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I 


JSC-10633 
REV. A 


ATMOSPHERIC SCIENCE FACILITY 


PALLET-ONLY MODE 


SPACE TRANSPORTATION SYSTEM PAYLOAD 


(FEASIBILITY STUDY) 


INTERIM TECHNICAL REPORT 


VOLUME 1 



EXPERIMENT SYSTEMS DIVISION 


5 






National Aeronautics and Space Administration 

LYNDON B. JOHNSON SPACE CENTER 

Houston^ Texas 

November 1 975 


(NASA-TM-X-7295e) ATPOSPHZHIC SCIENCi N76-24319 

FACILITY PALLET-CNLY ECDS SPACE 
TRANSPOETAIION SYSTEM PAYLOAD (FEASIfclLITY 

STUDY) , VOLUME 1 Interim Technical Report. Onclas 

(NASA) 493 p HC $‘»2.50 CSCL 22E G3/16 18414 


FOREWORD 

•* 

This study presents the interim results of an on-going evalua- 
tion study at the Johnson Space Center (JSC) of the pallet- 
only mode for Atmospheric, Magnetosphefic, and PTasmas-i n-Space 
(AMPS) payloads on the Space Transportation System (STS). The 
complete study is to address payload configurations for all 
AMPS disciplines, and also for payloads configured for selected 
disciplines. The configurations discussed in this interim re- 
port include the provisions for selected atmospheric science 
missions. The AMPS payload is being designed to conduct experi 
ments in areas of global remote sensing of the atmospheve, the 
ionosphere and the magnetosphere. Active perturbation by 
laser emissions, chemical reactions, or gas releases of the 
stratosphere, ionosphere, and magnetosphere will answer funda- 
mental scientific questions. Such answers are considered keys 
to a better understandi ng of man ' s total natural environment, 
his effects upon it, and its effects upon him. 

In the summer of 1973, the Space Science Board of the National 
Academy of Science convened a study at Woods Hole, Massachusett 
to explore the scientific uses of the Orbiter “/hich is a part 
of the STS. In this study, the discipline groups were asked to 
describe the scientific objectives of their respective disci- 
plines, to identify experiments or instruments that are both 
scientifically desirable and suitable for Orbiter operations, 
to determine v/hich mode of Orbiter use would be best suited to 
the operation of these instruments, to. outline a mission model 
and to make recommendations concerning their science and the 
Orbi ter . 


Discussion in the Woods Hole study report indicated that in the 
attempt to choose the best-rsui ted Orbiter mode, almost all dis- 
cipline groups were limited by the lack of detailed information 
being available at that time, which would enable them to deter- 
mine cost-eff ecti veness for the various modes . 


Although the presence of scientists in a pressurized Spacelab 
module appeared to be the preferred mode of operation for AMPS 
experiments which require manned control in real time, based on 
real time analysis of observational data, there were also experi- 
ments to be considered which might preclude the presence of the 
habitable Spacelab module and would require monitoring and con- 
trol from the Orbiter cabin (pallet-only mode). 

One of the recommendations by the Atmospheric and Space Physics 
Group in the Woods Hole study report indicated the need for the 
National Aeronautics and Space Administration (NASA) to study 
the relative merits of the two modes of operation: (1) press- 

urized habitable module with pallets; and (2) the pallet-only 
mode without the habitable Spacelab module. The recommended 
study v/as to include: (1) scientific payload weight, cost, 

available data rate, and (2) system coverage using active 
experiment control based on real time data evaluations. The 
current on-going study at JSC is partly the indirect result 
of this recommendation. 

The Woods Hole study report also noted that suitable instrumenta- 
tion for the AMPS payload could be available for the Orbiter 
missions if a program was to be started right away in certain 
crucial areas, and recommended that scientists should be selected 
as soon as possible to participate in the detailed scientific 
definition and development of the planned programs and instrumen- 
tation . 

The 49-member AMPS Science Definition Working Group (herein 
referred to as AMPS SDWG) was formed by NASA in the summer of 
1974. Definition and development studies by the working group, 
and also those being monitored by the working group, are being 
conducted in parallel with the definition and development of 
associated Orbiter systems because of the long lead time required 
for some of these, and so that AMPS payloads can be Included on 
early STS operational missions. 


The on-going AMPS study at JSC was begun in the Fall of 1974 to; 

(1) support the AMPS SDWG (2) to perfcrm an evaluation of the 
potential of the pellet-only mode for AMPS payloads, and (3) to 
define details of an atmospheric science payload configuration 
for the AMPS program. 

The JSC study required the establishment of a baseline which 
could be used for analysis for the evaluation of the pallet-only 
mode by specialists in the areas of systems hardware and soft- 
ware, instrumentation and sensors, data handling and processing, 
mission planning, crew procedures and timeline development, etc. 

It was necessary that the baseline be developed on criteria from 
the AMPS SDWG and include science objectives, candidate experi- 
ments, mission models, sensors and instrument definitions, and 
payload systems configurations. These criteria were represen- 
tative of, and compatible with, the plans and concepts being 
developed by the AMPS SDWG scientists and by the Orbiter and 
Spacelab designers. This interim report includes the results 
of that part of the study which is establishing this represen- 
tative configuration baseline. 

The JSC study team has worked with members of the AMPS SDWG, 
Marshall Space Flight Center (MSFC), Goddard Space Flight Center 
(GSFC) AMPS personnel, with JSC personnel who are associated with 
the definition and development of the Orbiter, and with other 
Orbiter and Spacelab related activities so that adjustments can 
be considered as the definition and development of each progresses. 

The cutoff date for technical input documentation to this In- 
terim Report was June 1 , 1975. Changes are to be expected 
during the course of the study which may affect the presenta- 
tions in subsequent reports. 


ii-3 


ACKNOWLEDGMENTS 


During the study to evaluate the feasibility of a Spacelab pallet- 
only configuration with which the AMPS Experiments Program could 
be conducted from the Orbiter, considerable background data were 
required. Numerous individuals in the scientific disciplines 
very generously answered the inquiries made of them and, in many 
instances, offered pertinent suggestions and provided detailed 
documentation on the program, the experiments, and the instruments 
i nvol ved . 

The members of the JSC AMPS feasibility study project wish to 
acknowledge the following individuals who provided the answers 
to many basic questions. 

H. R. Anderson, Rice University - Vector Magnetometer 
R. F. Benson, NASA/6SFC — Radio Frequency Sounder 

E. Brunner, University of Colorado - Normal Incidence Spectro- 
graph and Echelle Spectrograph 

G. R. Carignan, University of Michigan — Neutral Mass Spectrometer 

G. R. Carruthers, Naval Research Laboratory — Airglow Spectro- 
graph, UV Occultation Spectrograph, and Far UV TV System 

C. R. Chappell, NASA/MSFC - AMPS Program 

P. Cloutier, Rice University — Vector Magnetometer 

F. W. Crawford, Stanford University — High Frequency Quadrupole 
Probe 

T. N, Davis, University of Alaska — Optical Band Imager and 
Photometer System 

d. C. Gille, National Center for Atmospheric Research — Ultra- 
violet Spectrophotometer, and Limb Scanning Infrared Radiometer 

P. D. Feldman, Johns Hopkins University — Photoelectron/Secondary 
Electron Spectrometer, and UV-VIS-NIR Spectrometer/ Photometer 


M. D. Grossi, Smithsonian Astrophysi cal Observatory - ULF 
Antenna and Transmitter, and Doppler-Tracking Bistatic Sounder 

D. N. Heath, NASA/GSFC - XUV-UV Solar Intensity Monitor 

J. P. Heppner, NASA/GSFC — DC Electric Field 

J. R. Hickey, Eppley Laboratory — Eclectic Py rhel i ometer 

H. Hinteregger, Air Force Cambridge Research Laboratory - Solar 
EUV Spectrophotometer 

N. Kawashima, University of Tokyo ~ TAPAC and SEPAC Particle 
Accelerators 

H. C. Koons, The Aerospace Corpor .tion — Resonance Cone 
Techniques 

R. W. Kreplin, Naval Research Laboratory — X-Ray Flux Monitors 

A. L. Lane, Jet Propulsion Laboratory - Gas Release Module 

H. B. Liemohn, Batelle Memorial Institute — Triaxial Fluxgate 

E. Mangold, EPA Denver — 1972 Martin Report on ASF 

R. W. McEntire, Johns Hopkins University — Energetic Ion Mass 
Analyzer 

F. W. Mozer, University of California at Berkley — Incoherent 
Scatter Radar 

A. F. Nagy, Johns Hopkins University — Fabry-Perot Interferometer 

T. Obayashi, University of Tokyo — TAPAC and SEPAC Particle 
Accelerators 

D. L. Reasoner, NASA/MSFC - Ion Mass and Distribution Analyzer 

U. Samir, University of Michigan — Langmuir Probe 

O. Samson, University of Nebraska at Omaha — Flovn'ng Chamber 
Counter for the XUV-UV Solar Intensity Monitor 

J. M. Sellen, TRW Systems Group — Particle Accelerators and 
Diagnostics 

N. W. Spencer, NASA/GSFC — Neutral Temperature and Wind Drift 
Spectrometer 

N. P. Thekaekara, NASA/GSFC — Pyrhel iometer/Spectrophotometer 


CONTENTS 


Section Page 

1.0 INTRODUCTION 1-1 

1.1 AUTHORIZATION 1-1 

1.2 STUDY OBJECTIVES . 1-1 

1.2.1 ASSESSMENT : 1-1 

1.2.2 IDENTIFICATION 1-2 

1.2.3 DEFINITION 1-2 

1.2.4 PREPARATION. 1-2 

1.3 END PRODUCTS 1-2 

2.0 STUDY APPROACH 2-1 

2.1 NATURE AND SCOPE OF THE STUDY 2-1 

2.2 TECHNICAL APPROACH 2-1 

2.3 STUDY BASELINES 2-3 

2.3.1 JOHNSON SPACE CENTER ORGANIZATIONAL 

RESPONSIBILITIES 2-3 

2.3.2 DEFINITIONS 2-4 

2. 3. 2.1 Experiment 2-4 

2. 3. 2. 2 Prime Instrument 2-4 

2.3.2. 3 Substitute Instrument 2-4 

2. 3. 2. 4 Alternate Instrument 2-4 

2. 3. 2. 5 Core Instruments 2-5 

2.4 REFERENCE DOCUMENTATION AND BIBLIOGRAPHY . ... 2-5 

2.4.1 AUTHORITATIVE DOCUMENTS 2-5 

2.4.2 REFERENCE DOCL'^^ENTS 2-5 

2.4.3 INFORMATION DOCUMENTS 2-6 


iv-1 


Section Page 

2.5 ASSUMPTIONS 2-6 

3.0 SUMMARY 3-1 

3.1 GENERAL 3-1 

3.2 CONCLUSIONS 3-1 

3.2.1 FEASIBILITY 3-2 

3.2.2 MAJOR UNRESOLVED ISSUES 3-3 

3.2.3 TECHNICAL FOLLOW-UP REQUIREMENTS .... 3-4 

3.2.4 TRADE-OFF CONSIDERATIONS 3-5 

3. 2. 4.1 Scientific . . ' 3-5 

3. 2. 4. 2 Technical 3-6 

3. 2. 4. 3 Programmatic 3-6 

3.3 RECOMMENDATIONS 3-6 

3.3.1 ASF PAYLOAD SYSTEM DESIGN 3-6 

3.3.2 FOLLOW-ON STUDIES 3-6 

3. 3. 2.1 Sci enti f 1 c 3-8 

3.3. 2. 2 Technical 3-8 

3.4 COST 3-9 

4.0 SCIENTIFIC REQUIREMENTS 4-1 

4.1 ORBITER PAYLOAD ....... 4-1 

4.1 .1 INTRODUCTION 4-1 

4.1.2 OBJECTIVES 4-4 

4.1.3 EXPERIMENTS 4-5 

4.1.3 1 Background 4-5 

4. 1.3. 2 Experiment Summaries 4-5 

4.1.4 INSTRUMENTS (PRIME). . 4-18 


Sect! on 


Page 

4. 1.4.1 General 4-18 

4. 1.4. 2 Summary Descriptions 4-23 

4.1.5 OVERVIEW, ASF MISSION TIMELINE 4-30 

4.2 PARTICLE DETECTOR SUBSATELLITE (PDS) 

REQUIREMENTS 4-32 

4.2.1 INTRODUCTION 4-32 

4.2.2 REQUIREMENTS 4-32 

4.2.3 INSTRUMENT SUMMARY DESCRIPTIONS. .... 4-32 

4.3 SOLAR PHYSICS SATELLITE 4-36 

4.3.1 ASF SUPPORT INSTRUMENTS 4-36 

4.3.2 INTERFACES 4-36 

5.0 ASF SYSTEM DESCRIPTION AND INTEGRATION 5.1-1 

5.1 GENERAL 5.1-1 

5.1.1 ASF SYSTEM ELEMENTS . 5.1-1 

5.1.2 ASF SYSTEM CONFIGURATION 5.1-3 

5.1.3 ASF SYSTEM INTERFACES 5.1-7 

5.1.4 MAJOR SYSTEMS INTEGRATIONS ISSUES. ... 5.1-18 

5.2 FLIGHT SYSTEMS 5.2. 1-1 

5.2.1 THERMAL, STRUCTURAL AND MECHANICAL 

SUBSYSTEM (TSMS) 5.2. 1-1 

5. 2. 1.1 Introduction 5. 2, 1-1 

5.2.1 .2 Requi rements 5. 2.1-1 

5. 2. 1.3 Guidelines and Assumptions . . . 5. 2. 1-5 

5. 2. 1.4 Capabilities and Constraints . . 5. 2. 1-6 

5. 2. 1.5 Subsystems Description 5. 2. 1-9 

5. 2. 1.6 Analyses and Trade Studies. . . 5.2.1-39 


1v-3 


i 


Section Page 

5. 2. 1.7 Conclusions and 

Recommendati ons 5.2.1-51 

5.2.2 ELECTRICAL POWER AND DISTRIBUTION 

■SUBSYSTEM (EPDS) 5. 2. 2-1 

5. 2. 2.1 Introduction 5. 2. 2-1 

5. 2. 2. 2 Requirements 5. 2. 2-1 

5. 2. 2. 3 ASF Timelines and Power 

Usage 5.2 .2-4 

5.2. 2.4 Guidelines and Assumptions . , . 5. 2. 2 -7 

5.2.2. 5 Capabilities and 

Constrai nts 5. 2. 2-7 

5. 2. 2. 6 Subsystem Description ..... 5.2.2-12 

5. 2. 2. 7 Analyses 5.2.2-18 

5. 2. 2. 8 Conclusions and 

Recommendations . 5.2.2-24 

5.2.3 POINTING/CONTROL AND STABILIZATION 

SUBSYSTEM (PCSS) 5. 2. 3-1 

5. 2. 3.1 Introduction 5. 2. 3-1 

5.2.3. 2 Requirements 5. 2. 3-1 

5. 2. 3. 3 Guidelines and Assumptions . . . 5. 2. 3-3 

5 . 2 . 3 . 4 Capabilities and 

Constrai nts 5. 2. 3-6 

5. 2. 3. 5 Subsystem Description 5. 2. 3-9 

5.2. 3. 6 Analyses 5.2.3-39 

5 . 2 . 3 . 7 Conclusions and 

Recommendations . 5.2.3-40 

5.2.4 COMMAND AND DATA MANAGEMENT SUBSYSTEM 

{CDMS) 5. 2. 4-1 

5. 2. 4.1 Introduction . ......... 5. 2. 4-1 

5. 2. 4. 2 Requi rements 5. 2. 4-1 


iv-4 


Secti on 


Page 

5.2.4. 3 Guidelines and Assumptions . . . 5. 2. 4-2 

5. 2. 4. 4 Additional General 

Assumpti ons 5. 2. 4-3 

5. 2. 4. 5 Capabilities and 

Constraints 5. 2. 4-4 

5.2.4. 6 Subsystem Description 5. 2. 4-5 

5. 2.4.7 Analyses 5.2.4-36 

5. 2. 4. 8 Conclusions and 

Recommendati ons 5.2.4-48 

5.2.5 AFT CREW STATION CONFIGURATION 5. 2. 5-1 

5. 2. 5.1 Introducti on 5. 2. 5-1 

5. 2. 5. 2 Requirements 5. 2. 5-2 

5. 2. 5. 3 Guidelines and Assumptions . . . 5. 2. 5-3 

5. 2. 5. 4 Capabilities and 

Constraints 5. 2. 5-4 

5. 2.5.5 Aft Crew Station 

Configuration Description . . . 5. 2. 5 -9 

5. 2. 5. 6 Analyses and Trade Studies . . . 5.2.5-27 

5. 2. 5. 7 Conclusions and 

Recommendations ... 5.2.5-32 

5.2.6 PARTICLE DETECTOR SUBSATELLITE (PDS) . . 5.2. 6-1 

5. 2. 6.1 Introducti on 5. 2. 6-1 

5 . 2 . 6 . 2 Requirements . . 5 . 2 . 6 - I 

5. 2. 6. 3 Guidelines and Assumptions . . . 5. 2. 6-2 

5. 2. 6. 4 Capabilities and 

Constraints . 5. 2. 5-2 

5. 2. 6. 5 System Description 5. 2. 6-2 

5. 2. 6. 6 Analyses . . .... . . . . . . 5. 2. 6 -8 

5. 2. 6. 7 Conclusions and 

Recommendati ons 5.2.6-10 


i v-5 


Section Page 

5.3 GROUND SYSTEM 5.3-1 

5.3.1 GROUND PROCESSING OF FLIGHT DATA .... 5.3-1 

5. 3. 1.1 Introducti on 5.3-1 

5. 3. 1.2 Considerations 5.3-1 

5.3.2 GROUND SUPPORT, TEST AND CHECKOUT 

SUBSYSTEM 5.3-3 

5. 3.2.1 General 5.3-3 

5. 3. 2. 2 Requi rements 5.3-3 

5.3. 2.3 Guidelines and Assumptions . . . 5.3-3 

5.3. 2.4 Subsystem Description 5.3-4 

5. 3. 2. 5 Logistics and Transportation . . 5.3-25 

5.4 SUPPORT SYSTEMS . 5.4-1 

5.4.1 ORBITER.. 5.4-1 

5. 4. 1.1 Payload Placement . ...... 5.4-1 

5. 4.1. 2 Orbit Changes , 5.4-2 

5. 4. 1.3 Attitude Control. Maneuvering 

and pointing 5,4-2 

5. 4. 1.4 Communi cati ons 5.4-2 

5. 4. 1.5 Tracki ng 5.4-3 

5. 4. 1.6 Data/Gommand Interface and 

Processing 5.4-3 

5. 4. 1.7 Displays and Controls 5.4-4 

5. 4. 1.8 Remote Manipulator System . . . 5.4-4 

5.4.1 .9 Electrical Power . . . . , . . 5.4-5 

5.1.1.10 Structural /Mechani cal . . . . . 5.4-5 

5.1.1.11 Thermal Control . ....... 5 .4-6 

5.4.2 TRACKING AND DATA RELAY SATELLITE 

SYSTEM ..... 5.4-7 


Section 


! 


