
(8) Controls ascent engine burn time to obtain proper velocity for rendezvous orbit.
To perform its inertial guidance functions, the PGNCS employs an IMU containing accelerometers mounted on a gyro stabilized, gimbal-mounted platform. The IMU, three channels of the CDU, the pulse torque assembly (PTA), and the PSA form the ISS of the PGNCS.
2-1
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To perform its navigation functions, the PGNCS employs the LEM optical rendez¬ vous subsystem (LORS), which consists of an optical tracker located on the LEM and a luminous beacon located on the CSM. During the powered descent and landing phase, the PGNCS receives altitude and velocity data from the LR, which is used to update or check inertially derived data.
The LGC is a digital computer which serves as both the control element and the primary data processing element of the PGNCS. The LGC and the display and key¬ board (DSKY) form the computer subsystem of the PGNCS.
Figure 2-1 illustrates the signalflow and interface between the three PGNCS sub¬ systems.
2-3 LEM AND PGNCS AXES
Several sets of axes are associated with the LEM and PGNCS. Figure 2-2 illustrates these various orthogonal sets which are defined in the following paragraphs. Positive rotation about each axis is as defined by the right hand rule.
2-3.1 LEM SPACECRAFT AXES. The LEM spacecraft axes provide a reference for all other sets of axes and define the point about which attitude maneuvers are performed. The LEM spacecraft axes, designated XleM, YleM. ZleM. are referred to as the yaw, pitch, and roll axes respectively. The Xj_,EM axis points through the upper dock¬ ing hatch and the ZleM axis points through the forward hatch. The YleM axis is perpendicular to the XleM and the ZleM axes and can be considered to be pointing out of the astronaut’s right shoulder as he faces toward the forward portion of the LEM.
2-3.2 NAVIGATION BASE AXES. The navigation base provides a precise alignment of the IMU to the optical tracker and a means of attaching both units to the spacecraft. The navigation base is mounted to the LEM structure so that a coordinate reference system is formed by its mounting points. The Yj^b axis is defined by the centers of the two upper mounting points and is parallel to the YleM axis. The Xjsjg axis is defined by a line through the center of the lower mounting point, perpendicular to the yNB axis and parallel to the Xlem axis. The Zjyjg axis is mutually perpendicular to the Xnb and ynb axes and is parallel to the Zlem axis.
2-3.3 INERTIAL AXES. The inertial axes provide references for measuring changes in velocity and attitude. At zero degree, the inertial axes are parallel to the navigation base axes.
2-3. 3.1 Gimbal Axes. The gimbal axes (outer, middle, and inner) are the axes of the movable gimbals. The axes are defined by the intergimbal assemblies which provide each gimbal with rotational freedom. The attitude of the spacecraft with respect to the stable member is measured by the gimbal resolvers located in the intergimbal assemblies.
2-3. 3. 2 Stable Member Axes. The stable member axes (Xsm» Ysm. Zsm) provide a reference for aligning the inertial components and for defining the angular orientation of the inertial axes during flight.
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Figure 2-1. PGNCS Subsystems Interface, Block Diagram
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OPTICAL TRACKER AZIMUTH AXIS
Figure 2-2. LEM and PGNCS Axes
2-3. 3. 3 Accelerometer Axes. The accelerometer axes (Xa, Ya, Za) are the positive input axes of the accelerometers and are parallel to the stable member axes. Velocity changes are measured along the accelerometer input axes. This velocity data is used to determine spacecraft position and velocity.
2- 3.3. 4 Gyro Axes. The gyro axes (Xg, Y~, Z_) are the positive input axes of the stabilization gyros and are parallel to the stable member axes. If the attitude of the stable member is changed with respect to inertial space, the gyro senses the change about its input axis and provides an error signal to a servo loop which realigns the stable member to its original orientation.
2-4 INERTIAL SUBSYSTEM
The ISS performs three major functions. It measures changes in LEM attitude, assists in generating steering commands, and measures spacecraft velocity due to thrust. To accomplish these functions, the IMU provides an inertial reference consisting of a stable member with a three degree of freedom gimbal system and stabilized by
2-4
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three rate integrating gyros. Each time the inertial subsystem is energized, the stable member must be aligned with respect to a predetermined reference. During flight and prior to launch from the lunar surface, this alignment is accomplished by sighting the optical instrument on celestial objects.
Once the ISS is energized and aligned, any rotational motion of the LEM will be about the stable member, which remains fixed in space. Resolvers mounted on the gimbal axes act as angular sensing devices and measure the attitude of the LEM with respect to the stable member. These angular measurements are displayed by the FDAI and angular changes are sent to the LGC via the CDU.
The desired LEM attitude is calculated in the LGC and compared with the actual gimbal angles. Any difference between the actual and calculated angles results in the generation of attitude error signals by the ISS Chanels of the CDU which are sent to the FDAI for display.
Vehicle acceleration is sensed by three pendulous accelerometers mounted on the stable member with their input axes orthogonal. The signals from the accelerometers are supplied to the LGC which calculates the total vehicle velocity.
The modes of operation of the inertial subsystem can be initiated automatically by the LGC or by the astronaut selecting computer programs through the DSKY. The status or mode of operation is displayed on the DSKY.
For pruposes of explanation and description, the ISS is divided into functional blocks as shown in figure 2-3 and described in the following paragraphs.
2-4.1 STABILIZATION LOOP. The three stabilization loops (figure 2-4) maintain the stable member in a specific spatial orientation so that three mutually perpendicular 16 pulsed integrating pendulum (16 PIP) accelerometers can measure the proper compo¬ nents of LEM acceleration with respect to the coordinate system established by the stable member orientation. An input to the stabilization loops is created by any change in LEM attitude with respect to the spatial’ orientation of the stable member. With near zero gimbal angles, the inertia of the stable member tends to maintain the stable mem¬ ber in a fixed spatial orientation. Because of gimbal friction and unbalances, motion of the LEM structure relative to the stable member will produce a torque on the stable member which will tend to change its orientation. This change is sensed by the stabil¬ ization gyros. When the gyros sense an input, they issue error signals which are amplified, resolved, if necessary, into appropriate components, and applied to the gimbal torque motors. The gimbal torque motors then drive the gimbals until the stable member regains its original spatial orientation.
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1 5191
Figure 2-4. Stabilization Loop, Block Diagram
The stabilization loop consists of three pre-aligned Apollo II inertial reference integrating gyro (Apollo n IRIG) assemblies, a gyro error resolver, three gimbal servo amplifiers, three gimbal torque motors, three gimbals, and circuitry associated with these components. The inner gimbal is the stable member upon which the three stabili¬ zation gyros are mounted. The gyros are mounted with their input axes oriented in an orthogonal configuration. Movement of any gimbal tends to result in a movement of the stable member and rotation about the input axes of one or more of the stabilization gyros.
The stabilization loop contains three parallel channels. Each channel starts with a stabilization gyro (X, Y, or Z) and terminates in a gimbal torque motor. The torque motor drives the gimbals resulting in a movement of the stable member and a movement of the stabilization gyros. When a movement of the IMU support gimbal attempts to displace the stable member from its erected position, one or more of the stabilization gyros senses the movement and issues error signals. The phase and magnitude of the 3, 200 cps gyro error signal represents the direction and amount of rotation exper¬ ienced by the gyro about its input axis. The error signal is fed from the gyro signal
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generator ducosyn to the associated IRIG preamplifier, which is a part of the prealigned Apollo II IRIG assembly. Amplification of the error signal is required to achieve a high signal- to-noise ratio through the gimbal slip rings.
The amplified gyro error signals also represent motion of the stable member about its axis since the stable member axes (Xsm» Ysm» zSM)andthe gyro axes (X g, Yg, Zg) are parallel to one another. * If the middle and outer gimbal axes remain parallel with the stable member axes, then movement of the outer gimbal (a yaw movement of the LEM) is sensed by only the X gyro and movement of the middle gimbal (roll move¬ ment of the LEM) is sensed by only the Z gyro. Movement of the stable member about the inner gimbal axis (Ysm)> however, changes the relationship of the X and Z gyro input axes to the outer and middle gimbal axes. As a result, a movement of the middle or the outer gimbal is sensed by both X and Z gyros. The input required by the gimbal servo amplifiers to drive the gimbals and move the stable member back to its original position must be composed of components of both the X and Z gyros. The re¬ quired gimbal error signals are developed by the gyro error resolver. The gyro error signals, E(Xg) andE(Zg), are applied to the stator windings of the gyro error resolver. The rotor windings are connected to the inputs of the outer and middle gimbal servo amplifiers. Movement of the stable member about the inner gimbal axis (pitch move¬ ment of the LEM) changes the position of the resolver rotor relative to the resolver stator. This change corresponds electromagnetically to the change in the relationship of the stable member axes to the outer and middle gimbal axes. The outputs taken from the rotor are the required middle and outer gimbal error signals (Emg and E0g). Since the inner gimbal torque motor axis and the Y axis of the stable member are the same axis, the Y gyro error signal, E(Yg), is equal to the inner gimbal error signal, (Ejg), and is fed directly to the inner gimbal servo amplifier.
The three identical gimbal servo amplifier modules are located in the PSA and contain a phase sensitive demodulator, a filter, and a dc operational power amplifier. The phase sensitive demodulator converts either the 3, 200 cps gimbal error or 800 cps coarse align error, zero or pi phase, signals into a representative positive or negative dc signal. Thedc signal is filtered and applied to a dc operational amplifier with current feedback. The compensation network in the feedback circuit of the amplifier controls the response characteristics of the entire stabilization loop. The output of the dc amplifier has an operating range between +28 vdc and -28 vdc and drives the respec¬ tive gimbal torque motor directly in either angular direction.
The gain required for each stabilization loop differs. This difference compensates for the differences in gimbal inertia. The proper gain is selected by the connections to the gimbal servo amplifier module. A single torque motor is mounted on each gimbal at the positive end of the gimbal axis. The torque motors drive the gimbals to complete the stabilization loop.
* The Z gyro has its positive input axis aligned to the -ZgM axis but this is compensated for by reversing the polarity of the 3,200 cps excitation to the primary winding of the Z gyro signal generator ducosyn which causes the Z gyro error signal to be representative of the direction and amount of motion about the Zsm axis.
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The orientation of the stable member can be changed in either the coarse align, fine align, or IMU cage modes. Signals to reposition the gimbals are injected into the gimbal servo amplifiers from the CDU during the coarse align and IMU cage modes and into the stabilization gyros from the fine align electronics during the fine align mode. During the IMU cage mode and the coarse align mode, the reference signal for the demodulator in the gimbal servo amplifier is externally switched from 3, 200 cps to 800 cps.
2-4,2 FINE ALIGN ELECTRONICS. The fine align electronics (figure 2-5) provides torquing current to the stabilization gyros to change the orientation of the IMU gimbals during the fine align mode. The operation of the fine align electronics is controlled by the LGC.
The components of the fine align electronics are common to the three stabilization gyros. The fine align electronics provides torquing signals to the stabilization gyros one at a time on a time shared basis. The fine align electronics consists of a gyro calibration module, a binary current switch module, and a dc differential amplifier and precision voltage reference module, all located in the PTA.
151900
Figure 2-5, Fine Align Electronics - Computer Inputs
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The fine align electronics is enabled and controlled by LGC inputs to the gyro cal¬ ibration module. The LGC inputs consist of torque enable pulses, gyro select pulses, a torque set command, and a torque reset command. The fine align electronics is en¬ abled by the torque enable pulses. The torque enable pulses are a train of pulses three microseconds in width and occurring at 102.4 kpps. The torque enable pulses are applied through a relay driver to energize the torque enable relay in the cali¬ bration module. When the torque enable relay is energized, system 28 vdc is applied to the precision voltage reference (PVR) and regulated 120 vdc from the pulse torque power supply is applied to the dc differential amplifier and the binary current switch. The torque enable pulse train is received 20 milliseconds prior to any gyro set command.
The gyro to be torqued and the direction it is to be torqued is selected by the LGC by sending gyro select pulses to one of the six +A0or -A 6 inputs. (See figure 2-6. ) The gyro select pulse consists of a train of pulses three microseconds in width and occur- ringat 102.4 kpps. The pulse train activates a transistor switch network which controls current through the T+ or T- coils of the torque generator ducosyn in the gyro selected. The gyro select pulse train is received 312.5 microseconds (one LGC clock time at 3, 200 pps) prior to any torque set command.
