Spacecraft Thermal Control Handbook

Transcription

Spacecraft Thermal Control Handbook
Volume I: Fundamental Technologies
David G. Gilmore
Editor
Second Edition
The Aerospace Press • El Segundo, California
American Institute of Aeronautics and Astronautics, Inc. • Reston, Virginia
Contents
Preface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xi
Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xiii
Contributing Authors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .xv
Chapter 1 Spacecraft Systems Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1
Spacecraft Configurations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1
Earth Orbits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .7
Interplanetary Orbits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .11
Missions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .15
Chapter 2 Spacecraft Thermal Environments. . . . . . . . . . . . . . . . . . . . . . . .21
Environments of Earth Orbit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .21
Standard Earth Orbits. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .36
Environments of Interplanetary Missions . . . . . . . . . . . . . . . . . . . . . . . . . . . .48
Aerobraking Environments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .60
Launch and Ascent Environments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .63
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .67
Chapter 3 Thermal Design Examples . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .71
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .71
Spin-Stabilized Satellites . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .71
Three-Axis-Stabilized Satellites . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .72
Propulsion Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .73
Batteries . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77
Antennas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .79
Sun, Earth, and Star Sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .81
Cooled Devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .84
Solar Arrays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .86
The Huygens Probe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .87
System Overview The Hubble Space Telescope . . . . . . . . . . . . . . . . . . . . . . .95
Chapter 4 Thermal Surface Finishes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .139
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .139
Common Thermal Surface Finishes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .139
Causes of Thermal Surface Degradation . . . . . . . . . . . . . . . . . . . . . . . . . . . .143
Degradation Rates for Common Thermal Finishes . . . . . . . . . . . . . . . . . . . .152
LDEF Results. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .155
Electrical Grounding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .158
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .159
v
Contents
Chapter 5 Insulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Blanket Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Blanket Design Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Materials. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Provisions for Venting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Attachment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Provisions for Electrical Grounding. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fabrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Bakeout and Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
High-Temperature Blankets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Suggestions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Insulation for In-Atmosphere Applications . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
161
161
162
169
169
182
183
186
187
193
193
194
198
205
Chapter 6 Radiators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Passive Structure Radiators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Structural Panels with Heat Pipes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Body-Mounted Radiators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Deployable Radiators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Radiator Freezing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Radiator Effectiveness . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Experimental Radiators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
207
207
209
209
209
211
215
217
220
222
Chapter 7 Heaters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Heater Types. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Failure Modes of Mechanical Thermostats . . . . . . . . . . . . . . . . . . . . . . . . .
Circuits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Computer-Controlled Heater System Example . . . . . . . . . . . . . . . . . . . . . .
Radioisotope Heater Units . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
223
223
223
224
227
228
231
241
Chapter 8 Mountings and Interfaces. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Unit Conduction Cooling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Bolted-Joint Conductance without Interface Filler . . . . . . . . . . . . . . . . . . .
Bolted-Joint Conductance with Interface Filler . . . . . . . . . . . . . . . . . . . . . .
Complex Configurations and Special Topics. . . . . . . . . . . . . . . . . . . . . . . .
Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
247
247
247
260
275
284
320
323
vi
Contents
Chapter 9 Louvers. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Vane Louvers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Analysis of Vane Louvers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Designing Louvers for Operation in Sunlight . . . . . . . . . . . . . . . . . . . . . . .
Pinwheel Louvers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
331
331
332
335
346
349
352
Chapter 10 Heat Switches. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Heat-Switch Applications. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Heat-Switch Integration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Paraffin Heat Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Cryogenic Heat Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
353
353
354
355
357
366
370
Chapter 11 Phase-Change Materials. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Phase-Change-Material Applications. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Phase-Change Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
When To Use a PCM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
PCM Design Details. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
The PCM Design Process. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
373
373
377
380
383
402
402
Chapter 12 Pumped Fluid Loops . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Fluid-Flow Concepts and Basic Equations . . . . . . . . . . . . . . . . . . . . . . . . .
Forced Convection in Pipes and Tubes . . . . . . . . . . . . . . . . . . . . . . . . . . . .
System Hardware . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Analysis of a Fluid Loop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Computer Software for System Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . .
PFL Application . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
405
405
407
415
418
442
443
444
468
Chapter 13 Thermoelectric Coolers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Optimizations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Heat Load Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Interfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
XRT Focal-Plane TEC Mounting. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Design Development . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
473
473
473
475
476
478
478
479
480
vii
Contents
Power Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 481
Application Example . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 481
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 487
Chapter 14 Heat Pipes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Types of Heat Pipe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Heat-Pipe Applications and Performance . . . . . . . . . . . . . . . . . . . . . . . . . .
Heat-Pipe References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
LHPs and CPLs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Selecting a Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
489
489
490
496
500
501
502
502
518
521
Chapter 15 Thermal Design Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Spacecraft Project Phases. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Thermal Design/Analysis Process Overview . . . . . . . . . . . . . . . . . . . . . . . .
Fundamentals of Thermal Modeling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Thermal Design Analysis Example POAM . . . . . . . . . . . . . . . . . . . . . . . . .