\ 


Page 

5.4.3 SOLAR PHYSICS SATELLITE (SPS) 5.4-7 

5.4.4 SPACE TRANSPORTATION SYSTEM (STS) 

GROUND FACILITIES 5.4-9 

5.5 CONTAMINATION . 5.5-1 

5.5.1 INTRODUCTION 5.5-1 

5.5.2 STUDY APPROACH 5.5-1 

5.5.3 CONCLUSIONS 5.5-2 

5.5.4 RECOMMENDATIONS 5.5-2 

5. 5. 4.1 EMI Containi nation 5.5-2 

5. 5. 4. 2 Dust, 6as, Particulate 

Contamination 5.5-3 

5.6 STANDARDIZATION 5.6-1 

5.6.1 SUMMARY 5.6-1 

5.6.2 CONCLUSIONS 5.6-2 

6.0 MISSION OPERATIONS 6-1 

6.1 INTRODUCTION 6-1 

6.2 GUIDELINES AND ASSUMPTIONS 6-1 

6.3 OPERATIONS DESCRIPTION 6-4 

6.3.1 PREFLIGHT OPERATIONS 6-4 

6.3.2 FLIGHT OPERATIONS 6-4 

6.3.3 POSTFLIGHT OPERATIONS. 6-5 

6.4 PREFLIGHT OPERATIONS 6-5 


6.4.1 PALLET LEVEL INTEGRATION AND TEST. . . . 6-5 

6.4.2 PAYLOAD LEVEL INTEGRATION AND TEST ... 6-6 

6.4.3 VEHICLE INTEGRATION, TEST AND LAUNCH 

PREPARATION. . 6-7 


iv-7 


Section Page 

6.5 FLIGHT OPERATIONS 6-9 

6.5.1 LAUNCH AND MISSION ORBIT INJECTION ... 6-9 

6.5.2 ON-ORBIT OPERATIONS 6-9 

6. 5.2.1 Payload Preparation 

(Revolutions 1 through 10) . . . 6-10 

6. 5. 2. 2 PDS Deployment (Revolutions 

1 1 through 15) . . ! ^ ! i T . . 6-11 

6. 5. 2. 3 Experiment Operations 

(Revolutions 16 through 80) . . 6-13 

6.6 RE-ENTRY. DESCENT AND LANDING 6-35 

6.7 POST-LANDING 6-35 

7.0 ASF SYSTEM DEVELOPMENT STATUS AND REQUIREMENTS ... 7-1 

7.1 INTRODUCTION 7-1 

7.2 INSTRUMENTS 7-3 

7.2.1 PRIME 7-3 

7. 2. 1.1 Technical Considerations . ... 7-4 

7. 2. 1.2 Devel opment Schedule 

Requ i rements 7-7 

7 . 2 . 1 . 3 Conclusions 7-9 

7.2.2 SUBSTITUTE INSTRUMENT CONSIDERATIONS . . 7-11 

7. 2. 2.1 Introduction 7-11 

7. 2. 2. 2 Technical Considerations . ... 7-12 

7. 2. 2. 3 Development Requirements . . . . 7-12 

7. 2. 2. 4 Cone! usi ons 7-22 

7.2.3 SUPPORT SUBSYSTEMS 7-22 

7.2.4 GROUND SUPPORT AND TEST EQUIPMENT. . . . 7-25 

7.2.5 GROUND DATA HANDLING AND PROCESSING. . . 7-29 


i V-8 





1 


Section Page 

8.0 CONCLUSIONS 8-1 

8.1 INTRODUCTION 8-1 

8.1.1 TECHNICALLY FEASIBLE 8-3 

8.1.2 SCIENTIFICALLY FEASIBLE 8-3 

8.1.3 PROGRAMMATICALLY FEASIBLE 8-3 

8.2 FEASIBILITY ASSESSMENT 8-3 

8.2.1 SCIENTIFIC OBJECTIVES FULLFILLMENT ... 8-3 

8. 2. 1.1 Problems 8-5 

8. 2. 1.2 Impacts 8-7 

8.2.2 SUPPORT SUBSYSTEMS 8-7 

8. 2. 2.1 Concl us1 ons 8-7 

8. 2. 2. 2 Assessment 8-10 

8.2.3 PROGRAMMATIC FACTORS 8-19 

8. 2. 3.1 Schedule 8-19 

8.2.3. 2 Costs 8-19 

8.3 TRADE-OFF CONSIDERATIONS 8-19 

8.3.1 SCIENTIFIC 8-19 

8.3.2 TECHNICAL 8-21 

a. 3. 2.1 Approach 8-21 

8. 3. 2. 2 Structural . .... 8-22 

8. 3.2. 3 Thermal Control System 8-22 

8. 3. 2.4 Remote vs Direct Access 

Circuit Breakers . . 8-23 

8. 3. 2. 5 High Current Transmission 

Media 8-23 

8. 3. 2.6 AMS 8-23 


8. 3. 2. 7 Payload Specialist Work 

Station 8-24 

8 . 3 . 2 . 8 Instrument Sequence 

Initiation 8-24 

8. 3. 2. 9 Data Processing 8-24 

8.3.2.10 Mass Memory Operational 

Programs 8-24 

8.3.2.11 Data Compression 8-24 

8.3.2.12 Computer, Processor 8-25 

8.3.2.13 Subsatellite Retrieval .... 8-25 

8.3.2.14 Orbiter and Payload EMI 

Envi ronment 8-25 

8.3.2.15 Support Subsystem Equipment 

Trade-Off 8-26 

8.3.3 PROGRAMMATICS 8-27 

8.4 TECHNICAL FOLLOW-UP REQUIREMENTS 8-28 

8.5 UNRESOLVED MAJOR ISSUES ..... 8-30 

9.0 RECOMMENDATIONS 9-1 

9.1 ASF PAYLOAD SYSTEM DESIGN 9-1 

9.1.1 TDRS 9-1 

9.1.2 EPDS 9-1 

9.1.3 PCS 9-2 

9.1.4 CDMS 9-2 

9.1.5 AFT CREH STATION SUPPORT 9-3 

9.1.6 PDS 9-3 

9.2 FOLLOW-ON STUDY ... ..... 9-3 

9.2.1 UNRESOLVED MAJOR ISSUES 9-4 


Section Page 

9.2.2 ASF PAYLOAD SYSTEM DESIGN 9-6 

9. 2. 2.1 TSMS 9-6 

9. 2. 2. 2 EPPS 9-7 

9. 2. 2. 3 PCSS 9-8 

9. 2. 2. 4 CDMS 9-8 

9. 2. 2. 5 Aft Crew Station Support . . . . 9-10 

9. 2.2.6 m 9-10 

9.2.3 CONCEPTS OF STANDARDIZING. 9-10 

10.0 BIBLIOGRAPHY 10-1 

10.1 AUTHORITATIVE DOCUMENTS 10-1 

10.2 REFERENCE DOCUMENTS 10-2 

10.3 INFORMATION DOCUMENTS. 10-3 


Appendices ^ Page 

A ASF EXPERIMENT DESCRIPTIONS A-i 

B INSTRUMENT DESCRIPTIONS B-i 

C.l ORBITER ENVIRONMENTS ■ . . C.l-i 

C.2 ENVIRONMENTAL ANALYSES C.2-1 

,C.3 INTERFERENCE PROBLEMS FOR AMPS INSTRUMENTS C.3-i 

C.4 SPACE SHUTTLE PROGRAM OFFICE CORRESPONDENCE C.4-i 

D CENTRALIZING AND STANDARDIZING INSTRUMENT 

SUBSYSTEMS FOR THE ATMOSPHERIC, MAGNETOSPHERIC 

AND PLASMAS-IN-SPACE (AMPS) D-i 

E ASF PAYLOAD LOGISTICS AND TRANSPORTATION E-i 


iv-11 


TABLES 


Table Page 

3. 2. 4- 1 CANDIDATE SUBSYSTEM OPTIONS 3-7 

4. 1.4- 1 ASF INSTRUMENT IDENTIFICATION 4-19 

4. 1.4- 2 INSTRUMENTS ASSIGNED TO EXPERIMENTS 4-20 

4. 1.4- 3 ASF PRIME INSTRUMENT PERFORMANCE PARAMETERS . . 4-21 

4. 1.4- 4 ASF INSTRUMENT INTERFACE PARAMETERS 4-22 

4. 2. 2- 1 INTERFACE PARAMETERS 4-34 

4. 2. 2- 2 PDS PERFORMANCE PARAMETERS 4-35 

5. 1.1- 1 ASF PALLET INSTRUMENT CHARACTERISTICS 5.1-11 

5. 1.1- 2 ASF PARTICLE DETECTOR SUBSATELLITE SYSTEM 

CHARACTERISTICS 5.1-12 

5. 1.1- 3 ASF SUPPORT SUBSYSTEM CHARACTERISTICS ..... 5.1-13 

5. 2. 1- 1 AIM PACKAGING PARAMETERS FOR ASF 5. 2. 1-4 

5. 2. 1- 2 ATCS CONTROL CAPABILITY 5. 2. 1-7 

5. 2. 1- 3 SCIENTIFIC INSTRUMENT LOCATIONS 5.2.1-20 

5. 2. 1- 4 BOOM CANDIDATE MATERIALS • 5.2.1-42 

5. 2. 1- 5 SEPARATION AND EJECTION DEVICES 5.2.1-45 

5. 2.1- 6 WEIGHT AND BALANCE 5.2.1-49 

5. 2. 1- 7 WEIGHT AND BALANCE, PAYLOAD ON PALLETS AND 

PAYLOAD CHARGEABLE SUPPORT HARDWARE 5.2.1-50 

5. 2. 2- 1 ASF INSTRUMENT/SUBSATELLITE POWER 5.2. 2-3 

5. 2. 2- 2 ASF SUPPORT SYSTEM POWER 5. 2. 2-5 

5. 2. 2- 3 PAYLOAD POWER INTERFACE CHARACTERISTICS .... 5. 2. 2-9 

5. 2. 2- 4 EPOS EQUIPMENT 5.2.2-19 

5. 2. 2- 5 ASF AVERAGE POWER BY FLIGHT PHASE 5.2.2-20 

5. 2. 3- 1 ASF POINTING AND STABILITY REQUIREMENTS .... 5. 2. 3-4 


4 


Table Page 

5. 2. 3- 2 ATTITUDE POINTING ACCURACY - ORBITER 

REFERENCE SYSTEM 5. 2. 3-7 

5. 2. 3- 3 ORBITER THERMAL ATTITUDE CONSTRAINTS 5. 2. 3-8 

5. 2. 3- 4 CENTRAL VS DISTRIBUTED ATTITUDE REFERENCE 

SYSTEM 5.2.3-12 

5. 2. 3- 5 TYPICAL STAR TRACKER PERFORMANCE 

CHARACTERISITCS 5.2.3-17 

5. 2. 3- 6 SIZE, WEIGHT, AND POWER SUMMARY POINTING/ 

CONTROL AND STABILIZATION SUBSYSTEM 5.2.3-19 

5. 2. 3- 7 SUMMARY OF POINTING AND CONTROL SUBSYSTEM 

PRELIMINARY DATA REQUIREMENTS 5.2.3-24 

5. 2.3- 8 SYSTEM ERROR BUDGET (Ict) 5.2.3-41 

5. 2. 4- 1 REMOTE ACQUISITION UNIT (RAU) DATA OUTPUT 

CHARACTERISTICS 5.2.4-11 

5. 2. 4- 2 REMOTE ACQUISITION UNIT DATA INPUT 

CHARACTERISTICS 5.2.4-13 

5. 2.4- 3 COMPUTER CHARACTERISTICS 5.2.4-17 

5.2. 4- 4 TAPE RECORDER CHARACTERISTICS 5.2.4-19 

5. 2. 4- 5 CDMS EQUIPMENT CHARACTERISTICS 5.2.4-21 

5. 2. 4- 6 INSTRUMENT/CDMS INTERFACE LISTING 

(PALLET A-1) 5.2.4-31 

5. 2.4- 7 INSTRUMENT CDMS INTERFACE LISTING (PALLETS A-3 

AND A-4) 5.2.4-32 

5. 2. 5- 1 AMPS/ASF INSTRUMENT DISPLAY AND CONTROL 

REQUIREMENTS 5.2. 5-3 

5. 2. 5- 2 KEYBOARD TRADE-OFF COMPARISON 5.2.5-30 

5. 2. 6- 1 AE CAPABILITIES SUMMARY 5. 2. 6-3 

6. 5. 2- 1 ASF INSTRUMENT OPERATING TIMELINE 6-32 

6. 5. 2- 2 CREWMAN TASK TIMELINE 6-33 

7.1-1 SCHEDULE CONSTRAINTS. 7-2 


V-2 


Table Page 

7. 2. 1- 1 ASF PRIME INSTRUMENT DEVELOPMENT LEAD TIMES , . 7-8 

7. 2. 1- 2 ASF PAYLOAD DEVELOPMENT REQUIREMENTS VERSUS 

PROGRAM SCHEDULE HARD POINT 7-10 

7. 2. 2- 1 SUBSTITUTE INSTRUMENTS 7-13 

7. 2. 2- 2 DEVELOPMENT REQUIREMENTS PRIME INSTRUMENTS 

VERSUS CANDIDATE SUBSTITUTES 7-21 

7. 2. 2- 3 ASF DEVELOPMENT SCHEDULE REQUIREMENTS 

UTILIZING IDENTIFIABLE POTENTIAL SUBSTITUTE 
INSTRUMENTS 7-23 

7.2. 3- 1 SUPPORT SUBSYSTEM EQUIPMENT DEVELOPMENT 

STATUS 7-24 

7. 2. 3- 2 SUPPORT SUBSYSTEM DEVELOPMENT SCHEDULE 7-26 

8. 2. 2- 1 TSMS ISSUES 8-12 

8. 2. 2- 2 EPOS ISSUES 8-13 

8. 2. 2- 3 PCSS ISSUES 8-15 

8. 2. 2- 4 CDMS ISSUES 8-16 

8. 2. 2- 5 AFT CREW STATION SUPPORT ISSUES ........ 8-17 


v-3 


FIGURES 


I 


Figure Page 

2-1 ASF mission system 2-2 , 

4. 1.5-1 ASF mission timelines 4-31 

5. 1.1- 1 ASF systems operational relationship 5.1-2 

5. 1.1- 2 ASF pallet attachment locations 5.1-5 

5. 1.1- 3 ASF pointing system (concept) 5.1-8 

5. 1.1 - 4 AMPS pointing system stowed 5.1-9 

5. 1.1- 5 AMPS pointing system deployed 5.T-10 

5. 1.1- 6 ASF flight system interfaces 5.1-15 

5. 2. 1- 1 Pallet equipment hardpoints 5.2.1-10 

5. 2. 1- 2 Active thermal control system (ATCS) 5.2.1-12 

5. 2. 1- 3 Pallet cooling 5.2.1-13 

5. 2. 1- 4 Standard pallet configuration ... 5.2.1-17 

5. 2. 1- 5 AMPS pointing system operation 5.2.1-19 

5. 2. 1- 6 AMPS pointing system (APS) mount 5.2.1-21 

5. 2. 1- 7 Instrument/AIM thermal control 5.2.1-23 

5. 2. 1- 8 AMPS Instrument Module (AIM) 5.2.1-24 

5.2. 1- 9 Deploy-Retract system 5.2.1-26 

5.2.1- 10 Subsatellite retrieval and boom deployment. . . 5.2.1-28 

5.2.1- n STEM (single spool) 5.2.1-30 

5.2.1- 12 BI-STEM (twin spool) 5.2.1-30 

5.2.1- 13 STEM/BI-STEM element deployment comparison. . . 5.2.1-31 

5.2.1- 14 Chemically milled strip . 5.2.1-31 

5.2.1- 15 Subsatellite retention/ejection 

mechanization 5.2.1-34 


vi-1 


it 


i 


Figure Page 

5.2.1- 16 Subsatellite retention ejection mechanism . . . 5.2.1-35 

5.2.1- 17 Candidate pointing platform study concepts. . , 5.2.1-40 

5.2.1- 18 Z-axis CG location 5.2.1-52 

5.2.1- 19 X-axis CG location 5.2.1-53 

5.2.1- 20 Y-axis CG location 5.2,1-54 

5. 2. 2- 1 Energy/Power available to payload 5.2.2-10 

5. 2. 2- 2 ASF electrical power distribution system 

(EDPS) 5.2.2-13 

5. 2. 2- 3 ASF Particle Accelerator High Voltage Power 

System 5.2.2-16 

5. 2. 2- 4 ASF system power profile 5.2.2-22 

5. 2. 2- 5 Typical orbit revolution for peak electrical 

power 5.2.2-23 

5. 2. 3- 1 Pointing definitions 5. 2. 3-5 

5. 2. 3- 2' Central attitude reference concept 5.2.3-11 

5. 2. 3- 3 Centralized AMS signal plan 5.2.3-13 

5. 2. 3- 4 Altitude reference & pointing control 

functions 5.2.3-14 

5. 2. 3- 5 Interface and data flow diagram 5,2.3-21 

5. 2. 3- 6 Pointing and control interface and data flow 

chart 5.2.3-22 

5. 2. 3- 7 PCSS operational sequence ... 5,2.3-26 

5. 2. 3- 8 Inflight alignment sequence 5.2.3-31 

5. 2.3- 9 Block diagram for pointing and tracking 

simplified 5.2.3-33 

5.2.3- 10 Solar monitoring pointing and control 5.2.3-35 

5. 2. 4- 1 CDMS functional block diagram 5,2. 4-6 

5. 2. 4- 2 RAU mechanical configuration 5.2.4-14 

5. 2. 4- 3 ASF data time-lines . 5.2.4-37 


vi-2 


I 


Figure Page 

5. 2. 4- 4 ASF data rate requirements revolution 16. . . , 5.2.4-38 

5. 2. 4- 5 ASF data rate requirements revolution 17-25 . . 5.2.4-39 

5. 2. 4- 6 ASF data rate requirements revolution 26-31 . . 5.2.4-40 

5. 2. 4- 7 ASF data rate requirements revolution 32-35 . . 5.2.4-41 

5. 2.4- 8 ASF data rate requirements revolution 36-37 . . 5.2.4-42 

5. 2.4- 9 ASF data rate requirements revolution 38-42 . . 5.2.4-43 

5.2.4- 10 ASF data rate requirements revolution 43-47 . . 5.2.4-44 

5.2.4- 11 ASF data rate requirements revolution 48-79 . . 5.2.4-45 

5.2.4- 12 ASF data rate requirements revolution 80. . . . 5.2.4-46 

5.2.4- 13 ASF data rate requirements revolution 0 to 

90 5.2.4-47 

5. 2. 5- 1 Orbiter aft flight deck panel location code . . 5. 2. 5-4 

5. 2. 5- 2 Aft flight deck crew station 5. 2. 5-7 

5. 2.5- 3 Panel LIO 5.2.5-11 

5. 2. 5- 4 Panel Lll 5.2.5-12 

5. 2. 5- 5 Panel L12 payload specialist station 5.2.5-13 

5. 2. 5- 6 CRT graphics display format 5.2.5-14 

5. 2. 5- 7 CRT alphanumeric display format 5.2.5-15 

5. 2. 5- 8 Alphanumeric keyboard panel 5.2.5-17 

5. 2. 5- 9 Crew compartment, midsection plan view 5.2.5-25 

5.2.5- 10 Crew compartment, aft view . 5.2.5-26 

5. 2. 6- 1 Subsatellite configuration, component 

designations 5. 2. 6-4 

5. 2. 6- 2 Spacecraft coordinate system 5. 2. 6-7 

5. 2.6- 3 Subsatellite instrument layout. 5. 2. 6-9 

5. 3. 2- 1 AMPS EGSE assemblies. . 5.3-5 

5. 3. 2- 2 AMPS EGSE (payload integration site) 5.3-6 


Vi -3 


Figure Page 

5. 3. 2- 3 AMPS EGSE (launch site) 5.3-7 

5. 3. 2- 4 EGSE utilization 5.3-9 

5. 3.2- 5 AMPS simulator for experiments -. 5.3-11 

5. 3. 2- 6 Core segment simulator (CSS) 5.3-12 

5.3. 2- 7 Instrument handling concept 5.3-18 

5. 3. 2- 8 Concept of AMPS pallet handling frame 5.3-19 

5. 3. 2- 9 Horizontal ASF/AMPS payload installation/ 

removal (concept) 5.3-22 

5.3.2- 10 Standard payload maintenance and checkout 

station 5.3-23 

5.3.2- 11 Horizontal payload cannister concept 5.3-24 

5.3.2- 12 AMP/ASF payload vertical handling concept . . . 5.3-26 

5.3.2- 13 Pad payload vertical handling 5.3-27 

5.3.2- 14 Pad payload changeout .... 5.3-28 

5.3.2- 15 Pad payload vertical handling 5.3-29 

6.1- 1 ASF mission system 6-2 

6.2- 1 Normal Orbiter operations 6-3 

6.2- 2 ASF instrument pointing operations 6-3 

8-1 ASF mission system 8-2 


vi -4 


ACRONYMS AND ABBREVIATIONS 


A 

O 

A 

A&A 

ac 

A/D 

AE 

AGC 

Ah 

AIM 

AMPS 

AMS 

AMU 

A/N 

ANK 

APS 

APCS 

ASF 

AS TP 

ASPO 

ATCS 

ATS 

AZ 

BI-STEM 

bpm 

bps 

Btu 

B/U 

C 

CCIG 

CCTV 

CDMS 

CEP 

CDU 

CG 

cm 

COAS 


ampere(s) 
angstrom{ s ) 
alarm and advisory 
alternating current 
analog to digital 
Atmosphere Explorer 
Automatic Gain Control 
ampere-hour 

AMPS Instrument Module 

Atmospheric, Magnetospher i c , Plasmas in Spac 

Attitude Measurement . System 

Atomic Mass Unit 

al pha-numeric 

alpha-numeric keyboard 

AMPS Pointing System 

Attitude Pointing and Control System 

Atmospheric Science Facility 

Apollo Soyuz Test Project 

Apollo Spacecraft Program Office 

Active Thermal Control Subsystem 

Application Technology Satellite 

azimuth 

Bi (dual) Storable Tubular Extendable Member 

bits per minute 

bits per second 

British Thermal Unit 

backup 

Centigrade 

Cold Cathode Ion Gauge 

closed circuit television 

Command and Data Management Subsystem 

langmuir probe 

control and display unit 

center of gravity 

centimeters 

crewman optical alignment system 


V i i - 1 


CRT 

CSS 

C&W 

CUM 

CW 

D&C 

D/A 

dB 

dc 

deg/sec 

DEU 

DSKY 

ECS 

ED 

e. g. 
EGSE 
ELF 
EL 
EMC 
EMI 
EPDS 
ERNO 
E5D 
ESRO 
ESTEC 
EUV 
eV 
EVA 
F 

FCS 

FFK 

FHST 

FM 

FOV 

fps 

ft 

GARP 


cathode ray tube 
core segment simulator 
caution and warning 
cubic meter 
carrier wave 
displays and controls 
digital to analog 
deci bel 

direct current 
degrees per second 
display electronics unit 
display and keyboard assembly 
environmental control system 
experiment description 
for example 

electrical ground support equipment 
extremely low frequency 
el evati on 

el ectromagn eti c compati bi 1 i ty 

electromagnetic interference 

electrical power and distribution subsystem 

Entwickelung Ring Word {Engineering firm of the north) 

Experiment Systems Division 

European Space Research Organization 

European Space Technical Center 

extreme ultraviolet 

electron volt 

extra vehicular activity 

farad 

flight control system 

fixed function keyboard 

fixed heat star tracker 

frequency modulation (or modulated) 

field-of-view 

feet per second 

feet 

Global Atmospheric Research Program 


vi i -Z 


GEOS 

GFE 

GHz 

GMT 

GNg 

GN&C 

GPC 

GRA 

GSE 

GSFC 

GST 

HERD 

HF 

HPI 

hr 

HV 

Hz 

IC 

ID 

i . e. 

lEP 

IFOV 

IFRD 

IMU 

in 

I/O 

IPS 

i ps 

IR 

ISIS 

J 

JSC 

K 

kb 

kbps 

keV 

kg 


Geodetic Earth Orbiting Satellite 
government furnished equipment 
gi gahertz 

Greenwich Mean Time 

gaseous nitrogen 

guidance, navigation and control 

general purpose computer, or ground payload computer 

gyro reference assembly 

ground support equipment 

Goddard Space Flight Center 

gimballed star tracker 

high energy particle detector 

high frequency 

high performance insulation 

hour 

high voltage 
hertz 

integrated circuit 
instrument description 
that is 

instrument electronics package 

instantaneous field-of-view 

Instrument Functional Requirements Document 

inertial measurement unit 

inch(es) 

input/output 

Instrument Pointing System 
inches per second 
i nfrared 

International Satellite for Ionospheric Studies 
Joule 

Johnson Space Center 
degrees Kelvin 
kilobit 

kilo bits per second 
kilo electron volt 
kilogram{s) 


vi i-3 




kwds 


LIED 


l-Hp 

LLTV 


Ly-a 


MAIL 

Mbps 

MCC 

HCDS 

MECO 


MGSE 


mtnHg 

MPD 


mrad 


ki 1 ohertz 
kil ojouk 
ki 1 ometer ( s ) 

Kennedy Space Center 
kilowatt(s) 
kilowatt hour(s) 

kilo words (digital) data per second 

Launch Control Center 

low energy electron probe 

low energy ion detector 

liquid helium 

liquid hydrogen 

liquid nitrogen 

liquid oxygen 

low light television 

lunar module 

liquid nitrogen 

1 ine-of-s,i ght ( 1 oss-of-si gnal ) 

Lymman-Al pha 

meter ( s } 

multiple access 

triaxial fluxgate 

Mockup and Integration Laboratory 

mega bits per second 

Mission Control Center 

multifunction CRT display system 

main engine cutoff 

mission elapsed time 

mega electron volts 

mechanical ground support equipment 

Megahertz 

megajoule 

mi 1 1 imeter (s ) 

millimeters of mercury 

magnetopl asmadynamic 

magnetospheric and pi asmas-in-space 

milliradian(s) 

mi 1 1 i second 



vi i -4 


MS 

mV 

ym 

yF 

ys 

yv 

MSFC 

MSS 

MW 

N 

N/A 

NACE 

NASA 

NATE 

Ne 

NFFK 

nH 

NIR 

N/m^ 

NRL 

NIMBUS 

”2 

ns 

OBIPS 

OGO 

QMS 

OSO 

OFF 

OSSA 

PBI 

PCM 

PCSS 

PDS 

PES 

PI 

PM 

PMS 


Mission Specialist 
millivolt 
micrometer 
mi crofarad 
mi crosecond 
mi crovol t 

Marshall Space Flight Center 
mission specialist station 
megawatt 
Newtons 

not applicable 

neutral mass spectrometer 

National Aeronautics and Space Administration 

neutral atmosphere temperature 

Neon 

numeric fixed function keyboard 

nanohenry 

near infrared 

Newtons per square meter 

Naval Research Laboratory 

NIMBUS Spacecraft 

Nitrogen 

nanosecond 

Optical Band Imager and Photometer System 
Orbiting Geophysical Observatory (Spacecraft) 
Orbital Maneuvering System 
Orbiting Solar Observatory (Spacecraft) 

Orbiter Processing Facility 

Office of Science and Space Applications 

push button indicator 

pulse code modulation 

Pointing Control and Stabilization Subsystem 
Particle 

photoelectron spectrometer 

Principal Investigator 

phase modulation (or modulated) 

Performance Monitoring System (or Payload Mission 
Simulator) 


vii-5 


> 

I 

) 


i 


POK 

ppm 

pps 

PPU 

PS 

p . s . 

psi {a)(g) 

PSS 

QED. 

RAU 

R&D 

RCA 

RCS 

rf 

rms 

RMS 

RPA 

rpm 

rps 

R&QA 

RTOP 

SA 

S&AD 

SAIL 

SCATHA 

SCMR 

SDWG 

sec 

SEMIS 

SHP 

SIPS 

SOLRAD 

sps 

SPS 

sq cm 

sq m 

sr 

S/S 




page overlay keyboard 
parts per million 
pulses per second 
power processing unit 
Payload Specialist 
power supply 

pounds per square inch (absolute) (gauge) 

Payload Specialist Station 

Quick and Easy Design 

Remote Acquisition Unit 

rendezvous and docking 

Radio Corporation of America 

reaction control system 

radio frequency 

root mean square 

remote manipulator system 

retarding potential analyzer (planar ion trap) 

revolutions per minute 

revolutions per second 

reliability and quality assurance 

Research and Technology Objectives and Plans 

single access 

Science and Applications Directorate 
Shuttle Avionics Integration Laboratory 
spacecraft charging at high altitudes 
Surface Composition Mapping Radiometer 
Science Definition Working Group 
second 

Solar Energy Monitor in Space 
standard hardware program 
Small Instrument Pointing System 
SOLRAD Spacecraft (Explorer 44) 
samples per second 
Solar Physics Satellite 
square centimeter 
square meter 
steradian 

Subsystem (sample/seconds) 


v i i - 6 


STA 

Star Tracker Assembly 


STDN 

Space Tracking and Data Network 


STEM 

Storable Tubular Extendable Member 

STS 

Space Transportati on System 


ST 

star tracker 


TBD 

to be determined 


T&C/O 

test and checkout 


TORS 

Tracking and Data Relay Satellite 


TDRSS 

Tracking and Data Relay Satellite 

System 

Te 

temperature of electrons 


TM 

tel emetry 


T/R 

transmi tter/receiver 


TSMS 

thermal structural and mechanical 

subsystem 

TV 

television 


ULF 

ultra low frequency 


UV 

ul traviol et 


V 

volt 


VA 

vol tampere 


VAB 

Vertical Assembly Building 


Vac 

volts alternating current 


VAE 

airglow photometers 


Vdc 

volts direct current 


VIS 

visible 


VLF 

very low frequency 


W 

watt(s) 


XUV 

extreme ultraviolet 


Y-POP 

y/perpendicular to orbiter plane 


ZVV 

z/velocity vector 



v1i-7 


SYMBOLS 


Ar argon 

BeCu beryllium copper 

. magnetic field strength 
C carbon 

CH^ methane 

CO carbon monoxide 

dE/dX energy loss per unit length 

E/Q energy per unit charge 

H atomic hydrogen 

He helium 

HNOo nitric acid 

water 

wave number 
nitrogen 

gaseous nitrogen 
nickel 63 
nitric acid 
nitrogen dioxide 
nitrous oxide 
atomic oxygen 
ozone 

hydroxyl radical 
phosphorus 
budidium 
atomic number 
degree 

less than or equal to 
greater than or equal to 
energy resolution 
mass resolution 
A (M/Q)/(M/Q) resol ution or ratio of mass to charge 
AX/X wavelength resolution 

X 
Y 

IT pi 


H2O 

k 

N 

N2 

Mi 63 
NO 
NO2 
N2O 
0 

O 3 

OH 

P 

Rb 

Z 

0 

< 

AE/E 

AM/M 


wavel ength 
lO"^ gauss 


vii-8 


1.0 INTRODUCTION 


1.1 AUTHORIZATION 

The NASA, Johnson Space Center (JSC) was requested by the Office 
of Space Science (OSS) to submit a Research and Technology Opera- 
ting Plan (RTOP) entitled “Atmospheric, Magnetospheric , and 
PI asmas- In-Space (AMPS) Payload Definition Studies" which inclu- 
ded the study of the potential of the pallet-only mode for the 
AMPS project and the provision of conceptual designs for the 
AMPS Atmospheric Science Facility (ASF) payload which can be 
flown in the pallet-only mode. This report is submitted in 
response to that RTOP. 

1.2 STUDY OBJECTIVES 
1.2.1 ASSESSMENT 

Assess the potential of a 1981 AMPS mission in a pallet-only 
mode aboard the STS. This particular RTOP objective was inter- 
preted as requiring a study to address the following questions: 

a. Is it technically feasible to fly an AMPS mission in a pallet- 
only mode aboard the STS? 

b. If the pallet-only mode is feasible for AMPS, of what would 
the AMPS flight system consist and how would it be integrated 
and operated? What facilities would be required to support 
the AMPS program? 

c. What impact would AMPS pallet-only missions have on NASA 
resources such as cost, schedule, facilities? 

d. What major trade-off considerations would be applicable, and 
what options would be presented to Level I NASA management 
for assessment of overall potential. For example, schedule/ 
resources vs. scientific objectives/benefits. 


1.2.2 IDENTIFICATION 


Identify instrument designs and operational requirements for 
satisfying scientific objectives set forth by the NASA AMPS 
SDWG. 

1.2.3 DEFINITION 

Define a conceptual ASF system design in sufficient depth to 
serve as a baseline for both a Level I management start decision 
(cos t/schedul e/merit ) and a final design study. 

1.2.4 PREPARATION 

Prepare a JSC study report containing results, conclusions, and 
recommendations for transmittal to NASA Headquarters. 

1.3 END PRODUCTS 

The end products of the study are two reports. The first is an 
Executive Summary document that presents an assessment of the 
potential of the Orbiter pallet-only mode to satisfy AMPS require 
ments. The summary will include: 

a. Conclusions and recommendations. 

b. Description of the study baseline system. 

c. Significant technical and operational trade-off factors. 

d. Representative AMPS payload instrument complements. 

e. Payload development costs and schedules. 

f. Identification of experiment classes not applicable to pallet 
only operation . 

g. Identification of significant problem areas that need resolu- 
tion. 

h. Recommendations for further study. 


1-2 


? 


The second document is a technical report in two parts. The 
first part presents conceptual designs and specifications for 
an ASF and includes findings which disclose that an ASF can 
be fl in in the pallet-only mode. The second part is a feasibi- 
lity udy on t'le subject of flying a complete AMPS mission 
using the pallet-only mode. 


2.0 STUDY APPROACH 


2.1 NATURE AND SCOPE OF THE STUDY 

As previously mentioned, the principal objective of the study is 
to assess the economic and technical feasibility of employing a 
pallet-only mode for conducting AMPS experiments. The study plan 
is to develop a baseline incorporating the experiment and instru- 
ment descriptions provided by the AMPS SDWG. This baseline will 
be augmented by assumptions and judgments of scientists and 
engineers knowledgeable in the various phenomena and state-of- 
the-art instrumentation. That baseline, which includes experi- 
mental objectives, methodologies, instrumentation, experiment 
timelines, development schedules and costs is then used to 
assess the feasibility of a pallet-only mode. The results may 
be used for advance planning and decision-making that will pre- 
clude false starts and wasted resources in a stringent economic 
envi ronment . 

The AMPS system, of course, incorporates much more than the ASF 
payload, as depicted in figure 2-1. It includes not only the 
Orbiter with its scientific payload 'but also space-to-ground 
and ground-to-space communication and data link systems, inter- 
faces with other satellites, and supporting ground facilities. 

The scope of this study, however, primarily addresses the 
payload; giving substantive consideration only to those other 
system facets that are significantly impacted by the pallet-only 
mode operation. Cursory consideration is given to all system 
aspects to ascertain whether or not there may be such significant 
impacts. 

2.2 TECHNICAL APPROACH 

The approach employed, in essence, started with a set of Instru- 
ment Functional Requirements Documents (IFRD's) defined from 
inputs by the AMPS SDWG. Experiments were then defined by the 


2-1 




I 


JSC study group. The instrument characteristics and experiment 
requirements allowed definition of support subsystem requirements 
and subsequent translation into operational requirements which 
were integrated into a conceptual system and mission. This con- 
ceptual system was used as a baseline upon which to base a 
feasibility assessment. 

Close communications were maintained with many of the scientific 
investigators of the AMPS SDWG to assure a correct understanding 
of their experimental objectives and preferences in experiment 
operations and data handling. Their inputs were augmented by 
JSC engineering expertise to define a conceptual system con- 
sidered feasible, realizeable within a reasonable time frame, and 
capable of meeting a maximum portion of the ASF scientific 
objectives . 

The Magnetospheric and PI asmas-in-Space (MPS) portion of the 
total AMPS concept is not addressed in this ASF report because of 
the unavailability of information and study schedule limitations. 
A separate report will be prepared at a later date. 

2.3 STUDY BASELINES 

2.3.1 JOHNSON SPACE CENTER ORGANIZATIONAL RESPONSIBILITIES 

For the conduct of the program, areas of responsibility were 
established as detailed below: 

a. The Shuttle Payload Integration and Development Program 
Office will be responsible for project management of the 
AMPS pallet-only study activities and related inter-NASA 
Center interface functions. 

b. The JSC Science and Applications Directorate (S&AD) will 
be responsible for defining and interpreting science and 
experiment requirements and interfacing with the AMPS SDWG. 


2-3 


c. The JSC Experiment Systems Division (ESD) will be responsible 
for: 

(1) Instrument definition. 

(2) Systems aspect of pallet(s) requirements, integration, 
and hardware interfaces. 

(3) Overall study objectives. 

2.3.2 DEFINITIONS 

2. 3. 2.1 Experiment 

Experiment as used in this report is an orderly operation per- 
formed to acquire data that will provide certain desired 
scientific information. 

2.3.2. 2 Prime Instrument 

An instrument which has been ascribed by the Scientific Inves- 
tigator on the AMPS SDWG for a particular experiment or group of 
experiments . 

2. 3. 2.3 Substitute Instrument 

An instrument that is functionally similar to, but with different 
capability than, the prime instrument. It will be substituted 
for the prime instrument when the latter is not available for 
flight due to technical problems, schedule, cost, failure, or 
other reasons. 

2 . 3 . 2 . 4 Alternate Instrument 

That instrument which is dedicated to different experiment objec- 
tives (i.e., another experiment) for which it is the prime (or 
substitute) instrument. It will be used as an alternative when 
the previous experiment's instrument(s) is not available for 
flight or priority status has changed. 


2-4 


2. 3. 2. 5 Core Instruments 



i 


That set of instruments that is used for experiments in all 
three scientific disciplines, i.e., atmospherics, magnetospherics, 
and pi asmas “in-space . 

2.4 REFERENCE DOCUMENTATION AND BIBLIOGRAPHY 

During the course of this study, a great number of documents, 
reports, papers, and texts were used as reference material. In 
general, the material fell into three categories as follows. 

2.4.1 AUTHORITATIVE DOCUMENTS 

Documents which provide direct technical and programmatic infor- 
mation relative to the Space Shuttle Program, including publica- 
tions such as JSC-07700, Volume XIV, JSC-09310 through JSC-09325, 
and JSC Specification SL-E-0001 . The information provided by 
those documents and publications is intrinsic to all phases of 
this study, and no effort was made to cite the numerous references 
which were made. 

2.4.2 REFERENCE DOCUMENTS 

Reference documents include those documents and publications pub- 
lished by NASA, NASA contractors, and by or for other Government 
agencies, and which provide information and background data on 
spacecraft, projects, instruments, experiments, etc., including 
publications such as the various user's guides prepared for the 
unmanned spacecraft and satellites. In some instances, the pub- 
lication used was one prepared by the prime contractor for the 
vehicle, such as the Radio Corporation of America (RCA) publica- 
tion on the Atmosphere Explorer (AE) satellite. Except in the 
rare case where a reference was made on a specific aspect, no 
effort was made to cite the many areas from which the information 
was derived. 


2-5 




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2.4.3 INFORMATION DOCUMENTS 

Information documents include journals and other sources for 
scientific papers written on the theory and practice of experi- 
ments in the disciplines with which this report has been concerned. 
Many of the papers perused represented the work of the scientists 
who are members of the various working groups of the AMPS Program. 

All of the documents used have been listed in the Bibliography, 
section 10.0, under one of the three headings previously mentioned. 

2.5 ASSUMPTIONS 

A number of assumptions were made at the outset of the study to 
establish guidelines and common bases of reference for all 
study participants. These assumptions are listed below: 

a. The pallet-only mode may utilize up to five Spacelab pallets. 

As many as three pallets can be rigidly joined together. 

b. The study will define the instruments, support subsystems 
interfaces, and Orbiter related operational requirements for 
any free-flying, maneuverable satellites/subsatellites, and 
tethered satellites required to support the AMPS pallet-only 
project . 

c. Control of free-flying, maneuverable satellites, and tethered 
satellites necessary to support the ASF/AMPS payload will be 
effected from the Orbiter. 

d. Earth and/or sun synchronous satellites may be considered, 
if necessary, to support the ASF/AMPS pallet-only payload. 

When required, the instrument complement, supporting sub- 
systems and applicable interface requirements therewith will 
be defined during this study, 

e. The ASF/AMPS payload will be automated to the maximum extent 
possible. However, the design approach will not preclude 
man-in-the-1 oop when hardware complexity and/or cost prohibit 
the automatic mode. 


2-6 


Each experiment, instrument, and support subsystem v/ill 
utilize standard modular equipment for display and control 
mounting in the aft crew station payload console. Real time 
data displays, both onboard and downlink, will be provided 
as required. 

Extravehicular Activity (EVA) operations to service the cargo 
bay payload equipment will not be considered normal operating 
procedure. However, the equipment design will not preclude 
EVA operations. 

Utilization of Rendezvous and Docking (R&D) and Remote Manip- 
ulator System (RMS) capabilities will be normal procedure 
for the ASF/AMPS payload. 

Existing NASA Shuttl e/Spacel ab document guidelines will be 
followed where applicable. Programmed Spacelab equipment, 
excluding the manned modules, will be utilized wherever pos- 
sible. All European Space Research Organization/Entwickelung 
Ring Nord (ESRO/ERNO) supplied equipment will meet schedule, 
fit, and function requirements. 

The first flight opportunity for ASF/AMPS payload(s) equip- 
ment will be mid-1981. Although all prime instruments may 
not be available for several years, the ASF/AMPS pallet-only 
basic design will provide experiment instruments for the 
first flight opportunity. 

Wherever specific information is lacking, the study report 
will so state. 

Configuration management, safety, reliability, and quality 
control guidelines will be established to NASA specifica- 
tions for all ASF/AMPS equipment. Specifications tailored 
to environment, contractor history, experience, and develop- 
ment status of hardware will also be established. 

The ASF/AMPS project plan will be based on hardware require- 
ments of one engineering model, one qua! i fication model , 
one training model (control panels), one flight model, and 
critical component spares. 


2-7 


n. ASF and/or AMPS will be considered the prime payload in 

terms of priority for the use of Orbiter payload accommoda- 
tions during any ASF/AMPS mission. 

0. A TV system for scanning within the Orbiter payload bay 
will be supplied on the Orbiter. 

p. Many lower level, detailed technical assumptions relating to 
design approaches and operational philosophy necessary during 
this study are identified in appropriate sections of this 
report. 


3.0 SUMMARY 


3.1 GENERAL 

This study was initialized with a preliminary set of IFRD's 
developed by the AMPS SDWG from which Experiment Descriptions 
(ED' s) (appendix A) and Instrument Descriptions ( ID ' s ) (appendix B) 
were derived. The ED's and ID's are summarized in section 4.0. 

The prime instruments are packaged into four pallets in a physical 
and functional manner compatible with the STS capabilities and/or 
constraints and an Orbiter 7-day mission timeline (section 5.0). 

In section 6.0 operational compatibility is verified between the 
Orbi ter/payl oad and supporting facilities (Particle Detector 
Subsatellite (PDS), Solar Physics Satellite (SPS), Tracking and 
Data Relay Satellite System (TDRSS), Space Tracking and Data 
Network (STDN), Mission Control and Ground Data Processing 
facilities). Section 7.0 treats the development status and 
schedule requirements applicable to the ASF mission. Sections 
8.0 and 9.0 contain detailed treatments of the conclusions and 
recommendations resulting from this study. The abbreviations 
and acronyms used in this report are defined in a listing which 
is in the front matter of this report. 

3.2 CONCLUSIONS 

Many meaningful conclusions may be derived from results of this 
study; a study oriented toward assessing the potential of a 
1981 ASF pallet-only mode STS mission. The study involved much 
more than a go-no-go determination of scientific and technical 
feasibility. This mission-level approach, as opposed to merely 
evaluating a "flight package" concept, necessitated many tangen- 
tial studies into facility-level interface requirements. The 
study exposed programmatic factors not only of extreme signifi- 
cance to realistic management planning but also applicable to 
almost all missions utilizing the STS as a platform for scien- 
tific payloads. These factors influence each facet of this summary. 


The scope of the ASF study is depicted in figure 2-1 which illus- 
trates the major facility interfaces. 

3.2.1 FEASIBILITY 

In general, feasibility conclusions can be summarized as follows, 
but qualifications are in order subject to other factors presented 
in this summary section. 

The data required to satisfy the preliminary set of definitions 
of the atmospheric science objectives can be obtained, utilizing 
the pallet-only mode with the proper instrumentation. However, 
much refinement in the scientific requirements may significantly 
impact programmatic considerations, primarily in the areas of 
cost and schedule. 

Although the programmatic feasibility factors of cost, schedule, 
etc., can allow a wide latitude in trade-off considerations, the 
cost and schedule requirements to deliver certain prime instru- 
ments by 1981 are almost prohibitive. In addition, the costs to 
develop some instruments, considered prime at this time, could 
prove to be economically unfeasible. 

If the global coverage requirement is interpreted literally, the 
polar orbit missions required to accomplish this will not be 
possible until at least 1983 because of present schedules for 
availability of the western launch facility. However, partial 
global coverage would be possible, in the interim period, utili- 
zing the eastern launch facility. 