@ ♦ A0Z
I5I89C
Figure 2-6. Fine Align Electronics - Gyro Selection
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The torque set and reset commands are 3, 200 pps pulse trains containing pulses that are three microseconds in width. A 3, 200 pps pulse train will be present on the torque set line when any gyro is to be torqued. A 3,200 pps pulse train is present on the torque reset line at all other times. This ensures that the binary current switch is in the reset condition prior to receipt of a torque enable command from the LGC. When the gyro has been torqued the proper amount, a torque reset command is issued which causes the torque current to be cut off. The gyro select pulse train will be removed 312.5 microseconds after the torque reset command has been issued. The torque set and torque reset pulses are fed through a 1:2 step-up transformer in the calibration module to the set and reset inputs of the binary current switch.
The torque current from the binary current switch is applied through a tuned re¬ sistive-capacitive compensation network in the calibration module to make the torque generator ducosyn windings appear as a pure resistive load to the binary current switch. The torque current to the gyros is via the T±(common) line. Current will flow only through the selected torque generator coils, the current monitor resistor, and the scale factor resistor. The voltage drop developed across the scale factor resistor is used as a feedback to the differential amplifier to regulate the torquing current. The voltage drop across the current monitor resistor is applied to PTA test points for ex¬ ternal monitoring of gyro torque current.
When no gyro is being torqued, the binary current switch provides current flow through a dummy load resistor and through the current monitor and scale factor resis¬ tors. In this manner, the binary current switch maintains a continuous ilow of torque current. The dummy load resistor simulates the impedance of the torque generator coil and a compensation network.
The torque set and torque reset pulses trigger a flip-flop (bi-stable multivibrator) in the binary current switch (figure 2-7). If the flip-flop is in the +set condition, the +set condition will remain until a reset command resets the flip-flop. The out¬ puts of the flip-flop control two transistor switches. If the flip-flop is in the +set condition, the +set output is present at the base of the +torque current switch, causing the switch to turn on. The +torque current switch closes the path from the 120 volt supply through the current regulator to the proper T+ or T- winding of the selected gyro via the calibration module. If the flip-flop is in the -set condition, the -torque current switch will turn on and close the current path through the dummy load resistor.
The binary current switch used in the fine align electronics is identical to the one used in the accelerometer loops. The portion of the binary current switch used only for the accelerometer loops is disabled in the fine align electronics application. In the accelerometer loop application, current to the accelerometer T+ torque genera¬ tor coil is provided by the +torque current switch and current to the T- torque genera¬ tor coil is provided by the -torque current switch. Therefore, the +torque and -torque designations of the switches have significance. In the fine align electronics application the switch designations have no significance since current to both the T+ and T- coils of the gyro torque generators is provided by the +torque current switch while the -torque current switch provides only the dummy load current.
2-11
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ND-1021042
MANUAL
(OUTPUT FOP PI PA LOOPS ONLY) POSITIVE VELOCITY PULSE (P)
ISI80B
Figure 2-7. Binary Current Switch
The dc differential amplifier and PVR module (figure 2-8) maintains the cur¬ rent through the windings of the torque generator ducosyn at 84 milliamperes. The PVR is supplied with regulated 28 vdc and, through the use of zener diode circuits, develops an accurate 6 vdc for use as a reference voltage. The scale factor resistor in the calibration module also has 6 volts developed across it when 84 milliamperes of cur¬ rent flows through it. A comparison is made by the dc differential amplifier of the PVR 6 volts and the scale factor resistor 6 volts. Any deviation from the nominal 84 milli¬ amperes of torquing current will increase or decrease the voltage developed across the scale factor resistor and cause an output error signal from the dc differential ampli¬ fier. This error signal controls the current regulator in the binary current switch. The current regulator, which is in series with the torque generator coils of the selected gyro and the 120 vdc source, will maintain the torquing current at 84 milliamperes.
2-12
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MANUAL
I5I7SA
Figure 2-8. DC Differential Amplifier and Precision Voltage Reference
The current flow through the windings of the torque generator ducosyn causes the gyro float to rotate about the gyro's output axis. A + A 6 gyro select command from the LGC will allow torque current to flow through a T- torque generator coil which results in a positive rotation of the gyro float about the output axis. A - A 6 gyro select command produces a negative float rotation. * Float rotation results in an error output from the signal generator ducosyn. The error signal is applied to the stabilization loop to re¬ position the gimbals and the stable member. The change in gimbal angles is transmitted by the CDU read counters to the LGC.
* The positive input axis of the Z gyro is aligned to the -ZgM axis but this is com¬ pensated for by reversing the T+ and T- connections to the Z gyro torque generator ducosyn. A + A0Z gyro select command from LGC will cause a negative float rotation but since the polarity of the Z gyro signal generator is also reversed the gyro error signal will appear to represent a positive float rotation. The stabilization loops will then drive the gimbals in the desired direction.
2-13
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MANUAL
2-4.3 ACCELEROMETER LOOP. The three accelerometer loops measure the acceler¬ ation of the stable member along three mutually perpendicular axes and integrate this data to determine velocity. The velocity is used by the LGC to determine the LEM velocity vector. Figure 2-9 is a functional diagram of an accelerometer loop.
The three accelerometer loops contain three prealigned 16 PIP assemblies, three PIP preamplifiers, three ac differential amplifier and interrogator modules, three binary current switches, three calibration modules, three dc differential amplifier and precision voltage reference modules, a pulse torque isolation transformer, and asso¬ ciated electronics.
The three mutually perpendicular PIP’s are acceleration sensitive devices. When fixed in its associated accelerometer loop, the PIP becomes an integrating accelerom¬ eter. The PIP is basically a pendulum-type device consisting of a cylinder with a pen¬ dulous mass unbalance (pendulous float) pivoted with respect to a case. The axis of the pivots defines the PIP output axis. A signal generator is located at the positive end of the output axis to provide electrical output signals indicative of the rotational position of the float. A torque generator located at the other end of the float acts as a trans¬ ducer to convert electrical signals into mechanical torque about the float shaft. The accelerometer loop using a PIP is mechanized to operate in a binary (two state) mode.
In the binary mode, the PIP pendulum is continually kept in an oscillatory motion. Thus the two states: positive rotation or negative rotation. The rotation is accom¬ plished by continuously routing torquing current through the torque generator plus or minus windings.
I3I77A
Figure 2-9. Accelerometer Loop
2-14
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
The torque generator has two windings, one to produce torque (rotation) in a positive direction, the other to produce torque (rotation) in a negative direction. Only one winding will have current in it at anyone time. The torque winding selection is accom¬ plished by the setting of a flip-flop in the binary current switch (figure 2-7). When the loop is first energized, the interrogator sets the flip-flop to route the torquing current to one of the windings, which will rotate the float to null. As the float passes through null, the phase of the output signal of the signal generator changes, which causes the interrogator to issue pulses to reset the flip-flop in the binary current switch and thus route torquing current to the other torque winding. The float is then torqued in the opposite direction until the signal generator output again changes phase as the float passes through null which reinstates the cycle.
The output of the signal generator, after being amplified by the PIP preamplifier is interrogated 3200 times a second by the interrogate pulse. The binary current switch flip-flop can be reset only when the interrogate pulse is present and the signal generator output is of the proper phase.
The PIP pendulum motion is an oscillatory motion about its null point and can be measured in cycles per second. As is characteristic of every electro-mechanical loop, there exists some natural resonant frequency. The natural frequency is dependent upon float damping, signal and torque generator sensitivities, and other loop characteristics. In the case of the accelerometer loop this natural frequency is approximately 500 cps, and the pendulum oscillates at a frequency close to that. At a torque winding selection rate of 3200 pulses per second, the value of this frequency can be any value equal to 3200 -J- x where x is any even number.
Using the above ratio, it is possible for the pendulum to have a maximum frequency of 1600 cps (x equals 2). A frequency of 1600 cps means that for every torque selec¬ tion pulse, the torque current would be routed to the opposite torque generator winding. Solving the equation f = 3200 x, the frequency closest to 500 is 533-1/3. In this case; the value of x is six. Thus one complete pendulum cycle will occur during six torque selection pulses. Dividing the time for the six pulses into positive and negative rotations, it is seen that the PIP functions in a 3-3 mode (positive rotation for three torque selection pulses, negative rotation for three torque selection pulses).
The physical configuration of the PIP is such that the float, when moding in its 3-3 state and sensing no acceleration, rotates an equal angular distance on both sides of an electrical and mechanical null.
The2voltrms, 3200 cps, one phase signal generator excitation voltage is synchro¬ nized with the LGC clock. The signal generator has a center tapped secondary winding which provides a double ended output, one side having a zero phase reference with re¬ spect to the 3,200 cps excitation and the other side a pi phase reference. The center tap is connected to ground. The output signal is representative of the magnitude and direction of the rotation of the pendulous float about the output axis. The error signal is then routed to the preamplifier mounted on the stable member. The phase of the output signal from the preamplifier is -45° from the reference excitation. The phase shifted zero or pi phase signals from the preamplifier are applied as separate inputs to the ac differential amplifier and further amplified. The two signals are then sent to the interrogator.
2-15
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTIOl SYSTEM
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MANUAL
The ac differential amplifier and the interrogator are packaged in the same module which is located in the pulse torque assembly (PTA) (figure 2-10). The interrogator analyzes the ac differential amplifier outputs to determine the direction of the 16 PIP float movement and generates appropriate torquing commands. The two ampli¬ fied signals from the ac differential amplifier goto two summing networks and threshold amplifiers (represented in figure 2-10 by AND gates). Interrogate pulses (IP) are continuously being received by the interrogator from the LGC. An interrogate pulse is a two microsecond pulse occurring at 3,200 pps and timed to occur 135 degrees after the positive going zero crossing of the reference excitation. (See figure 2-11.) With this phasing, the interrogate pulse occurs at the 90 degree peaks of the phase shifted zero or pi phase input signals from the PIP preamplifiers. The interrogate pulse occurs at a positive 90 degree peak of the zero phase signal if the float angle is posi¬ tive and at a positive 90 degree peak of the pi phase signal if the float angle is negative. The zero and pi phase signals and the interrogate pulses are ANDed by the summing network and threshold amplifier. The gated outputs of the threshold amplifier are applied to a flip-flop as set or reset pulses. If the flip-flop is in the +set condition, a succession of set pulses will maintain the +set condition. The +set condition will remain until the float angle passes through null. At this time, a reset pulse is pro¬ duced to cause the flip-flop to go to the -set condition.
SWITCH I 1
Figure 2-10. AC Differential Amplifier and Interrogator Module
2-16
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
The outputs of the flip-flop are applied to two AND gates which are also driven by switch pulses received from the LGC. The switch pulses are a train of clock driven 3,200 pps pulses three microseconds in width, timed to occur three micro¬ seconds after the leading edge of the interrogate pulse. The flip-flop enables only one output gate at any switch pulse time. The outputs of the AND gates are called the TM + set pulse and the TM - set pulse.
The binary current switch (figure 2-7) utilizes the TM + set and TM - set outputs of the interrogator to generate 16 PIP torquing current. The TM + set and the TM - set pulses furnish the input to a flip-flop. If the flip-flop is in the +set condition, a succession of TM+set pulses will maintain the +set condition. The +set condition will persist until the float angle passes through null. The phase change will cause the flip-flop of the ac differential amplifier and interrogator module to reset to the -set condition. At this time a TM - set pulse is developed and causes the binary current switch flip-flop to go to the -set condition. The outputs of the flip-flop control two transistor current switches. If the flip-flop is in the +set condition, the +set output will be at the base of the +torque current switch and will turn it on. The +torque current switch closes the path from the current regulated 120 vdc supply through the PIPA calibration module to the 16 PIP T+ torque generator coils. If the flip-flop is in the -set condition, the -torque current switch will be turned on, closing the path through the T- torque generator coils.
An acceleration along the PIP input axis causes the pendulous mass to produce a torque which tends to rotate the float about the output axis. The torque produced by the acceleration is proportional to the magnitude of the acceleration. The acceleration produced torque aids and opposes the torque generator forces causing changes in the time required for the float to be torqued back through null. A change in velocity (AV) is the product of acceleration and incremental time (At), the torque is actually proportional to an incremental change in velocity ( AV).
tACCEL “ KlaAt - Kq AV
The float is already in motion due to loop torquing, therefore additional torque is required to overcome the acceleration torque and to keep the pendulum in its oscilla¬ tory motion. The additional torque is obtained by supplying torquing current for addi¬ tional time through one of the torque windings. The current at any one time is a con¬ stant, therefore the current must be present for a longer period of time. Thus to determine the amount of acceleration sensed by the PIP, it is necessary only to measure the length of time torquing current is applied to each torque winding.