Margins. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
TMM Computer Codes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Radiation Analysis Codes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
523
523
523
534
537
552
572
575
592
597
Chapter 16 Thermal Contact Resistance . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Contact Heat-Transfer Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Parameters Influencing Thermal Joint Resistance . . . . . . . . . . . . . . . . . . . .
Thermal Joint Resistance Models. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
The Effect of Oxidation on Thermal Contact Resistance. . . . . . . . . . . . . . .
The Effect of Interstitial Materials on Thermal Contact Resistance . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
599
599
600
602
603
623
626
636
Chapter 17 Precision Temperature Control . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
The Space Interferometry Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
The Hydrogen Maser Clock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
639
639
640
655
666
666
viii
Contents
Chapter 18 Space Shuttle Integration . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Engineering-Compatibility Assessment . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Safety Assessment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
The Cargo Integration Review . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Orbiter Payload-Bay Thermal Environment . . . . . . . . . . . . . . . . . . . . . . . .
Middeck Payload Accommodations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Ferry-Flight Accommodations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Optional Services . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
667
667
669
675
676
677
697
700
701
Chapter 19 Thermal Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Design Environments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Development Thermal Testing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Unit Thermal Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Subsystem and Payload Thermal Testing . . . . . . . . . . . . . . . . . . . . . . . . . .
System Thermal Testing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Launch Site Thermal Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
709
709
710
715
725
727
742
742
756
757
Chapter 20 Technology Projections. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Technology Drivers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Programmatic Concerns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Future Technologies and Innovations . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
759
759
760
761
761
786
787
Appendix A Surface Optical Property Data . . . . . . . . . . . . . . . . . . . . . . . . 791
Appendix B Material Thermal Properties . . . . . . . . . . . . . . . . . . . . . . . . . 803
Appendix C Thermally Conductive Filler Materials and Suppliers . . . . 819
Index . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 831
ix
Contributing Authors
K. Aaron, Jet Propulsion Laboratory, California Institute of Technology,
Pasadena, California
Chapter 17, Precision Temperature Control
J. Ambrose, Lockheed Martin, Sunnyvale, California
B. J. Anderson, NASA/MSFC, Huntsville, Alabama
Chapter 2, Spacecraft Thermal Environments
W. Batts, Computer Sciences Corporation, Huntsville, Alabama
V. Baturkin, National Technical University of Ukraine, formerly Kyiv
Polytechnic Institute, Kyiv, Ukraine
Chapter 8, Mountings and Interfaces
R. Bettini, B.F. Goodrich Aerospace, Danbury, Connecticut
Chapter 3, Thermal Design Examples
P. Bhandari, Jet Propulsion Laboratory, California Institute of Technology,
Chapter 12, Pumped Fluid Loops
G. C. Birur, Jet Propulsion Laboratory, California Institute of Technology,
Chapter 20, Technology Projections
L. Bledjian, The Aerospace Corporation, El Segundo, California
Chapter 11, Phase-Change Materials
D. A. Boyd, Smithsonian Astrophysical Observatory, Cambridge, Massachusetts
A. Chuchra, Swales Aerospace, Beltsville, Maryland
Chapter 13, Thermoelectric Coolers
J. F. Clawson, Jet Propulsion Laboratory, California Institute of Technology,
G. Cluzet, Alcatel, Velizy, France
R. L. Collins, The Aerospace Corporation, El Segundo, California
Chapter 15, Thermal Design Analysis
B. Cullimore, C&R Technologies, Littleton, Colorado
Chapter 14, Heat Pipes
G. M. DeVault, The Boeing Company, Houston, Texas
Chapter 18, Space Shuttle Integration
xv
J. Doenecke, Astrium, Friedrichshaten, Germany
M. Donabedian, The Aerospace Corporation, El Segundo, California
Chapter 5, Insulation
R. J. Eby, Orbital Sciences Corporation, Dulles, Virginia
Chapter 9, Louvers
D. Ferguson, Swales Aerospace, Beltsville, Maryland
C. Finch, BAE Systems, Basildom, United Kingdom
M. Fong, Lockheed Martin Corporation, Sunnyvale, California
Chapter 4, Thermal Surface Finishes
D. G. Gilmore, The Aerospace Corporation, El Segundo, California
Chapter 1, Spacecraft Systems Overview
Chapter 6, Radiators
Chapter 7, Heaters
Chapter 15, Thermal Design Analysis
D. F. Gluck, DFG Engineering, Albuquerque, New Mexico
Chapter 8, Mountings and Interfaces
G. E. Gurevich, Microcosm Inc., El Segundo, California
D. V. Hale, Lockheed Martin, Huntsville, Alabama