Although schedule and costs are a major factor, it is technically 
feasible to conduct an ASF mission in the pallet-only mode. Two 
of the technical factors which may affect technical feasibility 
in some areas are: (1) contamination from the STS, and (2) 

payload computer sizing. These factors influence the unresolved 
issues, follow-up requirements, and trade-off considerations as 
treated in this summary. 


3-2 


3.2.2 MAJOR UNRESOLVED ISSUES 

During the course of this studyj initial concepts and approaches 
were selected in the development of a pallet-only mode ASF mission 
utilizing the STS. Preliminary mission timelines resulting from 
limited definition of the experiment and instrument requirements 
were developed and subsequently updated. As appreciation of the 
Orbiter contamination environment developed, a particle detector 
subsatellite and boom-mounted equipment design were implemented. 

This resulted in a conceptual functional design considered tech- 
nically feasible, but with certain qualifications because of key 
assumptions developed along the way. Validity of some assumptions 
could not be fully verified. As a result, several potentially 
significant issues remain which warrant identification and require 
future investigation. 

a. Upon receipt of the updated and upgraded sets of AMPS/ASF 
experiment/instrument requirements from the SDWG, revised 
mission timelines will be needed to establish operational 
boundaries. These boundary timelines will then be used to 
complete the task of sizing the ASF system, followed by a 
reassessment of the ASF design concepts relative to the nev/ 
timeline. Particular emphasis will be given to the aft 

crew station, command and data management, power, and thermal 
subsystems for probable impacts. 

b. There is need to operate the particle detector instruments 
a relatively short distance away from the Orbiter to avoid 
an excessive contamination environment. The AE satellite 

was chosen to carry these instruments because it is operationally 
ready and the normal AE instrument complement requires 
minimal change. There are obviously many unresolved problems 
associated with this approach: 

(1) How do the above impacts compare with those of a tethered 
satellite or boom-mounted module? 


(2) Would it be feasible to modify the proposed subsatellite 
to remain in orbit and possibly be used for other scien- 
tific missions? 

(3) How practical is the boom concept to implement in view 
of the requirement for Orbiter attitude changes? 

Potential boom dynamics problems warrant further investigations 
related to technical, scientific, and operational factors. 

c. The AMPS/ASF pallet-only mode of operation has more instruments, 
more experiments, more automation and a much greater emphasis 
on data processing than any previous space payload. This pre- 
sents the need for a more detailed investigation of the Orbiter 
payload computer capability versus the forthcoming, upgraded 
requirements for AMPS/ASF experiments; more detailed than was 
possible within the scope of this study. This issue is 
addressed at length in paragraph 8.4. 

3.2.3 TECHNICAL FOLLOW-UP REQUIREMENTS 

Although study results indicate functional feasibility of this 
conceptual ASF payload design, more accurate capacity and sizing 
definitions are required in most areas. In order to refine the 
definitions, many details (not known originally) of the design 
and operation of the various instruments are required (i.e., 
detector and housing design for cryo-cooled instruments, and 
total payload data characteristics and timelines affecting data 
processing). A summary identification and priority of technical 
follow-up efforts resulting from this study are listed below. 

Details are contained in paragraph 8.3. 

a. Define in greater detail a comprehensive set of requirements 
for experiments, instruments, subsatellite and support sub- 
systems. This effort should include defining more detailed 
mission timelines for experiment, instrument and subsystem 
operations than that developed to date. 


3-4 


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b. Provide better and more comprehensive design and operational 
definitions for instruments and subsystems. 

c. Perform various analyses and trade-off studies to verify the 
preliminary selections or to update the design and operations 
with more optimum approaches. 

d. Generate preliminary design and operational specifications 
to be used as a basis for downstream development. 

e. Develop programmatic factors such as estimates of total 
program development, production, and operational costsj 
funding plans including expenditures by phases, allocation 
of resources, funding constraints and optional expenditure 
approaches; development, production and operational schedules 
including expected critical paths and availability of non-ASF 
support such as the Orbiter, the SPS, the TDRSS, etc,; develop- 
ment, production and operational plans for each major program 
element (e.g., flight hardware, flight software, ground support 
facilities and ground support software); and an analysis of 

the technical, cost, and schedule risks involved with full 
scale development. 

3.2.4 TRADE-OFF CONSIDERATIONS 
3.2.4. 1 Scientific 

The preliminary nature of the present scientific requirements 
plus the advanced state-of-art of many prime instrument concepts, 
present many potential trade-off areas. A detailed treatment 
is contained in paragraph 8.2.1 of this report. They can be 
broadly categorized in this summary as follows. 

a. Different techniques to derive the desired scientific 
information. 

b. Postponement of experiments requiring the instruments. 

c. Substitute instrument{s) which may affect optimized scientific 
goals. 


3-5 


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3. 2.4. 2 Technical 

Table 3. 2. 4-1 summarily lists the technical trade-off parameters 
which are comprehensively treated in paragraph 8.2.2 of this report. 

3. 2. 4. 3 Programmatic 

Many major trade-^off considerations of a programmatic nature are 
evident from this ASF pallet-only mode study. Information is 
available now for some; additional information is required for 
many others. Paragraph 8.2,3 and appendix C (4 parts) address 
this subject in more detail. The two major trade-off areas center 
around the STS contamination environment and the practicality of 
a 1981 launch requirement as opposed to a 1983-1985 launch date. 

3.3 RECOMMENDATIONS 

3.3.1 ASF PAYLOAD SYSTEM DESIGN 

Paragraph 9.1 of the text presents a detailed treatment and listing 
of specific recommendations for each major subsystem comprising 
the ASF pallet-only mode payload design. These recommendations 
incorporate an extensive use of Spacelab and Orbiter equipment 
and approaches. Although follow-on efforts are required to better 
refine the design concepts, the recommended configuration estab- 
lishes a feasible baseline from which tc initiate a preliminary 
system design study. 

3.3.2 FOLLOW-ON STUDIES 

Several unresolved major issues, identified above, must be addressed 
because they not only constrain technical effectiveness of this 
conceptual payload but they also involve major cost and schedule 
impacts to an ASF pallet-only mission(s). These issues and follow- 
on studies are treated in detail in paragraph 9.2. They are 
summarily listed as follows. 


3-6 


TABLE 3. 2. 4-1 . 


CANDIDATE SUBSYSTEM OPTIONS 


1. Cryogenic cooling 


L Thermal dissipation 


3. Large structural 

assembly installation 


. Circuit breakers 


5* High current 

transmission media 


6. Attitude measuring 
system 


7* Payload Specialist 
work station 

8. Experiment sequence 
ini ti ati on 

9. Data processing 

10. fiass memory operational 
programs 


11. Data compression 
{subsa tel 1 i te and 
fixed payload to 
Orbi ter ) 


12. Computer, processor 


14. Orbiter and payload 
EMI environment 


Support subsystem 
equipment selection 


Selected Approach 

Candidate Options 

Open loop 

Closed loop 

Payload coolant loop, Orbiter 
ATCS, Heat Radiator Kit 

Payload unique radiators 

Mounted on pallets 

Use Orbiter primary payload 
attachment points 

Remotely controlled 

Direct access (at crew 
s tati on ) 

Large gauge (4/0) wires 

Copper bus bars 

Centralized on Pallet 1, 
attitude transfer via optics 

Distributed star tracker, GRA 
on each AIM or APS 

Aft flight deck standard 
Orb i te r PSS 

Standard PSS plus mid-deck 
v.'Ork station 

Onboard control 

Ground control 

Onboard computer 

Ground facilities 

Temporary storage-reload from 
ground as programs are 
uti 1 i zed 

Permanent full mission 
programming capacity 

Conventional Bi-Phase 
Manchester II PCM and tape 
recorders 

Various high density systems 

Centralized experiment and 
subsystem (with backup) 

Distributed microprocessors 
plus less complex central 
p rocess or 

Retrieve subsatellite 
for subsequent reuse 

Leave subsatellite in orbit. 
Consider trade-offs bet.-,*een 
economics and operation 
CQmDlexi*_y, safety. Consider 
using on-staticn subsatellite 
for multiple ASF missions. 

Minimize payload generation 
and susceptibility through 
conventional design techniques 

Reduce Orbiter generation 
(e.g. change from structure 
to two wire return, increase 
shielding); adjust experiments 
to adverse environment 

Primarily Space Shuttle, 
Orbiter, Apollo 

Other existing or in develop- 
ment advanced, cost effective 
systems and ha rdware ; standard- 
ized modular designs 




3-7 






















3.3.2. 1 Scientific 

a. Using the upgraded ED's forthcoming from the AMPS SDWG, develop 
upgraded ASF mission timelines. The nev/ timelines* utilizing 
the new ED's and revised ID's, should be analytically exercised 
by the conceptual payload system to verify continuing feasi- 
bility of the payload concept with a more realistic ASF pallet- 
only mode mission. 

b. Choice of Instruments. Because of the unavailability of some ' 

ASF instruments for a mid-1981 launch date, it is recommended 

that a study be conducted v/ith the following objectives: 

r\ 

(1) Search for availability of instruments that can be used 
in lieu of those prime instruments presently described 
that cannot meet launch date and for which substitutes 
are not identified. Such instruments could be currently 
under development by either Government or industry, and 
could be completed in time to meet the scheduled launch 

date. Assess the impact to scientific value from the , 

use of substitute and/or alternate instruments. 

(2) Explore alternate means of acquiring desired scientific 
information without the use of those instruments that 
cannot meet launch date and for which there are no 
subst i tutes . 

(3) Assess scientific and cost impacts of flying certain 
experiments during 1981 and deferring others until 
requisite instruments are available. 

3. 3. 2. 2 Techni cal 

The specific follow-on studies recommended for the ASF payload « 

system design (described in paragraphs 9.2.1 and 9.2.2) are 
listed on the following page. 


3-8 


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a. EMI assessment. 

b. Particle contamination evaluation, 

d. Electrostatic charge assessment. 

d. Study the overall issue of the use of boojns, subsatell i tes , 
tethered satellites^ or other concepts to cope with problems 
posed by the operation of AMPS particles instruments. This 
study should encompass the following factors; 

(1) All Orbiter interfaces' (physical, operational, etc.). 

(2) Gross cost factors. 

(3) Scientific merit. 

(4) Program schedules, 

(5) Boom structural analyses. 

e. Those analyses, trade-offs, assessments, and definitions 
related to each subsystem and described in paragraphs 9. 2. 2.1 
through 9. 2.2. 5. 

f. Concept of standardizing (paragraph 9.2.3). 

3.4 COST 

All cost considerations related to this study are contained in 
the Executive Summary. 


4.0 SCIENTIFIC REQUIREMENTS 




4.1 ORBITER PAYLOAD 

4.1.1 INTRODUCTION 

Exploratory studies of the thermosphere during the last decade 
have provided the necessary information to describe gross 
features of structure, composition and variability of the 
region above 250 km. The region between about 120 km and 
200 km, where most of the extreme-ultraviolet (EUV) solar photons 
are absorbed, had not been studied extensively by in-situ satel- 
lite experiments until the flight of Atmosphere Explorer C (AE-C). 
The AE satellites are equipped to measure, simultaneously, the 
physical and chemical parameters of the neutral and ionized con- 
stituents, some of the airglow emissions, and the incident solar 
photon flux down to an altitude of 120 km. A significant 
improvement in our understanding of the structure and photo- 
chemistry of this region is expected to result from these mis- 
sions, leading to reasonably successful theoretical models of 
the structure of the upper thermosphere, and the upper E region 
and F region of the ionosphere. Most of the uncertainties in 
such models will likely be due to input parameters, such as 
reaction rates, cross sections, and absolute solar flux inten- 
sity. Improvement in our very limited present-day knowledge of 
the absolute intensity and variability of the solar EUV flux will 
result from the EUV spectrophotometer carried by the AE satel- 
lites, but a much needed increased data base, necessary for 
quantitative thermospheric and ionospheric calculations, will 
not be obtained before the Orbiter flights. 

The AMPS missions, and particularly the ASF mission, will provide 
a unique opportunity to study the basic processes in the areas of 
photochemistry, chemical kinetics, and atomic and molecular 
physics, that are of fundamental importance to the understanding 
of the evolution of planetary atmospheres as well as comet and 


4-1 


interstellar cloud formation. Experiments which cannot be per- 
formed in terrestrial laboratories can be conducted in the 
medium of space, for the unattenuated solar ultraviolet (UV) and 
x-ray flux can be utilized in excitation and ionization studies. 
Gas releases, either directly from the Orbiter or from a con- 
tainer some distance away, will permit the study of molecules and 
radicals found in the atmospheres of the planets or the major 
planets as well as the more complicated molecules suspected of 
being the parents of the commonly observed cometary species. 

In addition to photo excitation, electron excitation produced by 
the onboard electron accelerator can be used to produce multiple 
ionization and excitation of atomic species found in planetary 
nebulae. Laser fluorescence can then be used for the detection 
of long-lived metastable species. Electron impact cross sections, 
photo absorption cross sections, probabilities, ion and neutral- 
neutral reaction rates are examples of the type of atomic param- 
eters which can be determined. Photodissociation and 
photoionization lifetimes of cometary species can also be 
determined using either gas releases or an artifical comet (snow- 
ball) released in the vicinity of the Orbiter. 

Since a detailed knowledge of photochemi stry , dynamics, and 
energetics is essential to understanding the interrelationship 
of the atmospheric regions, it will be possible, for the first 
time, to treat the atmosphere in a unified manner. It will be 
of special interest to establish the relative. importance of dif- 
ferent energy sources to the behavior of the -atmosphere , e.g., 
solar radiation, wave energy from the lower atmosphere, and 
magnetospheric input including joule and energetic particle 
heating. The underlying troposphere is a source of natural and 
anthropogenic chemical species that enter into photochemical 
chains which are believed to have significant control over the 
composition of the stratosphere. A knowledge of the relationship 
between the minor constituent photochemistry and the energetics 


and dynamics of the stratosphere and mesosphere is essential for 
any significant improvement in our understanding of these regions. 

The Orbiter provides an unparalleled opportunity to conduct an 
investigation of the earth's atmosphere in the regions above the 
tropopause. These regions, the stratosphere, the mesosphere, and 
the thermosphere as far above the orbit as the Orbiter instru- 
mentation can acquire useful information, are very important to 
the understanding of atmospheric behavior. A large portion of 
the energy that originates from outside the atmosphere and becomes 
involved in the earth's atmospheric chemistry, physics, and 
mechanics is trapped, absorbed, or otherwise utilized in the 
regions above the tropopause. Probing directly into these 
regions has been done only with difficulty and for brief periods 
of time. Therefore, these regions are not well understood. 
However, the advent of the Orbiter provides the means to rectify 
the paucity of information about these regions that have such 
profound influence on the general terrestrial climate. 

The troposphere has been the subject of operational and research 
observations for many years and is currently being studied 
extensively as part of the Global Atmospheric Research Program 
(GARP). Operational instruments have been flown at altitudes up 
to 30 km above the land areas of the northern hemisphere and 
measurements have provided sufficient wind and temperature data 
to enable a meaningful understanding of the region. Satellite 
soundings, especially from the Nimbus satellites, have mapped 
stratospheric temj.erature to about 50 km and ozone distribution 
from 30 to 50 km. These data have contributed to a more 
detailed understanding of the dynamics of the stratosphere and 
are beginning to elucidate the overall ozone photochemistry 
scheme and the controlling transport processes. Experiments on 
the Nimbus F satellite are expected to map temperatures up to 
the lower mesosphere, and ozone and water vapor distributions 




to the stratopause. Later experiments on the Nimbus 6 satellite 
are expected to measure a, number of trace species from the tropo- 
pause into the mesosphere, although not always with the desirable 
vertical or horizontal resolution. 

It is now known that a very close coupling exists within the 
neutral atmosphere ionosphere-magnetosphere system, and that very 
complex interactive and feedback processes are present which 
involve mass, momentum, and energy transport, mostly along mag- 
netic field lines. Knowledge and understanding of these pro- 
cesses is minimal and will probably be so at the time of the 
Orbiter flights. 

4.1.2 OBJECTIVES 

The fundamental objectives of the ASF mission are to investigate 
the following. 

a. Composition and structure of the upper atmosphere. 

b. Dynamic and physical processes of the upper atmosphere. 

c. Interrelationships between the upper atmosphere and 
magnetosphere. 

d. Interrelationships between solar phenomena and the upper 
atmosphere. 

Fulfillment of the scientific objectives will require applica- 
tion of a number of instruments, including optical instruments, 
lasers, accelerators, and gas release devices. The ASF mission 
will use an array of such instruments operating concurrently, or 
in programmed sequences, to perform the observations and produce 
the data for studies of the upper atmosphere and of the correla- 
tion between upper atmosphere conditions and external influences. 
The instruments will permit separation of the temporal and '-ipatial 
aspects of the observed conditions and, t ough use of the accel- 
erators and gas release instruments, will initiate artificial or 
controlled perturbations of the ambient atmospheric constituents 
for observation and measurement. 


4-4 


4.1.3 EXPERIMENTS 
4. 1.3.1 Background 




I 


A series of fifteen atmospheric science experiments have been 
descri bed . 

a. Group D — Dynamics — Experiments to measure winds, tempera- 
ture, and diffusion of atmospheric constituents. 

b. Group C — Chemistry — Experiments to investigate photo- 
chemical reactions in the upper atmosphere. 

c. Group S - Structure — Experiments to investigate particle 
interactions in the upper atmosphere. 

Source data for the descriptions were obtained from papers and 
presentations by the scientists of the Atmospheric Science Sec- 
tion of the AMPS SDWG. The ED ' s are incorporated as appendix 
A of this report and include a statement of objective, a method 
of accomplishment, a list of instruments required, and the oper- 
ational timing for collection of data. 

The ED ' s are preliminary and, as a result, will undergo refine- 
ment and perhaps change as additional data becomes available. 
Nonetheless, these descriptions have been adequate for the pur- 
pose of establishing a baseline for performance of the ASF 
mission. 

4. 1.3. 2 Experiment Summaries 

The fifteen ASF exper’.,.ents are listed below and each is sum- 
marized in the subsequent paragraphs. 

Experiment Title 

AS-1 Identify Properties of Natural Tracers 

AS-2 Measure Winds and Temperature Fields 

AS-3 Profile the Atmosphere Temperature in the 

Region of 15 to 120 km Altitude 


4-5 


Experiment 


Title 


1 


f 


AS-4 

AS-5 

AS-6 

AS-7 

AS-8 

AS-9 

AS-10 

AS-n 

AS-12 

AS-13 

AS-14 

AS-15 


Determine the Thermal Structure and Dynamics 
of the Mesospheric and Lower Thermospheric 
Regions 


Determine the Eddy Diffusion Between the Alti- 
tudes of 85 km and 120 km 


Determine Atmospheric Interactions of Excited 
Radi cal s 


Determine Atomic and Molecular Oxygen Densi- 
ties Between 90 to 120 Kilometer Altitude 

Determine Solar Radiation Interaction with 
the Ambient Atmosphere 

Determine the Atmospheric Constituent Abun- 
dance Below 120 Kilometer Altitude 


Determine the Atmospheric Constituent Abun- 
dance Above 120 Kilometer Altitude 


Determine Change in the Ionospheric D Region 
Due To Seasonal Anomalies and Magnetic 
Storms 

Determine the Metallic Constituents in the 
Upper Atmosphere 

Evaluate Deposition of Meteoric Dust and 
Metallic Constituents 


Determine the Meteoric Production of Nitric 
Oxi de 


Investigate the Excitation Exchange Between 
Metastable Species and the Ambient 
Envi ronment 


4-6 


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a. Experiment AS-1 — Identify Properties of Natural Tracers: 

(1) Scientific Objective —To identify the properties of 
constituents which occur naturally in the atmosphere in 
order to determine suitability for scientific measure- 
ment and analysis. 

(2) Method — Radiation spectroscopy in the dovmv/ard and 
horizon looking directions will be used as a source of 
data. Wavelengths shorter than approximately 3 microm- 
eters will require a sunlit air column against a dark 
background, or active probing with a laser beam. 

Release of trace gases in the vicinity of the Orbiter 
will be investigated to determine the desirability for 
providing controlled concentrations of known tracers 
for calibration of instruments. 


(3) Instruments Required - 
Instrument No. 


118 

124 

126 

213 

532 


Title 


Limb-Scanning Infrared Radiometer 
Fabry-Perot Interferometer 
Infrared Interferometer 
Laser Sounder 
Gas Release Module 


b. Experiment AS-2 — Measure Winds and Temperature Fields: 

(1) Scientific Objective — To measure winds and temperature 
fields in the upper atmosphere, on a global scale, using 
natural tracers determined, in Experiment AS-1, to be 
suitable. 

(2) Method — Temperatures of gases may be derived from 
doppler broadening of emission lines, while scalar 
flow may be found from the doppler shifting of the 
same emission lines. 


4-7 


(3) Instruments Required — 


Instrument No . 
118 
124 
126 
213 


Title 

Limb-Scanning Infrared Radiometer 
Fabry-Perot Interferometer 
Infrared Interferometer 
Laser Sounder 


c. Experiment AS-3 - Profile the Atmosphere Temperature in the 
Region of 15 to 120 km Altitude: 

(1) Scientific Objective — To measure the vertical tempera- 
ture profile to differentiate from the horizontal tem- 
perature distributions found in Experiment AS-2. The 
resolution should be 1 km to 2 km of altitude. 


(2) Method — Horizon scanning will be used to collect data. 
Temperatures of gases will be derived from doppler 
broadening of emission lines. Active vertical sounding 
may be possible through use of laser probing to excite 
atmospheric sodium emissions. A nadir-pointing infra- 
red spectrometer will provide similar data by measure- 
ments of the shifts in the line of a carbon dioxide 
absorption edge. 


(3) 


Instruments 

Instrument 

118 

124 

126 


Requi red — 

No . Ti tl e 

Limb-Scanning Infrared Radiometer 
Fabry-Perot Interferometer 
Infrared Interferometer 


213 


Laser Sounder 


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d. Experiment AS-4 - Determine the Thermal Structure and Dynamics 
of the Mesospheric and Lower Thermospheric Regions; 

(1) Scientific Objective — To develop a relationship between 
the wind and temperature fields derived in experiments 
AS-2 and AS-3, and the inputs from excitation due to 
the solar radiation. 

(2) Method - The wind and temperature profiles derived 
through experiments AS-2 and AS-3 will be combined with 
measurements of the solar flux, Birkeland current, and 
particle precipitation. Energy balance calculations will 
be made through measurement of long wavelength infrared 
emissions. Estimates of the contribution of the albedo 
will be made from data collected by instruments carried 
on the Orbiter and the PDS. 


(3 ) Instruments Required - 
Instrument No. 


118 

122 

126 

1002 


Title 


Limb-Scanning Infrared Radiometer 
UV-VIS-NIR Spectrometer 
Infrared Interferometer 
Pyrhel 1 ometer/Spectrometer 


Supporting data from PDS. 


e. Experiment AS-5 — Determine the Eddy Diffusion Between the 

Altitudes of 85 km and 120 km: 

(1) Scientific Objective — To determine the rates of eddy 
diffusion, winds, and turbulence in atmospheric mixing 
phenomena . 

(2) Method — Measurements of winds and temperatures will be 
based on the doppler techniques used in experiments AS-2 
and AS-3. Diffusion will be measured by vertical pro- 
filing of selected constituents. Mea urements of gas 


4-9 



? 


I 


\ 


(3) 


release data and of data from the PDS instruments will 
be used as baseline for the reduction of data in this 
experiment. 

Instruments Required — 


Instrument No 
118 
122 
124 
126 
213 
532 


Title 

Limb-Scanning Infrared Radiometer 
UV-VIS-NIR Spectrometer/ Photometer 
Fabry-Perot Interferometer 
Infrared Interferometer 
(. .ser Sounder 
Gas Release Module 


Supporting data from PDS. 

f. Experiment AS-6 — Determine Atmospheric Interactions of 
Excited Radicals: 


(1) Scientific Objective — Significant portions of the total 
thermal energy in the upper atmosphere are believed to 
be held in molecular levels which do not possess radi- 
ative transitions. Excitation transfer by intermedi- 
ates such as carbon dioxide and hydroxyl radicals leads 
to radiative transfers. This experiment will ascertain 
the importance of the transfer mechanism in the overall 
thermal budget. 

(2) Method — Measurements of the hydroxyl vibrations levels 
will be made through induced fluorescence in the near 
UV spectral region. Collision excitation transfer will 
be estimated by comparing the measured level populations 
to theoretical populations in the absence of energy 
transfer. Principal measurements will be made through 
laser probing, supported by passive UV spectrography . 


4-10 


(3) Instruments Required - 




I 


Instrument No . 
116 
122 
213 

1 on 


Title 

Airglow Spectrograph 

UV-VIS-NIR Spectrometer/ Photometer 

Laser Sounder 

Ultraviolet Occultation Spectrograph 


Experiment AS-7 — Determine Atomic and Molecular Oxygen 
Densities Between 90 to 120 Kilometer Altitude; 


(1) Scientific Objective - To determine to a very high 

precision the densities of atomic and molecular oxygen 
as a function of geographic position (both global and 
small scale), season, time, solar dissociating flux, and 
other parameters (e.g., low-level auroral inputs, wind 
fields, etc.). 


(2) Method - Ai rgl ow emissions induced by the sun, particle 
precipitation, and laser probing v/ill be used. Solar 
and stellar occultation will assist in determining the 
density of atomic oxygen, and possibly of molecular 
oxygen. A precise determination of solar flux is 
essential for this experiment. The use of the SPS and 
PDS will provide data required to make a complete 
analysis . 

(3) Instruments Required — 

Instrument No. Title 


1 22 
124 
213 
1011 


UV-VIS-NIR Spectrometer/ Photometer 
Fabry-Perot Interferometer 
Laser Sounder 

Ultraviolet Occultation Spectrograph 


Supporting data from SPS and PDS. 


4-11 




h. Experiment AS-8 — Determine Solar Radiation Interaction with 

the Ambient Atmosphere: 

(1) Scientific Objective — To identify spectral transitions 
of long-lived metastable states which are pressure- 
quenched in ground based experiments. Baseline data 
will be provided for other experiments in the ASF. 

(2) Method — Clouds of neutral molecules of gas will be 
released from the Orbiter, and pre-ionized species may 
be released by the onboard accelerator instruments. 
Additional excitation may be produced by the onboard 
accelerator instrument, or may be provided by laser prob- 
ing and through use of electron beams. 

(3) Instruments Required — 


Instrument No. 

Title 

122 

UV-VIS-NIR Spectrometer/ Photometer 

124 

Fabry-Perot Interferometer 

213 

Laser Sounder 

303 

Electron Accelerator 

304 

Magnetopl asmadynami c (MPD) Arc 

532 

Gas Release Module 

534 

Optical Ban- Imager and Photometer 
System 

536 

Triaxial Fluxgate 

549 

Gas Plume Release 

550 

Level II Beam Diagnostics Group 


Supporting data from SPS. 


4-12 


Experiment AS-9 — Determine the Atmospheric Constituent 
Abundance Below 120 Kilometer Altitude: 

(1) Scientific Objective - To synoptically map the geographic 
distributions and vertical profile of atomic, molecular, 
and ionic abundance between the altitudes of 15 km and 
120 km. 

(2) Method — Airglow measurements will be used in the deter- 
minations for oxygen, nitrogen, and their compounds. 
Measurements in the UV. and infrared regions may.be 
required for determinations of ions and polyatomic 
species. Laser excitation may be useful in creating 

a promptly radiating state from weakly emitting species. 

( 3 ) Instruments Required — 


Instrument No. 

Title 

118 

Limb-Scanning Infrared Radiometer 

122 

UV-VIS-NIR Spectrometer/ Photometer 

124 

Fabry-Perot Interferometer 

126 

Infrared Interferometer 

213 

Laser Sounder 

1011 

Ultraviolet Occultation Spectrograph 


Supporting data from SPS and PDS. 

Experiment AS-10 — Determine the Atmospheric Constituent 
Abundance Above 120 Kilometer Altitude: 

(1) Scientific Objective — To synoptically map the geographic 
distributions and vertical profiles of atomic, molecular, 
and ionic abundance above the altitude of 120 km. 


(2) Method -- Data will be gathered primarily by occultation 
and by upward looking spectroscopy. The van Rhyn tech- 
nique of estimating spherical shell contributions based 
on changing angular absorption will be used to determine 
constituent distribution. 

(3) Instruments Required — 

Instrument No. Title 


122 

UV-VIS-NIR 

Spectrometer/ Photometer 

1 24 

Fabry-Perot 

Interferometer 

1011 

U1 traviol et 

Occultation Spectrograph 


Supporting data from SPS and PDS. 

Experiment AS-11 — Determine Change in the Ionospheric D 

Region Due to Seasonal Anomalies and Magnetic Storms: 

(1) Scientific Objective — To correlate the change in 
neutral composition, neutral species, temperature, ions, 
and particle flux to D Region propagation. 

(2) Method — Measurements of nitric oxide, water vapor, 
hydrated ions, ozone, and atomic oxygen abundance and 
temperature will be taken by spectrographic instruments 
and by laser probing techniques as used in experiments 
listed heretofore. Simultaneous measurements will be 
made of particle fluxes for correlation purposes and 

to assess their import to regions of high precipitation, 
such as the South Atlantic Anomaly. 




(3 ) Instruments Required 
Instrument No. 


Title 


UV-VIS-NIR Spectrometer/ Photometer 
Fabry-Perot Interferometer 
Infrared Interferometer 


Laser Sounder 


Supporting data from PDS. 

Experiment AS-12 — Determine the Metallic Constituents in 

the Upper Atmosphere: 

{1 ) Scientific Objective — To provide baseline data on the 
quantity and distribution of metals in the upper atmos- 
phere on a global basis. 

(2) Method — Spectrographi c data will form the primary source 
of information. Resonant backscatter from laser probing 
is a promising data source. Available techniques will 
probably measure only a limited portion of the total 
inventory of metals in the atmosphere. 


(3) Instruments Required 
Instrument No. 


Title 


Airglow Spectrograph 
UV-VIS-NIR Spectrometer/ Photometer 
Fabry-Perot Interferometer 
Infrared Interferometer 
Laser Sounder 

Ultraviolet Occultation Spectrograph 


Supporting data from SPS and PDS. 


4-15 


m. Experiment AS-13 - Evaluate Deposition of Meteoric Dust and 
Metallic Constituents: 


(1) Scientific Objective — To determine changes in the 
metallic content of the upper atmosphere due to meteor 
showers. This experiment will use the data resulting 
from Experiment AS-12. 

(2) Method ~ As in Experiment AS-12, spectrograph! c data 
will form the primary source of information, with 
resonant backscatter from the laser probing, if found 
to be a useful data source. The determination requires 
the advent of a substantial meteor shower subsequent to 
the results obtained in Experiment AS-12, with the 
measurements being made at the shower location. 

(3) Instruments Required — 

Instrument No. Title 


122 

124 

126 

213 


UV-VIS-NIR Spectrometer/ Photometer 
Fabry-Perot Interferometer 
Infrared Interferometer 
Laser Sounder 


Supporting data from SPS and PDS. 

n. Experiment AS-14 - Determine the Meteoric Production of 
Nitric Oxide; 


{!) Scientific Objective — To determine the amount of nitric 
oxide formed in the altitude region of 90 to 120 km by 
ionizing tracks of meteors. The measurements will be 
made after detecting the tracks with an onboard low 
light level television (LLTV) system. 


4-1 6 


(2) Method — Laser-induced fluorescence will be used for the 
quantitative determination of nitric oxide. It may be 
possible to quantitatively monitor the reactants spec- 
troscopical ly during the meteor shower, 

(3) Instruments Required — 

Instrument No, Title 


122 

124 

126 

213 

534 


UV-VIS-NIR Spectrometer/Photometer 
Fabry-Perot Interferometer 
Infrared Interferometer 
Laser Sounder 

Optical Band Imager and Photometer 
System 


0 . Experiment AS-15 — Investigate the Excitation Exchange 
Between Metastable Species and the Ambient Environment; 


(1) Scientific Objective — To study the quenching cross 
sections of metastable species at pressure levels and 
instrument volumes not available in ground laboratories. 

(2) Method — Gas clouds will be released as plumes or 
plasmoids and the energy transfer by the ambient 
photons and particle fluxes, or by active probing with 
electron beams, will be evaluated. Fluorescent decay 
will be observed with imaging devices and with the com- 
plement of spectrographi c instruments aboard the 
Orbiter. 