ACCELind = K2 E [(T+) - (T-)] At
2-17
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTIOl SYSTEM
ND-1021042
MANUAL
From the above identities, it is seen that torquing time (At) is proportional to the change in velocity (AV).
Ki AV = K 2 £ [(T+) - (T-)] At
K2
AV - - £
The time ( At), representative of the AV, is sent to the LGC in the form of P and N pulses (figure 2-9).
In addition to selecting the proper torque generator winding, the outputs of the binary current switch flip-flop also go to two AND gates where they are ANDed with the 3,200 cps data pulses from the LGC. The data pulse is three microseconds in width and is timed to occur two microseconds after the leading edge of the switch pulse. (See figure 2-11.) The data pulse and switch pulse are both 3,200 cps, therefore the LGC receives either a P pulse or an N pulse once every 1 -f- 3200 second. When the PIP is sensing no acceleration, the pendulum is oscillating at a frequency of 533-1/3 cps; and the LGC is receiving three P pulses and three N pulses once every cycle or once every 1 -h 533-1/3 seconds. The LGC contains a forward-backward counter which receives the velocity pulses and detects any actual gain in velocity.
The counter counts forward on the three P pulses and then backward on the three N pulses. The counter continues this operation and generates no AV pulses. With an acceleration input to the PIP, however, the loop no longer operates at the 3-3 ratio and the counter exceeds its capacity and reads out the plus or minus AV pulses which are then stored and used by the LGC. The additional pulses above the 3-3 ratio are representative of the additional torque supplied by the torque generator to compen¬ sate for the acceleration felt by the LEM. Each pulse indicates a known value of Av due to the loop scale factor.
The PIPA calibration module (figure 2-12) compensates for the inductive load of the 16 PIP torque generator ducosyns and regulates the balance of the plus and minus torques. The calibration module consists of two load compensation networks for the torque generator coils of the 16 PIP. The load compensation networks tune the torque generator coils to make them appear as a pure resistive load to the binary current switch. A variable balance potentiometer regulates the amount of torque developed by the torque generator coils. Adjustment of this potentiometer precisely regulates and balances the amount of torque developed by the T+ and T- torque generator coils. This balancing insures that for a given torquing current an equal amount of torque will be developed in either direction.
2-18
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
2-19
Figure 2-11. Accelerometer Timing
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
Figure 2-12. PIP A Calibration Module
The calibration module also includes a current monitor resistor and an adjustable scale factor resistor network in series with the torque generator coils. A nominal six volts is developed across the scale factor resistor network due to the torquing current and is applied as an input to the dc differential amplifier and precision voltage reference module. The voltage drop across the current monitor resistor is used for external monitoring purposes.
2-20
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
The dc differential amplifier and PVR are identical to the ones used in the fine align electronics. (See figure 2-8.) The dc differential amplifier and PVR module maintain the current through the ducosyn torque generator coils at 43 milliamperes. The PVR is supplied with regulated 28 vdc and, through the use of precision circuits, develops an accurate 6 volts for use as a reference voltage. The scale factor resistor in the calibration module also develops 6 volts when 43 milliamperes of current flows through it. A comparison is made by the dc differential amplifier of the PVR 6 volts and the scale factor 6 volts. Any deviation of the binary current switch torquing cur¬ rent from the nominal 43 milliamperes will increase or decrease the scale factor re¬ sistor voltage and result in an output error signal from the dc differential amplifier. This error signal controls the current regulator in the binary current switch. The current regulator, which is in series with the 120 vdc source and the ducosyn torque generator coils, will maintain the torque current at 43 milliamperes.
2-4.4 IMU TEMPERATURE CONTROL SYSTEM
The IMU temperature control system (figure 2-13) maintains the temperature of the stabilization gyros and accelerometers within the required temperature limits during both standby and operating modes of the IMU. The system supplies and removes heat to maintain the IMU heat balance with minimum power consumption. Heat is re¬ moved by convection, conduction, and radiation. The natural convection used during IMU standby mode changes to blower controlled, forced convection during IMU opera¬ ting modes. The IMU internal pressure is maintained between 3.5 and 15 psia to enable the required forced convection. To aid in removing heat, a water-glycol solution at approximately 45.0 degrees Fahrenheit from the spacecraft coolant system passes through the coolant passages in the IMU support gimbal.
2-4. 4.1 Temperature Control Circuit. The temperature control circuit maintains the gyro and accelerometer temperature. The temperature control circuit consists of a temperature control thermostat and heater assembly, a temperature control module, three IRIG end mount heaters, three IRIG tapered mount heaters, two stable member heaters, and three accelerometer heaters. The thermostat and heater assembly is located on the stable member and contains a mercury-thallium thermostat, a bias heater, and an anticipatory heater. Except for the bias heater, all heaters (a total of 12) are connected in parallel and are energized by 28 vdc through the switching action of transistor Q2, which completes the dc return path. The thermostat acts as a control sensing element and senses the temperature of the stable member.
2-21
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
When the temperature falls below 130 (±0.2) degrees Fahrenheit, the thermostat opens and transistor Q1 conducts and drives transistor Q2 to conduction. When tran¬ sistor Q2 conducts, current will flow through the twelve heaters. Because of the large mass of the stable member, its temperature will increase at a relatively slow rate as compared to the gyros, which have a heater in each end mount. The anticipatory heater improves the response of the thermostat to insure that the magnitude of the temperature cycling of the gyros and the accelerometers is as small as possible. When the temper¬ ature rises above 130 (±0.2) degrees Fahrenheit, the thermostat closes and the base of transistor Q1 is shorted to ground, cutting off transistors Q1 andQ2 and deenergizing the heaters. The thermostat has a 0.5 degree deadband which is the difference between contact closing with rising temperature and contact opening with falling temperature. The temperature control circuit will maintain the average of the gyro temperatures at 135 (±1. 0) degrees Fahrenheit and the average of the accelerometer temperatures at 130 (±1. 0) degrees Fahrenheit under normal ambient conditions. The temperature difference between the gyros and the accelerometers is adjusted by properly propor¬ tioning the amount of power in each heater. The balance is obtained by selection of resistor Rl.
During IMU operation, power is applied to the fixed accelerometer heaters to compensate for the additional heat supplied to the gyros by the gyro wheel motor heat dissipation. Power is also applied to a bias heater on the control thermostat. The bias heater supplies a fixed amount of heat to the control thermostat to maintain the proper absolute temperature level of the gyros and accelerometers. The amount of bias heat is controlled by the selection of resistor R5. The power for the fixed accelerometer heaters and the thermostat bias heater are the -90 degree and -180 degree outputs, respectively, from the 28 vac power supplies which are also used for gyro wheel power.
The 28 vdc heater power is applied to the heaters through the contacts of a safety thermostat which will provide protection against an extreme overheat condition in case a malfunction occurs in the temperature control circuit. The safety thermostat contacts open at 139.5 (±3.0) degrees Fahrenheit and close at 137 (±3) degrees Fahrenheit.
2-4. 4. 2 Blower Control Circuit. The blowers maintain IMU heat balance by removing heat. The blowers operate continuously during IMU operate modes. The blower control circuit shown on figure 2-13 is inoperative because the contacts of blower control relay K1 are bypassed.
The blowers are supplied from the -90 and -180 degree outputs of the 28 volt, 800 cps, 2.5 percent power supply which also provides gyro wheel motor power. Fused phase shift networks are associated with each blower so that excitation and control current can be supplied from the same source.
2-22
1 TEMP ALARM I THERMOSTAT ASSY
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ND-1021042 |
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LEM PRIMARY GUIDANCE, NAVIGATION, |
AND CONTROL SYSTEM |
MANUAL |
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i 1 1 1 MIDDLE . OUTER i GIM8AL | GIMBAL |
CASE |
LGC |
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|
i ! 1 |
I 1 GSNACE |
PSA | |
IMU AUXILIARY MODULE
Figure 2-13. IMU Temperature Control System
2-23/2-24
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
2-4. 4. 3 Temperature Alarm Circuit. The temperature alarm circuit monitors the tem¬ perature control system. The temperature alarm circuit consists of a temperature alarm thermostat and a temperature control module. If a high or low temperature is sensed by the temperature alarm thermostat located on the stable member a discrete is sent to the LGC and the IMU auxiliary module. When the temperature is within the normal range of 126.3 to 134.3 degrees Fahrenheit, 28 vdc is applied through the ther¬ mostat to the emitter of transistor Q1 causing the transistor to conduct. Transistor Q1 conducts through a grounding system in the LGC.
When the temperature falls below 126.3 degrees Fahrenheit, 28 vdc will be re¬ moved from transistor Ql, causing it to stop conducting and thus signaling the LGC of an alarm condition. When the temperature rises above 134.3 degrees Fahrenheit, 28 vdc will be applied directly to the base of the transistor as well as to the emitter. With 28 volts applied to both emitter and base, the base-emitter junction is no longer forward biased and the transistor stops conducting which signals the LGC of an alarm condition. There is no differentiation between a high or low temperature alarm. When the LGC senses a temperature alarm, it causes the IMU TEMP lamp and the PGNCS lamp to light. When the IMU auxiliary module receives a temperature alarm, it sends the information to telemetry.
2-4. 4. 4 External Temperature Control. External temperature control of the IMU is provided by GSE control heater circuits in the IMU which are controlled externally to the airborne equipment by the portable temperature controller or the temperature monitor control panel of the optics-inertial analyzer (OIA). The GSE control heater circuitry consists of a safety thermostat, six gyro heaters, two stable member heaters, three accelerometer heaters, temperature indicating sensors, and an IMU standby power sensor which disables the GSE when airborne power is on. The temperature indicating sensors act as the control sensing element of the external control and indi¬ cating circuitry. The heaters are connected in parallel. The six gyro temperature indicating sensors (two in each gyro) are connected in series to sense the average temperature of the gyros. The three accelerometer temperature indicating sensors (one in each accelerometer) are connected in series to sense the average temperature of the accelerometers. All of the GSE control heater circuitry is electrically indepen¬ dent of the airborne temperature control system and will not be used at the same time that the IMU temperature is being controlled by the airborne temperature control sys¬ tem. The GSE control heater circuitry cannot be used as a backup temperature control system during flight.
2-4.5 ISS MODES OF OPERATION. The ISS has four major modes of operation: IMU turn on, CDU zero, coarse align, and inertial reference. Submodes which will also be discussed are fine align, IMU cage, attitude error indication and display inertial data. An additional mode is the master reset condition which is available during laboratory testing only. All ISS moding is initiated and controlled by computer discretes to the CDU. (See figure 2-14.) To select an ISS mode of operation, the LGC can send a single discrete, a combination of discretes, or no discretes. The display inertial data func¬ tion utilizes the LORS channels of the CDU, therefore, a description of the discretes to the LORS channels of the CDU will also be presented.
2-25
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
2-4. 5.1 CPU Discretes. All LGC discretes issued to the CDU to initiate and control the various ISS modes or functions are 0. 0 (±2) vdc. LGC ground, applied through a 2, 000 ohm source impedance to the CDU mode module.
2-4. 5. 1.1 ISS CDU Zero. The ISS CDUzero discrete zeros or clears all three ISS CDU read counters simultaneously. It also inhibits the transmission of incrementing pulses to the read counters for the period of time the discrete is present. The CDU discrete will be present (minimum duration is approximately 400 milliseconds) for as long as the read counters are to be held at zero. The IMU is not disturbed by the CDU zero discrete.
2-4. 5. 1.2 ISS Enable Error Counter. The ISS enable error counter discrete enables all three ISS error counters simultaneously which allows them to accept incrementing pulses from the LGC. The error counters are normally cleared and inhibited. The ISS enable error counter discrete is used in conjunction with the coarse align enable discrete during the coarse align mode. The ISS enable error counter enable discrete is used alone when display of attitude error signals on the FDAI is required only.
2-4. 5. 1.3 Coarse Align Enable. The coarse align enable discrete enables a relay driver which energizes the coarse align and demodulator reference relays located in the PSA . This connects the coarse align error signal to the gimbal servo ampli¬ fiers and changes the reference voltage for the demodulator in the gimbal servo ampli¬ fiers from 3,200 cps to 800 cps. The discrete also enables the digital feedback pulses from the read counter to the error counter. The presence of the coarse align enable discrete and the absence of the enable error counter discrete also inhibit the increment¬ ing pulses to the read counter.
2-4. 5.1. 4 D /A Enable. The D/A enable discrete enables both LORSCDU error counters simultaneously. The error counters are normally cleared and inhibited. The LGC nor¬ mally provides positioning signals to the optical tracker through the LORS channels of the CDU. The D/A enable discrete, however, is also used in conjunction with the display inertial data discrete to allow the LGC to feed inertially derived velocity data through the LORS channels of the CDU to meter displays.