B. E. Hardt, The Aerospace Corporation, El Segundo, California
Chapter 9, Louvers
B. P. Harris, United Space Alliance, Houston, Texas
A. Hashemi, Lockheed Martin, Palo Alto, California
M. J. Hoover, Lockheed Martin, Huntsville, Alabama
C. G. Justus, Computer Sciences Corporation, Huntsville, Alabama
R. D. Karam, Orbital Sciences Corporation, Dulles, Virginia
Chapter 9, Louvers
xvi
T. T. Lam, The Aerospace Corporation, El Segundo, California
K. Lankford, Starsys Research Corporation, Boulder, Colorado
Chapter 10, Heat Switches
E. I. Lin, Jet Propulsion Laboratory, California Institute of Technology, Pasadena,
California
J. C. Lyra, Jet Propulsion Laboratory, California Institute of Technology,
Chapter 7, Heaters
J. F. Maddox, Smithsonian Astrophysical Observatory, Cambridge,
Massachusetts
M. B. H. Mantelli, Federal University of Santa Catarina, Florianopolis, Brazil
Chapter 16, Thermal Contact Resistance
E. M. Mattison, Smithsonian Astrophysical Observatory, Cambridge,
Massachusetts
C. R. Miller, NASA/JSC, Houston, Texas
M. Nikitkin, Swales Aerospace, Beltsville, Maryland
T. P. O’Donnell, Jet Propulsion Laboratory, California Institute of Technology,
M. J. O’Neill, Lockheed Martin, Huntsville, Alabama
B. Patti, European Space Agency, Leiden, Netherlands
J. Pecson, Lockheed Martin, Palo Alto, California
R. C. Prager, The Aerospace Corporation, El Segundo, California
H. A. Rotter, NASA/JSC, Houston, Texas
J. G. Santiago, Stanford University, Palo Alto, California
R. Serna, NASA/JSC, Houston, Texas
xvii
W. K. Smith, The Boeing Company, Houston, Texas
T. Stevenson, University of Leicester, Leicester, United Kingdom
Chapter 13, Thermoelectric Coolers
R. Stoll, B.F. Goodrich Aerospace, Danbury, Connecticut
W. K. Stuckey, The Aerospace Corporation, El Segundo, California
J. W. Stultz, Jet Propulsion Laboratory, California Institute of Technology,
Chapter 7, Heaters
T. D. Swanson, NASA Goddard Space Flight Center, Greenbelt, Maryland
G. T. Tsuyuki, Jet Propulsion Laboratory, California Institute of Technology,
B. Turner, BAE Systems, Basildom, United Kingdom
R. F. C. Vessot, Smithsonian Astrophysical Observatory, Cambridge,
Massachusetts
K. Vollmer, Astrium, Friedrichshaten, Germany
J. W. Welch, The Aerospace Corporation, El Segundo, California
Chapter 19, Thermal Testing
Y. Yoshikawa, Lockheed Missiles and Space Company, Sunnyvale, California
M. M. Yovanovich, University of Waterloo, Waterloo, Canada
Chapter 16, Thermal Contact Resistance
xviii
1 Spacecraft Systems Overview
D. G. Gilmore*
Introduction
During the past 40 years, hundreds of spacecraft have been built in support of scientific, military, and commercial missions. Most can be broadly categorized as
either three-axis-stabilized spacecraft, spin-stabilized spacecraft, or pallets; these
types are distinguished by their configurations, internal equipment, and thermalcontrol designs. This chapter is a brief overview of the characteristics of each of
these different types of spacecraft and the missions they support. Representative
thermal designs for each type are discussed in more detail in Chapter 3.
Spacecraft Configurations
The most common spacecraft configuration today is three-axis-stabilized. This
type of spacecraft is characterized by a body that is roughly box-shaped and by
deployable solar-array panels. Examples are the Defense Meteorological Satellite
Program (DMSP), the Japanese Earth Resources Satellite (JERS), and the Russian
communications satellite Gorizont, shown in Fig. 1.1. The bodies of these spacecraft are usually kept inertially stable except for a slow rotation induced about one
axis to keep the payload antennas or sensors continuously pointed toward Earth as
the satellite orbits. The solar-array panels are then counterrotated relative to the
spacecraft body to keep them inertially fixed on the sun. Some three-axis spacecraft, such as the European Infrared Space Observatory (ISO, Fig. 1.1), have
restrictions on attitude (the vehicle’s orientation relative to an inertial coordinate
system) or low power requirements that allow them to use fixed solar arrays that
do not rotate to track the sun.
A typical internal equipment complement for a three-axis-stabilized spacecraft
is shown in the exploded view of a Fleet Satellite Communications (FLTSATCOM) satellite in Fig. 1.2. The spacecraft is commonly referred to in terms of a
“payload” and a “bus,” or “platform.” The payload is the equipment that services
the primary mission—for example, a cloud-cover camera for a weather satellite or
an infrared (IR) sensor for a missile early-warning system. Since FLTSATCOM is
a communication satellite, the payload is the communications subsystem, which
consists of the antennas on the Earth-facing side of the vehicle and the communications electronics boxes mounted in the upper hexagonal compartment, as shown
in Fig. 1.2. The bus consists of all other spacecraft subsystems that support the
payload. These subsystems typically include
• Structures subsystem: the physical structure of the spacecraft, to which all
electronics boxes, thrusters, sensors, propellant tanks, and other components
are mounted
*The Aerospace Corporation, El Segundo, California.
1
2
Spacecraft Systems Overview
DMSP
Gorizont
ISO
JERS
Fig. 1.1. Three-axis-stabilized satellites.