(3) Instruments Required — 
Instrument No . 

116 
122 
124 


Title 


Airglow Spectrograph 

UV-VIS-NIR Spectrometer/ Photo me ter 

Fabry-Perot Interferometer 


4-17 


Instrument No. 

Title 

126 

Infrared Interferometer 

303 

Electron Accelerator 

304 

Magnetopl asmadynamic (MPD) Arc 

532 

Gas Release Module 

534 

Optical Band Imager and Photometer 
System 

536 

Triaxial Fluxgate 

549 

Gas Plume Release 

550 

Level II Beam Diagnostics Group 


4.1.4 INSTRUMENTS (PRIME) 

4 . 1 . 4. 1 General 

Fifteen instruments have been described in sufficient detail to 
evaluate the feasibility of their construction. The basis for 
the id's v/as the IFRD's prepared by the Atmospheric Science 
Section during meetings of the AMPS SDWG at HSFC. Preliminary 
information in the IFRD's was supplemented by discussions with 
the scientists who drafted them and with scientists at the JSC. 
These instruments have been termed "prime instruments" for the 
purpose of this report, and the ID's have been incorporated 
as appendix B. 

The 15 prime instruments are listed in table 4. 1.4-1 in numer- 
ical order with the names derived from the IFRD's. The matrix 
in table 4. 1.4-2 relates the prime instruments to the experiments 
described in appendix A. Performance parameters are listed 
in table 4. 1.4-3 and interface parameters are listed in table 
4.1 .4-4. 


TABLE 4. 1.4-1. - ASF INSTRUMENT IDENTIFICATION 


Instrument 

Number 

Instrument Name 

116 

Airglow Spectrograph 

118 

Limb-Scanning Infrared Radiometer 

122 

UV-VIS-NIR Spectrometer/ Photometer 

124 

Fabry-Perot Interferometer 

126 

Infrared Interferomete-jr 

213 

Laser Sounder 

303 

Electron Accelerator 

304 

Magnetopl asmadynamic Arc 

532 

Gas Release Module 

534 

Optical Band Imager and Photometer System 

536 

Triaxial Fluxgate 

549 

Gas Plume Release 

550 

Level II Beam Diagnostics Group 

1002 

Pyrhel 1 ome ter/ Spectrophotometer 

1011 

Ultraviolet Occultation Spectrograph 






as-15 X XXX X X X X X X X 
























^ 'j 







Si 


I 

ro 


TABLE 4. 1.4-3. - ASF PRIME INSTRUMENT PERFORMANCE PARAMETERS 


Instrument 1 

Instrument Range 1 


Resolution 


No* 

f^a^le 

Frequency Spectral 

Energy 

Dynamic 

Spatial 

Spectral 

Sensitivity 

S/N Ratio 

116 

Airglflw Spectrograph 

30D A to 2000 A 



COIlFA-l'’ 

cmiFA-5“ 

0,5 to 2.0 A 



TT8 

Limb Scanning IR 
Radiaraeter 

3 to 4P micrometers 


10'^ 


(TBD) 

5x10”'* w cm"’ 
SR*' m-‘ 

(TBO) 

112 

UV-VIS-NIR 

Spectrometer- Pho tameter 

10,000 A to 
10 micrometers 



12'’xlZ" 

10 A 

2 Pliotoelectrpns 
Raleigh ‘ Sec * 


124 

Fabry-Ffirot 

Interferometer 

1 to 

150 micrometers 



3 km 

o 

1 A 

5x10 detected photons 
Raleigh’® Sec"^ 


126 

Infrared Interferorreter 

1 to 150 micrometers 


10® 

0,5 cm"' 

0,5 cm*® 

10’®® w cm’^ 
SR-* micrometer 

100:1 

213 

Laser Sounder 

lOIlO A to 30000 A 



1 km 

0,1 mrad 

,001 A 



303 

Electron Accelerator 


1 to 30 Lev 

0 to 7 Amp 

10^ Di- 
vergence 




304 . 

Ma gnetopl asmady nam i c 
Arc ' 


100 to 
500 

10^ to 
2x10® Amp 

40‘* Di- 
vergence 

l50;j 



534 

Optical Band imoger 
and Photometer System 

Depends upon 
experiment 

m 

(TBI?) 

* 

NA 

(TBD) 

(TBD) 

536 

Trl axial Fluxgate 

<0.1 Hz 

Passive 

^10^ 



5«10"’ Gauss 


532 

Gas Release Module 

3DD A- 

1.2 micrometers 



2 to 3 
Degrees 

0.2 A o 
(a 1200 A 

(TBD) 

(TOD) 

549 

Gas Plume Release 



0-0.5 i 

moles fac. 

i 

60*^ 




550 

550 

Level I I 

Beam Diagnostics 


0 to 30 keV 

_ 

10 attip/cm^ 
+0 amp/cm^ 



10"* Amp cm"^ 

(TBD) 

1002 

Pyrhel iometer/ 
Spectrometer 

0.2-5. 0 mierjmeters 
0.25-2.6 misrameters 

RA 

125-145 
*10”’w cm“^ 

NA 

60 to 200 

O.U- 

100D 

1011 

UV Dcoul tation 
SpectrograpH 

Q. 03-0.2 
micrometers 

NA 

CTBD) 

NA 

0,4 A 

(TBD) 

(TBD) 


*Dapands an experiment 



quality; 




TABLE 4. 1.4-4. - ASF INSTRUMENT INTERFACE PARAMETERS 


1 InstruEiient ~| 

Physical dimensions (metric) f 


iwer 


Po 

inting 

Data [see 
TTr^Z r 

note) 

No» 

Name 

Length 

Width 

Height 

Volume 

V'eight 

Vac 

Vdc 

Watts 

Error 

Stability 

acienL 

D, A, F 

1 It. 

Rate 

D, A, Dis. 

Rate 

116 

Airglow 

Spectrograph 

2.0m 

0.6m 


0.56 

30 kg 


28 

1 

^0.5“ 

«15 arc sec 

F 

700 

frames 

A 

480 bps 

118 

IF Limb 
Scanning 

1.8m 

O.flm 


4.52 11!^ 

115 kg 

115 

400 cycle 


100 

fiO.5 

si 5 arc sec 

D 

12 kbps 

A 

460 bps 

122 

UV-VIS-NIR 

Spectrometer/ 

Photometer 

0.5m 

0.2m 

0.2m 

0.02 

16 kg 


28 

16 

^0.1 

s6 arc min 

D 

8 kbps 

A 

320 bps 


Fabry-Parot 

Interferometer 

0.6m 

0.4m 

0.3m 

HtM 

latM 

0.4m 

0,6 m\ 
D.OS i\T 
D.18 m3 

45 kg 


28 

14 

5l,0° 

TBD 




560 bps 

IjM 

Infrared 

Interferometer 

0.7m 

m 


0,45 

114 kg 

115 

4GO cycle 

■ 

Bi 





Hli 

200 bps 



{fl su 

bsystem 

5) 

9.3 

415 kg 

115 

400 cycle 


1.08 

fcw 

si .0° 

T8D 



D 

1 kbps 

303 

Electron 

Accelerator 




6.1 

740 kg 


28 

5kW avg 
lOkW max 

6“ 

I'’ sec"^ 

D , 

5 kbps 

A 

16 bps 

304 

Hagnetoplasma- 
dynamic Arc 




2.5 

630 kg 


28 

5kW avg 
10 kW max 

2“ 

r sec"^ 

D 

1 kbps 

A 

16 bps 

534 

Optical Band 
Imager and 
Photometer System 

0.9m 

0.9m 

3.1m 

3 m^ 

100 kg 



30 

2“ 

1“ min"^ 

A 

D 

4 mz 
? kbps 

Q 

TBD 

536 

Triaxial 

Floxgate 

Bocm or 
mounted 

subsat 

sensor 

ellite 

0.005 

5 kg 


28 

4 



D 

600 bps 

Combined 

w/scientific 

532 

Gas Release 
Module 




1.24 

49 kg 


28 

140 

1“ 

O.IS*^ sec"^ 

D,A 

77,5 kbps 

Combined 

w/scientif* 

c 

549 

Gas Plume 
Release 




0.12 

9 


28 

5-10 

N/A 

N/A 

Video ta 
^^3 sec/r 

pB 

eleasc 

A 

16 bps 

S50 

Level II Beam 
Diagnostics 

3 Sut 

isystems 


0.005 

23 


28 

20 

N/A 

N/A 

D 

6.5 kbps 

A 

12 bps 

lfl02 

Pyrhel iometer/ 
Spectrometer 

0.3m 

0,1m 

' D.3m 

0.01 

<10 kg 



10 

2.5^^ 

N/A 

0 

320 bps 

Combined 

w/scientifi 

c 

]01! 

UV OccnUation 
Spectrograph 

3tti 

Im 

Itn 

3 n? 

128 kg 


28 

ICO 

1 arc 

urn 

10-15 
arc sec 

F 

Ifps 

A 

1 

I™ 


NOTE: D = Digital; A - Analog; F « Film; Dis = Discrete 


t> 




























J 


f 


r 

! 


The id's for several of the prime instruments call for attributes 
which will require advancement of the start-of-the-art with an 
appropriate development program. The practicability of using 
more readily available instruments, termed "substitute instruments," 
was assessed in the light of scientific and program requirements, 
and the options for use of the substitute instruments are des- 
cribed in section 7.0 of this report. 

4. 1.4. 2 Summary Descriptions 

The 15 ASF prime instruments derived from the IFRD's are sum- 
marized below, and detailed technical descriptions are contained 
in appendix B. 

a. Airglow Spectrograph, Instrument 116. The Airglow Spectro- 
graph is used to collect data for the study of upper atmos- 
phere emissions and absorptions in the vacuum UV range of 

O Q 

300 A to 2000 A, The instrument provides high spectral 
and spatial resolution in the collection of data, which is 
recorded on film in the form of spectrograms. The range 
of observations extends from zenith to nadir. The instru- 
ment has two configurations, either of which can be selected 
in flight, with one configuration having a f ield-of-view 
(FOV) of 5° square, and the second having a FOV 1° square. 

The operating volume of the instrument is 0.56 cubic meters, 
and the operating weight is 30 kilograms. 

b. Limb Scanning Infrared Radiometer, Instrument 118. The Limb 
Scanning Infrared Radiometer is a cryogenic multi-channel 
instrument which acquires data to permit measurement of trace 
species and evaluation of the vertical distribution of trace 
gases in the altitudes up to approximately 120 kilometers. 

The spectral range of operation is from 3 to 40 urn. Twelve 
detectors are incorporated into the instrument which is com- 
pletely encased in a dewar housing to maintain the cryogenic 
operating temperature of 77 K; the detectors are cooled to 


4-23 


4 K for operation. The operating volume of the instrument is 
is 0.9 cubic meters, and the operating weight is 115 kilo- 
grams plus 185 kg for cryogen dewar and associated plumbing. 

c. UV-VIS-NIR Spectrometer/Photometer, Instrument 122. The 
UV-VIS-NIR Spectrometer is used to obtain measurements of 
natural and induced atmospheric and ionospheric emissions in 
wavelengths ranging from 0.11 pm to 1 pm. The instrument is 
comprised of four small spectrometers of Ebert-Fastie config- 
uration, although as many as eight such instruments can be 
incorporated into the main spectrometer. The use of multi- 
ple spectrometers permits the simultaneous observation of 
several spectral features. Photomultiplier tubes are used 

as detectors. The operating volume of the spectrometer is 
0.02 cubic meters and the operating weight is 16 kilograms. 

d. Fabry-Perot Interferometer, Instrument 124. The Fabry-Perot 
Interferometer collects data which enable measurements to be 
made of doppler velocity and of temperature in the mesosphere 
and thermosphere using selected atomic line emissions in the 
UV, visible, and near infrared spectral regions. The large 
size of the etalons (i.e., 25 cm in diameter) permits high 
resolution and high etendue photometric studies of line and 
band emissions between the wavelengths of 0.2 pin and 10 pm. 

o 

The overall spectral range of the instrument is 2000 A to 
10 pm. The instrument operates in one of three different 
modes, interferometer, photometer, or radiometer, each pro- 
viding different sensitivity and different FOV. Mechaniza- 
tion of the instrument allows selection of operating mode 
during flight. The operating volume of the interferometer 
is 0.1 cubic meters and the operating weight is 45 kilograms. 

e. Infrared Interferometer, Instrument 126. The Infrared Inter- 
ferometer acquires data in the spectral region ranging from 

1 pm to 150 pm, in three descrete intervals. The instrument 
incorporates interchangeable filter/beam-splitter/detector 


4-24 





J 


combinations to cover each of the three spectral ranges; the 
combinations are assembled into the instrument prior to 
flight and are not changeable during flight. The instru- 
ment is cryogenical ly cooled to 77 K, and all components 
are enclosed within a dewar structural casing to maintain 
the requisite temperature during operation. The detector 
units are further cooled to 4 K for maximum sensitivity. 

The telescope is pointed at areas between the nadir and the 
horizon for collection of data. The operating volume of the 
instrument is 0.45 cubic meters and the operating weight is 
114 kilograms plus 186 kg for cryogen dewar and plumbing. 

Laser Sounder, Instrument 213. The Laser Sounder enables 
studies to be made of the composition, structure, and dynamics 
of the atmosphere through backscatteri ng and absorption of 
the laser beam. The primary area of concern is the upper 
atmosphere in the nadir direction from the Orbiter. The in- 
strument consists of the laser emitter and the receiving 
interferometer, as major components. The laser is a tunable 

o 

dye laser which operates over the spectral range of 1000 A 

o 

to 30000 A, and which has an output energy of one joule, a 
pulse duration of ten nanoseconds, and a pulse rate of one 
per second. The interferometer section receives the returned 
energy through a 2-meter aperture Cassegrain telescope, which 
directs the energy to an array of ten Fabry-Perot etalons 
which separate the beam into discrete spectral bands which 
then impinge on the photomultiplier tube detectors. 

The size of the interferometer telescope, i.e., 2-meter aper- 
ture, presents a serious problem in the accommodation array 
of instruments in the Orbiter payload bay, for the dimension 
far exceeds the available envelope for the Laser Sounder. 

The aperture has been reduced to'^O.S meter diameter for 
accommodation purposes. The operating volume of the instru- 
ment is 9.21 cubic meters, and the operating weight is 


4-25 


f 

! 


415 kilograms; with the reduction of telescope aperture, 
the operating volume is reduced to 5.51 cubic meters, and 
the operating weight is 395 kilograms. 

g. Electron Accelerator, Instrument 303. This instrument is a 

subsystem of the AMPS Particle Accelerator System and will be 
used to: (1) study the excitation of upper atmospheric and 

ionospheric constituents, (2) map the magnetic field lines 

of the earth, (3) determine ionospheric electric field 
magnitude and direction, and (4) study the plasma wave 
excitation in the ionosphere. It consists of an electron gun 
with variable energy and current output up to 30 keV and 7 
amperes respectively. Operation of the electron beam can be 
continuous direct current (dc), pulsed, or modulated (up to 
10 MHz). Energy storage for high intensity pulsed operation 

5 

is accomplished with a 10 joule, 500 volt capacitor bank. 

The operating volume for this instrument is 6.1 cubic meters 
and the operating weight is 740 kilograms.* 

h. Magnetopl asmadynamic (MPD) Arc, Instrument 304. The MPD Arc 

is a subsystem of the AMPS Particle Accelerator System. It 
will be used to: (1) study the excitation of upper atmos- 

pheric and ionospheric components, (2) trace and map the 
earth's magnetic field lines, (3) modify the conductivity in 
certain regions of the ionosphere, and (4) generate plasma 
waves in the very low frequency/extremely low frequency (VLF/ELF) 
regimes. The instrument consists of a low voltage plasma gun 

(up to 500 volts) with a discharge current up to 2 x 10 amperes. 
Energy storage for high intensity pulses is accomplished with 

5 

a 10 joule, 500 volt capacitor bank. The operating volume of 
this instrument is 2.77 cubic meters and the operating weight 
is 630 kilograms.* 

*N0TE : Weights indicated apply if either instrument is flown 

without the other instrument. If both instruments 303 and 304 are 
flown, the combined weights will approximate 781 kilograms due to 
the common usage of certain power components. 




} 


Gas Release Module, Instrument 532. This instrument will be 
used to study photoexcitation and photoionization of various 
species exposed to solar radiation. In addition, it will be 
utilized to study the decay of excited species including 
metastable states. Gas will be admitted to an excitation 
chamber which is viewed by a monochromator and a quadrupole 
mass analyzer. The chamber will be exposed to the full un- 
attenuated solar flux. Ion masses in the range of 1 atomic mass 
unit (amu) to 100 amu can be measured and by the use of three 
monochromators (one selected and mounted before flight), wave- 
lengths in the ranges 300 A to 1500 A, 1100 A to 4500 A and 

o 

4000 A to 1.2 pm will be measured. Metastable states will be 
measured by a free gas release to space which is viewed by the 
monochromator. The operating volume of this instrument is 1.42 
cubic meters and the operating weight is 49 kilograms. 

Optical Band Imager And Photometer System (OBIPS), Instru- 
ment 534. The OBIPS obtains monochromatic images of airglows 
due to natural aurora and atmospheric perturbation experi- 
ments such as chemical releases and high energy electron 
injections. The optical bandwidth is just sufficient to pass 
the radiation of a particular molecular band. The configura- 
tion depends upon the mission. The typical configuration 
has two LLTV's and two photometers operating at two different 
wavelengths. The TV's are used to point the narrow field 
photometers and the latter give accurate radiometric read- 
ings. A very large baffling system precedes the lens in 
order to block extraneous radiation and obtain data of 
faint airglows despite sunlight scattered by the earth's 
atmosphere. The optical band is determined by filters which 
are interchangeable. The operating volume of this instrument 
is 0.52 cubic meters and the operating weight is 100 kilograms. 

Triaxial Fluxgate, Instrument 536. The objectives of this 
instrument are to: (1) study the natural hydromagnetic wave 

propagation, (2) probe the ultra low frequency (ULF) noise 
generated by the Orbiter, (3) study noise generated by 


4-27 


f 


controlled discharge from the ULF antenna, and (4) determine 
the magnetic environment of the Orbiter as a safety measure 
during accelerator operation. Because of lov; electromagnetic 
field interference requirements, the instrument will be sub- 
satellite or boom mounted. The sensors are orthogonally 
mounted coils on high permeability cores. Sensors will re- 
quire about 0.003 cu m volume. The operating weight of this 
instrument is 5 kilograms. 

l, Gas Plume Release {AMPS Particle Accelerator System Level I 
Diagnostic), Instrument 549. The Gas Plume Release will be 
used for optical tailoring and alignment of the particle beam 
from the Electron Accelerator {Instrument 303). The Gas 
Plume Release system resides within the volume of the elec- 
tron accelerator and consists of a gas storage system from 
which gas can be released from four jets. Interaction of 
either the ion or electron beam with the gas will allow a 
visual observation of the profile of the beam. The operating 
volume of this instrument is 0.12 cubic meters and the 
operating weight is 9 kilograms. 

m. Faraday Cup Probe/Retarding Potential Analyzer/Cold Plasma 
Probe/{AMPS Particle Accelerator System Level II Diagnostic 
Group), Instrument 550. This group of instruments will be 
utilized to define the energy, beam intensities and profiles 
of the Electron Accelerator {Instrument 303) and to determine 
the rise in potential of the Orbiter with respect to the 
ambient plasma during accelerator firing. The Faraday Cup 
Probe is a cylindrical cavity current collector and will be 
utilized to determine the spatial profiles and intensities 

of the beams. The retarding potential analyzer will deter- 
mine beam energy and will operate up to 30 keV. The cold 
plasma probe is a passive floating potential probe and will 
be used to measure Orbiter charge build-up. The operating 
volume of this instrument is 0.005 cubic meters and the 
operating weight is 23 kilogram*:. 


4-28 


1 


n. Pyrhel iometer/SpectrophotOi.i *’r, Instrument 1002. The two 
instruments are combined in ^ne small package with a single 
data output. The design is called the Solar Energy Monitor 
in Space (SEMIS). The system is optimized for accurate quan- 
titative measurements. The pyrhel iometer is the thermopile 
type, modified from a commercially available design which is 
used as a standard radiation detector. The range of radia- 
tion detected is 0.2 pm to 5.0 pm. The spectrophotometer 
views solar radiation reflected from a diffuse plate, thus 
no scanning of the sun is necessary. The radiation is dis- 

^ persed by a quartz Littrow monochrometer. A beamsplitter 

divides the radiation into two spectra which are detected with 
a photomultiplier and lead sulfide detector. Ten minutes 
is required for a scan. The spectral range is 0.25 pm to 
2.6 pm, but by changing to sapphire optics it is expected to 
go to 4 pm. The operating volume of this instrument is 0.01 
cubic meters and the operating weight is 10 kilograms. 

0 . Ultraviolet Occultation Spectrograph , Instrument 1011. As the 
sun or a star appears to approach the limb of the earth, at 
certain wavelengths molecules and free radicals absorb radia- 
tions. This instrument measures the absorption, so the con- 
centration as a function of altitude may be calculated. The 
initial value of the radiation is obtained when the sun or 
star is at a distance from the limb. A series of spectra 
are obtained with the sun or star at different distances from 
the limb. Two configurations are used, one for stellar and | 

the other for solar occultation. Stars have a better con- 
tinuum but the sun has a stronger signal. A Cassegrain tele- 
scope focuses on the slit of the spectrograph. For stellar 
occultation, the telescope is large. A concave grating focuses 
the spectrum on an opaque photocathode. The photoelectrons 
are emitted in the direction of the incident radiation, ac- 
celerated by an electric field, focused by a magnetic field 
and impinged upon film with a thick emulsion of the type made 

■I ■ 

; 


I 

J 


4-29 


for recording nuclear particles. The spectral range is 

0 0 o 

300 A to 2000 A. The resolution is approximately 0.4 A. 

The operating volume of this instrument is 1.66 cubic meters 
and the operating weight is 125 kilograms. 

4,1.5 OVERVIEW, ASF MISSION TIMELINE 

Operational timing of the instrument in each experiment is in- 
cluded in the ED's, and graphically depicted for the ASF payload 
in the timeline shown in figure 4. 1.5-1. Although the timeline 
was developed without regard to whether the mission would be polar 
or low inclination orbit, global coverage from high inclination 
as well as low inclination orbits is required to satisfy the ASF 
experimental objectives. Since the west coast launch site which 
is required for polar orbits will not be completed until after 
1981, the early ASF flights will be flown in low inclination 
orbits with the result that not all of the experiment objectives 
will be achieved on early flights. The objectives of many experi- 
ments may, however, be completely achieved on orbits of low 
i ncl i nation . 


























4.2 PARTICLE DETECTOR SUBSATELLITE (PDS) REQUIREMENTS 

4.2.1 INTRODUCTION 

A subsatellite will be used as the platform on which the particle 
detection instruments will be mounted. The instruments will 
provide the necessary particle data in support of the experiments 
being conducted by the ASF. This subsatellite will be of the 
AE type. The subsatellite configuration and a description of the 
subsatellite are contained in paragraph 5.2.6 of this report. 
Operations of the subsatellite are contained in section 6.0, 

4.2.2 REQUIREMENTS 

The functional requirements of the PDS are the following. 

a. Measure energy of electrons, protons and plasma potentials. 

b. Measure energy levels, drift velocities, temperature, mass 
and quantity of ions. 

c. Measure mass of neutral particles. 

d. Measure gas temperature and density. 

e. Detect upper atmosphere emissions in spectral lines at 
specific wavelengths and within ranges of wavelengths. 

f. Measure the instantaneous components of the magnetic field 
vector. 

4.2.3 INSTRUMENT SUMMARY DESCRIPTIONS 

The subsatellite instrument complement required to support the 
ASF experiments is listed in this section with comments on their 
use. The instruments are described in Radio Science, Volume 8, 
Number 4, April, 1973, Special Issue: The Atmosphere Explorer 

Satellite. Interface and performance parameters are listed in 
tables 4. 2. 2-1 and 4. 2. 2-2. 


4-32 


a. (CEP) Cylindrical Electrostatic Probe — Low energy electrons 
and plasma potentials at levels from 0 to 20 eV. 

b. (RPA) Planar Ion Trap — Ion drift velocities^ temperature, 
mass and quantity) 

c. (PES) Photoelectron Spectrometer — Electrons with energy 
ranges from 2 eV to 500 eV. 

d. (LEID) Low Energy Ion Detector — H , He , and 0 with energy 
levels ranging up to 10 KeV. 

e. (LEE) Low Energy Electron Detector — Electrons with energy 
ranges from 200 eV to 25 KeV. 

f. (NACE) Neutral Mass Spectrometer — Mass values for neutral 
particles from 1 to 47 amu. 

g. (NATE) Neutral Atmosphere Temperature — Gas temperature 
measurement using Nitrogen (N 2 )* 

h. (HEPD) High Energy Particle Detector — Covering the ranges 
of energetic electrons and protons in the range from 25 

up to 10 MeV. 

h. (VAE) Air glow Photometer — Detecting the upper atmosphere 

emissions in the spectral lines at 3371, 4278, 5200, 5577, 

0 o 

and 6300 A, and in the band from 7319 to 7330 A. 

i. (CCIG) Cold Cathode Ion Gauge — Gas density measurement. 

j. (MAG) Triaxial Fluxqate Magnetometer — Measure the instan- 
taneous vector components of the local magnetic field. 


The data obtained by the above instruments will provide all the 
information concerning the particle, electron and ion, environ- 
ment that is of direct concern in the analysis of the data from 
the atmospheric science experiments. 


4-33 







-35 


TABLE 4. 2. 2-2. - PDS PERFORMANCE PARAMETERS 


j Instrument 

Instrument Range 

No. 

Name 

Frequency/ Spectral 

Energy 

CEP 

Cylindrical Electrostatic Probe 

NA 

0 to 20 eV 

RPA 

Planar Ion Trap 

1 to 40 amu 


PES 

Photoelectron Spectrometer 

NA 

2 to 500 eV 

LEID 

Low Energy Ion Detector 

“f* 

H ,He ,0 

up to 10 KeV 

LEE 

Low Energy Electron Detector 


200 eV to 25 KeV 

NACE 

Neutral Mass Spectrometer 

I to 64 amu 


NATE 

Neutral Atmosphere Temperature 

500 to 5000 K 


HEPD 

High Energy Particle Detector 


25 KeV to 10 MeV 

VAE 

Airglow Photometer 

3371 ,4278,5208 , 
5577,6300,7519-7330 A 


CCIG 

Cold Cathode Ion Gauge 

10"^ to 10~® torr 


MAG 

Triaxial Fluxgate Magnetometer 

NA 













4.3 SOLAR PHYSICS SATELLITF 

4.3.1 ASF SUPPORT INSTRUMENTS 

Solar radiation, both v/ave and particulate matter, into the 
atmosphere is the prime energy input to which the atmosphere 
dynamics respond. The particulate input can be measured by 
instrumentation on the PDS discussed in the preceding section. 
This instrumentation is needed for measuring other experimental 
parameters . 

In the case of electromagnetic energy emanating from the sun, 
however, all necessary data can be obtained from a SPS which 
is planned for late 1970's deployment. This effort, which is 
being planned by the Solar Physics Working Group, has as an 
objective; the detailed investigation of solar phenomena on an 
instantaneous, as well as multi-year basis. Use of the data 
from this satellite will eliminate the need for extensive solar 
instrumentation on the Orbiter, leaving space and support facil- 
ities for other needed instrumentation. All that is required 
on the Orbiter is relatively simple instrumentation to be used 
to calibrate the data from the SPS. A Pyrhel i ometer/Spectrometer 
(Instrument 1002), provides this capability. 

4.3.2 INTERFACES 

Data from the SPS can either be received and processed on the 
Orbiter in real time, or received and processed on the ground for 
subsequent correlation with other experimental data. 

The sampling rate from the satellite is not critical. One sample 
of data at all wavelengths each minute appears sufficient. 
However, the optimum sa.Ttpling rate can be found only by examining 
experimental data to determine how rapidly the changes occur. 

The data rate depends upon the number of wavelengths and energy 
intervals sampled. 



4-36 


5.0 ASF SYSTEM DESCRIPTION AND INTEGRATION 


5.1 GENERAL 

5.1 .1 ASF SYSTEM ELEMENTS 

The ASF System elements are the flight, ground, and support 
systems. The flight system consists of; (1) the instruments, 
(2) the PDS, and (3) the support subsystems. For the purpose of 
this study, the ASF ground system consists of: (1) the ASF pay- 
load and ground support equipment, and (2) the ASF unique data 
handling facility. The support systems are part of the national 
space program inventory of facilities shared by all payloads. 
These include: (1) the Orbiter, (2) the TDRSS, (3) the SPS, 
and (4) the STS ground facilities. 

Figure 5. 1.1-1 shows the interrelationship among the major ASF 
system elements discussed in this report during each operational 
^hase of the ASF flight system. 

Test and integration of the ASF payload will occur at various 
levels (pallet, integrated payload, and integrated Orbiter). 

The basic ASF ground support hardware and software will be 
required together with simulators at each level. The ASF ground 
support equipment (GSE) will be utilized at both Kennedy Space 
Center (KSC) and at the western launch facility to support the 
prelaunch and launch activities. For prelaunch support, the 
ASF GSE will be integrated into the Orbiter Processing Facility 
(OPF) and the Vertical Assembly Building (VAB). For launch 
support, the ASF GSE will be integrated into the Launch Control 
Center (LCC). After the payload is returned from orbit, it is 
removed from the Orbiter at the OPF and is refurbished and 
retested in an ASF dedicated facility. The ASF GSE will be 
required to support operations during this phase. 



7 


tn 


I 

rsD 



lEGEND: 


GSE GROUND SUPPORT EQUIPMENT 

LCC ■ LAUNCH CONTROL COMPLEX 

MCC MtSSlQN CONTROL COMPLEX 

□PF OROITER PROCESSING FACILITY 

STDN- SPACE TRACKING & DATA NETVl'ORK 

TORSS TRACKING & DATA RELAY SATELLITE SYSTEM 

VAD ■ VERTICAL ASSEMBLY BUILDING 


Figure 5. 1.1-1 


ASF sys-;enib operational relationship. 


t 


■ 

































During the flight phase, the ASF payload primarily operates 
automatically, sequencing the experiment and support subsystem 
operations and processing the instrument data through the 
onboard ASF computers. However, Orbiter vehicle and crew opera- 
tions are required to support the missions. The Orbiter vehicle 
will orient the payload to approximately the right direction, 
providing a stable platform from which the payload pointing and 
tracking system can operate, and will change the orbit to 
rendezvous with (‘he subsatellite for retrieval operations. The 
payload specialise (PS) will initiate and interrupt preprogrammed 
experimental sequences, check initial conditions, perform manual 
operations, make decisions for off-nominal conditions, and 
perform real time updat. and changes to sequences. 

The ASF flight system depends upon: (1) Support of the SP3 

to provide critical correlative experimental data, (2) Mission 
Control Center (MCC) to provide monitoring and diagnostic 
support, and (3) ASF ground facilities to provide the required 
data cataloging, segregating, storage and dissemination 
required for billions of bits of data. The communication links 
between the ASF payload onboard the Orbiter, or the SPS, and 
the grour'd facilities will be provided by the STDN and the 
TDRSS. The communication betv/een the ASF and the Orbiter is 
provided through the attached payload interface. These inter- 
faces are shown in figure 2-1. 

5.1.2 ASF SYSTEM CONFIGURATION 

a. Configuration - Figure 2-1 shows the ASF flight system 

configuration including the instruments onboard the Orbiter 
and the PDS. 


Instrument placement in the ASF pallet-only mode study is based 
on the optical sensor pointing requirements which are too severe 
for the Orbiter reaction control system (RCS), on the necessity of 



avoiding mutual interference between instruments, and on the 
desire to keep similarly operating instruments together. The 
order of the platforms in the Orbiter payload bay is dictated 
by the clearances required to permit full articulation of the 
pointing structures. 