2-4. 5. 1.5 Display Inertial Data. The display inertial data discrete energizes relays which switch the dc output from the digital to analog (D/A) converter in the LORS channels of the CDU to the LEM velocity meters.
2-4. 5. 1.6 LORS CDU Zero. The LORS CDU zero discrete clears both LORS read counters simultaneously and inhibits the transmission of incrementing pulses to the read counters. This discrete is not used for any ISS flight modes or functions but can be used for CDU test functions.
2-26
PSA
•3,200 CPS
IMU
0C SIGNAL (NOT USED FOR LEM)
AC SIGNAL TO FOAI
I
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LEM PRIMARY GUIDANCE, NAVIGATION, |
AND CONTROL SYSTEM |
ND-1021042 MANUAL |
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rLGC |
CDU
D/A CONVERTER (800 CPS WOOER NETWORK ANO (JEMODULATOR )
t tl t t t t
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ERROR C< |
UNTER |
|||||||
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IL25* |
5.6* |
2 0* |
1.4* |
1 |
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.17* |
0.8* |
044* |
|
8 2 |
7 2 |
6 2 |
5 2 |
i |
L3 |
2 2 |
2 |
0 2 |
+ A0C
+A0G
COARSE AND FINE ERROR SIGNALS
|
11 A/D SVf |
CH |
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SELECTION |
[OGIC |
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(COARsJ |
Rnd |
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FINE SY!I |
pMS) |
ERROR
DETECTOR
t t t t t t t 1 1 1 t t t t t r
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READ COUN-j |
[r (*) |
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180* |
90* |
45* |
225* |
11.25* |
5.6* |
2.8* |
• 1.4 |
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.17* |
08* |
044* |
022* |
40* |
ro O |
|
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15 2 |
14 2 |
13 2 |
12 2 |
II 2 |
10 2 |
9 2 |
8 2 |
2'j |
.7 |
6 2 |
5 2 |
4 2 |
3 2 |
2 2 |
2 |
0 2 |
A0G
•A0G
- AflG
EACH PULSE=022*X 2=0445(60 SEC'
+ A8G
5
MODE
LOGIC
28 VDC ISS OPERATE - WV-
MODE
LOGIC
RATE SELECT AND UP-DOWN LOGIC
|
6 |
4 KPPS |
|
|
8 |
00 CPS |
+ PI
EACH PULSE=20 SEC X 2=40 SEC
PHASE PULSES FOR TIMING
MOOE
LOGIC
1
COARSE ALIGN ENABLE (CA)
ISS ENABLE ERROR COUNTER (EEC)
+ Aflc] ^
I 160 SEC/PULSE > 3200 PPS MAX RATE
— A0CJ
CLOCK 51.2 KPPS
( 6,400 PPS MAX RATE )
CDU-ZERO (Z)
J L _
Figure 2-14. ISS-CDU Moding
2-27/2-28
t
4
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
2-4. 5. 2 IMU Turn On Mode. The purpose of the IMU turn on mode is to drive the gimbals to their zero position and hold them there. (See figure 2-15.) The IMU turn on mode is initiated upon closure of the ISS OPERATE circuit breaker and allows for a 90 second gyro run up period. The ISS OPERATE circuit breaker routes 28 vdc IMU operate power through the deenergized contacts of the ISS turn on control relay, located in the IMU auxiliary assembly module, to the cage relays. The 28 vdc IMU operate power is also routed through the same deenergized contacts to the LGC as a continuous turn on delay request discrete. The cage relays energize and, in turn, cause the coarse align relays to be energized. The cage relays route the IX gimbal resolver sine winding outputs through the contacts of the coarse align relays to the respective gimbal servo amplifiers. The gimbal servo amplifiers drive the gimbals until the resolver signals are nulled. The operation of the caging loops is discussed further in the IMU cage mode description.
Upon receipt of the ISS turn on delay request discrete, the LGC sends the ISS CDU zero discrete and the coarse align enable discrete to the CDU for a minimum period of 90 seconds. The CDU zero discrete clears the read counters and inhibits the incrementing pulses to the read counters. The coarse align enable discrete provides a redundant means of energizing the coarse align relays.
A second set of deenergized contacts on the ISS turn on control relay routes a ground to the time delay circuit of the pulse torque power supply which inhibits the operation of the power supply and thus prevents accelerometer pulse torquing during the 90 second turn on period. This allows time for the accelerometer floats to become centered and the gyro wheels to run up prior to torquing.
After the 90 second delay has been completed, the LGC sends the ISS turn on delay complete discrete. The ISS turn on delay complete discrete acts through a relay driver to energize and latch in the ISS turn on control relay. Energizing the ISS turn on control relay deenergizes the cage relay, removes the ISS turn on delay request discrete, and removes the inhibit from the pulse torque power supply. The computer program can then place the ISS in the inertial reference mode by removing both the CDU zero and the coarse align enable discretes, or it can initiate the coarse align mode by re¬ moving only the CDU zero discrete and sending the ISS enable error counter discrete. The IMU turn on circuit will be reset whenever 28 vdc IMU operate power is turned off.
2-4. 5. 3 IMU Cage Mode. The IMU cage mode is an emergency backup mode which allows the astronaut to recover a tumbling IMU by setting the gimbals to zero. (See figure 2-15.) During this mode, the IX gimbal resolver sine winding outputs are fed through the CDU to the gimbal servo amplifiers to drive the gimbals until the resolver signals are nulled.
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MANUAL
The IMU cage mode is initiated when the astronaut presses the IMU CAGE switch. The switch is held until the gimbals settle at the zero position (five seconds maximum). The gimbal position may be observed on the FDAI. The IMU CAGE switch routes a 28 vdc discrete signal to the LGC and to the cage relays located in the PSA. (See figure 2-15.) The cage discrete energizes the cage relays, which in turn, cause the coarse align relays, the demodulator reference relay, and a relay in the gimbal servo amplifiers to energize. The relay in the gimbal servo amplifiers switches in addi¬ tional capacitance into the RC compensation networks to tune them for 800 cps operation. The demodulator reference relay changes the gimbal servo amplifier demodulator refer¬ ence signal from 3,200 cps to 800 cps. The cage relays switch the IX gimbal resolver sine winding outputs through the energized contacts of the coarse align relays into the corresponding gimbal servo amplifier inputs. The gimbal servo amplifiers drive the gimbals until the resolver signals are nulled.
Upon receipt of the IMU cage discrete, the LGC will discontinue sending the error counter enable discrete, the coarse align enable, the display inertial data discrete, and the incrementing pulses to the CDU. The LGC will also discontinue sending torquing commands, if any are in process, to the fine align electronics.
After the IMU CAGE switch is released, the LGC will allow the read counters to settle and will then place the PGNCS in an attitude control mode. During the time the IMU cage discrete is present and while the read counters are settling, the NO ATT lamp on the DSKY is lighted.
The cage mode will also be entered automatically if the IMU is turned on when the LGC is off or in standby mode. During the normal turn on sequence, the closure of the ISS OPERATE circuit breaker will route 28 vdc through the deenergized contacts of the ISS turn on control relay to the cage relays. The cage relays energize and cage the gimbals. After the 90 second turn on time delay has been completed, the LGC will send the ISS turn on delay complete discrete which will energize the ISS turn on control relay which, in turn, deenergizes the cage relays. If, however, the LGC is off or in standby when the IMU is turned on, the ISS turn on control relay will remain deenergized and the ISS will remain in the IMU cage mode.
If the IMU cage mode is entered as a result of an IMU turn on with the LGC off or in standby, the ISS can be placed in the inertial reference mode by allowing 90 seconds for gyro runup then pressing the IMU CAGE switch. The IMU CAGE switch will energize and latch in the ISS turn on control relay which removes the 28 vdc which had been energizing the cage relays. With the ISS turn on control relay latched, the cage relays will deenergize and remain deenergized when the IMU CAGE switch is released. Deenergizing the cage relays causes the coarse align relays to be deenergized which connects the gyro error signals to the respective gimbal servo amplifiers. The stabilization loops will maintain the stable member inertially referenced to the orientation established by the caging loops.
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MANUAL
1
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MANUAL
2-4. 5. 4 ISS CPU Zero. The purpose of the ISS CDU zero mode is to clear and inhibit the three ISS CDU read counters. (See figure 2-14.) The mode is initiated by the LGC sending the ISS CDU zero discrete. The presence of the discrete is maintained for as long as the read counters are to be held at zero.
2-4. 5.5 Coarse Align Mode. The purpose of the coarse align mode is to change the orientation of the gimbals by LGC command. The change in gimbal orientation is accom¬ plished by feeding the CDU error counter computer pulses equal to the required change in gimbal angles. The mode is initiated when the LGC sends the coarse align discrete and, after a short delay, the ISS error counter enable discrete to the three ISS portions of the CDU. When the mode is entered, the CDU read counter will be in the process of repeating the gimbal angle and supplying angular data to the LGC and will continue to do so for the duration of the coarse align mode. The LGC, knowing the actual gimbal angle registered in the read counter, calculates the desired amount of change in gimbal angle required to reposition the gimbal to the desired angle and converts this change into a number of ±A 9C pulses to be sent to the error counter. The ±A 6C pulses are sent to the error counter at a rate of 3, 200 pps in bursts of 80 milliseconds duration. Bursts of pulses are sent 600 milliseconds apart. EachA0c pulse is equal to a change in gimbal angle of 160 arc seconds. The error counter, having been enabled, accepts the pulses and counts up or down, as necessary, until all the pulses have been registered.
The digital information in the error counter is converted into an 800 cps, amplitude modulated, analog error signal by the ladder decoder in the D/A converter module. The ladder decoder signal is summed with a feedback signal and applied through a mixing amplifier located in the D/A converter module to the gimbal servo amplifiers to drive the gimbals to the desired angles. The function of the feedback signal and the mixing amplifier will be discussed later. The output of the mixing amplifier, referred to as the coarse align error signal, is applied to the gimbal servo amplifiers through the con¬ tacts of the coarse align relays located in the PSA. The coarse align relays, which are energized by the coarse align enable discrete acting through a relay driver, switch the input of gimbal servo amplifiers from the gyro preamplifiers to the coarse align error signal output of the ISS D/A converter. The demodulator reference relay is also energized by the coarse align enable discrete and switches the reference frequency of the demodulator in the gimbal servo amplifiers from 3,200 cps to 800 cps. The coarse align enable discrete also energizes a relay in the gimbal servo amplifiers which switches in additional capacitance into the amplifier’s compensation networks to tune them for 800 cps operation.
As the gimbals are driven, ±A0cpulses, representing the change in actual gimbal angle, are generated by the read counter and applied to the error counter. TheA06 pulses are also equal to 160 arc seconds and act to decrease theA0c pulses registered in the error counter. The error counter output to the ladder decoder, therefore, represents the difference between the desired amount of change in gimbal angle and the amount of change actually accomplished. When the error counter reaches a null and the gimbals stop moving, the actual gimbal angle has changed by an amount equal to the total value of the ±A0C pulses sent by the LGC to the error counter.
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MANUAL
The rate at which the gimbals are driven is limited to prevent damage to the gyros and to assure that the read counter can track the gimbal angle accurately. The rate of gimbal movement is limited by feeding back the C DU fine error signal [sin 16 (0-0)] to the input of the mixing amplifier located in the D/A converter module. The CDU fine error signal is out of phase with the output of the ladder decoder and has an amplitude proportional to the difference between the actual gimbal angle (0) and the angle in the read counter (0). The fine error signal is applied through a voltage limiting cir¬ cuit to the summing junction of the mixing amplifier where it is summed with the 800 cps ladder decoder output signal. The D/A converter ladder decoder output is applied to the mixing amplifier through a scaling amplifier and a voltage limiting diode net¬ work. The scaling amplifier controls the signal gain to produce a scale factor of 0.3 volt rms per degree. The output of the mixing amplifier will be at a null when the D/A converter ladder decoder output, after limiting, is equal to the fine error feedback signal. The fine error signal will be a constant value only when the gimbal and the CDU are going at the same rate and with the gimbal angle leading the CDU angle. Since the CDU is limited to counting at one of two speeds, the gimbals will be limited to a rate equal to one of these two speeds. During the coarse align mode, the CDU is limited to a high counting speed of 6.4 kpps and a low counting speed of 800 cps. At all othertimes, the high counting speed is 12.8 kpps.