• Electrical power/distribution subsystem (EPS or EPDS): the equipment used to
generate and distribute electrical power to the spacecraft, including solar
arrays, batteries, solar-array controllers, power converters, electrical harnesses,
battery-charge-control electronics, and other components
• Telemetry, tracking, and command subsystem (TT&C): The electronics used to
track, monitor, and communicate with the spacecraft from the ground. TT&C
equipment generally includes receivers, transmitters, antennas, tape recorders,
and state-of-health sensors for parameters such as temperature, electrical current, voltage, propellant tank pressure, enable/disable status for various components, etc.
• Attitude/velocity control subsystem (ACS or AVCS): The devices used to sense
and control the vehicle attitude and velocity. Typical components of the ACS
system include sun and Earth sensors, star sensors (if high-precision pointing
is required), reaction or momentum wheels, Inertial Measurement Units
(IMUs), Inertial Reference Units (IRUs), and the electronics required to process signals from the above devices and control satellite attitude.
3
• Propulsion subsystem: Liquid and solid rockets or compressed-gas jets and
associated hardware used for changing satellite attitude, velocity, or spin rate.
Solid rockets are usually used for placing a satellite in its final orbit after separation from the launch vehicle. The liquid engines (along with associated
plumbing lines, valves, and tanks) may be used for attitude control and orbit
adjustments as well as final orbit insertion after launch.
• Thermal-control subsystem (TCS): The hardware used to control temperatures
of all vehicle components. Typical TCS elements include surface finishes,
insulation blankets, heaters, and refrigerators.
Many of these subsystem components are shown in the drawing of FLTSATCOM in Fig. 1.2.
The second category of spacecraft is spin-stabilized. These are less common
than the three-axis-stabilized type and have been used mostly for relatively highaltitude missions in geosynchronous or Molniya orbits (p. 9). Some spinning satellites, however, are used in low-altitude orbits. A typical “spinner,” Intelsat VI, is
shown in Fig. 1.3. As the category name implies, these satellites achieve attitude
stability by spinning like a top. Each spins at approximately 15 rpm about the axis
of a cylindrical solar array. In the case of Intelsat VI, the communications payload
is mounted on a large shelf, which is despun relative to the rest of the spacecraft so
that it points continuously at Earth.
A spinner has the same basic subsystems as a three-axis-stabilized spacecraft:
structures, EPS, TT&C, ACS, propulsion, and TCS. Usually, the payload is contained entirely on the despun section, while most of the other subsystems are on
the spinning side. Some types of spinners, however, such as the Defense Support
Program satellites (DSP; Fig. 1.4), do not have a despun shelf. In the case of DSP,
the payload, an IR telescope/sensor, spins with the rest of the satellite; the rotation
of the vehicle provides a scanning motion for the sensor.
A pallet is technically a collection of one or more payloads plus some limited
support services, such as power distribution, data recording, or telemetry sensors.
Pallets may be anything from a small experiment mounted to the side of a host
spacecraft to a large structure containing many instruments and mounted in the
payload bay of the space shuttle. The principal difference between the pallet and
other spacecraft is that the pallet is not able to function autonomously, but instead
relies on the host vehicle for ACS, EPS, and TT&C support.
The Experiment Support System (Fig. 1.5) is a typical pallet system. It consists
of a rather large structure that supports a half-dozen experiments and an equipment compartment containing power distribution, command processing, and data
recording equipment. The pallet is mounted in the space-shuttle payload bay, and
the shuttle provides ACS, EPS, and TT&C functions. In addition to the pallet
itself, there is a command monitor panel mounted in the crew compartment to
allow the astronauts to control the operation of the experiments on the pallet.
Because of the support provided by the shuttle, the pallet does not have propulsion, ACS, EPS, or TT&C subsystems, and it is incapable of operating on its own
in space.
4
27
42
45
17
47
49
34
39
38
35
52
36
32
31
41
40
30
44
3
18
14
43
48
51
28
46
50
33
29
15
53
56
8
36
37
36
2
12
7
4
6
10
26
24
11
22
1
10
55
25 20 23 21 2 5 54 9
13 15 16 19
Fig. 1.2. Exploded view of FLTSATCOM. (Legend on facing page. )
5
Legend for Fig. 1.2.
Communication System
UHF Transponder
Attitude and Velocity
Control
1.
2.
3.
4.
5.
6.
7.
8.
9.
Solar array drive assembly
Sun sensor assembly
Earth sensor assembly
Control and auxiliary electronics
Spinning Earth sensor assembly
Reaction wheel assembly
Coarse sun sensor assembly
Earth sensor electronics
Nutation damper assembly
Electrical Power/
Distribution System
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
Battery assembly
Power control unit
Converter, spacecraft equipment
Converter, communications no. 1
Converter, communications no. 2
Converter, transmitter
Payload switching unit no. 1
Payload switching unit no. 2
Solar panel assembly
Electrical integration assembly
Telemetry, Tracking, and
Command System
S-Band Command Group
20. S-band receiver
21. Decrypter KIR23 (2 required)
22. Command unit
S-Band Telemetry Group
23. S-band telemetry transmitter
24. PCM encoder
S-Band Antenna Group
25. S-band diplexer
26. RF coaxial switch
27. S-band antenna
28.