The instruments for the ASF mission are arrayed in the pay- 
load bay in four groups, each on a separate pallet. Two pallets 
are fitted with an AMPS Pointing System (APS), while the remain- 
ing pallets have the instruments or facilities mounted on non- 
maneuverable accommodations. The pallets are identified (for 
the purposes of this report) numerically from the forward end 
of the payload bay. The forv/ard edge of Pallet A-1 is at 
Station Xq 685.5 and the aft edge of Pallet A-4 is at Station 
Xq 1157.5. Figure 5. 1.1-2 depicts the general pallet arrange- 
ment within the payload bay. 

The first pallet has a steerable platform carrying instruments 
213 (Laser Sounder), 532 (Gas Release Module), 534 (Optical 
Band Image and Photometer System), 1002 ( Pyrhel iometer/Spectro- 
meter), 1011 (Ultraviolet Occultation Spectrograph) and 550 
(Level II Beam Diagnostic). Instrument 1002 will be used to 
verify calibrations of an identical device on the sun-synchronous 
SPS. Only one revolution should be required for data acquisition 
to verify the calibration. Instrument 532 will require only 
one revolution for each of the types of gas to be released. 

Instrument 534 will be used with the accelerator package when ^ 

it is being operated. Instrument 213 is the principal sensor 
on the platform operating for as much of the time as possible. 
Instrument 1011 is used on a time available basis when either 
the sun or a UV rich star is in the proper orientation. 

The second pallet carries the PDS. This location is a compromise 
between the desired forward location of a light package, 

f ■ 

1 

I 


5.1-4 




LOAD RETENTION BEAM 
LOWER ATTACH POINTS 


Figure 5. 1*1-2* -ASF pallet attachment locations. 


I 


i 

i 


r 


platform maneuvering requirements, and access by the vehicle 
RMS which is required for subsatellite recovery. 

The third pallet has a second steerable platform which carries 
instruments 116 (Airglow Spectrograph), 118 (Limb Scanning 
Infrared Radiometer), 122 (UV-VIS-NIR Spectrometer/Photometer ) , 
124 (Fabry-Perot Interferometer), and 126 (Infrared Interfer- 
ometer). With the exception of Instrument 116, this package 
will generally be pointed at the same atmospheric feature for 
simultaneous data collection. Instrument 116 will be used less 
frequently than the others and would have been assigned to the 
first pallet/platform, if mounting space were available in that 
position. 

The fourth pallet contains the particle accelerators, an accel- 
erator beam diagnostic, a magnetometer for pointing, instruments’ 
303 (Electron Gun), 304 (Magnetopl asmadynami c Arc), 536 (Tri- 
axial Fluxgate), and 549 (Gas Plume Release). These instruments 
are assigned to the rearmost pallet because of their total 
weight, so that a favorable vehicle center of gravity (CG) may be 
mai ntai ned . 

The APS, which is symmetrically located on the floor of pallets 
A-1 and A-3 has a central column which contains the mechanism 
for the deployment and retraction of the instrument modules 
and which forms the axis for the coarse azimuthal rotation of 
the modules. At the upper end of the central column, tv^o 
identical yokes accommodate the ASF Instrument Modules (AIM), 
one on each side of the column. The yokes pivot to provide 
vertical rotation of the modules and have the capability for 
almost full-circle rotation. One AIM is installed in each yoke, 
on provisions which allow for fine azimuthal rotation of up to 
five degrees either side of the nominal. The yoke-mounting 
provision is the primary interface of the AIM with the APS. 


5.1-6 


t 

i 


I 


Figure 5. 1.1-3 depicts the concepts of the APS and the AIM, 

The detailed descriptions of APS and AIM are contained in 
paragraph 5.2.3 of this report. 

Figure 5. 1.1-4 shows the instruments installed in the stowed 
condition and figure 5. 1.1-5 shows the instruments in the 
operational configuration. 

The ASF support subsystems consist of the following: 

a. Thermal, Structural and Mechanical Subsystem (TSMS). 

b. Electrical Power and Distribution Subsystem (EPOS). 

c. Pointing Control and Stabilization Subsystem (PCSS). 

d. Command and Data Management Subsystem (CDMS). 

The support equipment and operations required at the aft crew 
station are also discussed in later sections of this report. 

The ASF flight system instrument complement onboard the Orbiter 
and onboard the PDS, and the support subsystems onboard the 
Orbiter are listed in tables 5. 1.1-1 through 5. 1.1-3. The 
support subsystems onboard the PDS are the basic AE subsystems 
and will not be discussed in great detail in this report. 
Instruments are discussed in detail in section 4.0 and appendix 
B of this report. 

5.1.3 ASF SYSTEM INTERFACES 

The ASF flight system interfaces are illustrated in figure 
5. 1.1-6. These interfaces include those within the ASF payload, 
those between the ASF payload and the Orbiter, and those between 
the ASF payload and other systems which are linked with the ASF 


5.1-7 














Instruments/Item 


Instrunrents 

ne-AIrglaw Spectrograph 
na-LiTjib Scanning IR Radiometer* 
122-UV-VI.S-HIR Spectrometer 
1 24-Fabry-Perot Interferometer 
126-IR Interfere .titer* 

2 ' J-Laser Sounder 

c Emi tter/col 1 imator 
Capacitor Bank 

• Interferometer 
■ Electronics 

303- Electron Accelerator 
« Power Unit 1 

» Capacitor Bank 

• Power Unit 2 
f Accelerator 

304- Magnetopl asmadynami c (MPD) Arc 

• Power Unit {share with 303} 

fl Capacitor Bank(share with 303) 

• Arc Generator 
532-Gas Release Module 

• Gas System 

ft Excitation Chamber 
ft Monochromator 
ft Mass filter 
t Electronics 

534-Optical Band Imager and Photometer System 
f TV Cameras 
ft Photometers 
536-Tria3tial Fluxgate 

549- Gas Plume Release 

550- Faraday Cup Retarding Potential Analyzert 
(RPA) Cold Plasma Probe 

• Faraday Cup Probe 
i RPA 

ft Cold Plasma Probe 
1002 Pyrhel iometer and Spectrophotometer 
loll UV Occultation Spectrograph 
ft Telescope 
ft Spectrograph 
ft Solenoid 


Total Instruments 


PALLET INSTRUMENT CHARACTERISTICS 


qty 

To tal 
Weight 
kg(LB) 

Qperati ng 
Power 

(Wa tts-Ave ) 

Unit 5ize-L(or D)LnWxH or 
Diam X L Meters (Ft, ) 

1 

30{66} 

10 

0.6 Diam x 2.0 L (0.3Z Diam x 6,46 L) 

1 

300(662) 

TOO 

0.8 Diam X 1.8 L (2.58 Diam x 5.83 L) 

1 

J6{35} 

16 

0.5 > D.2 0.2 (1.62 x 0.65 - 0.65) 

1 

45{99) 

14 

0.3 Diam x 0.6 L (0,97 Diam x 0.94 L) 

■ 

300(662) 

25 

0,9 Qiam x 0.7 L (2.91 Diam x 2.26 L) 

■ 

100(221) 

Ik 

1.0 X 1.0 X 2-0 (3.23 X 3.23 - 6.46) 

■■ 

250(552) 

- 

1,0 X 1,0 X KO (3,23 X 3.23 x 3.23) 

1 

5Q(ltO) 

25 

2.0 Diam x 1.0 L (6.46 Diam x 3,23 L) 

1 

15(33) 

50 

0,2 X 0.2 X C.2 (0.65 x 0-65 * 0.65) 



5k 


1 

45(99) 


0,5 X 1,0 X 0.5 (1.62 X 3,23 x 1,62) 

1 

540(1393) 


0.5 X 3,0 x 1,5 (1.62 X 9.69 x 4.84) 

1 

110(243) 


1.0 X 1.0 X 0.5 (3.23 X 3,23 « 1.62) 

1 

41(91) 


3.0 X 1,0 1.0 (9. 69 X 3,23 x 3,23) 



5k 


1 

41(91) 


2.0 X 0.3 X 0.6 (6.46 x q, 97 1,62) 



140 


1 

23(51) 


l.D X 0. 5 X 0. 3 (3, 23 x 1.62 x 0- 97) 

1 

2(5) 


0.5 X 0.3 X 0,1 (1.62 X 0,97 « 0,32) 

1 

11(25) 

1 

1,8 X 1,0 0*5 (5.82 X 3,23 « 1,62) 

1 

9(205 


0.5 Qiam x 0.8 U (1,62 Piam x 2.50 L) 

1 

3(7) 


0.3 X 0,3 X 0.3 (0.97 x 0.97 x 0.97) 


2»50(221 ) 



? 


2Q( ea) 

0.2 X 0,2 K 1,3 (0.65 X 0,65 x 4,27) 

2 


5(ea) 

Within camera envelope 

1 

3(7) 

4 

0.1 X 0,1 X 0.1 (0.32 X 0,32 X 0-32) 

1 

9(20) 

5 

Contained within 303 envelope 



10 


1 

9(20) 


0.1 X 0.1 X 0.1 (0,32 X 0. 32 0.32) 

1 

9(20) 


0.2 X 0.2 X 0.2 (0.65 x 0.65 x 0.65) 

1 

5(11) 


0.1 X 0.1 - 0.1 (0.32 X 0-32 X 0.32) 

■ 

10(22) 

10 

0,3 X 0,3 X 0,1 (Q.97 K 0.97 x 0.32) 

D 

100(221 ) 

- 

1.0 Diam x 2.0 L (3.23 Diam x 6,42 L) 

IH 

20(44) 

1 

1.0 X 0.3 X 0,3 (3.23 * 0.97 x 0,97) 

iH 

5(11) 

100 

Within telescppe enveippe 

IH 

2203[4a5S} 




5 . 1 -n 












TABLE 5. 1.1-2. -ASF PARTICAL DETECTOR SUBSATELLITE 
SYSTEM CHARACTERISTICS 


Item 

Qty 

Total 

vieight 

kg (lb) 

Operating 

power 

(Watts-Ave) 

Physical Layout 

1. 

Cylindrical Electrostatic 
Probe (CEP) 

1 

1.9 (4.2) 

5 


2, 

Photoelectron Spectrometer 
(PES) 

2 

2x4.1 (18.2) 

5 


3. 

Lovj Energy Ion Detector 
(LEIO) 

2 

2x6.0 (26.5) 

10 


4. 

High Energy Particle 
Detector (HEPD) 

2 

2x7.0 (30.8) 

6 


5. 

Low Energy Electron 
Detector (LEE) 

2 

2x4.2 (18.4) 

5 

See Figure 5. 2. 6-3 

6. 

Airglow Photometer (VAE) 

1 

8,6 (19.0) 

4.5 


7. 

Tri axial FI uxgate 
Magnetometer (HAG) 

1 1 

1.2 (2.6) 

3.1 


8. 

Planar Ion Trap (RPA) 

1 

5.1 (11.2) 

6 


9* 

Neutral Mass ' ' 'ctrometer 
(NACE) 

1 ' 

1 

8.3 (18.3) 

18 


10. 

Neutral Atmospheric Temp. 
(NATE 

1 

9.2 (20.3) 

17.5 


IK 

Cold Cathode Ion Gauge 

1 

2.5 (5.5) 

1.5 


Total Instruments 


79.3 (175.0) 

81.6 


Satellite Structure and 
Support Equipment 


662.8 (1372.9) 

150 


Total 

Subsatel 1 i te 


702.1 (1547,9) 

231.6 





■/ 








TABLE 5. 1.1-3. -ASF SUPPORT SUBSYSTEM CHARACTERISTICS 




Thermal , Structural , Mechanical Subsystem 

a. Pallet A-1 

(1) Pallet Structure 1 

(2) APS 1 

(3) Boom & mechanism 1 

(4) Cold pi ate » thermal capacitor 1 se 

b. Pallet A-2 

^0) Pallet structure 1 

(2) Subsatellite launch platform 1 

c. Pallet A-3 

(1) Pallet Structure 1 

(2) APS 1 

d. Pallet A-4 

(1) Pallet Structure 1 

(2) Boom S mechanism 1 


(3) Cold plate* thermal capacitor 

e» Active thermal control loop, 
pump, coolant 

f. Cryogenic coolant storage and 

distribution system 

0) Tank, plumbing, valves 
(2) Cryogen 

g. Igloo Structure 

Total TSMS 


2, Electrical Power & Distribution Subsystem 

a. Emergency battery (igloo) 

b. OC/AC inverter (igloo) 

c. Power control box (igloo) 

d. Secondary power dist, box (igloo) 

e. Pallet distribution box (pallets) 

f. Harnesses 

(1 } 4/0 gauge 
(2; 4 gauge 

(3) 10 gauge 

(4) 20 gauge 


Total EPOS 


3. Pointing Control & Stabilization Subsystem 
a* Gyro reference assembly (pallet A-3) 

b. Star tracker assembly (palletA-3) 

c. Sun sensor (pallet A-1) 

d. Optical alignment measuring device 

(pallet 1/pallet A-3) 

e. Signal processing electronics 

(pallet 1/pallet A-3) 


Total PCSS 


428 (945) 
1100 (2426) 
28 (62) 

39 (86) 

428 (945) 
10 ( 22 ) 

42S (945) 
110? (2426) 

42B (945) 
3 (7) 

39 (86) 
105 (232) 


4x52 (456) 
93 (204) 
55 (121) 



Operating 

power 

(Watts-Ave) 

Unit size - meters (ft) 

L - length, W - v/idth, D = depth 
diam = diameter, H = height 


See figure 5. 2, 1-1 

200 

See figure 5. 2. 1-6 

4 

Boom - 0.D79 diam x 18, OL (C.26 diam 
X 59.10L) 

- 

See figure 5. 2. 1-3 

- 

See figure 5. 2. 1-1 

_ 

See figure 5. 2. 1-1 

2Q0 

See figure 5.2.K6 

- 

See figure 5. 2. 1-1 

4 

Boom - 0,013 diam x 20. OL (0.043 diam 
X 65.61) 

- 

See figure 5. 2. 1-3 

200 

0,76 diam (2,5 diam) 


! 

0,95 diam x K5L (3.12 diam x 4.92L) 
(internal ) 


4494 (9909) 


78 (172) 


6 (13) 

500 

6 (11) 

20 

6 (13) 

10 

8x4,5 (102) 

8x10 

290 (640) 
40 (88) 

- 

55 (122) 

- 

41 (90) 

- 

567 (1251) 


30 (66) 

100 

3x11 (73) 

75 

13 (29) 

10 

12 (26) 

30 

20 (44) 

40 

108 (239) 



D.46Lx0.37Hx0.24H (1 .5QLx 1 .20WxQ.a0H) 
0 . 4QLxO , 25WxO . 1 5 H (1.311x0. 82WxO . 49 H ) 
0.31LxQ.15Wx0.13H (KOOLxO.5DUxO.42H) 
0.31Lx0.15Wx0*13H (1 .aQLx0.50Wx0.42H) 
G.25Lx0.I5UxO.13H (0.83LxO, 50Wx0, 42H) 

1831 (60QL) 

183L (600L) 

1097L (3600L} 

8,536L (28,0001) 


0.l8Lx0,25Wx0.20H (0.59Lx0.82Hx0.66H) 
0.60Lx0.21Wx0.21H (K97LxO.69WxO.69H) 
0.15Lx0.30Wx0.20H *'?»49LxQ.99Wx0.66H) 
0.l0Lx0.25WxQ.lQH (0.33Lx0.82Wx0. 33H) 

0.15Lx0.40Wx0.35H (0.49LxK3iWxK 15H) 


_»y •' 


5.1-13 























TABLE 5, 1.1-3. -ASF SUPPORT SUBSYSTEM CHARACTERISITCS - Concluded 


Item 

Qty 

Total 
weight 
kg (lb) 

Operating 

power 

(Watts-Ave) 

Unit size - meters (ft) 

L = length, W = v/idth, 0 = depth 
diam = diameter, H = height 

4. Conmand & Data Management, Subsystem, 
Displays and Control 





a. 

Computer (igloo) 

3 

3x32 (21 D) 

3x245 

0.50Lx0.26Wxl).20H (1 .64Lx0.85Wx0.65H) 

b. 

I/O unit (igloo) 

2 

2x32 (140) 

2x210 

0.50LxO.26Wxlt.2QH (1 .64LxO.85HxO.65H) 

c. 

Mass memory (igloo) 

1 

27 (60) 

35 

0.46Lx0.31Wx0.24H (1 .50Lxl .QDWx0.8H) 

d. 

C&W electronics unit (igloo) 

1 

4 (8) 

25 

0 . 23LxO . 1 3WxO . 1 0H ( 0 . 75LxO . 42WxO . 33H ) 

e. 

ASA electronics unit (igloo) 

1 

4 (8) 

40 

0.23Lx0.13Wx0.10H(0.75Lx0.42Ux0.33H) 

f* 

Remote acquisition unit 
(1) aft crew station 

3 

3x3 (20) 

3x30 

0 . 23Lx0 . 1 2HxO . 09H (0 . 71 LxO . 39HxO . 30H ) 


(2) igloo 

3 

3x3 (20) 

3x30 

0 . 23LxO . 1 2WxO . 09H (0.71 LxO . 39WxO . 30H ) 


(3) pallet A-1 

8 

8x3 (53) 

8x30 

0.23Lx0.12Wx0.09H (0.71Lx0.39Mx0.30H) 


(4) pallet A-2 

2 

2x3 (13) 

2x30 

0.23Lx0.12Wx0.09H (0.71 LxO. 39HX0.30H) 


(5) pallet A-3 

7 

7x3 (46) 

7x30 

0 . 23LxO . 1 2UxO . 09H ( 0 . 71 LxO . 39HxO. 30H ) 


(6) pallet A“4 

6 

6x3 (40) 

6x30 

0 . 2 3LxO . 1 2Hx0 . 09H ( 0 . 71 LxO . 39WxO . 30H ) 

9- 

Tape recorder (aft crew station) 

2 

2x11 (50) 

45 

0. 33LxO.33UxO.15H (1 .08Lxl .08W.-0.49H) 

h. 

Cathode ray tube {aft crew station) 

2 

2x12 (54) 

2x90 

0 . 26HxO . 1 9HxO .300 (0 . SSHxO . 62HxO . 990 ) 

i . 

Keyboard (aft crew station) 

s^ace) 

1 

2x3 (12) 

210 

0 . 48Ux0 . 1 8Hx0 . 330 ( 1 . 57HxO . 59Hx 1 . 080 ) 


Control & display unit (aft crew 
station) 

16 (36) 

0.46Lx0.2SWx0.20H (1 .50Lx0.83Wx0.67H) 

k. 

PSS control & display panel 
(aft crew station) 

3 

3x36 (240) 

260 

0 . 48Wx 0 . 53HxO .150 ( 1 . 57Wx 1 . 74HxO . 490 ) 

Total CDMS & D&C 


459 (1010) 



5. Mission Kits 





a. 

Radiator panels 

2 

87 (193) 

- 


b. 

Electrical energy 

2 





(1) O2 tank -F 0^ 

2 

2x511 (2254) 

- 

U22 d'am (4. 00 diam) 


(2) Kg tank + Hg 

2 

2x198 (874) 

• 

L32 diam (4.33 diam) 

c. 

QMS Kit 

1 

1351 (2978) 

- 


Total - ASF Flight System 


11,389 (25,113) 


















T-4 

UMBILICAL 


OmmAL PAGE IB 
OF POOR QUALrra 


DISPLAYS AND CONTROLS 
INPUT/OUTPUT 
CAUTIDN AND WAR NINO 
ALARM AND ADVISORY 
MISSION specialist STATION 
PAYLOAD specialist STATION 
PERFORMANCE MONITORING SYSTEM 
REMOTE ACQUISITION UNITS 
GYRO REFERENCE ASSEMBLY 
SIGNAL PROCESSING ELECTRONICS 
STAR TRACKER ASSEMBLY 




Figure 5. 1.1-6.— ASF Flight System 


r 











/ 

f - 


i 


1 


STA. X- 33 , 197 £ mm (1307 IN) 



ITION AND WARNING 
.RM AND ADVISORY 
SION STECIALIST STATION 

load specialist station 

FOHMANCe MONITORING SYSTEM 
lOTE ACQUISITION UNITS 
10 REFERENCE ASSEMBLY 
^AL PROCESSING ELECTRONICS 
R TRACKER ASSEMBLY 


* ■ — POWER, RETURN LINES 

ft e SJ oftical path 

hMm FLUID AND GAS LINES 
ty//X ORBITER FURNISHED EQUIPMENT 


1-6.- ASF Flight System Interfaces 


5.1.1-15 


C7'. 





1 


[ 


I 


payload through the Orbiter rf communication link or, prior to 
launch, through the l^ndline umbilicals. 

a. Payload Interfaces. The interfaces within the payload 

include the following. 

(!) The structural and mechanical interfaces for the hard- 
mounted instruments and support equipment, 

(2) The structural and mechanical interfaces for the APS 
and the instruments mounted on the APS. 

(3) The structural and mechanical interfaces for the booms, 
the boom actuators and the instruments mounted on the 
booms . 

(4) The structural and mechanical interfaces for the stowed 
PDS. 

(5) The thermal interfaces for active thermal dissipation 
and for cryogenic cooling. 

(6J The avionics interfaces between the instruments/support 
equipment and the data and command system. 

(7) The avionics interfaces between the attitude measuring 
system and the APS. 

(8) The interfaces between the electrical power distribu- 
tion points and the instrument/support subsystems. 

b. Hardware, Software and Operational Interfaces. The hardware, 

software, and operational interfaces between the ASF pay- 

load and the Orbiter include those for the following. 

(1) Orbiter maneuvering. 

(2) Orbiter orbit change. 

(3) Subsatellite range and range rate measurements. 

(4) Subsatellite retrieval using Orbiter RMS. 

(5) Installation of pallets to Orbiter standard payload 
attach points. 


5.1-16 


(6) Fluid lines to and interconnects with the T-0 umbilical 
through the station Xo 33,197.8 mm {1307 in) bulkhead 
servi ce panel . 

(7) Fluid lines to and interconnects with the T-4 umbilical 
at station Xo 21 ,209 mm {.835 in). 

(8) Active thermal control heat exchanges interface. 

(9) Electrical power from Orbiter fuel cells through the 
station Xo 17,653 mm (695 in) interfaces, 

(10) Data and command interfaces between the ASF payload 
igloo and the PS and MS stations, and the Orbiter Per- 
formance Monitoring System (PMS) through the station 
Xo 14,630.4 mm (576 in) bulkhead service panels. 

(11) Caution and Warning (C&W) and Alarm and Advisory (A&A) 
interfaces between the payload igloo and the Orbiter 
C&W electronics and the PMS through the station Xo 
14,630.4 mm (576 in) service panels. 

(12) Closed circuit television (CCTV) monitor and control 
interfaces between payload bay camera.s and the aft 
flight deck through station Xo 14,630.4 mm (576 in) 
bulkhead service panels. 

(13) Audio communication interface between the PS and 
other Orbiter crew members at the aft crew station. 

c. ASF Flight System/STDN/inRSS Interfaces. The interface 

between the ASF flight system and STDN or TDRSS is provided 
by the Orbiter by interleaving the data presented to the 
Orbiter rf communication signal processors through the 
station Xo 14,630.4 mm (576 in) bulkhead service panels. 


5.1-17 


I 

I 



f 


Communication with the MCC and other mission control and data 
processing facilities are provided through these interfaces. 

d. Fluid, Avionics and Power Interfaces (Ground). The fluid, 
avionics and power interfaces between the ASF flight system 
and the ground facilities after the ASF payload is mated 
with the Orbiter are provided through the T-4 prelaunch and 
T-0 launch umbilicals. 

5.1.4 MAJOR SYSTEMS INTEGRATION ISSUES 

a. Payload. Within the payload element, the major integration 
issues were as follows. 

(1) How best to install all the instruments on the pallets 
such as to meet the ASF experimental and operational 
requirements within the known constraints. 

(2) How a practical system could be developed to provide 
the required pointing for the instruments. 

(3) How a practical cryogenic cooling system could be 
.developed for instruments 118 and 126. 

(4) How payload data could be processed to the maximum 
extent possible onboard. 

(5) How to maximize subsatellite, and support subsystem 
operati ons . 

(6) How the payload could minimize the EMI, electro- 
static buildup, and contamination generated to 
allow valid experiment measurements to be made. 

b. Payl oad/Orbi ter/Crew . The integration issues betv/een the 
payload and the Orbiter vehicle and crew were as follows. 

(1) Whether the Orbiter vehicle attitude control accuracy 
would be adequate for instrument pointing. 


5.1-18 


(2) How the payload thermal dissipation could be kept 
within the Orbiter Active Thermal Control Subsystem 
(ATCS) capability. 

(3) Whether the EMI generated by the Orbiter could 
prevent valid experiment measurements to be made. 

(4) Whether the payload specialist station (PSS) space 
allocation would be adequate for ASF displays and 
control s . 

(5) Whether one PS could perform the required func- 
tions without overload. 

(6) The number of PS's required to provide 24 hours/ 
day coverage. 

(7) Whether the data rate handling capability of the 
Orbiter communication system was adequate to handle 
the onboard experiments and the deployed PDS. 

(8) The ASF payload failures, which could create a safety 
problem, and the best way to handle these failures. 

c. Payload/MCC. The issues between the ASF payload and the 
MCC, ground data handling facilities included the following. 

(1) The functions to be performed by MCC to support ASF 
data processing and mission operations. 

(2) The best way to handle the large quantity of data 
transmitted to the ground. 

d. Payload/Test, Integration and Launch. The major integration 
issue between the ASF payload and the test, integration and 
launch base facilities was the practicality of providing 
meaningful test and calibration of these extremely sensi- 
tive instruments under earth environments. 

Each of the integration issues identified, which was unique 
to the ASF payload, was evaluated to determine functional 


5.1-19 


1 




feasibility {i.e., can the required function be performed?) 
and as many as possible within the study constraints were 
evaluated further to determine implementation feasibility 
{i.e., can all these functions be implemented by practical 
hardware and software?). The issues of how best to install 
all instruments on the pallets and how to provide an ASF 
pointing system were evaluated in detail. The other issues 
involving the payload elements alone were addressed only from 
a functional feasibility standpoint. Based on existing systems 
which have proven to be capable of performing the same or 
similar types of functions, it was concluded that these func- 
tions could be performed for the ASF payloads. The impact of 
sizing and capacity will be established during the next phase 
of study. 

e. Instrument Arrangement. A number of factors were weighed 

in the determinations whicn resulted in the instrument arrange- 
ment depicted in figure 5. 1.1-4. For each instrument, the 
factors included the following. 

(1) Operating weight. 

(2) Operating volume. 

(3) Instantaneous Fiel d-Of-View (IFOV), 

(4) Pointing requirements. 

(5) Scanning requirements. 

(6) Temperature control requirements. 

Certain instruments do not have a requirement for maneu- 
vering, for either pointing or stabilization. Accordingly, 
those instruments were grouped to be mounted directly to a 
pallet instead of on an APS. Similarly, certain instruments 
must operate at a distance from the Orbiter, which dictated 
mounting on the subsatellite. The two groupings can be noted 


5.1-20 


k 

I 


I 




in figure 5. 1.1-4. The remaining instruments were grouped in 
the four AIM units installed on Pallet A-1 and Pallet A-3. 

f. Selection of Pallets. The selection of pallets for installa- 
tion of the two APS assemblies was based on two factors. 

(1) The swept volume of an AIM while it is maneuvered. 

(2) The need for visibility of, and access to, the sub- 
satellite during the separation and retrieval opera- 
tions. 


g. Overall Arrangement of Pallets. The overall envelope through 
which an AIM unit is maneuvered precluded the use of Pallet 
A-4 for an APS, for the envelope of the Orbital Maneuvering 
System (QMS) kit was encroached under certain combinations 
of the AIM azimuth and elevation settings. Similarly, the 
maneuvering space required by the AIM units precluded locat- 
ing the APS assemblies on adjacent pallets, i.e., A-1 and A-2, 
or A-2 and A-3. These considerations alone seemed to dictate 
use of the first and third pallets for installation of the 
APS assemblies. The requirement for visibility of the sub- 
satellite was the final determinant in locating the ASF 
instruments since placement of the subsatellite on Pallet A-2 
permitted continuous viewing of that vehicle while the RMS 
was manipulated in the separation and retrieval operations. 

The final arrangement, as depicted in figure 5. 1.1-4, is: 

Pallet A-1 - ASF Pointing System 
Pallet A-2 - Subsatellite 
Pallet A-3 - ASF Pointing System 
Pallet A-4 - Non-maneuverabl e Instruments 


T 

. f 

/ 

! 


h . 


Selecting Pointing System. On the issue of selecting a 
practical pointing system, the Instrument Pointing System 
(IPS) was evaluated due to the ground rule that Spacelab 
equipment was to be used, if possible. However, several 




5.1-21 


operational features made the use of the IPS unsuitable. The 
primary objections were the necessity of decoupling the point- 
ing system payload for the vehicle launch and recoupling after 
attaining orbit and the lack of multiple pointing from the one 
system. It was not considered feasible to initially hard 
mount the experiment instrumentation and then attempt to 
install it, either with the RMS or a special, volume consuming 
apparatus on the pointing system, after the Orbiter was in 
orbit. Further, the need for more than one pointing direction 
for the various instrument clusters would have required more 
pallet space than was available if the IPS were used. Also, 
the IPS cannot be tested under one g conditions due to its 
gimbal suspension design. 

The two APS assemblies have equal capability from the stand- 
points of instrument accommodation and operating precision and 
accuracy. Because of this, the requirements of. the individual 
instruments for accuracy of pointing and tracking did not 
greatly influence the location of instruments in one AIM or 
another, A facet of the pointing requirement which did receive 
consideration was that of co-alignment of instruments for 
participation in the experiments. Where a high order of co- 
alignment precision has been specified, the instruments have 
been co-located in the same or adjacent AIM units. 

The capabilities of the APS to provide accurate pointing, 
tracking, and stabilization are described more fully in 
paragraph 5.2.3 of this report. 

Overall System Considerations. In the area of the interface 
between the ASF payload and the Orbiter vehicle and crew, 
detailed assessments were made of most of the issues 
mentioned. These are discussed in paragraphs 5.2.1 through 
5.2.5. The one major issue which was not resolved was whether 
the EMI generated by the Orbiter would allow meaningful 


5.1-22 


experimental results. Preliminary assessment of instrument 
susceptibility and expected Orbiter EMI background indicates 
that conventional Electromagnetic Compatibility (EMC) design 
approaches should be adequate to prevent EMI problems. (See 
section 5.5). However, this issue is one which could impact 
not only the ASF pallet-only mode but could raise feasibi- 
lity questions with every payload which has instruments and 
equipment sensitive to high levels of electrical and magnetic 
interference fields. Further study is planned in this area 
after the sensitivities of the instruments are further defined. 

In the area of interface between the ASF payload and the ground 
facilities, the full functional role of the MCC and other 
mission control facilities was not fully evaluated since the 
approach taken for the study was to perform as much of the data 
processing and mission operations onboard the payload as 
was considered practical. The question of the best way to 
handle the large amount of data handled on the ground was also 
not fully addressed from an implementation standpoint and 
should be further assessed during the next study phase. 

The major issue for the area of interface between the ASF 
payload and the test, integration and launch base facilities 
is one which is not unique to the ASF program, Al,l payloads 
which have sensitive, precision instruments with thresholds 
far below the background levels of magnetic or electric fields, 
particle contamination, etc., created by the earth-bound 
environment, or which cannot operate in the sea level atmos- 
phere, will be subject to the same test and verification 
problems. Comprehensive analyses to identify the error 
sources which can affect the precision and thresholds of these 
instruments, and great care in selecting design to minimize 
these error sources, can assure successful experimental 
results. 


5.2 FLIGHT SYSTEMS 


5.2.1 THERMAL, STRUCTURAL AND MECHANICAL SUBSYSTEM (TSMS) 

5. 2. 1.1 Introduction 

The objective of this phase of the study was to show the feasi- 
bility of installing and servicing all ASF instruments, sub- 
satellite, and support equipment on multiple ESRO furnished 
equipment pallets within the operational and environmental 
requirements and constraints imposed by the instruments and 
support equipment. 

To meet the objective it was necessary: (1) that the instruments 

be grouped according to complementary operations and other 
specific experimental requirements, (2) that an IPS be developed, 
(3) that boom and actuation concepts be selected, (4) that a 
subsatellite retention and ejection concept be defined, and (5) 
that instrument, subsatellite and support equipment installation 
and layout design be performed. 