If the gimbals are moving at a faster rate than the rate at which the CDU is counting, the fine error signal will increase, causing a retarding torque to be developed by the gimbal servo amplifier. If the gimbals are moving at a rate slower than the rate at which the CDU is counting, the fine error signal will decrease, causing the gimbal servo amplifier to apply an accelerating torque to the gimbals. By adjusting the gain of the fine error signal into the mixing amplifier, the gimbal drive rate is limited to either 35. 5 degrees per second (6.4 kpps CDU counting rate) or 4.5 degrees per second (800 cps CDU counting rate).
2-4. 5.6 Inertial Reference Mode. The inertial reference mode provides a coordinate reference system on which attitude and velocity measurements and calculations may be based. During the inertial reference mode, the stable member is held fixed with respect to an inertial reference by the stabilization loops. The ISS CDU read counters provide the LGC with changes in gimbal angles with respect to the stable member. The ISS is in the inertial reference mode during any operating period in which there is an absence of moding commands. During the inertial reference mode, the fine align electronics is inhibited and the ISS CDU error counters are cleared and inhibited.
2-4.5.7 Fine Align Mode. The purpose of the fine align mode is to reposition the stable member to a fine alignment by torquing the gyros. The fine align mode is actually a gyro torquing function accomplished during the inertial reference mode. The torquing current to the gyros is provided by the fine align electronics located in the PTA. The fine align electronics is enabled and controlled by LGC pulses sent directly to the fine align electronics. The LGC does not send command discretes to the ISS CDU’s during this mode. The ISS is in the inertial reference mode prior to the enabling of the fine align electronics and returns to that mode when the fine align electronics is disabled.
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MANUAL
The fine align electronics torques the gyros on a time shared basis. The LGC sends four types of pulse trains to the fine align electronics. The first pulse train sent is the torque enable command which enables the fine align electronics. The second pulse train is a gyro select command which selects a particular gyro and the direction it is to be torqued by means of a switching network which closes the current path through the proper torque ducosyn coil. The third and fourth types of pulse trains are the torque set and torque reset commands which control a binary current switch to start and stop the current flow through the selected torque ducosyn coil. The amount of current flow through the torque ducosyn coils is precisely controlled at a fixed value. The amount of gyro torquing to accomplished is determined by the amount of time.torque current is applied, that is, the time duration between the receipt of the torque set and the torque reset commands. This time duration is calculated by the LGC programs and may be based on optical alignment measurements. The torque ducosyn displaces the gyro float, causing the ducosyn signal generator to apply an error signal to the sta¬ bilization loop. The stabilization loops drive the gimbals to reposition the stable mem¬ ber. Upon completion of the torquing, the stable member remains fixed in its inertial reference and in fine align mode until the torque enable command is removed, after which the ISS remains in inertial reference mode until further change is commanded.
During the fine align mode, the ISS error counters remain cleared and inhibited. The read counters continue to repeat the gimbal angles and send angular data (±A0G) to the LGC.
2-4. 5. 8 Attitude Error Indication. The attitude error indication mode supplies attitude error signals to the FDAI. The attitude error indication mode is initiated when the LGC sends the ISS error counter enable discrete to the CDU. The LGC will calculate the difference between the actual gimbal angles and the correct angles and convert this into the number of ±A©C pulses to be sent to the error counter. The error counter, having been enabled, accepts the pulses and counts up or down until it has registered all the pulses.
As the actual gimbal angles increase or decrease, ±Aqg pulses, which represent the change in actual gimbal angle, are generated by the read counter and applied to the error counter. The ±A0G pulses cause the error counter to count up or down and, in effect, register the difference between the actual gimbal angle and the desired gimbal angle. The digital information in the error counter is converted into an 800 cps, ampli¬ tude modulated, analog error signal by the ladder decoder in the D/A converter. This ac signal is applied through a scaling amplifier to the FDAI.
2-4. 5. 9 Display Inertial Data. The display inertial data mode permits the LGC to pro¬ vide inertially derived forward and lateral velocity signals through the digital to analog section of the LORS channels of the CDU to the LEM velocity display meters. The display inertial data mode is used during the last phases of the LEM powered descent.
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LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
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MANUAL
The display inertial data mode is requested by the astronaut closing a switch on the main control panel. (See figure 2-16.) The mode is initiated by the LGC sending the display inertial data discrete to the LORS channels of the CDU. The display inertial data discrete acts through a relay driver to energize relays which connect the D/A converter dc error signal outputs to the LEM velocity display meters. After a brief delay to allow for relay pull in time, the LGC sends the D/A enable discrete followed by incrementing pulses to the error counters. The LGC sends ±Aec pulses representing LEM forward velocity (motion along the Zlem axis) to one error counter and ±A0c pulses representing LEM lateral velocity (motion along the vehicle Ylem) to other error counter. The read counters will not send incrementing pulses to the error counters; therefore, the only information registered in the error counters will be the ±Aoc pulses.
The digital information registered in the error counter is converted into an 800 cps, amplitude modulated, analog signal by the ladder decoder in the D/A converter. This signal is converted into a positive dc analog signal by a phase sensitive demodulator circuit also located in the D/A converter. The dc analog signal is applied through the energized relay contacts to the LEM velocity display meters. As the velocity changes, as calculated by the LGC, representative ± A0c pulses will continue to be sent to the error counter, causing it to count up or down and thereby changing the D/A converter dc signal to the display meters.
L_
0/A CONVERTER (800 CPS LADDER NETWORK AND DEMODULATOR)
t t t t t t t t t
|
ERROR COUNTER |
||||||||
|
w.zi |
5 6* |
2.8* |
14* |
.7* |
.35* |
.17* |
0.8* |
44* |
|
28 |
2? |
2 6 |
2 5 |
24 |
2 3 |
22 |
2 1 |
2 0 |
MOOE
LOGIC
D6PLAY INERTIAL DATA REQUEST
D/A ENABLE
+ A0C ~ A 0 cj
I
160 SEC
| _ LGC _
15637
Figure 2-16. Display Inertial Data Mode
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LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
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MANUAL
2-4.5.10 Master Reset Condition (Test Area Only). The purpose of the master reset condition is to establish preselected standard operating modes in both the airborne equipment and the GSE. The master reset condition is operable in the ISS test con¬ figuration only. The master reset condition is initiated when the MASTER RESET pushbutton on the test control panel of the OIA is pressed.
The effects of establishing a master reset condition are dependent upon the particular ISS level of test, power mode status, et cetera, at the time the MASTER RESET push¬ button is pressed. With the ISS STANDBY pushbutton selected, but prior to pressing the ISS OPERATE pushbutton, the MASTER RESET pushbutton will cause the simultaneous closure of the IMU stabilization loops. During the first 90 seconds after pressing the ISS OPERATE pushbutton, the MASTER RESET pushbutton is disabled. Ninety seconds after pressing the ISS OPERATE pushbutton, the MASTER RESET push¬ button is enabled and, if selected, simultaneously performs the following operations: causes the coarse align mode to be commanded, places the gimbals under gimbal positioner control, and removes all IMU caging signals. The master reset condition also discontinues all LORS mode commands and commands the LORS channels of the CDU to repeat the LORS angles.
2-4.6 ISS POWER SUPPLIES.
The ISS power supplies convert the +28 vdc prime LEM power into the various dc and ac voltages required by the ISS. The power supplies are the pulse torque power supply; the -28 vdc power supply; the 800 cps, 1 percent power supply; the 800 cps, 5 percent, 2 phase, power supply; and the 3,200 cps power supply. The pulse torque power supply is in the PTA and the remaining power supplies are in the PSA.
The +28 vdc prime power is supplied by the LEM electrical power system through the ISS OPERATE circuit breaker. All ac power supplies are synchronized to the LGC clock by means of computer pulses. Thedc supplies, using multivibrators as ac sources for transformation, are also synchronized to the LGC. Synchronization is accomplished by a multivibrator which will free runata lower frequency without the computer pulses, assuring operation of the ISS power supplies in the event of an LGC failure.
2-4. 6.1 Pulse Torque Power Supply. The pulse torque power supply (figure 2-17) provides 120 vdc to the three binary current switches and three dc differential ampli¬ fiers in the accelerometer loops and the binary current switch and dc differential amplifier in the stabilization loop fine align electronics. The pulse torque power supply also provides three individual 28 vdc outputs to the accelerometer loop PVR’s, 20 vdc to the three accelerometer loop ac differential amplifier and interrogator modules and the associated binary current switches, and -20 vdc to the ac differential ampli¬ fier and interrogator module in the accelerometer loops.
The -20 vdc output is derived from the -28 vdc power supply by using a zener diode as a voltage divider and regulator. The output is regulated at -20(±0.8) vdc.
The 20 vdc output is derived from 28 vdc prime power by the use of a three tran¬ sistor series regulator which maintains the output voltage at 20 (±0.55) vdc.
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LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
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MANUAL
I523I-B
Figure 2-17. Pulse Torque Power Supply
The 120 vdc and 28 vdc PVR outputs are derived from a multivibrator, a power amplifier, and a rectifier and filter. A 12.8 kpps synchronizing pulse is received from the LGC through a buffer transformer in the pulse torque insolation transformer assem¬ bly and is applied to an amplifier-inverter. The output of the amplifier-inverter is applied to a multivibrator-chopper causing it to be synchronized at 6,400 cps. A tran¬ sistorized time delay circuit is incorporated into the emitter circuits of the multi¬ vibrator to provide a turn on time delay of approximately 350 milliseconds. During the 90 second IMU turn on mode, 0 vdc is applied through the turn on circuits of the IMU auxiliary assembly module to the time delay circuit which inhibits the 120 vdc and
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LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
28 vdc PVR supplies. The multivibrator-chopper output is applied to the primary of a transformer which has 28 vdc prime power applied to its center tap. The secondary of the transformer, which is also center tapped, is coupled to a two stage push-pull power amplifier which operates from 28 vdc prime power. The output of the power amplifier consists of a transformer with four secondary windings; one with center tap return for the 120 vdc power supply, and one each for the X, Y, and Z accelerometer loop 28 vdc PVR supplies. The 120 vdc power supply consists of a full wave rectifier whose output is filtered, regulated, and again filtered. The 28 vdc power supplies are identical and consist of a full wave bridge rectifier whose output is filtered, regulated, and again filtered. The PVR time delay circuit inhibits the operation of the regulator in each 28 vdc PVR circuit to provide a six to eight second time delay in the 28 vdc PVR outputs.
2-4. 6. 2 -28 VDC Power Supply. The -28 vdc power supply provides input power to the three gimbal servo amplifiers in the stabilization loops and to the pulse torque power supply to generate -20 vdc foruseinthe accelerometer loops. The -28 vdc power supply consists of a pulse amplifier-inverter, a multivibrator-chopper, a power amplifier, and a rectifier and filter. (See figure 2-18.) The 25.6 kpps synchronization pulse input is amplified and inverted for use in synchronizing the multivibrator-chopper at 12.8 kcps. The multivibrator-chopper output is applied to the primary of a transformer which has 28 vdc prime power applied to its center tap. The secondary of the transformer, which is also center tapped, is coupled to a push-pull power amplifier. The output of the amplifier is transformer coupled to a full wave rectifier and filter whose positive side is referenced to ground to provide a -27.0 (±1.0) vdc output.
2-4. 6.3 800 CPS Power Supply. The 800 cps power supply (figure 2-19) consists of four modules; an automatic amplitude control, filter, and multivibrator; a 1 percent ampli¬ fier; and two 5 percent amplifiers. The 1 percent amplifier provides IMU gimbal resolver excitation, gimbal servo amplifier demodulator reference, and FDAI and autopilot reference. The two 5 percent amplifiers provide gyro wheel excitation, IMU blower excitation, and accelerometer fixed heater power. The 1 percent amplifier also provides the input to one of the 5 percent amplifiers whose output is phase shifted -90 degrees. The output of this 5 percent amplifier is applied to the second 5 percent amplifier whose output is also phase shifted -90 degrees, or -180 degrees from the output of the 1 percent amplifier. The outputs of the 1 percent amplifier and the 5 percent ampli¬ fiers are applied to their respective loads through the IMU load compensation network which provides a power factor correction.
Zero and pi phase, 800 cps pulse trains from the LGC synchronize the multivibrator at 800 cps. In the absence of the synchronizing pulses, the multivibrator will free run at between 720 and 800 cps. The output of the multivibrator controls the operation of the chopper and filter circuit. The filtered chopper output is applied to the 1 percent ampli¬ fier. The output of the 1 percent amplifier, in addition to its direct uses, is a feedback signal to the automatic amplitude control circuit. The positive peaks of this feedback signal are detected and added to a dereference signal. The sum is filtered and provides a dc bias to the multivibrator driven chopper. The bias controls the amplitude of the chopped signal.