29.
30.
31.
32.
33.
34.
35.
36.
37.
38.
39.
40.
41.
42.
43.
44.
45.
46.
47.
48.
Preamp/downconverter/IF limiter no.1
IF filter limiter no. 2
Processor receiver/synthesizer
Repeater receiver
Command receiver/synthesizer
Oven-controlled crystal oscillator (2)
AF processor
UHF command decoder
UHF transmitter Navy low power
UHF transmitter Navy high power
UHF transmitter (DODWB)
UHF transmitter (AFNB)
UHF transmitter filter
UHF multicoupler filter assembly
Transmit antenna assembly
Frequency generator
Receiver filter
UHF receive antenna assembly
Signal distribution unit no. 1
Signal distribution unit no. 2
Passive hybrid
SHF Transponder
49.
50.
51.
52.
FB processor
SHF receiver
SHF transmitter
SHF antenna
Propulsion System
53.
54.
55.
56.
Propellant tank
Fill and drain valve
Thruster assembly
Apogee kick motor
6
Telemetry and command antenna
C-Band multiple
beam antennas
K-Band spot
beam antennas
Thermal radiator
Spinning structure
and electronics
Earth coverage
antennas
Orbit and attitude
control propulsion
system
Communications
equipment
Traveling wave
tube amplifiers
Solar cell arrays
Nickel-hydrogen
batteries
Fig. 1.3. Intelsat VI satellite.
Fig. 1.4. DSP satellite.
Earth Orbits
Grad
Far U.V.
7
Grad sample box
Grad electronics
CIRRIS
HUP
QINMS
URA
CIRRIS pallet
Tape recorder
2 places
Fig. 1.5. Experiment Support System.
Another spacecraft configuration worth noting here is that of upper stages.
Although they are not spacecraft per se, upper stages may be of a similar level of
complexity, and they may contain some of the same subsystems. They are
included in this handbook because upper-stage thermal control after separation
from the booster is quite similar to the thermal control of spacecraft.
Upper stages are generally used to raise a spacecraft to a higher operational
orbit from the relatively low orbit to which the booster delivers it. The duration of
their missions varies from a few hours to several days. Upper stages can use solid,
liquid, or cryogenic propellants. The Inertial Upper Stage (IUS, Fig. 1.6) is an
example of a solid-propellant upper stage that can be used in conjunction with
either the space shuttle or expendable boosters. The IUS itself has two stages; the
first is generally used to put the spacecraft into a highly elliptical transfer orbit,
and the second is fired at transfer-orbit apogee (the point in the orbit with the
greatest altitude above the planet surface) to make the orbit circular at the higher
altitudes. Like a satellite, the IUS has structures, EPS, TT&C, ACS, propulsion,
and thermal-control subsystems.
Earth Orbits
A variety of orbits are used for different types of Earth-oriented missions. The
most common orbits, in order of increasing altitude, are low Earth (LEO),
Molniya, and geosynchronous (GEO). These are drawn to scale in Fig. 1.7. The
following section briefly describes these orbits, and a more detailed discussion of
orbit parameters can be found in Chapter 2.
Orbits whose maximum altitudes are less than approximately 2000 km are generally considered low Earth orbits. They have the shortest periods, on the order of
an hour and a half. Some of these orbits are circular, while others may be somewhat elliptical. The degree of eccentricity is limited by the fact that the orbit is not
much larger than Earth, whose diameter is approximately 12,760 km (Fig. 1.7).
The inclination of these orbits, which is the angle between the plane of the equator
8
Thrust vector control
potentiometer
Spacecraft interface plane
Equipment support section
Nozzle
Solid rocket
motor
Interstage structure
Extendable exit cone
Solid rocket motor
Control
system
Reaction engine module
SRM-2 safe and ARM (2)
T34D medium-gain
antenna
STS antenna (optional)
Battery (3)
Battery (3)
Thrust vector control
Computer
Transponder
Fuel tank (optional)
Thrust vector control A
Power distribution unit
Environmental
measurement unit
Power transfer unit
Decrypter (optional)
Signal conditioning unit
STS antenna
Power amplifier
Power amplifier (optional)
Diplexer
Inertial measurement
unit
Star scanner
STS antenna
Computer
Pyro switching unit
Transponder (optional)
Fuel tank
Reaction engine mode
Encrypter (optional)
Signal interface unit
Signal conditioning unit
Spacecraft connectors
T34D interface unit
DC-DC converter
(optional)
Encrypter (optional)
STS antenna (optional)
Fuel tank
Diplexer (optional)
Fig. 1.6. Inertial upper stage.
Earth Orbits
9
Molniya
Geosynchronous
Low Earth orbit
Fig. 1.7. Orbit types.
and the plane of the orbit, can vary from 0 deg to greater than 90 deg. Inclinations
greater than 90 deg cause a satellite in LEO to orbit in a direction opposite to
Earth’s rotation. Low Earth orbits are very often given high inclinations so that the
satellite can pass over the entire surface of Earth from pole to pole as it orbits.