Analyses were conducted, alternative candidate concepts were 
assessed, and a baseline conceptual configuration was established. 
The instrument pointing, boom and boom deployment, and sub- 
satellite retention and ejection systems received greater emphasis 
than did other areas since the more significant questions of 
feasibility involve these areas. Also, the implementation tech- 
niques for these areas differ significantly from the ERNO approach 
where similar requirements apply. 

5. 2. 1.2 Requi rements 

The following functional requirements apply to the ASF TSMS. 

a. Thermal Control. The subsystem will provide for the ASF 
instruments and support equipment the capabilities for: 


(1) active thermal dissipation, (2) passive thermal control, 
(3) heating, and C4) cryogenic cooling. 

The active thermal control system must have the capability 
of dissipating payload thermal energy resulting from the 
use of the following levels of electrical power. 

(1) 5.3 kW average over the entire mission. 

(2) A maximum average of 6.9 kW during any given orbit. 

(3) 9.0 kW maximum for 15 minutes each orbit from revolutions 
32 through 47. 

Instruments 213, 303 and 304 impose the greatest demand on 
the active thermal control system since they use the highest 
level of power (1.1 kW, 5.0 kW and 5.0 kW, respectively). 

The detectors for instruments M8 and 126 must be cooled to 
4K and portions of the instrument housings must be cooled to 
at least 77K. Although the instruments will be designed to 
be compatible with the cryogenic cooling requirements (e.g., 
the housings will be of dewar construction), the TSMS must 
provide cryogen storage, distribution and gas exhaust 
facil ities . 

b. Structural and Mechanical. The subsystem will provide for 
the installation of 15 ASF instruments on the equipment 
pallets. In addition, the subsystem will provide for the 
installation of the following support equipment on one or 
more of the pallets in the payload bay. 

(1) An APS capable of pointing instruments in the desired 
direction with a high degree of accuracy, 

(2) One APS control electronics for each APS. 

(3) An Attitude Measuring System (AMS) consisting of a gyro- 
reference assembly, three star tracker assemblies (fixed 


5, 2. 1-2 



head ) 3 or one glmballed star tracker assembly and a 
processing electronics assembly. 

(4) Autocollimators, porro prisms* optical flats and tvn’st 
sensors for precise attitude transfer betv^een pallets. 

(5) Booms and boom actuator mechanisms. 

(6) Subsatellite retention, ejection mechanisms. 

(7) A thermal coolant loop pump and heat exchanger on one or 
more pallets. 

(8) Up to 8 remote acquisition units (RAU's) per pallet. 

(9) A pov/er distribution box on each pallet. 

(10) A pressurized equipment module (igloo). 

Special installation requirements for instruments are included in 
the id's and include individual instrument pointing and track- 
ing accuracy, and requirements for accurate co-alignment of tv/o 
or more instruments. These are summarized in table 5. 2. 1-1. In 
addition, requirements exist to have one diagnostic (Instrument 
550) scan particle accelerator output to determine beam charac- 
teristics, and one instrument (Instrument 536) to be located 
such that the influence of the Orbiter in relationship to the 
earth's magnetic field is within acceptable limits as defined 
i n the ID's. 

No special requirements exist for installation or location of 
support equipment other than AMS equipment. The gyro-reference 
assembly and star tracker assembly reference axes must be aligned 
within a few seconds of arc to each other and to the APS reference 
axes. Optical attitude reference transfer media (porro prisms, 
twist sensor and optical flats) must also be aligned within a few 
seconds of arc to each other and the APS and AMS reference axes. 


5. 2. 1-3 


TABLE 5. 2, 1-1,- AII^ PACKAGING PARAMETERS FOR ASF 


AIM Instrument 
No ♦ No . 





1.0 X 1.0 X 2.0 

1.0 X 1.0 X 1,0 

2.0 diam x 1.0 

0.2 X 0.2 X 0*2 


1.0 X 0.5 X 0.3 

0,5 X 0.3 X 0.1 
1.8 X 1.0 X 0,5 
0.5 diam x o.8 
0.3 X 0.3 X 0.3 



0.6 diam x 2,0 
0.8 diam x i .8 


122 0.5 X 0,2 X 0.2 

124 0.3 X 0.5 X 0.7 

126 . 0,9 diam x o.7 

Startracker 
and Gyro 


Weight 

(kg) 


100.0 

250.0 


Temperature 


A«2mc? Coalignment 


0.2 X 0.2 X 1.3 

0,9 X 0,9 X 1.8 

0,1 X 0.1 X 0.1 

0,2 X 0.2 X 0,2 

0,1 X 0.1 X 0,1 


0,3 X 0.3 X 0,1 
1,0 diam X 2.0 

1,0 X 0,3 X 0.3 



300,0* 

63.0 

Total 

424.0 
AIM 

:3A + B 
Total 

754.0 


-20C to +50C 


4K to 28K 


0,1^ to 3° 


0 . 1 ° ® 


±0.017° (S) 
±0.017° ® 

±0.1° (D 


DTE: Instruments marked ® or ® are respectively coaligned with each other. 

Includes weight of cryogenic dewar. 


% K>0B. 


5 . 2 . 1-4 
























i 


I 


The equipment will be in close proximity to instruments being 
serviced. This factor is especially critical for the high voltage, 
high power converters, the cryogenic coolant systems, and the AHS. 
Physical and environmental interference of instrument or other 
support equipment will be minimized and those requiring active 
cooling will be located on cold plates. 

The subsatellite installation must provide for reliable separation 
while providing mechanical integrity during the launch and the 
reentry environments. The separation mechanism must impart a 
relative rate of 20 cm/sec to the subsatellite. 

The extended booms with instruments attached must be capable of 
withstanding Orbiter maneuvering accelerations and decelerations 
without damage. The booms must also maintain instrument attitudes 
within 0.5 degrees during orbit limit cycling operation and during 
instrument scanning operations. 

Safety considerations dictate that any part of the payload wb-'ch 
could fail is to be securely latched or any part which could pre- 
vent closure of the payload doors is to be capable of being 
jettisoned . 

5. 2. 1.3 Guidelines and Assumptions 

In addition to the general guidelines and assumptions listed in 
paragraph 2.3.4, the following guidelines and assumptions unique 
to the TSiiS were used in this study. 

a. ERNO designed, ESRO furnished equipment including the equipment 
pallets will be utilized to the maximum extent possible. 

b. The Orbiter attitude control system will be capable of pro- 
viding coarse pointing to an accuracy of within 2°. 


5. 2. 1-5 


c. Normal Orbiter orientation during ASF missions will be the X-X 
axis tilted 45° to the earth's radius vector (nose up or dov/n) 
with the payload bay forward in the direction of flight. 
Attitude changes will be made from this position for specific 
experiments . 

5. 2. 1.4 Capabilities and Constraints 

The following capabilities and constraints apply to the ASF TSHS. 
a. Orbiter. The ATCS for the Orbiter and payload consists of: 

(1) Radiators mounted on the interior of the payload bay 
doors which deploy upward when the doors are open. 

(2) Heat exchangers and coolant pumps provided in the Orbiter. 

(3) Heat exchangers, thermal capacitors, and coolant pumps 
provided on the pallets for the experiment payload. The 
ATCS is available to the payload during all mission 
phases, including ground operations. 

The Orbiter ATCS will provide a baseline on-orbit payload heat 
rejection of up to 21,500 Btu/hr (6.3 kW) with the payload bay 
doors open and coolant temperatures of 7.2®C maximum to the pay- 
load and 54.4°C returned from the payload (see table 5. 2. 1-2). 


The on-orbit heat rejection capability can be increased to 29,000 
Btu/hr (8.5 kW) by the addition of payload chargeable radiator 
kits provided that the Orbiter cabin is appropriately powered 
down. Coolant temperatures will be 7.2°C to the payload and 
40°C returned from the payload. The ATCS will provide an ascent 
(after main engine cutoff (MECO)), on-orbit, entry and post- 
landing heat rejection capability of 5,200 Btu/hr (1.2 kW) with 


5. 2. 1-6 


TABLE 5.2.1-2,- ATCS CONTROL CAPABILITY 


Mission Phase 

Payload Heat 
Rejection 

Coolant Temperature 
“F (“0 


Btu/hr (kW) 

In 

Out 

Payload doors open 

21,500 (6.3) 

45 (7.2) 

130 (54.4) 

(with additional 
radiator kits) 

29,000 (8.5) 

45 (7.2) 

104 (40.0) 

Ascent (post MECO) 
in-orbit, entry and 
post landing 

5,200 (1.2) 

45 (7.2) 

100 (37.8) 

Ground cooling 
with four thermal 
capacitors- 
15 min/3 hours 

29,000 (8.5) 
(12.4) 

45 (7.2) : 

104 (40.0) 

Cold plates 

(1.0) 


(10 to 30) 


5. 2. 1-7 









the payload bay doors closed and coolant temperatures of 7.2‘’C to 
the payload and 37.8“C returned from the payload. Within 15 
minutes following touchdown, ground cooling will be available to 
the Orbiter; with ground cooling and with the Orbiter cabin 
appropriately cooled down, the ATCS will provide a payload heat 
rejection capability of 29,000 Btu/hr (8.5"kW) with the payload 
bay doors closed and coolant temperatures of 7.2°C maximum to the 
payload and 40“C returned from the payload. 

The payload heat exchanger will be designed so that any of the 
following can be selected (by the payloads) as a payload coolant: 
water. Freon 21, Flutec PP50. 

The payload side of the payload heat exchanger is being designed 
with two coolant passages. The payload may use either or both of 
these passages. Each of the payload coolant passages is being 
sized for a maximum delta pressure of 6 psia with 9,072 kg/hr 
(2,000 Ib/hr) of Freon 21 and a maximum operating pressure of 
200 psia. 

b. Payload. The physical accommodation capability of a single 
pallet segment is as follows: 

(1) The overall payload carrying capability of a single 
pallet segment is about 3500 kg (5500 kg, multiple 
pallets) (uniformly distributed over the pallet) with 

a c.g. limitation between 250 mm above the pallet floor 
line and the Orbiter bay horizontal centerline at station 
1 ^ = 1016 cm (400 in.). 

(2) A single pallet provides 36 m volume above the floor. 

(3) The floor panel of a single pallet segment provides about 
17.0 m^ (183 ft^) of mounting area, which is available 
for mounting payload equipment. 

If the equipment exceeds the floor. panel load capability, it can 
only be mounted on standard equipment hard points. Provisions 


for 24 such hard points are located on each pallet segment on the 
inner surface at the intersections of the frames and longitudinal 
members (figure 5. 2. 1-1). 

Each hard point provides a spherical nut with 36 mm or 45 mm 
diameter metric thread. They are bolted to the pallet structure 
and have a dynamic load carrying capability of: 

direction 28,547 N ( 6,418 lb) 

Yq direction 18,443 N ( 4,146 lb) 

direction 75,046 N (16,871 lb) 

Load carrying capability is equal for all hard points regardless 
of their locations. 

The ESRO furnished coldplates are designed for 1 kW maximum. The 
coolant temperature will be between 10*C and 32°C for a pallet- 
only configuration depending on total heat load on the loop and 
the location of the coldplate. The coldplates are connected in 
series. 

5. 2. 1.5 Subsystems Description 

5 . 2 . 1 . 5 . 1 Thermal Control System 

The thermal control system consists of: (1) an active thermal 

control, (2) a passive thermal control, and (3) instrument 
cryogenic cooling systems. 

5. 2. 1.5. 1.1 Active Thermal Control Subsystem (ATCS) 

a. Cooling. The present instrument definition is not refined to 
the point that exact heating or cooling requirements can be 
defined. Some instrument de.signs may use direct radiation to 
space for cooling, others may have minimal cooling require- 
ments such that heat may be directly coupled into the pallet 


ORIGINAL PAGE IS 
OF POOR QUALITY 




Figure 5. 2. 1-1. — Pallet equipment hardpoints. 







structure. For those requiring significant thermal dissipa- 
tion, the ATCS may be used. The Orbiter ATCS will include a 
heat exchanger to allow transfer of payload thermal energy to 
the Orbiter system. 


Fluid circulation through the payload side of the heat ex- 
changer w'll be provided by a pump kit chargeable to the 
payload (figure 5. 2. 1-2). The freon coolant loop is designed 
to accommodate up to eight cold plates and up to four thermal 
capacitors to take up peak heat loads. The cold plates and 
thermal capacitors used for ASF missions will be those fur- 
nished by ESRO for Spacelab. 


The cold plates are mounted to the pallet floor panels as 
shown in figure 5. 2. 1-3. These panels fit the 48“ section of 
the pallet. It is thus possible to mount all eight cold 
plates on one pallet segment or to distribute them over 
several pallet segments in (TBD) configuration. Since the 
coolant loop plumbing can be changed at the integration site, 
it is in principle also possible to mount cold plates in 
other positions. Isolators are used to thermally isolate 
cold plates from the pallet panels. All cold plates ar'.^ 


connected in series. 


Thermal capacitors can be mounted to cold plates to accomo- 
date peak heat loads. It is not necessary, however, to mount 
peak load generating equipment directly to a thermal capac- 
itor. The thermal capacitor can also be mounted to a dif- 
ferent cold plate in the coolant loop. The size of a thermal 
capacitor is 750 x 500 x 52 mm. The capacitors are designed 
to accommodate the maximum permissible peak heat load of 
12.4 kW for 15 minutes every three hours when all four are 


used . 




5.2.1-11 





I 


f 




REDUNDANT ASSEMBLIES 


Figure 5. 2. 1-2. —Active thermal control system (ATCS) 


5.2.1-12 




AS INDICATED, COLD PLATES MOUNT 



BOLTS GO THROUGH MATCHING HOLES 
IN COLD PLATE, THERMAL CAPACITOR AND 
STAND OFF AND SCREW INTO THE PALLET PANEL. 



COLD PLATE MOUNTING 


Figure 5. 2. 1-3. — Pallet cooling. 


5.2.1-13 








Cold plates and thermal capacitors provide the same standard 
mounting hole pattern as the pallet panels. Payload equipment 
are mounted to the pallet and cold plates (and thermal capaci- 
tor if used) with titanium bolts with isolating washers to 
limit heat transfer to the pallet panels. 

Thermal contact is established in the area around the bolts. 
Th" design provides heat transfer capability of 13 W/°C per 
bolt with a conductance of 1 W/°C. The maximum capability per 
cold plate is 1 kW. 

The mechanical load carrying capability of cold plates is 
limited by the load carrying capability of the pallet panels. 

b. Heating. Some of the ASF instruments and subsystem equipment 
may require controlled heating to maintain temperatures above 
the structure to which they are attached or above the ambient 
payload bay environment. Each instrument or equipment requir- 
ing heating will provide it internal to its package using 
conventional techniques (thermistor bridges, proportional or 
pulsed drivers, wound wire heating element). The power 
required for these circuits are included in the individual 
equipment allocations. 

5. 2. 1.5. 1.2 Passive thermal control 


Passive thermal control may be necessary to minimize temperature 
extremes of the pallet structure. Control techniques include 
thermal coatings, high performance insulation (HPI) blankets and 
thermal isolation between pallet-mounted equipment and pallet 
structure. When heat transfer to the pallet is desirable, equip- 
ment can be mounted directly to the structure, resulting in the 
equipment following closely the pallet structure temperature. 


5.2.1-14 


f " 

•as V 







The thermal covers for the top surfaces of the pallet are flat HPI 
blankets the same size as the pallet structural panels. They are 
designed to be easily installed and removed so that the amount of 
exposed surface may be varied from mission to mission. 

5. 2. 1.5. 1.3 Cryogenic cooling system . 

Two sensors, the Limb Scanning Infrared Radiometer (Instrument 118) 
and the Infrared Interferometer (Instrument 126), require cryogenic 
cooling of the instrument and optical telescopes. Specific heat 
loads are not defined at this time and further development will be 
required before a systems analysis can be performed. However, 
the requirement to cool large optical telescopes in the range of 
60 to 100 cm diameter and the complete instrument housing rules 
out the feasibility of using state-of-the-art closed loop refriger- 
ant systems at the required temperature of 4K. Therefore, looking 
only at the detector array heat load and estimating this to be in 
the order of 2 v/atts, a supercritical helium cryogenic system will 
suffice for these sensors. The 0.83 m (33") diameter, 74.8 kg 
(165 lb.) storage tank system used on the Apollo program for the 
lunar module (LM) vehicle will provide an adequate supply of 
helium if expanded through a joule Thompson expander to provide 
the 4 K temperature. 

Multiples of these tankage systems could provide the additional 
cooling capacity for the instrument enclosure and optical tele- 
scopes. Liquid nitrogen or solid Ne may be better cryogens for 
the 77K and 20K temperatures of the telescopes and enclosures. 
However, these could be stored in the available tankage designs. 

5. 2. 1.5. 2 Structural /Mechani cal System 

The structural /mechanical system consists of the following 
standard items which will be supplied by ESRO: 

a , Pall ets . 

b. Pallet panels with threaded inserts as required. 


5.2.1-15 


Cold plates with standard mounting holes for sensors, 
d. Thermal capacitors. 

e- Hard point mounting provisions for large instruments, cryogenic 
storage systems, and APS as required. 

f . Igloo. 


In addition to the ESRO items, the following Items are included 
in the structural /mechanical system: 

a. The APS and its truss structural mount. 

b. The satellite mounting and deployment system. 

c . The booms and boom mounting system . 

d . All auxiliary brackets and mounting provi sions plus restraint 
systems as required to provide structural integrity to the 
system. , 

The pallet cross section is U-shaped and of aeronautic-type con- 
struction. It provides for hard points for mounting heavy 
experiments and a large panel surface area to accommodate lighter 
payload elements. Pallet segments are modular (3 m nominal length) 
and can be flown independently or i nterconnected . As many as 
three pallets can be interconnected and supported by one set of 
r etqnti on f i tf i ngs * 

The pall et structure accommodates experiment and equipment for 
direct exposure to space. The general structural configuration at 
the pallet is shown in figure 5. 2. 1-4, The pallet provides mount- 
ing support for the experiments either directly on the inner ski n 
panel s , or as mission-dependent equipment through specific hard 
points for better dispersion of concentrated loads. 

5.2.1 .5.2.1 Pallets A-1 and A-3 

Th e structur a 1 and mechani cal : co nfiguration of pal 1 ets A - 1 and 


5.2.1-16 



A-3 are similar. Each has an APS mounted to the basic structure 
to provide precision pointing for the ASF instruments. The 
instruments installed on the APS are contained in AIM's to pro- 
vide a convenient integrated package which can be gimballed 
about two axes as shown in figure 5. 2. 1-5. Each APS is com- 
prised of two AIM's, independently controlled. Table 5. 2. 1-3 
lists the instruments contained in each AIM for each of the APS's. 
Other equipment installed on pallets A-1 and A-3 are as follows. 

a. Pallet A-1 

(1) Four experiment RAU ' s . 

(2) Four subsystem RAU ' s . 

(3) Sun Sensor. 

(4) APS control electronics. 

(5) Boom and boom actuation mechanisms. 

(6) Power distribution box. 

(7) Optical alignment transfer devices. 

b. Pallet A-3 

(1) Four experiment RAU ' s . 

(2) Three subsystem RAU ' s . 

(3) Porro prisms and optical flats for attitude reference 
transfer from the AMS on pallet A-1, 

(4) APS control electronics. 

(5) Power distribution box. 

(6) Attitude measuring system including a gyro reference 
assembly, 3 star tracker assemblies, AMS processing 
electronics, and optical sensors for attitude reference 
transfer. 

c. AMPS Pointing System. Each APS consists of two pointing 
platforms each mounted on a separate pallet as shown in 
figures 5. 1.1-3, 5. 1.1-4, 5. 1.1-5 and 5. 2. 1-6. These may 



TABLE 5.2, SCIENTIFIC INSTRUMENT LOCATION 


> 

( 


APS 

1 (Pallet A-1 ) 

AIM lA 

AIM IB 

213 - Laser Sounder 

1011 “ Ultraviolet Occulation Spectrograph 

532 - Gas Release Module 

534 - Optical Band Imager and Photometer 
System 


1002 - Pryhel iometer/Spectrophotometer 


550 - Level II Diagnostics (on boom) 


APS 3 (Pallet A-3) 


AIM 3A 

AIM 3B 

116 - Airglow Spectrograph 

122 - Ultraviolet-Visible-Near IR 


Spectrometer 

118 - Limb Scanning Infrared 

124 - Fabry-Perot Interferometer 

Radiometer 

126 - Infrared Interferometer 









I 




operate independently or in unison with each other. In 
addition, each of the two AIM's on each platform may operate 
at different pitch angles to one another and may point within 
10 degrees of one another in yaw. All four AIM's are of the 
same size, therefore instruments may be interchanged as 
pointing requirements change. 

The primary characteristics that dictate the instrument group- 
ing are the instruments' weight, envelope, FOV, pointing 
accuracy, scanning requirements, and temperature control. 

These are tabulated in table 5. 2. 1-1. The AIM's housing 
these instruments provide thermal control, and contamina- 
tion and acoustic protection. End covers are remotely 
operated to provide protection against contamination of optics 
and sensors. Each heated AIM is lined with multilayer 
aluminized kapton insulation. Heaters to maintain instrument 
temperatures to within ±2° C are mounted to the inner surface 
of the blankets. For those instruments requiring cooling, 
cold plates are mounted to the inner side of the AIM's. These 
in turn interface with the payload bay heat exchanger fluid 
loop (figure 5. 2. 1.7). The exterior surfaces of the AIM's 
are finished according to the thermal control desired. The 
coolant lines pass through the gimbals before interfacing with 
the payload bay heat exchanger fluid loop. 

The experiment AIM's are attached to the yoke which provides 
the interface to the APS (figure 5. 2, 1-8). The yoke will 
accept AIM's of variable length and width. However, at the 
present all AIM's are 1x2x3 meters (3.28 x 6.56 x 9.84 
feet). Lugs are provided to accept takeoff and landing 
restraint latches. The AIM's, with their associated sensors, 
are attached to the frame. Each of the two yokes are 
attached to the up-down (pitch) fine pointing drive. The two 
drives allow each AIM to be pointed independently of each 
other from 0 to 90° in pitch orientation with an accuracy 


5.2.1-22 


-23 


tn 

ro 


TO 

ASF 

ATCS 

(SEE 


PALLET, 
COLD 
PLATE 













potential better than one arc second, This capability would 
require sensors with greater accuracy than those planned for 
ASF. The tv/o yokes revolve as a unit about the^ axis of the 
depl oy-retract column to provide a coarse pointing maneuver- 
ability of 180° and an accuracy of six arc minutes. The slew 
rate of the combined gimbal system is 2 degrees/second. 

The depl oy-retract system elevates (figure 5.2. 1-9) the 
telescoping central column through a ball screw jack from the 
stowed position to the fully extended position 2.15 meters 
(7.05 ft) high. This position allows the AIM's to be slewed 
within the gimbal envelope without Orbiter dimensional inter- 
ference. Microswitches indicate full extension and retrac- 
tion of the column. In the retracted position, eight solenoid 
actuated mechanical latches between the yoke and pallet mount 
prevent motion. Individual microswitches indicate latch 
position. In the event a latch fails in the extended position 
(gimbal frame locked) the individual latch may be separated 
with an explosive squib. Conversely, if one latch fails in 
the retracted position, two of the four are sufficieuc to 
enable the AIM to survive landing loads. 


Direct drive brushless “pancake" dc torque motors are selected 
for the fine pointing gimbals because of their inherent 
frictionless character! Stic and reliability. Brush type 
motors are selected for coarse pointing because of their high 
torque/weight and volumetric efficiency. 

The total weight of the entire pointing system, not including 
instruments, is 1,100 kg (2,420 lbs). Each platform is capable 
of mounting two AIM's with instruments weighing 465 kg 
(1025 lb) per AIM. 



5.2.1-25 





In the event a malfunction occurs in the pointing system which 
prevents the column from being retracted and/or locked in the 
landing position, the column may he separated from the mount- 
ing structure with explosive bolts and ejected from the pay- 
load bay through a spring mechanism. 

The following three modes of operation are possible with the 
APS. 

(1 ) The AIM's may be pointed using acquisition and fine 
tracking sun or star trackers. 

(2) A preprogrammed subroutine may be initiated in the igloo 
payload computer through a PS keyboard entry. This 

is typically used to drive the APS in performing a 
raster scan utilizing the boom mounted Level II Diagnostics 
(Instrument 550) . 

(3) A two-axis displacement "joystick" provicTes a manual 
fine pointing capability. The attitude of the APS will 
be displayed on the cathode ray tube (CRT) at the 

PSS. 

Instrument 550 Boom. Instrument 550 (Faraday cup, retarding 
potential analyzer, cold plasma probe) is installed on a 
furlable boom which is attached to the AIM IB as shown in 
figure 5.2.1-10. The purpose of this boom is to allow the 
instrument to measure the particle energies and the exhaust 
beam plasma potential to establish Instrument 303 beam 
characteristics. These data will be used to support experi- 
ments using Instrument 303. In use, after the boom is ex- 
tended. Instrument 550 is positioned by the APS at the desired 
elevation above the particle accelerator. Initiation of a 
preprogrammed subroutine accomplishes a raster scan of the 
accelerator beam field by yawing of the APS while the boom is 
extended or retracted to maintain the desired elevation. 


5.2.1-27 
















The instrument must be capable of being completely removed 
from the accelerator beam except when calibration Is being 
performed. Also the structure used to mount Instrument 550 
must have minimum impact on accelerated particle beam 
characteristics. 

The Storable Tubular Extendable Member CSTEM) design has been 
selected for the ASF application. The STEM is a thin strip of 
metal heat treated into a circular overlapped cross section 
(figure 5.2.1-11). The bending strength of a STEM element is 
almost equivalent to that of a seamless tube to the same 
diameter and wall thickness. The element is stored on a drum 
by a flattening and rolling process, and very long lengths of 
tubular structure may be extended or retracted by rotating 
the drum in the appropriate direction. 

A further development of this principle is a mechanism that 
employs two diametrically opposed "underlapped" elements as 
shown in figure 5.2.1-12. These BI-STEM elements are stored 
on two drums instead of one. This configuration offers 
several advantages over the STEM; the natural tendency of the 
STEM to warp because of the high compressive stresses built 
in during fabrication is eliminated and two drums instead of 
one allows a more compact deployment package (figure 5.2.1-13). 
The perforated BI-STEM boom is fabricated from precision 
rolled beryllium copper chosen for its excellent heat transfer 
characteristic combined with high strength-to-weight ratio. 

For Instrument 550, the 0.036 cm {.014”) thick x 22.1 cm 
(8.7") wide strip is rolled and heat treated to form a 7.92 cm 
(3,12") diameter x 18 meters (59 ft) long tube. The tubes 
are then flattened back into strips and wound on spools for 
maximum compactness. 

The geometric and thermal configuration controls the rate of 
absorption of heat on opposite sides of the boom, thus pro- 


5.2.1-29 






INCHES 


Figure 5.2.1-13. - STEM/BI-STEM element 
deployment comparison. 




Figure 5.2.1-14. -Chemically milled strip 




5.2.1-31 



ducing equal thermal expansions on opposite walls and avoid- 
ing thermal bending. A unique perforation pattern produced 
by chemical milling allows a selected amount of solar 
radiation to pass through the near wall of the boom and 
impinge on the far wall. This produces the same proportion- 
ality of inside to outside exposure for all incident angles 
of the sun. The perforation pattern consists of small 
circular holes arranged in a double helix pattern (figure 
5.2.1-14). The ratio of absorptivities (outside to inside) 
is equal to the fractional area of wall cut out for windows. 
Polished silver plating is used on the outside while a black 
oxide coating is used on the inside. 

The deployment mechanism provides a positive drive for both 
extension and retraction and contains a simple mechanism for 
joining the seams. A position potentiometer allows the boom 
for Instrument 550 to be precisely extended during raster 
scanning (figure 5.2.1-10). The wire harness is simultane- 
ously deployed or retracted through the center of the boom. 

The combined weight of boom plus deployer is 28 kg (61.7 lb) 
for Instrument 550. 

In the event the boom cannot be retracted, an explosive device 
is used to separate and eject the boom and allow closure of 
the payload bay doors. 

Perforated BI-STEMS, similar to the one previously described, 
have been manufactured and successfully flown on various 
spacecraft by several aerospace firms. Therefore minimal time 
and effort are necessary to produce a flight-qualified unit 
meeting the desired specifications. 

5. 2. 1.5. 2. 2 Pallet A-2 

The PDS, which is deployed soon after mission orbit is achieved, 

is installed on Pallet A-2. In support of the subsatel 1 i te , the 


5.2.1-32 


j' following equipment are algo mounted on Pallet A-2. 

a. Experiment RAU. 

b. Subsystem RAU. 

c. Power distribution box. 

d. Deployment mechanisms. 

e. Latch/unlatch mechanisms. 

t? 

The following discussion is limited to the mounting, deployment, 

^ and latch/unlatch mechanisms for the subsatellite. The sub- ' 
satellite is discussed in section 5.2.6 and the other equipment 
installed on Pallet A-2 are discussed in their respective sections. 

Since the subsatellite is to be reused it is imperative that the 
deploy/retrieval/retention operations present a minimal possi- 
bility of damaging the subsatellite. Therefore, the subsatellite 
mount and grab ring must be designed with this objective in mind. 
Also, the retrieval operation must be as simple and foolproof as 
possible to prevent damage to adjacent structures by the sub- 
satellite and remote manipulator arm while allowing retrieval to 
be accomplished in a minimum time period. 

The configuration of a collet containing a cold gas velocity 
separation device is depicted in figures 5.2.1-15 and 5.2.1-16. 

The mechanism carries the tensile load of the subsatellite in the 
locked position. When the system is "armed" an explosive squid 
shears out a metal slug in the isolation valve and admits 
25.8 X 10® N/m^ (3750 psig) GN^ to the inlet of the pilot valve. 

* Energizing the "eject" switch fires the electrical harness 

guillotine and separates the wiring which was used to power and 
monitor the subsatellite. One hundred milliseconds later, gas 
pressure is introduced Into the cylinder bore through the ener- 
gized solenoid actuated pilot valve, causing the piston to move 
toward the collet and allowing the collet fingers to spring 

t . : 


5,2.1-33 


cn 

ro 


I 

CO 



r 


i 


SJBS^TELL'^E 
IMPACT <bu{?PACE 


coi_i_er - 

!P.V>iQ 


striker /detekit 


COLLET 


SUBS stellite 
•STRUCTueE. 


SRAB collar 



GUIDE RAIL 
Q PEACES 


- n/oumt come 


Figure 5. 1-15. - Subsatel 1 i te retention/ejection mechanization 







inward due to their stored strain energy. The piston continues 
forward until it contacts the surface of the subsatellite where 
further expansion of the trapped gas causes the subsatellite to 
separate at the required velocity of 20 cm/sec. The subsatellite 
grab collar continues on teflon guide rails until free of the 
mount. The rails assure liftoff in a precise direction. The vent 
hole uncovered by the piston allows the trapped gas to escape. 

During retri eval , the subsatellite grab collar is grasped with the 
remote manipulator clav/ and positioned on the mount cone. The 
tapered mount cone assists in maneuvering the subsatellite to the 
proper location on the pallet. When the three sensors at the top 
of the mount are simultaneously contacted, the grab cone is fully 
seated on the mount. Gas is then automatically admitted to the 
retract side of the piston through the solenoid pilot valve, 
retracting the collet piston and expanding the collet fingers, 
thereby locking the subsatellite on the mount. A. micro-switch 
indicates piston positions assuring that the piston is fully 
retracted and the collet locked. 

5-2. 1.5. 2. 3 Pallet A-4 

Pallet A-4 is utilized to mount the Electron Accelerator (Instru- 
ment 303), MPD Arc (Instrument 304), Triaxial Fluxgate (Instru- 
ment 536), and the Gas Plume Release (Instrument 549). 

In addition to the instruments, the following support equipment 
are installed on Pallet A-4. 

a. Four experiment RAU's. 

b. Two subsystem RAU's. 

c. Power distribution box. 

d. Boom and boom actuator mechanism. 

Instruments 303 and 304 are hard mounted to the pallet. Since 
these instruments are high power users (5 kW, average) provisions 


5.2.1-36 





will be made to install these instruments on cold plates and on 
thermal capacitors. 