2-39
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
20 VDC PRIME POWER
FROM LGC
25 6KPPS 0 PHASE
PULSE
AMPL-
INVERTER
|
« |
||
|
MULTI- |
||
|
VIBRATOR |
||
|
CHOPPER |
||
|
FULL |
|
|
WAVE |
|
|
RECTIFIER |
|
|
8 FILTER |
— 27.5 VDC
I5230C
Figure 2-18. -28 VDC Power Supply
28vll.5% 800 CPS 0* PHASE
20 V + 5.0% 800 CPS -90* PHASE (A PHASE)
28V t 7.5% 800 CPS - 180* PHASE (8 PHASE)
152288
Figure 2-19. 800 CPS Power Supply
2-40
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
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MANUAL
The 1 percent amplifier is push-pull in operation with transformer coupled input and output and with overall voltage feedback for gain and distortion control.
The two 5 percent amplifiers are identical in operation. The amplifiers are push-pull and have transformer coupled inputs and outputs. The input transformer primary center tap is connected to the input signal low. The input signal high is applied directly to one side of the primary winding and is also applied through a phase shift network to the other, or out of phase, side of the primary. A feedback signal from the secondary of the output transformer is also applied to the out of phase side of the input transformer primary where it is mixed with the phase shifted portion of the input signal. This mixing results in a -90 degree phase shift in the secondary of the input transformer. The output of the first 5 percent amplifier is used as an input to the second 5 percent amplifier to provide an additional -90 degree phase shift.
2-4. 6. 4 3,200 CPS Power Supply. The 3,200 cps power supply provides excitation voltage for the signal generator and the magnetic suspension portions of the IRIGand PIP ducosyns. The 3, 200 cps output is also used as a reference for the demodulator in the gimbal servo amplifiers.
The excitation voltage to the signal generators requires both voltage stability and phase stability. To accomplish this stability, the excitation voltage power transmission to the stable member is through a step down transformer on the stable member which reduces the slip ring current and, therefore, voltage dropeffects due to slip ring, cable, and connector resistance. In addition, each wire connecting the output of the transformer to the input terminals of each PIP is cut to exactly the same length. The voltage level at the primary of the transformer is fed back to the power supply and is compared to a voltage reference.
The 3, 200 cps power supply (figure 2-20) consists of an amplitude control module and a 1 percent power amplifier. The amplitude control module contains an automatic amplitude control circuit, a multivibrator, a chopper, and a filter.
The 3, 200 pps pulse trains of zero degree phase and 180 degree phase synchronize a multivibrator. The output of the multivibrator controls the operation of the chopper circuit. The output of the chopper is applied to the 1 percent power amplifier. The 28 volt rms output of the amplifier is transmitted through the slip rings to the trans¬ former on the stable member where the voltage is stepped down to 2 volts for the accel¬ erometer ducosyns and 4 volts for the gyro ducosyns. A sample of the 28 volt level at the primary of the transformer is fed back through the slip rings to the input of the automatic amplitude control circuit. The positive peaks of the feedback signal are detected and added to a dc reference signal. The sum is filtered and provides a dc bias to the chopper circuit. The dc bias controls the amplitude of the chopper output to the filter.
2-5 LEM OPTICAL RENDEZVOUS SUBSYSTEM
This paragraph will give a functional description of the LORS and it will link the LORS operations to systems level operations. This paragraph will be supplied when information is available.
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LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
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MANUAL
I - 1
[ IMU ]
I5229C
Figure 2-20. 3,200 CPS Power Supply 2-6 COMPUTER SUBSYSTEM
The computer subsystem (CSS) is the control and processing center of the PGNCS. It consists of the LGC and a DSKY. The CSS processes data and issues discrete outputs and control pulses to the PGNCS and other LEM systems. The LGC is a parallel digital control computer with many features of a general purpose computer. As a control computer, the LGC aligns the IMU, positions the optical tracker, and issues control commands to other LEM systems. As a general purpose computer, the LGC solves the guidance and navigation equations required for the LEM mission. In addition, the LGC monitors the operation of the LEM, including the CSS.
The main functions of the LGC (see figure 2-21) are implemented through the ex¬ ecution of the programs stored in memory. Programs are written in a machine language called basic instructions. A basic instruction contains an operation (order) code and a relevant address. The order code defines the data flow within the LGC, and the relevant address selects the data that is to be used for computations. The order code of each instruction is entered into the sequence generator, which controls data flow and pro¬ duces a different sequence of control pulses for each instruction. Each instruction is followed by another instruction. In order to specify the sequence in which consecutive instructions are to be executed, the instructions are normally stored in successive memory locations. By adding the quantity one to the address of an instruction being executed, the address of the instruction to be executed next is derived. Execution of an instruction is complete when the order code of the next instruction is transferred to the sequence generator and the relevant address is in the central processor.
2-42
- DATA FLOW
- CONTROL SIGNALS
4-28VDC
PRIMARY
INPUT
|
POWER |
||
|
SUPPLY |
4- 4 V, 14V
(TO all
FUNCTIONAL
AREASI
FROM
SPACECRAFT
MODE / STATUS SIGNALS ENGINE SIGNALS TELEMETRY
FROM
INERTIAL
SUBSYSTEM
GIMBAL ANGLES (CDU)
VELOCITY INCREMENTS (PIPA)
FROM {
LEM OPTICAL TRACKER ANGLES(CDU) - -»
RENDEZVOUS <
SUBSYSTEM | MODE DISCRETES - »
STABILIZATION AND CONTROL SYSTEM
TRANSLATION, ROTATION ,
AND IMPULSE SIGNALS
INPUT
NTERFACE
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
FROM PONCS
MODE AND CAUTION SIGNALS
1 E YBOARD INPUTS
DISPLAY
AND
KEYBOARD
MODE AND CAUTION SIGNALS
\ TO PGNCS ' AND
SPACECRAFT
DISPLAY DATA AND CAUTION SIGNALS
INPUT
CHANNELS
SEOUENCE
GENERATOR
INPUT
LOGIC
CENTRAL
PROCESSOR
OUTPUT
CHANNELS
REAL time
KEYRUPJ _AND HNDRPT
INCREMENTAL PULSES
PRIORITY
CONTROL
OUTPUT
INTERFACE
OUTPUT SYNC SIGNALS
I
- TIMER
TIMING SIGNALS (TO ALL AREAS)
MASTER CLOCK AND ENGINE CONTROL SIGNALS
RADAR TIMING SIGNALS DOWNLINK DATA
CDU DRIVE PULSES PIPA INTERR 8 SWITCH GYRO DRIVE PULSES
.CDU DRIVE PULSES TIMING SIGNALS
TIMING SIGNALS
PITCH, YAW. ANO ROLL SIGNALS
TO SPACECRAFT
TO INERTIAL SUBSYSTEM
TO
LEM OPTICAL RENDEZVOUS SUBSYSTEM
TO PSA TO
STABILIZATION AND CONTROL SYSTEM
Figure 2-21. Computer Subsystem, Block Diagram
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The central processor consists of several flip-flop registers. It performs arithmetic operations and data manipulations on information accepted from memory, the input channels, and priority control. Arithmetic operations are performed using the ONE's complement number system. Values of 14 bits, excluding sign, (up to 28 bits during double precision operations) are processed with an additional bit produced for overflow or underflow. All operations within the central processor are performed under control of pulses generated by the sequence generator (indicated by dashed lines in figure 2-21). In addition, all words read out of memory are checked for correct parity, and a parity bit is generated within the central processor for all words written into memory. The LGC uses odd parity, that is, all words stored in memory contain an odd number of ONE's including the parity bit. The central processor also supplies data and control signals through the output channel's and provides interface for the various spacecraft subsystems.
The LGC has provision for nine program interrupts. These interrupts are: T6 RUPT, T5 RUPT, T4 RUPT, T3 RUPT, KEYRUPT, UPRUPT, DOWNRUPT, RADAR, and HNDRPT. The T6 RUPT through T4 RUPT programs are initiated by the LGC. The DOWNRUPT program is initiated at the completion of every parallel-to-serial conversion for downlink operation. The remaining interrupt programs are initiated by external inputs to the LGC. The KEYRUPT programs are initiated when a DSKY pushbutton is depressed or when priority control receives a signal (discrete bit) from the optical tracker to indicate a sighting. The UPRUPT and RADAR programs are initiated when a complete UPLINK word (used for unmanned flights) is received. The HNDRPT program is initiated as soon as a hand controller is moved out of detent by the astronaut.
Before a priority program can be executed, the current program must be inter¬ rupted; however, certain information about the current program must be preserved. This information includes the program counter contents and any intermediate results contained in the central processor. The priority control produces an interrupt request signal, which is sent to the sequence generator. This signal, acting as an order code, causes the execution of an instruction that transfers the current contents of the program counter and any intermediate results to memory. In addition, the control pulses transfer the priority program address in priority control to the central processor, and then to memory through the write lines. As a result, the first basic instruction word of the priority program is entered into the central processor from memory, and execution of the priority program is begun. The last instruction of each priority program restores the LGC to normal operation, provided no other interrupt request is present, by trans¬ ferring the previous program counter and intermediate results from their storage loca¬ tions in memory back to the central processor.
Certain data pertaining to the flight of the LEM is used to solve the guidance and navigation problems required for the LEM mission. This data, which includes real time, acceleration, and IMU gimbal angles, is stored in memory locations called counters. The counters are updated as soon as new data becomes available. An incrementing process which changes the contents of the counters is implemented by
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priority control between the execution of basic instructions. Data inputs to priority control are called incremental pulses. Each incremental pulse produces a counter address and a priority request. The priority request signal is sent to the sequence generator, where it functions as an order code. The control pulses produced by the sequence generator transfer the counter address to memory through the write lines of the central processor. In addition, the control pulses enter into the central processor the contents of the addressed counter to be incremented.
Real time plays a major role in solving guidance and navigation problems. Real time is maintained within the LGC in the main time counter of memory. The main time counter provides a 745. 65 hour (approximately 31 days) clock. Incremental pulses are produced in the timer and sent to priority control for incrementing the main time counter.
The LEM mission requires that the LGC clock be synchronized with the KSC clock. The LGC time is transmitted once every second by downlink operation for comparison with the KSC clock.
Incremental transmissions occur in the form of pulse bursts from the output channels to the CDU, the gyro fine align electronics, the RCS of the spacecraft, the optical tracker and the radar. The number of pulses and the time at which they occur are controlled by the LGC program. Discrete outputs also originate in the output channels under program control. These outputs are sent to the DSKY and various other subsystems. Continuous pulse trains originate in the timing output logic for synchronization of other systems.
The uplink word from the LEM telemetry system (unmanned flights) is supplied as an incremental pulse input to priority control. As this word is received, priority control procudes the address of the uplinkcounter in memory and requests the sequence gener¬ ator to execute the instructions which perform the serial-to-parallel conversion of the input word. When the serial-to-parallel conversion is completed, the parallel word is transferred to a storage location in memory by the uplink priority program. The uplink program also retains the parallel word for subsequent downlink transmission. Another program converts the parallel word to a coded display format and transfers the display information to the DSKY.
The downlink operation of the LGC is asynchronous with respect to the LEM telemetry system. The telemetry system supplies all the timing signals necessary for the downlink operation. These signals include start, end, and bit sync pulses.
Through the DSKY, the astronaut can load information into the LGC, retrieve and display information contained in the LGC, and initiate any program stored in memory. A keycode is assigned to each keyboard pushbutton. When a keyboard pushbutton on the DSKY is depressed, the keycode is produced and sent to an input channel. A signal is also sent to priority control, where it produces both the address of a priority program stored in memory and apriority request signal, which is sent to the sequence generator. This operation results in an order code and initiates an instruction for interrupting the program in progress and executing the KEYRUPT priority program stored in memory.
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A function of this program is to transfer the keycode, temporarily stored in an input channel, to the central processor, where it is decoded and processed. A number of keycodes are required to specify an address, or a data word. The program initiated by a keycode also converts the information from the DSKY keyboard to a coded display format. The coded display format is transferred by another program to an output channel and sent to the display portion of the DSKY. The display notifies the astronaut that the keycode was received, decoded, and processed properly by the LGC.
2-6.1 PROGRAMS. An LGC program performs such functions as solving guidance and navigation problems, testing the operation of the PGNCS, and monitoring the operation of the LEM. Such a program consists of a group of program sections that are classified according to the functions they perform. These functions are defined as mission functions, auxiliary functions, and utility functions. (See figure 2-22.)