This coverage is important for weather and surveillance missions.
One particular type of low Earth orbit maintains the orbit plane at a nearly fixed
angle relative to the sun (Fig. 1.8). The result of this is that, on every orbit, the satellite passes over points on Earth that have the same local time, that is, the same
local sun-elevation angle. Because Earth rotates beneath the orbit, the satellite
sees a different swatch of Earth’s surface on each revolution and can cover the
N
N
etc.
Rev 1
Rev 2
Fig. 1.8. Sun-synchronous orbit.
entire globe over the course of a day. The ability to see the entire surface of Earth
at the same local sun angle is important for weather observation and for visualsurveillance missions. This type of orbit is known as sun-synchronous and is discussed in more detail in Chapter 2. Sun-synchronous orbits may be positioned so
that satellites always see points on Earth at a specific time, anywhere from local
sunrise/sunset to local noon. They are often known as “noon” or “morning” orbits.
The next higher type of common orbit is known as Molniya. These orbits are
highly elliptical (apogee 38,900 km, perigee [the point in the orbit with the lowest
altitude above the planet surface] 550 km) and highly inclined (62 deg). They provide good views of the north polar region for a large portion of the orbit
(Fig.1.9). Because the satellite travels very slowly near apogee, it has a good view
of the polar region for up to eight hours out of its 12-hour period. A constellation
of three satellites in Molniya orbits can provide continuous coverage of the northern hemisphere for missions such as communication with aircraft flying over the
polar region.
The highest common orbit type is geosynchronous. These orbits are circular and
have very low inclinations (< 10 deg). They have an altitude of 35,786 km. Their
distinguishing characteristic is a period matching Earth’s rotation, which allows a
satellite to remain over the same spot on Earth at all times. This characteristic is
valuable for a wide variety of missions, including weather observation, communication, and surveillance.
One final useful observation is that most Earth-orbiting satellites travel through
their orbits in a counterclockwise motion as seen from above the north pole. They
move in this direction to take advantage of the initial eastward velocity given to the
the satellite as a result of Earth’s rotation (approximately 1500 km/h at the Kennedy
6h
4h
2h
1h
30 min
0
Fig. 1.9. Molniya orbit.
Interplanetary Orbits
11
Space Center). To travel the orbit in the opposite direction would require the
booster to overcome the initial 1500 km/h eastward velocity before starting to
build up speed in a westerly direction. This requirement would significantly affect
booster size and allowable payload weight.
Orbits used in interplanetary missions range from simple, direct planet-to-planet
transfer orbits to complicated trajectories involving close flybys past multiple
planets on the way to a final destination. Lunar transfer orbits, such as those used
on the Apollo program (Fig. 1.10), offer direct, minimum-energy transfer to the
moon. Similar direct transfers are usually used for missions to Mars or Venus, as
shown in Fig. 1.11. Spacecraft going to the outer planets often take advantage of
gravity assists from flybys past other planets along the way. In a flyby, the spacecraft enters the gravitational field of a planet it is passing, and it achieves a net
acceleration as a result of the planet’s own velocity. This gravitational “slingshot”
effect allows for either a smaller, lower-cost launch vehicle or the accommodation
of more payload equipment mass. The Venus-Venus-Earth-Jupiter Gravity Assist
trajectory (VVEJGA) of the Cassini mission to Saturn is shown in Fig. 1.12, and
Table 1.1 summarizes the key orbital parameters for the planets of our solar system. The wide range of environments encountered in a Cassini-type trajectory can
complicate the spacecraft thermal design process; this idea is discussed in subsequent chapters.
Ascent stage
jettison (rev 34)
Trans-Earth
injection (rev 46)
LM ascent
CM/SM separation
190-km Earth
parking orbit
CM
splashdown
S-IVB restart
Translunar injection
Spacecraft
separation
S-IVB evasive
maneuver
tory
rajec
Docking
Tra
LM insertion
(17 by 83 km)
Insertion
LM powered descent
initiation (rev 14)
CSM/LM separation
Launch
jectory
nar tra
Translu
r th t
ns-Ea
Hybrid
trajectory
transfer
S-IVB impact
S-IVB residual
trajectory
propellant dump
CSM
111 km
LM landing
CSM orbit
circulation
(rev 12)
111 km
CSM LM 106
by 311 km
(1st 2 orbits)
CSM LM 13 by
106 km (12 orbits)
Fig. 1.10. Lunar transfer orbits (NASA). Spacecraft modules: CM, command module;
CSM, command-service module; LM, lunar module; SM, service module; S-IVB, Saturn IVB.
Venus
Sun
Earth
Mars
Fig. 1.11. Minimum-energy direct transfers used for missions to Mars or Venus.