The Gas Plume Release (Instrument 549) is located internal to the 
Electron Accelerator and is used for the determination of accel- 
erator-produced electron and ion beam flux densities and emergence 
angles by means of optical observations of the excitation of the 
released gas. 

In order to map the earth's magnetic field it is necessary to 
extend the Triaxial Fluxgate (Instrument 536) a sufficient dis- 
tance from the Orbiter to negate the magnetic interference of the 
vehicle. A furlable boom is used to accomplish this task. The 
boom is hard mounted to allow deployment at a 45° angle from the 
Orbiter Z axis in the Y-Z plane as shown in figure 5.2.1-10. 

The basic design of the boom for Instrument 536 is similar to that 
described in paragraph 5, 2. 1.5. 2.1 for Instrument 550. The boom 
for Instrument 536 utilizes the same BI-STEM technique and 
material (beryllium copper). The material is 0.005 cm (0.002") 
thick X 3.56 cm (1.4") wide and forms a 1.27 cm (0.5") diameter 
X 20 meters (66 ft) long tube. The combined weight of the boom 
and deployment mechanism is 2.4 kg (5.3 lb). 

Automatic limit switches indicate full extension and retraction of 
the boom. As with the boom for Instrument 550, explosive devices 
are used to separate and eject the boom if the retraction 
mechanism fails, 

5. 2,1, 5. 2. 4 Igloo 

The igloo is a pressurized vessel containing support equipment for 
experiments on pallet-only mode Spacelab missions. It is being 


^Spacelab Payload Accommodations Handbook, ESTEC SLP/2104, 
May 1975, 


5.2,1-37 


developed by the ERNO consortium under the direction of ESRO/ESTEC 
with the intent that it also be provided as a standard off-the- 
shelf module for other pallet-only mode users. 

The igloo provides a convenient centralized location for those 
equipment which must service all of the pallet-mounted instru- 
ments. It is mounted off the edge of the front pallet, therefore 
does not take up pallet space required by the instruments. It 
also provides a pressurized environment for laboratory type equip- 
ment not designed to operate in vacuum. Most, if not all, 
interfaces between the pallet-mounted equipment and the Orbiter 
interfaces at station Xo 14,630.4 mm (576 in) will be provided 
through the igloo. 

For ASF mission applications, the following CDMS and EPOS equip- 
ment will be mounted within the igloo for the pallet-only mode: 

a. CDMS equipment. 

(1) 3 computers. 

(2) 2 Input/Output (I/O) units. 

(3) 1 mass memory. 

(4) 3 subsystems RAU's. 

(5) 1 payload C&W logic electronics. 

(6) 1 experiment A&A electronics. 

b. EPOS equipment. 

(1) Experiment inverters. 

(2) 1 emergency battery. 

(3) 1 power control box. 

(4) 1 secondary power distribution box. 

The design of the igloo is such that no changes are necessary for 
ASF missions. Necessary Vfiring and ducting are permanently 
installed and the built-in environmental control system will 


provide an environment of 15 to SQ^C with a heat rejection capa- 
bility of 1.5 kW (5,115 Btu/hr). 

Connectors for power supply and data lines are provided at the 
removable bulkhead. 

The usable volume for equipment accommodation is approximately 
0.7 cu. m. (20.6 cu ft) and equipment weights up to 290 kg 
(641 lb) can be accommodated. The equipment will be mounted on 
platforms which are adjustable in their relative position to 
accommodate various sizes of equipment. 

5. 2. 1.6 Analyses and Trade Studies 

5. 2. 1.6.1 AMPS Pointing System 

Several different concepts were investigated for pointing plat- 
forms. Among these was the ERNO IPS utilizing ring gimbal versus 
inside out and suspended control moment gyro controlled platforms 
(figure 5.2.1-17), the Ball Brothers, Small Instrument Pointing 
System (SIPS) was used only for the purpose of establishing con- 
ceptual feasibility, mainly because of its versatility. Other 
systems available in the time frame for ASF missions (1981) will 
be considered during subsequent phases. 

Instruments contained within AIM lA may be pointed at different 
elevation angles to those contained within AIM IB. In addition, 
APS 1 (AIM's lA and IB) may be pointed independent of APS 3 
(AIM's 3A and 3B). The variety of pointing choices between 
instruments is obvious using this method. Also since the AIM 
modules are of the same size, instruments may be interchanged as 
requirements vary. Figure 5. 1.1-5 shows the wide range of 
pointing which can be accomplished with the AIM's. 


5.2.1-39 


SOFT MOUNTED, CMG-CONTROLLED PALLET 



TYPICAL RING GIMBAL SYSTEM 



INSlDE-OUT GIMBAL SYSTEM 


Figure 5.2.1-17. - Candidate pointing platform study concepts 


5.2.1-40 


5.2.1 ,6.2 Booms 


The BI-STEM boom configuration was chosen over the Quasi-Biconvex j 
fiberglass coilable lattice, and articulated lattice because of 
its inherent thermal bending stability. One of the design 
requirements is a static deflection angle of less than 0.5 degree 
to allow determination of the position of the Instrument 550 
through the interrogation of the attitude of the boom base. 

Solar radiation on one side of the boom combined with the deep 
space view on the other creates a severe differential thermal 
bending problem. The geometric and thermal configuration of the 
perforated BI-STEM allows solar radiation to pass through the near 
wall of the boom and impinge on the far wall thus producing equal 
thermal expansion of opposite walls and avoidance of thermal 
bending. Table 5. 2. 1-4 compares candidate materials. 

The boom payload (Instrument 550) will deviate from its theoreti- 
cal position while being scanned because of the acceleration force 
imposed on it. If the error angle is limited to 0.5 degree, the 
actual position of Instrument 550 may differ 15.7 cm (6.2") from 
the position measured by the pointing platform (APS), In a 
weightless environment the acceleration causing this force and 
thus the deflection is 0.152 m/sec^, 0.5 ft/sec^. 

The scan period is defined as the time to traverse the maximum 
accelerator beam field. The boom is accelerated at the maximum 
rate which will not exceed the 0.5 degree allowable deviation 
error to a point midv/ay across the 7 0 meter scan field where it 
will then be decelerated at the same rate to zero for the com- 
pletion of one scan line. The total traverse time for one scan 
line is found to be 29,6 sec. Therefore, the beam may be 
accurately scanned in either of two methods: 

a. Since acceleration is theoretically constant, deflection is 
constant and may be calibrated out. 

b. The boom may be positioned and the accelerator discharged. 

The boom is then moved to a new position and the accelerator 


5.2.1-41 


Materials 

Yield Strength/ 
Density- @ 70®F 
in X 10-^ 

Yield Strength/ 
Density @ 400° F 
in X 10 

Thermal 

Conductivity BTU 
in/Ft3 -Hr°F 

Thermal 

Expansion . 

(in/in°F)X10^6 

Magnetic 

Beryl 1 i urn 
Copper 25 

(605) 

(537) 

(750) 

(9.3) 

No 

Stainless Steel 






17-7 PH 

(656) 

(585) 

(146) 

(9.5) 

Yes 

PH15-7MO 

(710) 

(656) 

(146) 

(8.5) 

Yes 

Maraging Steel 
300 

(1003) 

(865) 

(138) 

(5.6) 

Yes 

Titanium 

Ti-6A14V 

(938) 

(812) 

(50) 

(4.9) 

No 

Inconel -X 

(419) 

(386) 

(83) 

(7.0) 

No 

Aluminum 






7075-T6 

(660) 

(300) 

1 

(1580) 

(13.1) 

No 










again discharged. This will provide a dot matrix rather 
than a line matrix. 

No time period has been allowed for a boom settling-out period 
because of the unknown damping characteristics of the boom. 

The curvature of the boom is found by writing an expression for 
the strain energy due to bending and thermal gradients and finding 
the curvature required to make it a minimum. 

Assumptions : 

(1) Window pattern distributes radiation to back side of boom 
regardless of orientation. 

(2) Axial temperature variation along boom surface is < 10°F at 
any instant. 

(3) Inside surface coating reflects diffusely. 

(4) Conductivity along seam is same as elsewhere. 



er Js 
2K‘t 




where : 



e 

r 


V = 
t 


a 


0 


radius of curvature due to solar irradiation 

coefficient of thermal expansion 
radius of boom 
solar radiation flux 

effective conductivity of boom considering effect 
of hole pattern 

strip thickness 

fractional window area of holes 
solar absorptivity of outer surface 





= solar absorptivity of inner surface. 

0 = angle between boom axis and solar flux. 

If a perforation or v/indow area is chosen such that 
a 

A = — then thermal bending can be eliminated. 

W ^ 

for outer surface = 8 percent for polished silver. 

for inner surface = 95 percent for flat black 
then = 8 percent. 

But, assume half the radiation passes through holes in 
the backside of the boom, 

then A^ = 16 percent. 

Test results have shown actual thermal bending to be 
very close to theoretical calculations, especially if 
degradation of the silver plating is accounted for. 

5. 2. 1.6. 3 Subsatellite Separation 

The devices normally used for separation include linear explosives 
(flat linear shaped charge, mild detonating fuse, primer cord, and 
various encapsulated designs), explosive bolts and nuts, V-band 
clamps, ball locks, pin pullers, and cable cutters. The devices 
normally used to perform an ejection function and/or obtain 
separation velocity include springs, thrusters, retro-rockets, 
and hot or cold gas systems. A limited evaluation of the advan- 
tages and disadvantages of these candidate systems as related to 
the subsatellite mount requirements is presented in table 5. 2. 1-5. 
The selection of separation devices quickly narrows down to the 
ball -lock and the collet mechanisms which are the only mechanisms 
which can be reused. The obvious advantage is the combination of 


5.2.1-44 


J* 




O 



cn 

ro 


1 

C71 


TABLE 5. 2. 1-5, -SEPARATION AND EJECTION DEVICES 




Separation Devices 

Ejection Devices I 


Linear 

Explosive 

Contained 

Linear 

Explosive 

Explosi ve 
Bol ts/Nuts 

flail 

Lock 

V- 

Band 

Pin 

Puller 

Cable 

Cutter 

Collet 

Mechanism 

Springs 

Thruster 

Rockets 

Hot/ 

Cold 

Gas 

Contained 

Cold 

Gas 

load capability 

E 

G 

E 

G 

E 

G 

G 

E 

- 


_ 



Uniform load 

E 

E 

G 

G 

E 

P 

P 

G 

- 

- 

_ 



iiinimum shock 

P 

P 

G 

E 

E 

P 

G 

E 

P 

G 

E 

E 

E 

Minimum impulse 

P 

P 

P 

E 

E 

E 

E 

E 

- 


_ 



Minimum tipoff 

P 

P 

G 

E 

G 

E 

E 

E 

G 

P 

G 

E 

E 

No contamina- 
t1 on 

P 

E 

G 

E 

E 

E 

E 

E 

E 

G 

P 

G 

E 

flo debris 

P 

E 

G 

E 

G 

E 

E 

E 

E 

E 

P 

E 

E 

Kaintainabili ty 

P 

P 

G 

E 

E 

G 

G 

E 

E 

P 

P 

P 

G 

Reusabi 11 ty 

P 

P 

P 

E 

P 

P 

P 

E 

E 

P 

P 

P 

G 

Safety 

P ' 

P 

G 

E 

t 

E 

E 

E 

E 

P 

P 

P 

G 

High 

reli abili ty 

G 

1 

G 

E 

E 

E 

E 

P 

E 

E 

P 

E 

G 

G 

Minimum v;eighC 

p ; 

P 

E 

G 

G 

E 

G 

G 

P 

G 

E 

G 

G 

Minimum volume 

G 

G 

E 

G 

G 

G 

E 

G 

1 P 

P 

E 

G 

G 

Survival of 
temp extremes 

P 

P 

G 

E 

G 

6 

6 

E 

G 

G 

P 

P 

6 

Survival of 
radi ation 

P 

P 

G 

E 

E 

E 

E 

E 

E 

P 

a 

. G 

G 

Minimum AV 

- 

- 

- 

- 

- 

- 

- 

- 

E 

E 

E 

E 

E 

Predictable AV 


- 

- 

- 

- 

- 


- 

P 

1 P 

G 

i ^ 

E 


: Note^ E =* Excellent 
D - Good 
P = Poor 





separation and retrieval /retention mechanisms in one unit. Of 
these two, the collet offers the higher load bearing capability. 
This factor is significant in that only one mechanism in the 
center of the mount is necessary for separation/retention. 

Selection of an ejection device eliminated thruster, rockets, and 
expelled hot/cold gas because of debris and contamination. Con- 
tained cold gas was chosen over a spring because of the spring's 
higher weight and shock characteristics and lesser AV predict- 
ability. Furthermore, a device to retract the spring prior to 
retrieval would be necessary with this concept. 

The fingers ^ the collet are analyzed as a cantilevered beam with 
an initial deflection and an axial load equivalent to controlled 
crash conditions. 


The equation for the total stress for a trapezoidal finger cross 
section is found to be: 


^total ^axial ^ ^initial 


total 


W6 


TTt(Dg - t - 0.019 in) 



where 

W = weight of the satellite 

G = crash "G" load in +X direction = 9 

D = outside diameter = 2.54 cm (1") 

D. = inside diameter = 2.22 cm (0.875") 

E " modulus of elasticity =19.3 x 10^^ N/M^ (28 x 10^ psi) 
n = number of fingers =8 


5.2.1-46 



L = finger length = 5,08 cm (2") 

S = initial finger deflection = 0.254 cm (0.100") 




The stress analysis of the collet fingers for a 6,124 kg (13,500 
pound) load shows a stress of 7.5 x 10^ N/M^ (108 kpsi). Using 
4340 steel with ultimate strength 12.6 x 10^ N/M^ (182 kpsi), 
the margin of safety is: 


M.S. 


128 

76 




0.69 


This margin of safety is more than adequate considering the fact 
that the finger v/as analyzed as a perfect cantilever neglecting 
the effect of the collar. 


The collet mechanism, compressed gas and actuator are analyzed in 
three steps: 

a. Isentropic expansion of gas after pilot valve actuation. 

b. Unlocking collet mechanism (Initial 0.635 cm (0.250") 
movement of piston). 

c. Power stroke (continued movement of piston against satellite 
surface). 

The subcritical mass flow of a perfect gas is: 



I 


I 

I * 


5.2.1-47 





where: 

A = piston area - 5.08 cm^ (0.785 in^) 

= tank volume - 0.0018 m^, 0 = 15.2cm {113 in^, 0=6") 

= tank pressure - 2.58 x 10^ N/M^ (3750 psig) 

w = gas mass flow-gram/sec (pounds/sec) 

K = gas constant 

The piston velocity using the impulse-momentum relationship is 



P„AAt + I 
r 


0 


where: 

At = time for satellite separation, sec 

Ig = initial impulse of piston, N sec (pound sec) 

= satellite mass, kg (slugs) 

Pj, = pressure at release conditions, N/M (pounds/inch ) 

V = velocity of piston, m/sec (ft/sec) 

H 

The pneumatic analysis shows that operating from a 25.8 x 10® 
N/M^ (3750 psi), 12.5 cm (6“) diameter sphere, the desired 
separation velocity of 20 cm/sec is achieved. 

Adequate GNg pressure remains for approximately 10 latchings 
after satellite retrieval. 

5. 2. 1.6. 4 Mass Properties 

Tables 5. 2. 1-6 and 5. 2. 1-7 show the overall mass properties of 
the ASF payload. All pallet-mounted hardware is included. 
Items of relatively small mass (RAU's, heat exchangers, etc.) 
have been included in the mass properties of the pallet. 


5.2.1-48 


TABLE 5.2. 1-6.- HEIGHT AND BALAHCE 


Coinponent 


Pallet 1 + Igloo 
Igloo electronics** 

APS 

AIM lA 
AIM IB 
RAU (0) 

J-Box (Z) 

ColdpTateSt thermal capacitors (2) 
Sun sensor 


Pallet 2 
Subsatel 1 f te 
Launch mechanism 
RAU (2) 

J-Box (2) 

Cr>ogenic Tanks (4) 
5 tructure 


Pallet 3 

Co Opiates, thermal capacitors (2) 
J-Bex (2) 

RAU (7) 

APS 

AIM 3A 
AIM 3B 

Star Tracker, gyros, etc. 

Totals 



Orbiter Coordinates 


Pallet 4 

Coldplates, thermal 
J-Box (2) 

RAU {6) 

Power unit 
Capaci tor bank 
High voltage 
Instrument 304 
Instruments 303/549 
Instrument 536 

Totals 

Pallet 1 
Pallet 2 
Pallet 3 
Pallet 4 

ASF Totals 


thermal capacitors (10) 


944 


6,044 

3,285 

5.214 

3,144 


433. 

.0 

743, 

9 

18, 

,895 

318. 

3 

668. 

8 

16, 

.998 

1100. 

.0 

1 744. 

,5 

18, 

,910 

464. 

0 

744. 

,5 

18, 

,910 

Z86. 

0 

744. 

.5 

18, 

.910 

21. 

6 

744. 

.5 

18, 

.910 

4. 

0 

691 . 

.5 

17, 

,564 

39. 

,0 

786. 

8 

19, 

.985 

25. 

.0 

’ 796. 

,0 

20, 

.218 

2740. 

.9 

736. 

,6 

IB, 

,710 

428. 

0 

862. 

.5 

21 . 

,9C8 

702. 

,0 

892. 

.0 

22, 

.657 

10. 

,0 

Q92. 

.0 

22. 

»657 

5, 

,4 

875, 

.0 

22, 

.225 

4, 

.0 

809, 

.5 

20 

.561 

295. 

,0 

B45, 

.0 

21 

.463 

45. 

.0 

864, 

,5 

21. 

.958 

1489. 

,4 

873, 

,1 

22 

>177 

42B, 

.0 

980. 

.5 

24, 

.90S 

39, 

,0 

1022. 

.8 

25, 

,979 

4. 

.0 

927, 

.5 

23, 

,559 

18. 

.9 

980, 

.5 

23 

.905 

1100 

.0 

980, 

.5 

24 

.90S 

330, 

.0 

930, 

.5 

24 

,905 

361. 

,0 

980, 

.5 

24 

.905 

B3 

,0 

977 

.5 

24 

.829 

2363. 

.9 

981 , 

.0 

24 

.917 

428, 

.0 

1098. 

.5 

27, 

.902 

794, 

.0 

1098. 

.5 

27 

.902 

4. 

.0 

1045. 

.5 

26 

.566 

16, 

.2 

1104. 

.2 

26, 

.047 

45, 

.0 

1066, 

.0 

27, 

.076 

540. 

.0 

1104, 

.4 

28 

.052 

no 

.0 

1104, 

.4 

28 

.052 

40, 

,5 

1089, 

.8 

27 

.681 

40, 

.5 

1104, 

.4 

28 

,052 

7. 

.5 1 

1118. 

.2 

28. 

.402 

1425, 

.7 

noQ. 


27: 

.943 

2 740. 

9 

736, 

.6 

18 

,710 

1489, 

.4 

B73, 

.1 

22: 

.177 

2363. 

9 

981 , 

.0 

24. 

.917 

1425. 

.7 ; 

1100. 

.1 

27. 

.402 

8019, 

,9 

838. e 


22 

.825 


in 

mm 

i n 

mm 

5.6 

142.0 

359.1 

9,121 


-36.0 914.0 

^36.Q 914.0 


5.7 145.0 

0 0 
0 0 
0 0 
0 0 
-35.0 -914.0 
36.0 914.0 

52.5 1334.0 


0 0 
-9.5 -241 .0 
19.7 -500.0 
9.0 229.0 

-19.7 -500.0 
-19.7 -500.0 
10.2 259.0 

-37.4 -950.0 


I* (Deployed) 


412.0 10,465 

493.7 12,540 

493.7 12.540 

493.7 12,540 


416.4 10,577 


9,721 

9,121 

9,093 

8,814 

10,660 

9,830 

10,660 

10,660 

10,173 

9,949 

9,121 

9,093 

8,814 

8,814 

9,520 

9,030 

9,779 

11,270 

10,759 

11,781 

9.263 


493.7 12,540 

412.0 10,465 


427.8 10,867 


416.4 10,577i 

427.8 TO, 867-' 

1404,3 10,269 


**Tlifs fncludesJ 3 computers (31.8 kg each), 2 I/O units (31.8 kg each), Mass Memory (27,3 kg), 3 RAU (2.7 kg each), 
caw logic (3.6 kg). Power distribution (6.0 kg). Power control (5,0 kg), Emergency battery (78 kg), 400 Ht inverter 
(110.3 kg), and AfeA logic (3.6 kg). 


is 

-POOil Qir4X,riY 


5 . 2 . 1-49 










TABLE 5. 2. 1-7. - WEIGHT AND BALANCE, PAYLOAD ON PALLETS AND PAYLOAD 

CHARGEABLE SUPPORT HARDWARE 


Orbiter Coordinates 


Component 


Weight 



lb 

kg 

in 

mm 

in 

mm 

in 

mm 

in 

mm 

ASF Pallets 

17,687 

8,019.9 

898.6 

22,825 

0.7 

18 

383.1 

9,731 

404.3 

10,269 

Radiator Panels* 

193 

87.5 

1213.1 

30,813 

0 

0 

472.8 

12,009 

404.7 

10,279 

Mission Kit-Og Tanks 

2,254 

1 ,022.2 

1115.5 

28,334 

0 

0 

300.8 

7,640 



-Hg Tanks 

874 

396.3 

949.3 

24,112 

0 

0 

300.8 

7,640 



OMS Ki t 

2,978 

1,350.6 

1249.5 

31,737 

0 

0 

388.0 

9,855 



Payload Specialist 
Station 

405 

185.0 

540.5 

13,729 

-53.5 

1359 

i 

446.0 

11,328 



ASF Mission Total 

24,391 

11,061.5 

959.8 

24,379 

-0.4 

-10 

374.9 

9,523 

389.7 

9,899 


*This includes the two aft panels only. 

Mote; Total weight does not include cable harnesses 


















1 




The Orbiter Imposes strict CG location constraints to allow an 
aborted launch condition, in addition to enabling a safe landing. 
Figures 5.2.1-18, 5.2.1-19, and 5.2.1-20 show the location of 
the composite CG within the CG limitations imposed by the Orbiter. 
The shaded areas of each envelope indicate the launch condition 
constraints when overall Orbiter mass is greater than the landing 
mass (expendables, non-returnable satellites, etc.) 

The coordinate system utilized in tables 5. 2. 1-6 and 5. 2. 1-7 and 
figures 5.2.1-18, 5.2.1-19 and 5.2.1-20 is that of the Orbiter; 
Station Zq = 1016 cm (4-00 in) is the geometric centerline of 
payload bay envelope and the Station Xq - 1478.3 cm (582 in.) is 
the forward edge of the envelope. 

As shown in the three figures, the ASF composite payload CG falls 
well within the Orbiter constraints. 

5.2.1 .7 Conclusions and Recommendations 


5. 2. 1.7.1 Conclusions 

Study results indicate that in the area of active thermal control, 
the capabilities of the ASF coolant system and the Oru^ter ATCS 
exceed the expected ASF thermal dissipation (29,500 Btu/hr capa- 
bility versus 24,000 Btu/hr expected). The ASF coolant system 
will utilize the ERNO designed cold plates, thermal capacitors, 
pumps and the Orbiter heat exchanger. Conventional passive thermal 
control techniques will provide greater flexibility in the design 
and allow better control of dissipation paths. 

Cryogenic cooling of instruments 118 and 126 presents the greatest 
challenge for thermal control. The study results indicate that 
a closed loop cryogenic system is not practical on the basis of 
electrical power required if the instruments' housings are to be 
cooled (more than 5 kilowatts of power are estimated to be 
required). An open loop system requires considerably less 


5.2.1-51 






WEIGHT {x 1000) 

WiAXIlViUIVi DESIGW 



Figure 5.2.1-19. 


X-axis CG location. 


-54 


WEIGHT (x 1000) 
KG LBS 



Figure 5.2.1-20 





10 


20 


Y-axis CG location. 



electrical power but will require significantly greater quantities 
of cryogen. The amount of cryogen required depends on the heat 
load and the duration of the operation. Further effort is 
required to establish heat loads when the instrument designs are 
esfabl i shed. 

The mechanical and structural aspects of the ASF payload were 
evaluated to some considerable depth with preliminary designs 
being established for instrument pointing, boom and boom actua- 
tion and subsatellite retention and ejection. Results indicate 
that the ASF instruments, subsatellite and support equipment can 
be installed and serviced within the operational and environmental 
requirements and constraints defined using the modified Ball 
Brothers pointing system, the perforated BI-STEH boom configura- 
tion and the integral collet/GN 2 subsatellite retention/ejection 
system. 


5. 2. 1.7. 2 Recommendations 

The following recommendations apply to the TSMS: 

a. Perform a detailed thermal analysis establishing heat loads 
for instruments 118 and 126 when instrument designs are 
further defined. Calculate the flow rate of cryogen required 
to cool the detectors and the housings and determine the 
volume and weight of cryogen required from the flow rate 
required and the instrument operating time. 

b. Analyze the effect of the cryogen discharge cloud from the 
open loop cooling system on the operation of ASF instruments. 
Evaluate methods of reducing the effects of the discharge 
(e.g., use of Orbiter vent or dump lines to route exhaust 
gases away from instruments). Alternative means of cooling 
may have to be evaluated if contamination from open loop 
cooling cannot be tolerated. 

c. Perform a trade study on installing the APS directly to the 
Orbiter payload attach trunions instead of using the equip- 
ment pallets. Weight would be reduced but structural and 


5.2.1-55 









5.2.2 ELECTRICAL POWER AND DISTRIBUTION SUBSYSTEM (EPOS) 

5.2. 2.1 Introduction 

The study objective for the ASF EPOS was to determine if the 
power needs of the ASF payload could be met within the power, 
energy and thermal dissipation constraints of the Orbiter 
electrical power and ATCS. 

Using the mission operational timeline and the power requirements 
for the individual instruments, support subsystems and the 
subsatellite, average and maximum operational power levels were 
defined and total electrical energy required for a 7-day mission 
was established. A conceptual ASF EPOS utilizing ERNO designed 
equipment to the maximum extent possible was developed. 

Study results show that utilizing two Orbiter energy kits and 
time phasing the high power users, the total energy and maximum 
power required for the ASF payload can be provided (with 
substantial margin) within the Orbiter energy, power and thermal 
dissipation constraints. 

5. 2. 2.2 Requirements 

The power and energy requirements used to size the EPOS were 
derived from the following sources. 

a. id's (appendix B). 

b. AE satellite descriptions (used to size ASF subsatellite 
requirements). 

c. Support subsystem descriptions. 

d. ASF mission timelines (see figure 4. 1.5-1 in section 4.0). 


5. 2. 2-1 


5. 2. 2. 2.1 Functional 


> 


The EPOS will provide the following functions to the ASF instru- 
ments and the support subsystems from prelaunch to postlanding. 

a. Primary electrical povier, 

b. Secondary electrical pov/er. 

c. Emergency electrical power. 

d. Power conversion, inversion and conditioning. 

e. Power distribution, control and overload protection. 

f- Data for subsystem status verification, test, maintenance, and 
diagnostic support. 

The EPOS will provide full operational capability before and after 
the first failure with no degradation of power quality. Levels 
and time duration of power dropouts and transients during switch- 
over from primary to secondary power sources will be minimized. 

After a second power source failure, the EPOS will provide 
sufficient power to maintain the ASF payload in a safe condition. 

5. 2. 2. 2. 2 Performance 

Power requirements for the ASF scientific instruments, subsatel- 
lite (before deployment) and the sup;^ort subsystems were identified. 
Table 5. 2. 2.-1 lists the power required by each of the instruments 
and the subsatellite. The column titled "System Input" lists the 
levels and types of voltages required of the Orbiter or ASF 
primary power sources by each of the instruments. The basic 
elements (emitter, cathode) of the Laser Sounder (Instrument 213), 
Electron Accelerator (Instrument 303), and the Magnetoplasma- 
dynamic (MPD) Arc (Instrument 304) instruments require high 
voltage; high power not directly available from the Orbiter or 
ASF primary power sources. These special power requirements are 
listed under the heading "Element High Voltage/Power Input". 


5. 2. 2-2 


TABLE 5. 2. 2-1. -ASF INSTRUMENT/SUBSATELLITE POWER 


Voltage 


Power (Watts) 


Instrument 


Subsatell i te ' 


Systein^^ ^ 
Input 


28 Vdc 

115 Vac, 400 Hz 
28 Vdc 
28 Vdc 

115 Vac, 400 Hz 


28 Vdc 
28 Vdc 

28 Vdc 
28 Vdc 
28 Vdc 
28 Vdc 
28 Vdc 
28 Vdc 
28 Vdc 

28 Vdc 


El ement ^ 


Operating 


High voitage/power 
Input 

Standby 

Average 

Peak 


10 

10 

n 


15 

100 



16 

16 

16 


14 

14 

14 


10 

25 

25 

5kV pulses 

no 

1.1 K 

TBD 

(10~® to 10'^ sec) 




30 kV dc 

400 

5 K 

10 K 

500 V pulses 

50 

5 K 

10 K 

(TBD sec) 





120 

120 

140 


10 

50 

190 


4 

4 

4 


5 

5 

10 


5 

10 

20 


3 

10 

10 


100 

100 

100 


200 

300 

300 


Input required from Arbiter or ASF primary power source. 

Special high voltage, high power input not available directly from Orbiter 
or ASF primary power source. 

Predeployment power only. 










standby power represents that required for instrument warmup and 
preoperati onal status checks. Many of the instruments do not 
operate continuously during the mission as discussed subsequently 
in this section. The power listed under the heading "Average'* 
represents the average power over only the time the instrument 
is actually performing its experimental operations. Thus, the 
average and peak power for many instruments are identical. The 
peak power differs from the average operating power for some of 
the instruments since they are operated in a pulsed or modulated 
mode, or periodic control of devices and elements such as relays, 
solenoids, motors, actuators, valves, etc. is required^ 

The subsatellite utilizes the Orbiter primary power sources until 
shortly before deployment. The power requirement listed reflects 
only this predepl oyment power. 

Table 5. 2. 2-2 lists the voltages and power required by support 
subsystems; the APS and the thermal control systems. The 
support subsystems will require power during virtually the 
entire mission from prelaunch to postlanding. A peak power of 
700 watts for each APS is required for only one or two minutes 
during the reorientation of the platforms. An average of 
400 watts is required for the remainder of the operations to 
maintain tracking and stabil ization. A continuous level of 
200 watts for the thermal control system (including cryogenic 
cooling and the freon coolant systems) is required from mission 
stai't to experiment completion. 

5 . 2 . 2 . 3 ASF Timelines and Power Usage 

The ASF timelines shown in figure 4. 1.5-1 in section 4.0 are used 
in this study to determine worst case energy requirements. During 
the first 15 revolutions, the instruments will require little or 
no power. The only power which might be required during this 
period for the instruments is the standby power during revolution 


\i 


TABLE 5. 2. 2-2.- ASF SUPPORT SYSTEM POWER 


Support Systems 

Input 

Power ( 

Watts ) 

Vol tage 

Average 

Peak 

Pointing, Control and Stabilization 

AMPS Pointing System (2 systems) 

Command and Data Management 

Aft Crew Station* 

Electrical Power and Distribution 

Cryogenic Cooling and Active 
Thermal Control System 

28 Vdc 
28 Vdc 
28 Vdc 
28 Vdc 
28 Vdc 

28 Vdc 

230 

400 ea. 
1915 
61 0 
580 

200 

4335 

230 

700 ea. 
3805 
610 
580 

200 

6825 




*Power to payload unique equipment at th 
including display and control equipment is no 
but must not exceed that allocated. Energy u 
ment, however, is charged to payload. 

e aft crew 
t charged t 
sed by thes 

s tati on 
0 payload 
e equip- 











15 for preparation and status checks. The experiments will be 
performed from revolution 16 through revolution 80 with some 
instruments powered continuously during this period and others 
sequenced as shown. 

The support subsystems will be powered continuously in the time 
span shown in figure 4. 1.5-1. 

The pointing systems will require peak power only during reorienta- 
tion for 1 to 2 minutes. During the tracking or hold mode a 
constant level of power is required. 