2-6. 1. 1 Mission Functions. Mission functions are performed by program sections that implement operations concerned with the major objectives of the LEM mission. These operations include erecting the IMU stable member and coarse aligning it to a desired heading prior to separating the LEM from the CSM and fine aligning it after separation. In addition, the mission functions include computation of spacecraft position and velocity during coasting periods of the flight by solution of second-order differential equations which describe the motions of a body subject to the forces of gravity.
2-6. 1.2 Auxiliary Functions. Auxiliary functions are executed at the occurrence of certain events, requests, or commands. These functions are performed by program sections that provide a link between the LGC and other elements of the PGNCS. This link enables the LGC to process signals from various devices and to send commands for control and display purposes. In addition, the auxiliary functions implement many and varied operations within the LGC in support of the LEM mission functions.
ITnput/output
L
1 [memory
I I
J L
1
J
40383
Figure 2-22. Program Organization
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2-6. 1.3 Utility Functions. Utility functions are performed by program sections that coordinate and synchronize LGC activities to guarantee orderly and timely execution of required operations. These functions control the operation of the LEM mission functions and schedule LGC operations on either a priority or a real-time basis. The utility functions also translate interpretive language to basic machine language which allows complex mathematical operations such as matrix multiplication, vector addition, and dot product computations to be performed within the framework of compact routines. In addition, the utility functions save the contents of registers A and Q during an inter¬ rupt condition and enable data retrieval and control transfer between isolated banks in the fixed- switchable portion of fixed memory.
2-6.2 MACHINE INSTRUCTIONS. The LGC has three classes of machine instructions: regular, involuntary, and peripheral (table 2-1). Regular instructions are programmed and are executed in whatever sequence they have been stored in memory. Involuntary instructions (with one exception) are not programmable and have priority over regular instructions. One involuntary instruction may be programmed to test computer opera¬ tions. No regular instruction can be executed when the LGC forces the execution of an involuntary instruction. The peripheral instructions are used when the LGC is connected to the peripheral equipment. During the execution of any peripheral instruction, the LGC is in the monitor stop mode and cannot perform any program operation.
2-6.2. 1 Regular Instructions. Four types of instructions comprise the regular instruc¬ tion class. They are the basic, channel, extracode, and special instructions. Basic instructions are used most frequently. The instruction words stored in memory are called basic instruction words. They contain an order code field and an address field. Special instructions have predefined addresses and order codes; basic instructions have only predefined order codes. The special instructions are used to control certain operations in the LGC. For example, one special instruction is used to switch the LGC to the extend mode of operation. This mode extends the length of the order code field and converts basic instruction words to channel or extracode instruction words. Channel instructions can only be used with input-output channel addresses. Extra code instructions perform the more complex and less frequently used arithmetic oper¬ ations.
Regular instructions can also be functionally subdivided into the following:
(1) Sequence changing.
(2) Fetching and storing.
(3) Modifying.
(4) Arithmetic and logic.
(5) Input- output.
(6) Editing.
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Table 2-1. Instruction Classes
|
Class |
Type |
Control |
|
Regular |
Basic Extracode Channel Special |
Program |
|
Involuntary |
Interrupt Counter |
Priority |
|
Peripheral |
Keyboard Tape |
Operator |
The sequence changing instructions alter the sequence in which the instructions stored in memory are executed. One group, called transfer control instructions, changes the program path as defined by the programmer. The other group, called decision making instructions, branches to alternate program paths in response to predefined conditions.
The fetching and storing instructions move data, without alteration, from one loca¬ tion to another. One group, called copy instructions, provides a non-destructive transfer of data from memory to the central processor. Another group, called exchange in¬ structions, transposes data between memory and the central processor. One instruc¬ tion provides a nondestructive transfer of data from the central processor to memory.
The modifying instructions alter the next instruction to be executed by changing the contents of the order code field, address field, or both.
The arithmetic and logic instructions perform numerical computations. One group, called the basic arithmetic instructions, performs addition, subtraction, multiplication, and division in the ONE'S complement number system. Another group, called the add and store instructions, performs single or double precision addition and transfers the resultant from the central processor to memory. The incrementing instructions incre¬ ment a signed quantity, increment its absolute value, or diminish its absolute value by one. One instruction performs subtraction in the TWO's complement number system for angular data and one instruction performs the Boolean AND operation.
The input-output or channel instructions link the interface circuits to the central processor. One group, called read instructions, transfers the total or partial contents of any channel (register) location to the central processor either directly or accompanied by the Boolean AND, OR, or EXCLUSIVE OR operation. Another group of instructions transfers all new or partially new information to any channel location in the same manner.
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The editing or special instructions are address-dependent and control the operation of the program. One special instruction, as mentioned previously, controls the extend mode of operation. Other instructions prevent a program from being interrupted or shift and cycle data to the left or right.
2-6. 2. 2 Involuntary Instructions. Involuntary instructions contain two types of instruc¬ tions: interrupt and counter. The interrupt instructions use the basic instruction word format just as the regular instructions do; however, the interrupt instructions are not entirely programmable. The contents of the order code field and the address field are supplied by computer logic rather than the program. The counter instructions have no instruction word format. Signals which function as a decoded order code specify the counter instruction to be executed and the computer logic supplies the address. The address for these instructions is limited to one of 29 counter locations in memory.
There are two interrupt instructions. One instruction initializes the LGC when power is first applied and when certain program traps occur. The other interrupt in¬ struction is executed at regular intervals to indicate time, receipt of new telemetry or keyboard data, or transmission of data by the LGC. This interrupt instruction may be programmed to test the computer.
There are several counter instructions. Two instructions will either increment or decrement by one the content of the counter location using the ONE'S complement number system. Two other instructions perform the same function using the TWO's complement number system. Certain counter instructions control output rate signals and convert serial telemetry data to parallel computer data.
2-6. 2. 3 Peripheral Instructions. There are two types of peripheral instructions. One type deals' with memory locations and the other type deals with channel locations. The peripheral instructions are not used when the LGC is in the LEM. They are used when the computer is connected to peripheral equipment during subsystem and preinstallation system testing. The peripheral instructions are not programmable and are executed when all computer program operations have been forcibly stopped. These instructions are used to read and load any memory or channel location and to start the computer program at any specified address. The peripheral instructions and counter instructions are processed identically.
2-6.3 TIMER. The timer generates the timing signals required for operation of the LGC and is the primary source of timing signals for all LEM systems.
The timer is divided into the areas indicated in figure 2-23. The master clock fre¬ quency is generated by an oscillator and is applied to the clock divider logic. The divider logic divides the master clock input into gating and timing pulses at the basic clock rate of the computer. Several outputs are available from the scaler, which further divides the divider logic output into output pulses and signals which are used for gating, for generating rate signal outputs, and for accumulating time. Outputs from the divider logic also drive the time pulse generator which produces a recurring setof time pulses. This setof time pulses defines a specific interval (memory cycle time) in which access to memory and word flow take place within the computer.
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GATING AND TIMING PULSES
PULSE AND SIGNAL OUTPUTS
TIME PULSES
STOP
40686*
Figure 2-23. Timer, Block Diagram
The start- stop logic senses the status of the power supplies and specific alarm conditions in the computer and generates a stop signal which is applied to the time pulse generator to inhibit word flow. Simultaneous with the generation of the stop signal, a fresh start signal is generated which is applied to all functional areas in the computer. The start- stop logic and subsequent word flow in the computer can also be controlled by inputs from the Computer Test Set (CTS) during pre-installation systems and sub¬ system tests.
2-6.4 SEQUENCE GENERATOR. The sequence generator executes the instructions stored in memory. The sequence generator processes instruction codes and produces control pulses which regulate the data flow of the computer. The control pulses are responsible for performing the operations assigned to each instruction in conjunction with the various registers in the central processor and the data stored in memory.
The sequence generator (figure 2-24) consists of the order code processor, com¬ mand generator, and control pulse generator. The sequence generator receives order code signals from the central processor and priority control. These signals are coded by the order code processor and supplied to the command generator. The special purpose control pulses are used for gating the order code signals into the sequence generator at the end of each instruction.
The command generator receives instruction signals from priority control and peripheral equipment and receives coded signals from the order code processor. The command generator decodes the input signals and produces instruction commands which are supplied to the control pulse generator.
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Figure 2-24. Sequence Generator, Block Diagram
The control pulse generator receives twelve time pulses from the timer. These pulses occur in cycles and are used for producing control pulses in conjunction with the instruction commands. There are five types of control pulses: read, write, test, direct exchange, and special purpose. Information in the central processor is trans¬ ferred from one register to another by the read, write, and direct exchange control pulses. The special purpose control pulses regulate the operation of the order code processor. The test control pulses are used within the control pulse generator. The branch test data from the central processor changes the control pulse sequence of various instructions.
2-6.5 CENTRAL PROCESSOR. The central processor, figure 2-25, consists of the flip-flop registers, the write, clear, and read control logic, write amplifiers, memory buffer register, memory address register, and decoder and the parity logic. All data and arithmetic manipulations within the LGC take place in the central processor.
Primarily, the central processor performs operations indicated by the basic in¬ structions of the program stored in memory. Communication within the central pro¬ cessor is accomplished through the write amplifiers. Data flows from memory to the flip-flop registers or vice-versa, between individual flip-flop registers, or into the central processor from external sources. In all instances, data is placed on the write lines and routed to a specific register or to another functional area under control of the write, clear, and read logic. This logic section accepts control pulses from the sequence generator and generates signals to read the content of a register onto the write lines and to write this content into another register of the central processor or to another functional area of the LGC. The particular memory location is specified by the content of the memory address register. The address is fed from the write lines into this register, the output of which is decoded by the address decoder logic. Data is subsequently transferred from memory to the memory buffer register. The decoded address outputs are also used as gating functions within the LGC.
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Figure 2-25. Central Processor, Block Diagram
The memory buffer register buffers all information read out or written into memory. During readout, parity is checked by the parity logic and an alarm is generated in case of incorrect parity. During write-in, the parity logic generates a parity bit for informa¬ tion being written into memory. The flip-flop registers perform the data manipulations and arithmetic operations. Each register is 16 bits or one computer word in length. Data flows into and out of each register as dictated by control pulses associated with each register. The control pulses are generated by the write, clear, and read control logic.
External inputs through the write amplifiers include the content of both the erasable and fixed memory bank registers, all interrupt addresses from priority control, control pulses which are associated with specific arithmetic operations, and the start address for an initial start condition. Information from the input and output channels is placed on the write lines and routed to specific destinations either within or external to the central processor. The CTS inputs allow a word to be placed on the write lines during system and subsystem tests.
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2-6.6 PRIORITY CONTROL. Priority control is related to the sequence generator in that it controls all involuntary or priority instructions. The priority control processes input-output information and issues order code and instruction signals to the sequence generator and issues twelve-bit addresses to the central processor.
The priority control (figure 2-26) consists of the start, interrupt, and counter in¬ struction control circuits. The start instruction control initializes the computer if the program works itself into a trap, if a transient power failure occurs, or if the interrupt instruction control is not functioning properly. The computer is initialized with the start order code signal, which not only forces the sequence generator to execute the start instruction, but also resets many other computer circuits. When the start order code signal is being issued, the T12 stop signal is sent to the timer. This signal stops the time pulse generator until all essential circuits have been reset and the start in¬ struction has been forced by the sequence generator. The computer may also be initial¬ ized manually when connected to the peripheral equipment and placed into the monitor stop mode. In this mode, the time pulse generator is held at the T12 position until the monitor stop signal is released.
SEQUENCE
GENERATOR
CENTRAL
'PROCESSOR
SEQUENCE
GENERATOR
SPACECRAFT
4 0689
Figure 2-26. Priority Control, Block Diagram
#
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The interrupt instruction control can force the execution of the interrupt instruction by sending the interrupt order code signal to the sequence generator and the twelve bit address to the central processor. There are ten addresses, each of which accounts for a particular function that is regulated by the interrupt instruction control. The interrupt instruction control links the keyboard, telemetry, and time counters to program oper¬ ations. The interrupt addresses are transferred to the central processor by read control pulses from the sequence generator. The source of the keyboard, telemetry, and time counter inputs is the input-output circuits. The interrupt instruction control has a built- in priority chain which allows sequential control of the ten interrupt addresses. The decoded interrupt addresses from the central processor are used to control the priority operation.