Orbit of
Jupiter
Venus swingby
Apr. 21, 1998
June 20, 1999
Orbit of
Saturn
Saturn arrival
July 1, 2004
Orbit of Earth
Deep space maneuver
Dec. 2, 1998
Orbit of Venus
Earth swingby
Aug. 16, 1999
Launch
Oct. 6, 1997
Jupiter swingby
Dec. 30, 2000
Fig. 1.12. VVEJGA trajectory. (Courtesy of NASA)
13
In some interplanetary missions, aerocapture maneuvers (Fig. 1.13) are used to
slow the spacecraft and place it in orbit around a planet. This process involves sending the spacecraft close enough to the planet so that it actually passes through the
upper reaches of the planet’s atmosphere. Friction in the atmosphere slows the
vehicle to a velocity that is below the planet’s escape velocity. Injecting the spacecraft
into orbit around the planet at just the right altitude and direction is critical to
avoid its being either excessively heated or deflected back into interplanetary space.
Several orbits around the planet may be required to gradually lower the orbit altitude.
Table 1.1. Planetary Orbit Parameters
Mercury
Venus
Earth
Moon
Mars
Jupiter
Saturn
Uranus
Neptune
Pluto/Charon
Orbit Semimajor
Axis (AU)
Min. Distance
from Sun (AU)
0.3871
0.7233
1.000
1.000
1.524
5.20
9.54
19.18
30.06
39.44
0.3075
0.7184
0.9833
0.9833
1.381
4.95
9.01
18.28
29.80
29.58
Max. Distance Equatorial Radius
from Sun (AU)
(km)
0.4667
0.7282
1.0167
1.0167
1.666
5.45
10.07
20.09
30.32
49.30
Atmosphere
Circularization burn
Fig. 1.13. Aerocapture maneuvers.
2425
6070
6378
1738
3397
71,300
60,100
24,500
25,100
3200
(Pluto)
In some cases, a similar process helps minimize the use of propellant when certain
kinds of orbit changes are required during a spacecraft’s orbital mission. Aerocapture maneuvers create significant heat loads that must be addressed in the thermal
design process.
Some rather unique orbits rely on balances between centrifugal and gravitational forces among multiple bodies. The Italian-French mathematician Josef
Lagrange discovered that in cases where one body orbits around a much larger
one, such as the moon around Earth or Earth around the sun, the centrifugal force
and the two gravitational forces balance each other at five points. A body located
precisely at any of these points will therefore remain there unless perturbed. These
points, known as the Lagrange points, are designated L1 through L5, as shown in
Fig. 1.14. L1, L2, and L3 are so unstable that, for a body positioned at any of
them, a slight perturbation can knock the body out of equilibrium and send it on
its way. The other two points, L4 and L5, are stable enough for a body positioned
at either one to return to equilibrium if perturbed. For the unstable Lagrange
points, a spacecraft can be placed in a small, fairly stable orbit around the point
that requires little in the way of corrective maintenance maneuvers. The Solar and
Heliospheric Observatory (SOHO) is placed at the Earth-sun L1 point; the Microwave Anisotropy Probe (MAP) satellite and the Next Generation Space Telescope
are considering the Earth-sun L2 point as a possible home.
L4
L1
L3
L5
Fig. 1.14. Lagrange points.
L2
Missions
15
Missions
A wide variety of missions are supported by the three general types of spacecraft
platforms discussed earlier. The type of mission will dictate the orbit, the payload,
and, in some cases, the platform. Typical missions include communication, scientific observation, weather monitoring, navigation, remote sensing, surveillance,
and data relay. This section briefly describes each of these missions.
The most common mission for both commercial and military satellites is communication; there are currently 294 operating communication satellites in orbit.
Thuraya and Singapore Telecom-1 (ST-1, Fig. 1.15) are commercial communication satellites. “Comsats” relay radio, telephone, television, or data signals from
one point on Earth to another. These satellites are usually, but not always, in highaltitude geosynchronous orbits, where they remain over the same point on Earth at
all times. Communication can be provided between any two points on the side of
Earth to which the satellite has a direct view. Communication between two points
Thuraya
Iridium
Astrium
Fig. 1.15. Comsats.
on opposite sides of Earth, however, requires the use of multiple satellites with
crosslinks between them. Both Thuraya and ST-1 are typical communication satellites that do not have crosslink capability. Iridium (Fig. 1.15) is a satellite constellation that has crosslinks and is able to provide communication between any
two points on Earth.
Weather monitoring is another mission common to civilian and military space
programs. The DMSP spacecraft (Fig. 1.16) is a typical low-altitude weather satellite. It carries visual and IR cameras that continuously photograph cloud patterns, as well as secondary sensors, such as Special Sensor Microwave Imager/
Sounder (SSMIS), that can monitor phenomena such as surface wind speeds, soil
moisture content, and precipitation rates. Low-altitude weather satellites are usually in sun-synchronous orbits. This allows them to scan the entire surface of
Earth at the same local sun angle over the course of a day. High-altitude weather
satellites, such as NASA’s GOES (Geostationary Operational Environmental Satellite, Fig. 1.16), are usually in geosynchronous orbits that allow them to continuously photograph one entire hemisphere of Earth.