The counter instruction control is similar to the interrupt instruction control in that it links input-output functions to the program. It also supplies twelve-bit addresses to the central processor and instruction signals to the sequence generator. The instruction signals cause a delay (not an interruption) in the program by forcing the sequence gen¬ erator to execute a counter instruction. The addresses are transferred to the central processor by read control pulses. The counter instruction control also has a built-in priority of the 29 addresses it can supply to the central processor. This priority is also controlled by decoded counter address signals from the central processor. The counter instruction control contains an alarm detector which produces an alarm if an incremental pulse is not processed properly.
2-6.7 INPUT-OUTPUT. The input-output section accepts all inputs to, and routes to other systems all outputs from, the computer. The input-output section (figure 2-27) includes the interface circuits, input and output channels, input logic, output timing logic, and the downlink circuits.
START SYNC AND
TO OUTPUT INTERFACE
DOWNLINK WORD
KEYCOOE TO DSKY
STABILIZATION AND CONTROL SYSTEM OUTPUTS
PGNCS OUTPUTS
RADAR ACTIVITY OUTPUTS INERTIAL SUB¬ SYSTEM OUTPUTS
TIMING SIGNALS TO PGNCS AND SPACECRAFT
I5808B
Figure 2-27. Input-Output, Block Diagram
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Most of the input and output channels are flip-flop registers similar to the flip-flop registers of the central processor. Certain discrete inputs are applied to individual gating circuits which are part of the input channel structure, typical inputs to the chan¬ nels include keycodes from the DSKY and signals from the PGNCS proper and other LEM systems. Input data is applied directly to the input channels; there is no write pro¬ cess as in the central processor. However, the data is read out to the central processor under program control. The input logic circuits accept inputs which cause interrupt sequences within the computer. These incremental inputs (acceleration data from the PIPA’s, et cetera) are applied to the priority control circuits and subsequently to asso¬ ciated counters in erasable memory.
Outputs from the computer are placed in the output channels and are routed to specific systems through the output interface circuits. The operation is identical to that in the central processor. Data is written into an output channel from the write lines and read¬ out to the interface circuits under program control. Typically, these outputs include outputs to the stabilization and control system, the DSKY, the PGNCS, et cetera. The downlink word is also loaded into an output channel and routed to the LEM spacecraft telemetry system by the downlink circuits.
The output timing logic gates synchronization pulses (fixed outputs) to the PGNCS and the LEM spacecraft. These are continuous outputs since the logic is specifically powered during normal operation of the computer and during standby.
2-6.8 MEMORY. Memory (figure 2-28) consists of an erasable memory with a storage capacity of 2048 words and a fixed core rope memory with a storage capacity of 36, 864 words. Erasable memory is a random-access, destructive-readout storage device. Data stored in erasable memory can be altered or updated. Fixed memory is a nondestructive storage device. Data stored in fixed memory is unalterable since the data is wired in and readout is nondestructive.
Both memories contain magnetic-core storage elements. In erasable memory, the storage elements form a core array; infixed memory, the storage elements form three core ropes. Erasable memory has a density of one word per 16 cores: fixed memory has a density of eight words per core. Each word is located by an address.
In fixed memory, addresses are assigned to instruction words to specify the sequence in which they are to be executed; blocks of addresses are reserved for data, such as constants and tables. Information is placed into fixed memory permanently by weaving patterns through the magnetic cores. The information is written into assigned locations in erasable memory with the CTS, the DSKY, uplink, or program operation.
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CENTRAL
PROCESSOR*
REGISTER G (CENTRAL PROCESSOR)
40691
Figure 2-28. Memory, Block Diagram
Both memories use a common address register (register S)and an address decoder in the central processor. When resister S contains an address pertaining to erasable memory, the erasable memory cycle timing is energized. Timing pulses sent to the erasable memory cycle timing then produce strobe signals for the read, write, and sense functions. The erasable memory selection logic receives an address and a decoded address from the central processor and produces selection signals which permits data to be written into or read out of a selected storage location. When a word is read out of a storage location in erasable memory, the location is cleared. A word is written into erasable memory through the memory buffer register (register G) in the central pro¬ cessor by a write strobe operation. A word read from a storage location is applied to the sense amplifiers. The sense amplifiers are strobed and the information is entered into register G of the central processor. Register G receives information from both memories.
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The address in register S energizes the fixed memory cycle timing when a location in fixed memory is addressed. The timing pulses sent to the fixed memory cycle timing produce the strobe signals for the read and sense functions. The selection logic receives an address from the write lines, a decoded address and addresses from register S, and produces selection signals for the core rope. The content of a storage location in fixed memory is strobed from the fixed memory sense amplifiers to the erasable memory sense amplifiers and then entered into register G of the central processor.
2-6.9 POWER SUPPLIES. The two power supplies (figure 2-29) furnish operating volt¬ ages to the LGC and the DSKY. Primary power of 28 vdc from the spacecraft is applied to both power supplies. Regulator circuits maintain a constant output of +4 volts and +4 volts switched from one supply, and +14 volts and +14 volts switched from the other. The regulator circuits are driven by a sync signal input from the timer, each power supply having a different sync frequency. During system and subsystem tests, inputs from the CTS can be used to simulate power supply failures.
The standby mode of operation is initiated by pressing the standby (STBY) pushbutton on the DSKY. During standby, the LGC is put into a RESTART condition and the switch- able +4 and +14 voltages are switched off, thus putting the LGC into a low power mode where only the timer and a few auxiliary signals are operative.
VOLTAGE
ALARM
FRESH
START
+ 4VDC (SWITCHED)
+I4VDC
(SWITCHED)
OSCILLATOR
■M4VDC -
+ 4VDC (SWITCHED) +14 VDC (SWITCHED)
SCALER + 4VDC +I4VDC
OSCILLATOR
ALARM
SCALER
ALARM
FRESH
START
SCALER
FAIL
FILTER IN
+ 4VDC ( SWITCHEO )
+ I4VDC (SWITCHED)
WARNING
INTEGRATOR
FILTER
OUT
Figure 2-29.
Power Supplies, Block Diagram
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The voltage alarm circuits monitor the +28, +14, and +4 volt outputs and produce an LGC restart signal (Fresh Start) should any of the voltages deviate from nominal by more than a predetermined amount. The oscillator alarm produces an LGC restart signal (Fresh Start) if the oscillator fails or if the LGC is in the standby mode. The scaler alarm circuit monitors the scaler output of the timer and generates a fail signal if the scaler output fails. The warning integrator monitors certain operations and gen¬ erates an LGC warning signal (Filter Out) if these operations are frequently repeated or prolonged.
2-6. 10 DISPLAY AND KEYBOARD. The DSKY is located below the center panels of the cockpit display and control panels.
The DSKY (figure 2-30) consists of a keyboard; a relay matrix with associated de¬ coding circuits, displays, mode and caution circuits; and a power supply. The keyboard, which contains several numerical, sign, and other control keys, allows the astronaut to communicate with the LGC. The inputs from the keyboard are entered into an input channel and processed by the LGC.
COMPUTER OUTPUT CHANNEL 10
KEYCOOE INPUTS TO COMPUTER
CAUTION SIGNALS TO PGNCS
MODE AND CAUTION SIGNALS TO PGNCS ANO SPACECRAFT
15106
Figure 2-30. Display and Keyboard (DSKY), Block Diagram
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The inputs entered from the keyboard, as well as other information, appear on the displays after processing by program. The display of information is accomplished through the relay matrix. A unique code for the characters to be displayed is formed by fifteen bits from output channel 10 in the LGC. Bits 12 through 15 are decoded by the decoding circuits, and, along with bits 1 through 11, energize specific relays in the matrix which causes the appropriate characters to illuminate . The information displayed is the result of a keycode punched in by the astronaut, or is computer-controlled information. The display characters are formed by electroluminescent segments which are energized by a voltage from the power supply routed through relay contacts. Specific inputs from the PGNCS are also applied, through the LGC to certain relays in the matrix through output channel 10 of the LGC. The resulting relay- controlled outputs are caution signals to the PGNCS.
The mode and caution circuits accept direct input signals from channels 11, 12, and 13, without being decoded. The resulting outputs can give an indication to the astronaut on the DSKY and route the output signal to the PGNCS and spacecraft.
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Chapter 3
PHYSICAL DESCRIPTION
3-1 SCOPE
This chapter describes the physical characteristics of the components which com¬ prise the LEM PGNCS. The PGNCS components and their locations within the LEM are listed in table 3-1 and illustrated in figure 3-1. The locations of PGNCS component modules are also illustrated and the module functions described.
3-2 PGNCS INTERCONNECT HARNESS
The PGNCS interconnect harness (composed of harnesses A and B) interconnects the components of the PGNCS and provides the electrical interface between the PGNCS and other LEM systems. The IMU and PTA are interconnected by harness B. Harness A interconnects the PSA, CDU, LGC, and signal conditioner. The two harnesses are connected to each other by vehicle cables. Table 3-II lists the harness connectors and the components or cable to which they are mated.
Table 3-1. LEM PGNCS Components
|
Component |
Part Number |
Location |
|
CDU |
2007222-041 |
Mounted on coldplate on center section of after crew compart¬ ment wall. |
|
PGNCS interconnect harness |
6014515-011 |
|
|
Harness A |
6014506 |
Attached to rear wall of after crew compartment |
|
Harness B |
6014507 |
Located in IMU compartment. |
(Sheet 1 of 2)
3-1
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
Table 3-1. LEM PGNCS Components
|
Component |
Part Number |
Location |
|
IMU and PTA |
6007001-011 |
|
|
IMU |
2018601-011 |
Bolted to after end of nav base. |
|
PTA |
6007000-011 |
Mounted on coldplate on^LEM structure immediately aft of IMU/nav base complex. |
|
LEM guidance computer group |
6003001-021 |
|
|
DSKY |
2003985-031 |
Mounted to front wall of crew compartment below LEM display and control panel. |
|
LGC |
2003100-021 |
Mounted on coldplate on upper section of after crew compart¬ ment wall above CDU. |
|
Luminous beacon |
Mounted on coldplate on CSM on upper bulkhead of service module. |
|
|
Nav base |
6899950-011 |
Bolted to LEM structure in unpressurized compartment above astronauts1 heads. |
|
Optical tracker |
Bolted to forward end of nav base and extending outside LEM. |
|
|
PSA |
6007200-011 |
Mounted on coldplate on lower section of after crew compart¬ ment wall below CDU. |
|
Signal conditioner |
Attached to top of PSA. |
(Sheet 2 of 2)
3-2
PTA
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
3-3
Figure 3-1. Location of LEM PGNCS Components
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
Table 3-II. PGNCS Harness Interconnections
|
PGNCS Harness Connector |
Component |
Component Connector |
|
Harness A |
||
|
56P1 |
Signal conditioner |
- |
|
56P2 |
PSA |
45J19 |
|
56P3 |
CDU |
40J53 |
|
56P4 |
LGC |
05A1J51 |
|
56P5 |
LEM spacecraft harness |
J221 |
|
56P6 |
LEM spacecraft harness |
J220 |
|
56P7 |
LEM spacecraft harness |
J219 |
|
56P8 |
LEM spacecraft harness |
J222 |
|
56P9 |
LEM spacecraft harness |
J218 |
|
56P10 |
LEM spacecraft harness |
J217 |
|
56P11 |
LEM spacecraft harness |
J223 |
|
56P12 |
LEM spacecraft harness |
J224 |
|
56P13 |
LEM spacecraft harness |
J215 |
|
56P14 |
LEM spacecraft harness |
J216 |
|
Harness B |
||
|
56P15 |
LEM spacecraft harness |
J226 |
|
56P16 |
LEM spacecraft harness |
J227 |
|
56P17 |
LEM spacecraft harness |
J228 |
|
56P18 |
LEM spacecraft harness |
J225 |
|
56P19 |
PTA |
35A2J19 |
|
56P20 |
IMU |
35A1J2 |
|
56P21 |
IMU |
35A1J1 |
|
56J1 |
LEM spacecraft harness |
P230 |
3-4
LEM PRIMARY GUIDANCE, NAVIGATION, AND CONTROL SYSTEM
ND-1021042
MANUAL
3-3 NAVIGATION BASE ASSEMBLY
The navigation base assembly (nav base), figure 3-2, is a lightweight mount which supports, in critical alignment, the IMU and optical tracker. The nav base, constructed of one inch diameter, aluminum alloy tubing, weighs approximately three pounds. It consists of a center ring supporting four legs which extend from either side. The ring is approximately 14 inches in diameter and each of the four legs is approxi¬ mately ten inches long. The IMU is mounted to the ends of the four legs on one side of the ring, and the optical tracker is mounted to the