Navigation constitutes a third type of spacecraft mission. For the United States,
this mission is currently fulfilled by one satellite program, NAVSTAR-GPS (Global Positioning System). The GPS system includes a constellation of 24 satellites
in 12-hour circular orbits. Each GPS satellite (Fig. 1.17) continuously broadcasts
a signal that can be picked up by small receivers on the ground, in aircraft, or even
DMSP
Unyawed
velocity
vector
+YG
–XG
Apparent
sun motion
95°
Solar
array
Celestial
sensor
Sun
shade
0° +Z
G
–ZG
GOES
Precision
mounting
platform Earth
sensor
assembly
+XG
Sun
sensor
unit
Earth
center
Fig. 1.16. Weather satellites.
Thruster
Propulsion
module
Equipment
support
module
–YG
Missions
GPS
17
GLONASS
Fig. 1.17. Positioning satellites.
in another satellite. If three or more GPS satellites are visible at any one time,
the receiver can determine its own position and velocity to within 1 m and 0.1 m/
sec. Russia also operates a system of positioning satellites, known as GLONASS
(Global Navigation Satellite System, Fig. 1.17), that are located in similar orbits.
A next-generation navigation satellite program, aptly named Galileo, is also currently planned by the European Space Agency.
Surveillance is a general category for satellites whose mission is to monitor various activities on Earth. This surveillance can be in the form of IR sensors to
detect missile launches, radar to track aircraft or ships, visual observation of
ground activities, or intercept of radio transmissions. Satellites designed to support each of these different missions have markedly different configurations.
Space Imaging’s Ikonos (Fig. 1.18) is a commercial optical-surveillance satellite.
It provides 1-m panchromatic and 4-m color resolution digital imagery of Earth’s
Helios
Ikonos
Fig. 1.18. Surveillance satellites.
surface. Its photos are used for mapping, urban planning, and environmental assessment. Helios (Fig. 1.18) is a national optical-surveillance satellite operated by France.
The DSP spacecraft shown in Fig. 1.4 is an example of an IR surveillance satellite. The payload is an IR telescope that detects and tracks missiles by the heat
emitted from their rocket plumes. The detectors in the telescope are cooled to
approximately 150 K by a cryogenic radiator with a helium-coolant loop. The
entire satellite rotates at 6 rpm to provide a scanning motion that sweeps the linear
detector array across Earth’s surface. Ground software reconstructs the sweep into
an Earth image with all heat sources displayed. DSP provides the United States
with its first warning of missile launches.
Space Based Radar (SBR, Fig. 1.19) is an example of a radar-surveillance satellite. Spacecraft proposed for this program are quite large, with antenna dimensions on the order of 30 m. They would be developed to track aircraft and ships,
with some designs being proposed to track missiles and individual warheads for
defense applications. Radarsat, a remote-sensing satellite program led by the
Canadian Space Agency, is also shown in Fig. 1.19.
Relay satellites support another type of mission similar to that of communication satellites except that the communication link is between the ground and a second satellite (Fig. 1.20). Such links eliminate the need for ground stations spaced
throughout the world, and they provide continuous contact with satellites in any
orbit. An example of a relay satellite is NASA’s Tracking and Data Relay Satellite
System (TDRSS), shown in Fig. 1.20. TDRSS is used to provide ground-toground and ground-to-satellite links and to communicate with shuttle astronauts.
Most Earth-orbiting scientific satellites need go no higher than low Earth orbit
to accomplish their missions. Astronomical satellites, such as the Earth Observing
Dual rotation
solar array
Solar array
Component
enclosure
Flight
path
Phased
array
Bus module
SAR
antenna
MLI blanket
Payload module
SBR concept
Solar array
Radarsat
Fig. 1.19. Radar satellites.
Missions
19
Fig. 1.20. TDRSS relay.
System (EOS) and the Hubble Space Telescope (Fig. 1.21), need only get above
Earth’s atmosphere to conduct their observations. A low-altitude orbit is an advantage for programs like EOS, whose mission is to study Earth. Some missions, like
the Russian Granat X-ray and gamma-ray observatory (Fig. 1.21), do require
high-altitude Earth orbits. There are also, of course, missions that require interplanetary scientific spacecraft to leave Earth’s orbit entirely. These programs, such
as Cassini (Fig. 1.21), sometimes must follow complicated trajectories through the
solar system to get to their final destination.
EOS
Hubble Space Telescope
Granat
Cassini
Fig. 1.21. Scientific satellites.
Remote-sensing missions are accomplished by satellites such as the U.S. Landsat, the French SPOT (Systeme Pour l’Observation de la Terre), and the European
ERS (Earth Resources Satellite) (Fig. 1.22). These vehicles gather images in a
variety of wavelengths. This information is used to manage crops and other Earth
resources and to support environmental and global change research. For this kind
of mission, the satellites are usually placed in sun-synchronous polar orbits at an
altitude of approximately 830 km.
Landsat 7
Coarse
sun sensors
ETM+
Instrument
aperture
Cooler
door
(open)
Full
aperture
calibrator
+Z
Solar array
(canted 20° from +Z axis)
Earth sensor
assembly
Gimballed
X-band
antennas
+Y
–Velocity
+X
Nadir
SPOT
ERS
Fig. 1.22. Remote-sensing satellites.

Spacecraft Thermal Control Handbook

Transcription

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