The instruments which use the greatest amount of power are the 
Laser Sounder (Instrument 213), the Electron Accelerator (Instru- 
ment 303) and the Hagnetop 1 asmadynami c Arc (Instrument 304). 

The Laser Sounder will operate over much of the orbit to provide 
maximum global coverage. The voltage pulse applied to the 
emitter is 10 nanosecond minimum in duration and is applied at 
a one pulse per second rate. The Electron Accelerator operates 
in either a dc , a pulsed, or a modulated mode. The voltage 
applied to the cathode will be 1 to 30 K Vdc and the electron 
current will be controlled by controlling the grid voltage. 

The MPD Arc will operate in a pulsed mode with the voltage 
applied to the cathode being 1 to 10 milliseconds in duration. 

The duty cycle will be determined by the maximum allowable 
Orbiter power drain. 

Instruments 118 and 126 require cryogenic cooling. Cooling starts 
some period prior to initial use of the instruments and continues 
as long as the instruments are operating. Instruments 213, 303 
and 304 will require active thermal control using the freon loop 
and cold plates. Other equipment may also be tied in to this 
coolant system. 


5. 2. 2-6 


In addition to the general guidelines and assumptions listed in 
para. 2.3.4, others unique to the EPDS were used during the 
study. These are the following. 

a. The ASF EPDS will provide centralized processing and distri- 
bution for both instrument and support subsystem primary 
power. A single point ground will be provided for the 
instruments. Subsatellite power will be provided from this 
centralized system until just prior to deployment. 

b. Primary input power to the ASF will be 28 Vdc + 4 Vdc, and 
three phase 115 Vac, 400 Hz. 

c. Special power conditioning (conversion, inversion, regulation) 
will be provided by the using equipment. 

d. In addition to the primary power source, backup and emergency 
sources will be provided. The emergency source will be used 
only in safing the ASF payload in the event both primary and 
backup sources fail. 

5. 2. 2. 5 Capabilities and Constraints 

The Orbiter v/ill provide electrical power to the ASF payload 
during all phases of the mission. The primary constraints are; 

(1) maximum power capability of the Orbiter fuel cell, (2) avail- 
able energy, and (3) heat dissipation. 

5. 2. 2. 5.1 Interfaces 

Four interfaces for payload access to Orbiter power will be 
provided. Primary and secondary interfaces are located at station 
Xq 17653 (695 in) on the right hand side just below the longeron. 
Two interfaces v/ill be provided at station 33197 (1307 in) at 
the aft bulkhead. Power will normally be supplied to only one 
of the four interfaces at a time. However, more than one outlet 
can be used by the payload at the same time as long as these 



separate outlets are not tied together within the payload and 
provisions are made such that no single failure or series of 
failures within the payload systems can cause loss or degradation 
of Orbiter power. 

5. 2. 2.5. 2 Orbiter Fuel Cells 

Each of the three Orbiter fuel cells will provide up to 12 kW of 
power. However, the maximum power available to the payload depends 
on which fuel cell Is used and what phase of operation Is Involved. 
Table 5. 2. 2-3 shows the power levels available at each outlet for 
each operational phase, and the constraints Imposed on the payload 
and the Orbiter. 

5 . 2 . 2 . 5 . 3 Baseline Orbiter Pov/er 

The baseline Orbiter power system can provide 50 kWh of energy to 
the payload for a 7-day mission. Energy above 50 kWh from 
the Orbiter power subsystem may be available without adding kits 
if the mission duration Is less than seven days. As figure 5. 2.2-1 
shows, for a 6.5-day mission, the Orbiter can supply the payload 
with 1 kW of continuous power (156 kWh). For energy levels 
exceeding these, up to five cryogenic energy kits which are payload 
chargeable can be utilized. Each of these kits consists of one 
O 2 tank and one tank. 

5. 2. 2. 5. 4 Kits 

Up to five kits can be Installed In the payload bay below the 
payload and outside the payload envelope. Full Installation 
provisions for the first kit are allocated In the baseline Orbiter 
although the weight of this first kit (and every other kit) is 
chargeable to the payload. Additional kits can be Installed 
within the payload envelope If this becomes necessary. 


5-2. 2-8 

























OL- 














5. 2. 2. 5. 5 Orbiter Power Dissipation 

The maximum allowable Orbiter power dissipation is constrained 
by the heat rejection capability of the radiators. The baseline 
Orbiter ATCS provides a capability of 21,500 Btu/hr heat rejec- 
tion for the payload on-orbit with the doors open. This limits | 

the payload to a power level of 6.3 kW. This capability can be I 

increased to approximately 29,000 Btu/hr (8.5 kW) by the addition of | 

a payload chargeable radiator kit. If the dedicated fuel cell j 

is used with the Orbiter in a powered down condition, the payload 1 

can use up to 12 kW peak power for a maximum duration of 15 
minutes every 3 hours. The electrical potential at the primary 
payload interface (dedicated fuel cell interface at 17653 
(695 in) will be a minimum of 27 volts with a 12 kW load. 

5. 2. 2. 5. 6 Backup Mode 

In a backup mode (one of the three Orbiter fuel cells inoperative), 
the backup interface at station (695) from the main Orbiter bus 
will supply a minimum of 27 volts at 8 kW peak power and 5 kW 
average power during on-orbit payload operations. The aft payload 
power interfaces located at station X^ (1307) are supplied by the 
Orbiter aft local buses. The minimum potential at these interfaces 
will be 24 volts at 1.5 kW average power and 2.0 kW peak power per 
bus. 

5 . 2 . 2 . 5 . 7 Aft Flight Deck Equipment 

For payload unique equipment located at the aft flight deck (within 
the crew compartment) such as displays and control panels, tape 
recorders, etc., the power required is not included in the alloca- 
tions shown in table 5.2. 2-3. This power is included in the Orbiter 
baseline allocations. However, the electrical energy required by 


5.2.2-n 


these equipment at the aft flight deck is chargeable to the \ 

payload. The power allocations for these payload unique aft 1 

flight deck equipment by phases are: | 

I 

a. Ground operations (GSE power) and on-orbit operations - 750 | 

watts average, 1000 watts peak, 

b. Prelaunch (Orbiter internal power), ascent, descent and 
postlanding (Orbiter internal power) - 350 watts average, 

420 watts peak. Peak power is limited to two minutes each 
phase . 

5 . 2 . 2 . 5 . 8 Other Orbit e r Systems 
Other Orbiter power s,yst'^:ri characteristics are: 

a. Peak-to-peak ripple fcr Orbiter electrical power is 0.9 volts 
or less over a broadband of frequencies for both aft flight 
deck and payload bay interfaces. 

b. Orbiter fuel cells have no voltage control requirements for 
loads up to 2 kW except that the voltage will not exceed 
40 volts. 

c. A two wire power/return interface is provided to payloads. 

However, the Orbiter uses a distributed structure return system 
for its own loads. Up to 400 amperes of dc current can flow 
through Orbiter structure during on-orbit operations. 

5. 2. 2.6 Subsystem Description 

Figure 5. 2. 2-2 describes the ASF EPOS equipment and interfaces with 
the Orbiter power system and with the ASF payload. 

5. 2. 2. 6.1 Energy Sources 

The ASF EPOS utilizes the Orbiter fuel cells as the primary source 
of its power. Preliminary assessment indicates that the ASF energy 
requirement (897 kWh) far exceeds the 50 kWh energy allocated by 
the baseline Orbiter electrical power system under worst case 


5 . 2 . 2-12 
































conditions. Therefore two energy kits (two 0 ^ ^nd two H 2 tanks) 
are included in the ASF baseline EPOS configuration. Primary 
power is obtained from the dedicated fuel cell (cell 3) and the 
second outlet at station 17653 (695 in) is used to provide a 
backup power source. In the event power is lost from both 
Orbiter outlets, an emergency source will be available to provide 
power to the ASF payload safety critical functions such as the 
cryogenic tank pressure monitors. The energy required for these 
functions is expected to be a small fraction of that required for 
normal operation and relatively inefficient sources such as silver- 
zinc batteries can be considered. A 28 Vdc silver-zinc battery of 
the same type used on Spacelab (and also on a number of space- 
crafts and boosters) has been selected for the baseline ASF EPOS. 

5.2. 2. 6. 2 Power Conversion, Inversion and Distribution 

The baseline ASF EPOS includes the 2.5 kVA, 30, 115 V, 400 Hz 
inverter provided for Spacelab. Instruments 118, 126 and 213 
presently require this ac power source. However, the final 
instrument designs could include individual inverters with little 
impact on total development or unit cost since qualified, flight 
proven power supply designs are available. 

Centralized dc to dc converters and regulators are not included in 
the EPOS baseline since the current approach is that power conver- 
sions and regulation requirements will be satisfied by using 
equipment with internal provisions. However, the ERNO designed 
dc to dc converter will meet most of the ASF regulated dc power 
requirements and is considered an acceptable option. These 
converters would be located on the individual pallets in the 
standard Spacelab location or under the pallet floor if available 
floor space becomes a factor and the converters do not require 
active cooling. Overall program cost and weight differentials 
between the two approaches should be minor. A variety of available 
converter designs can be used for ASF applications, therefore 
little development cost is involved. Some weight and unit cost 


5.2.2-14 



savings may accrue if centralized rather than dedicated converters 
are used. However, due to the relative conversion efficiencies, 
the overall programmatic trade-offs are not expected to be 
si gni f icant . 

The Laser Sounder (Instrument 213), the Electron Accelerator 
(Instrument 303), and the Magnetopl asmadynamic Arc (Instrument 
304) all require high level voltage sources to operate (see table 
5. 2. 2-1). These high level voltages will be provided to the 
accelerator and the MPD arc through the special high voltage 
power supply shown in figure 5. 2.2-3. A separate power source 
v/ill be required for the Laser Sounder due to the distance 
between the two groups of instruments- Orbiter 28 Vdc power is 
converted to 500 Vdc by power processing unit 1 and this voltage 
is used to charge a large (10 joule, 0.8 Farad) capacitor bank. 
The output of the capacitor bank is converted by a second power 
processing unit to the high voltages (30 kV) required to operate 
the Electron Accelerator. The MPD Arc will use the 500 Vdc output 
of the capacitor bank directly through a solid state switch. 

Fundamental issues involving the development and utilization of 
the high voltage, high power system are how best to accomplish 
the following. 

a. Generate required voltages. 

b. Minimize voltage attenuation and power losses within power 
conversion and transmission media. 

c. Provide required insulation, minimize corona effects. 

d. Contain generated conducted and radiated EMI. 

Circuit breakers and pov/er sv/itches will be provided to isolate 
the ASF central bus from the Orbiter power sources and the ASF 
pov/er busses from the individual instruments and equipment. As 
shown in figure 5. 2. 2-3, circuit breakers are also provided in 
the baseline Orbiter to protect the Orbiter power sources. 


5.2.2-15 


ELEMENTS 


SUBSYSTEM 

A POWER PROCESSING UNIT {PPU) I (28V/400A/ /50av/23A} 

B LOW VOLTAGE/HIGH CAPACITANCE (LV/HC) BANK {5Q0V/,8F) 

C PPU II (500V/400A/ /30.000V/10A), DRAINAGE RESISTOR (Rq), 

BUFFER CAPACITOR (Cgl, HIGH VOLTAGE SWITCH 
D SOLID STATE SWITCH (SSS I), MAGNETOPLASMADYNANIIC 

(MPD) ARC (5a0V/200,G0QA) 

E ELECTRON ACCELERATOR (30,000V 7A) 


Figure 5. 2. 2-3. —ASF Particle Accelerator 
High Voltage Power System. 


5.2.2-16 










} 


These circuit breakers and those in the ASF igloo will be used to 
provide redundant means of isolating the two Orbiter power outlets 
used by the ASF payload. 

Further effort is required to establish the criteria for selection 
of circuit breakers over a combination of relays and fuses for 
overload protection and to determine the operational and safety 
requirements that dictate which circuit breakers can be remotely 
controlled and which should be located at the aft crew station. 

In addition to the circuit breakers, the need for overload protec- 
tion within each equipment or instrument should be assessed. 

The EMI filters will be required to protect the Orbiter and ASF 
power systems from conducted interference effects and to reduce 
the effect of Orbiter power and ground system noise and transients 
on the ASF system. 

The two wire power/return interface provided by the Orbiter will 
be utilized by the ASF power system. A single point return bus 
for the ASF payload will be provided in the igloo. Each pallet 
power distribution bus will have a return bus dedicated to the 
instruments and equipment on the pallet. The Orbiter power system 
uses vehicle structure as its dc return. It is expected that as 
much as 400 amperes of dc current can flow through the Orbiter 
structure during mission operations. As part of the on-going 
EMC evaluation, the possibility of structure noise feeding into 
the ASF power system through the by-pass capacitors and its 
impact on the payload operations must be assessed. 


Each power and return bus in the individual pallet power distri- 
bution boxes interfaces directly with the respective centralized 
busses in the igloo. In addition, the individual pallet dc power 
busses are connected in series (as are the return busses) for 
redundancy purposes. 


5.2.2-17 




The connections between the dc power busses are made through 
normally opened switches to provide isolation, if required. 

Preliminary assessments indicate that to keep line voltage drops 
to less than 10 percent of power voltage (2 to 3 volts for dc 
power) wire sizes of up to 4-0 (0.04 ohms/310 m {1000 ft)) will 
have to be used for the primary Orbi ter-to-pay 1 oad interface and 
for the high power users such as instruments 303 and 304 assuming 
individual harness runs (to load and back) are less than 62 
meters (200 ft). Together with insulation, this size of wire 
will measure about 1/2 inch in diameter and weigh over 1.5 kg/ 
meter (1 Ib/ft). Other wires used will range in size from 4 to 
20 gauge. 

5 . 2 . 2 . 6 . 3 EPPS Equipment Characteristics 

Table 5. 2. 2-4 summarizes pertinent characteristics of the ASF EPPS 
equi pment. 

5. 2. 2. 7 Analyses 

Worst case analyses of the total ASF payload power and energy levels 
were performed using the power and timeline requirements discussed 
in paragraph 5. 2.2.2. Table 5. 2.2-5 shows the power level required 
for each major phase of the mission requiring significant changes 
in instrument or associated equipment operations. Two hours 
before lift-off, a transfer is made from ground support power to 
internal Orbiter power. From insertion into mission orbit until 
orbit revolution 16, the crew makes preparation for the start of 
the experimental operations. During this period, only the support 
subsystems are assumed to be pov/ered. During revolution 15, the 
power to instruments other than those associated with the particle 
accelerators is turned on and the platform pointing closed loop 
servo system is powered. The cryogenic cooling systems for 
instruments 118 and 126 are turned on during revolution 7. During 
revolutions 16 through 80, the experiments are conducted. 


5.2.2-18 


TABLE 5.2.2 


EPOS EQUIPMENT 


1. Dc-ac 30, 2.5 kva> 115 V, 
400 Hz Inverter 


2. Power Distribution Box 


3. Energy Battery 


4, Power Control Box 


5. Harnesses (including 
connectors ) 

a. 4-0 Gauge-183m (600 ft) 

b. 4 Gauge-163tn {600 ft) 

c. 10 Gauge-109m (3600 ft) 

d. 20 6auge-8536m 
(28,000 ft) 


6. Energy K1t 

a. O 2 ctnd tanks 
(and fittings) 

b. Cryogenic Peactants 

• 

■ 


ASF EPOS TOTAL 


2 eac 


28 (62 


290 

{640) 

40 

(88) 

55 

(122) 

41 

(90) 


626 

(1381) 

708 

(1564) 

82 

(182) 

1983 

(4378) 


Prior Use 


Spacel ab 


30 (66) 5x10 Spacelab 


78 (172) Spacelab 


5 (11) I 20 Ispacelab 


Modifications 
Requi red 


(TBD) 


(TBD) 


Orbi ter Mission Ki ts 


Capacitor bank and power processing unit vjeights included in instruments 303 and 213 
wf=i gbts . 

(1) Inverter inefficiency 
























































i 


From revolution 81 through 95, pov;er to the instruments are 
turned off. Orbiter maneuvers are performed, the orbit is changed 
for rendezvous vMth the PDS. The subsatellite is retrieved on 
about revolution 95. During revolutions 95 through 112, the 
Orbiter and payload systems are prepared for the return phase. 

From the timelines and the power requirements of the individual 
instruments and the support systems, average pov/er levels required 
by flight phase were established as shown in table 5. 2. 2-5. The 
maximum power required is 6882 watts during revolutions 43 through 
47 which i '■ 42 percent less than the 12 kW maximum available from 
the dedicated fuel cell. 

The average power by phase v/as integrated over the entire mission 
as illustrated in figure 5.2. 2-4 to establish the total energy 
required. The 1730 kWh (50 basic and 1680 kit) of energy available 
from the Orbiter provides a 48 percent margin over the 897.3 kWh 
required by the ASF payload for the 7-day mission. 

The heat dissipation capability of the Orbiter using the payload 
chargeable radiator kit limits power levels to 8.5 kW average during 
the mission and a peak of 12 kW for 15 minutes every 3 hours. 

Figure 5. 2. 2-4 shows that the ASF average power required over 
periods greater than 15 minutes (6.9 kW during revolutions 43 
through 47) results in a 19 percent power margin. Figure 5. 2. 2-5 
shows the peak power required by the ASF payload during a typical 
orbit. The maximum power required by the ASF payload is approxi- 
mately 9 kW for a period of about 15 minutes every 1-1/2 hours 
(one revol uti on ) . 

Although the integration time for the ASF peak power is one-half 
that used to define Orbiter constraints (1-1/2 hours compared to 
3 hours) as indicated earlier, the total integrated power (energy) 
over each orbit is well within the 8.5 kW specified. 


5.2.2-21 


§ g V. 

^ s? ■ 

F 

j3 


TOTAL ENERGY » B97.3 KWH 
AVERAGE POWER « 5^ KW 


MAXIMUM AVERAGE ALLOWED (MAXIMUM OVER AN EXTENDED PERIOD, > 15 MINUTEST 


AVERAGE = B.3 K WATTS 


B5.D KWH 

47.7 KWH 

39.6 KWH 

69.3 KWH 

51,6 KWH 

302.7 KWH 

r- n 1 

/'i* 










REVOLUTION 37 (TYPICAL) 


SUNLIGHT 


DARK 


12 KW LIMIT FOR 15 MINUTES EVERY 3 HOURS 



INSTRUMENTS 
118, 126 & 213 


SUPPORT SUBSYSTEMS (EXCLUDING AFT CREW STATION) 


T 

10 


AMPS POINTING SYSTEMS 12) 


20 30 40 


T“ 

50 


THERMAL CONTROL SYSTEM 


60 70 80 


90 


TIME (MINUTES) 


Figure 5 . 2 . 2-5 Typi ca 1 orbit revolution for peak electrical power 


5. 2. 2.7. 2 High Voltage Source Efficiency 

The particle accelerators baselined for the ASF mission are an 
Electron Accelerator (Instrument 303) and a Magnetopl asmadynamic 
Arc (Instrument 304). Both of these instruments require high 
power and high voltage levels. The required power is provided 
through the use of a 0.8 Farad 500 volt dc capacitor bank fed by 
a dc converter (see figure 5. 2. 2-3). 

A capacitor bank with these characteristics is capable of an 

5 

energy storage of 10 joules and would weigh approximately 540 kg, 
and have a volume of approximately 2 cu m. 

5 

Although the total storage capability of the capacitor bank is 10 
joules, all of this energy is not available for useful energy in 
the accelerator beams. In the case of both the Magnetoplasma- 
dynamic Arc and the Electron Accelerator, efficiency losses in 
both the power conversion and in the guns need to be considered. 

Analysis performed on this study indicates that with a power 
converter interval impedance of 4 ohms, and the capacitor bank 
value of 0.8 Farads, the time required to charge the capacitor 
bank is 14.6 seconds and the efficiency of the high voltage source 
is above 70 percent. 

5 - 2 . 2 . 8 Conclusions and Recommendations 
5 . 2 . 2 . 8 . 1 Conclusions 

The most significant conclusion relative to the ASF EPOS is that 
vn'th logical time sharing by the high power users, there is every 
indication that the Orbiter power and energy constraints can be 
met with adequate margin. The heat radiator kit provided by the 
Orbiter will probably be required. 

On a worst case basis, the total energy required by the ASF payload 
is 897 kWh. Since the baseline Orbiter payload support is only 


50 kWh of energy, two energy kits with 1680 kWh additional energy 

capability will be Included as part of the EPOS baseline. 

5. 2. 2.8. 2 Recommendations 

Results of the study have led to the following recommendations. 

a. More fully develop the EPOS concepts in the areas of power 
control, conditioning, conversion and inversion. Establish 
whether ac power should be provided from a central ASF bus 

or if it can be more effectively provided by the using equip- 
ment. Establish criteria defining the use of circuit breaker 
vs. relays, remote vs. aft flight deck circuit breakers, fuses 
or other overload protection in individual loads. 

b. Identify the safety critical functions which require power 
redundancy. Establish power levels required and perform 
trades/studies to select the most effective power source. 

c. Evaluate the total impact of using extremely high power levels 
on EMC, heat dissipation, required sizes for power and return 
lines, common impedances and conducted interference effects, 
insul ati on , etc. 

d. Evaluate the possible impact on payload operations of Orbiter 
vehicle structure noise coupling through the EMI by-pass 
capacitors i..to the ASF system. 




5.2.3 POINTINS/CONTROL AND STABILIZATION SUBSYSTEM (PCSS) 

5.2.3. 1 Introduction 

The objective of this phase of the study was to establish the con- 
ceptual feasibility of providing precision pointing, tracking and 
stabilization for the ASF instruments. The approach was to eval- 
uate the Orbiter attitude accuracy capability relative to the in- 
strument requirements and to develop a dedicated ASF/PCSS concept 
if the Orbiter capability fell short of these requirements. 

The scope of this phase was limited to: (1) defining the tech- 

niques for pointing and control, (2) defining a conceptual design 
approach, and (3) determining the hardware requirements and func- 
tional interfaces required for pointing/control and stabilization. 

No attempt was made during this study to perform dynamic simula- 
tions and evaluations of the control laws or to analyze the PCSS 
performance. A secondary goal was to research the State-of-the-art 
hardware that can meet the pointing and stability requirements thus 
minimizing development time and cost. Other studies for advanced 
pointing systems are being conducted for payloads that require a 
high degree of pointing accuracy and ^ability. However, these 
are not included in this study. 

The study showed that the Orbiter attitude control and stabiliza- 
tion capabilities are not adequate to meet experimental needs. A 
PCSS concept was develop'ed which consists of two major elements: 

(1) the AMS, and (2) the APS. This section discusses the AMS in 
detail and describes the integrated AMS/APS operations. The 
details of the APS configuration are provided in Section 5.1. 

5. 2. 3. 2 Requi rements 

The ASF payload consists of instruments that require stellar, solar, 
and earth pointing. The ASF experiments require the pointing of 


5. 2. 3-1 


one or more FOV's at a target such as the solar disc, the nadir, 
or along a specific direction. In addition, the FOV or line-of- 
sight (LOS) must have pointing stability that permits experiment 
measurements to be made without distortion. 

Pointing accuracy requirements are usually functions of the instru- 
ment FOV and of target and experiment data characteristics. Sta- 
bility requirements, however, depend on the resolution capability 
of the experiment instruments; that i , the sensitivity of these 
instruments to a LOS rate. 

In general, the stellar instruments usually require large gimbal 
angles, long exposure times, and stringent pointing and stability 
accuracies. The solar instruments remain sun-centered or search 
the surface of the sola*^ disk. The earth looking instruments 
usually require high gimbal rates for tracking earth based targets 
and the use of the Orbiter for maintaining an earth oriented 
attitude with the payload bay toward the nadir. Thus, the atti- 
tude control and pointing system must be capable of meeting these 
various types of requirements. 

5. 2. 3. 2.1 Functional 

Assessment of the Orbiter pointing and stabilization capabilities 
(see paragraph 5. 2.3. 4) indicates that an independent pointing/ 
control and stabilization system is required for the ASF payload. 
The prime mission functions which this system must perform to sup- 
port the experiments are payload reference axes attitude determi- 
nation, pointing/control (target tracking) and stabilization. 

Other functions include providing data for downlink telemetry and 
for onboard display and processing, power conditioning and control 
within PCSS equipment, and data to be used for failure detection 
and isolation. 


5. 2. 3-2 


5,2.3. 2. E Performance 


> 


i 

i 

i 


The principal source for the pjinting accuracy, pointing stability, 
and rate stability requirements is the ASF ID documents (see 
appendix B). A summary of these requirements is listed in 
table 5. 2. 3-1. The definition of these errors and a graphical 
presentation is illustrated in figure 5. 2. 3-1. The justification 
for the pointing accuracy and stability rate requirement lies 
with the scientific community and/or payload users. The remain- 
der of the study is based on that data as defined by the users 
in the id's. 

The instrument stability requirements defined in the ID's vary 
from .003°/sec. to 36°/sec. The instruments pointing accuracy 
requirements vary from 1 minute of arc to 6°. 

5. 2. 3. 3 Guidelines and Assumptions 

Following are the guidelines and assumptions used for this phase 
of the study. 

a. Pointing accuracy knowledge of better than 0.1“ must be 
provided by the payload AMS. 

b. Those instruments which have a requirement of 2“ but whose 
operations {such as TV monitors) are controlled by the crew 
can be hard mounted to the pallet. The crew will use visual 
means to keep instrument LOS on target. 

c. The Orbiter will be operating v/ithin iLs minimum deadband 
(±0.1°) and minimum rate (0.01°/sec) to provide the payload 
with the least vehicle motion. 

d. The AMS will be placed on the reference base requiring the 
greatest accuracy and stability. 

e. The LOS of the cluster of instruments on a given platform 
will be boresighted to a common LOS. 


5. 2. 3-3 


I 


table 5. 2. 3-1* - ASF POIMTIMG AND STABILITY REQUIREMENTS 



Instrument Requirements^ 

Orbiter Capability^ j 

Accuracy 

Stability Rate 

fiicct^rstcy 

Stability 

stability Rate 

AMPS Pointing System (APS): Pallet A-1 
AMPS Instrument Module (AIM) lA 

±1.0° 

±1° 

±0.017° 

±2.5° 

±2.0° 

N/A 

(TBD) 

O.T5°/sec 

0.0Q3°/sec 

(TBO) 

0.017°/sec 

N/A 

±2° 

±0.1° 

0.0T°/sec 

Instrument No. Title 

213 Laser Sounder 

532 Gas Release Module 

AMPS Instrument Module (AIM) IB 

1011 UV Occultation Spectro- 

graph 

1002 Pyrohel iometer/spectro- 

photometer 

534 Optical Band Imager and 

Photometer System 

550 Level IT Beam Diagnostics 

Subsa tell ite: Pallet A-2 

Instrument No. Title 

(TBD) 






AMPS Pointing System (APS): Pallet A-3 
AMPS Instrument Module - AIM 3A 


(TBD) 

O.OOAVsec 

(TBD) 
(TBD) 
0, 05V sec 

±2° 

±0.1° 

0.01°/sec 

Instrument No. Title 

116 Airglow Spectrograph 

178 Limb Scanning IR Radiometer 

AMPS Instrument Module - AIM 3B 

122 UV/VIS/NIR Spectrometer 

124 Fabry-Perot Interferometer 

126 Infrared Interferometer 

Hard Mounted: Pallet A-4 

Instrument No. Title 

303 Electron Accelerator 

304 Magnetpplasmadynamic (MPD) Arc 

536 Triaxial Fluxgate 

549 Gas Plume Release 

±6° 

±2° 

±0.5° 

N/A 

r/sec 

l°/sec 

ROLL ‘^36°/sec 
N/A 

±2° 

1 

to.r 

1 

I- 

[ ±0.01°/ sec 

l 


^ ID Requirements 

^JSC-07700, Vol. XIV* Rev, C, dated July 3, 1974 

ASF - Atmospheric Science Facility 
N/A - Not applicable 


5,2, 3-4 


page is 


jOPPOOR 































POiNTrNG 

ERROR 

EWVELOPE 


POINTING DEFINITIONS 

POINTING ERROR - DEFINES THE TOTAL FRRDR THAT CAN 
BE tolerated BV THE INSTRUMENT OR PAYLOAD, H 
NORMALLY IS THE RSS OF THE POINTING ACCURACY AND 
POINTING stability, 

POINTING ACCURACY DEFINES HOW CLOSE TO THE 
DESIRED TARGET THE INSTRUMENT MUST INITIALLY POINT. 
IT USUALLY IS ASSOCIATED WITH AN INSTRUMENT FIELD 
OF VIEW fFOVI. OR A PARTICULAR AREA AROUND THE 
INSTRUMENT CENTERLINE ERRORS THAT CONTRIBUTE TO 
POINTING ACCURACY ARE USUALLY OF THE STATIC TYPE 
AND RESULT FROM ITEMS SUCH AS MISALIGNMENT, EN 
CODER READOUT, ETC. AND CONSEQUENTLY. POINTING 
ACCURACY IS often REFERRED TO AS A BIAS. 

POINTING STABILITY - DEFINES HOW CLOSE THE INSTRU 
MENT MUST STAY TO THE POINT AT WHICH IT WAS 
INITIALLY POINTED. ERRORS THAT CONTRIBUTE TO THE 
STABILITY ARE NORMALLY OF THE DYNAMIC TYPE SUCH 
AS VIBRATION DISTORTIONS, ELECTRONIC DRIFT. VEHICLE 
DRIFT, GIMBAL MOUNT DRIFT. ETC 

STABILITY DURATION “ DEFINES TIME DURATION DURING 
WHICH STABILITY MUST BE HELD. IT USUALLY IS ASSO 
CIATED WITH INSTRUMENT EXPOSURE TIME OR EXPERI 
MENT SEQUENCE OBSERVATION TIME 



GLOBAL ILLUSTRATION OF DEFINITIONS 


POINTING 




TIME ILLUSTRATION OF DEFINITIONS 


Figure 5. 2. 3-1. — Pointing definitions. 








f. The AMS will have to be updated once per orbit in order to 
maintain the 60 arc second accuracy. Time of update will be 
approximately five minutes. 

g. The operation of the APS will be primarily computer controlled; 
however, the fine pointing of certain instruments will be man- 
ually performed by the crew. The crew will have to activate 
the system operations through keyboard-entered computer pro- 
grams such as: 

(1) Initial alignments. 

(2) Update or realignments. 

(3) Tracking programs. 

(4) Calibration, etc. 

h. The solar instrument group will use a sun sensor. 

i. Instrument 1011 should have a sensor (star or sun) in the 
optical train. The output of this sensor would be available 
as an input into the control loop and operate as a closed loop 
sensor around the target star or sun. 

5. 2. 3.4 Capabilities and Constraints 

The Orbiter avionics system provides pointing and control capability 
through the use of its guidance, navigation, and control (GN&C) 
subsystem. The Orbiter GN&C subsystem consists of an inertial 
measurement unit (IMU), star trackers, and a flight control system 
(including vernier and large reaction control system thrusters). 

The Orbiter vehicle has the capability of attaining and maintain- 
ing desired inertial, local vertical, and earth surface pointing 
attitude within the accuracy defined in table 5, 2. 3-2 and Orbiter 
thermal attitude constraints defined in table 5. 2. 3-3. The Orbiter 
IMU may be initially aligned to 0-1® with a drift rate of 0.1° per 
hour while other errors in the flight control subsystem contribute 


5.2. 3-6 


TABLE 5. 2.3-2. - ATTITUDE POINTING ACCURACY - 
ORBITER REFERENCE SYSTEM 


Reference 

Atti tude 
accuracy 
(3a) 

Attitude 
degrati on 
(3a) 

Duration 
between 
al i gnments 

Inertial 

±0.4° 

0.1 °/Hr 

1 . 5 Rev 

Cel estial 

±0.24° 

0 

Not 

Appl i cabl e 

Earth Target 

±0.4° 

0.1°/Hr ^ 

1 Rev 







TABLE 5. 2. 3-3. - ORBITER THERMAL ATTITUDE CONSTRAINTS 


Orblter Orientation 

Hold capability 
(Hours ) 

Pre-entry thermal 
conditioning re- 
quirement (Hours) 

Any 

A. Other than inertial 

>160 

<12 

hold 

B. 3-axis inertial 

Cycles of 6-hour holds 
followed by 3 hours of 
thermal conditioning 
for worst thermal 
attitudes 

<7 

holds 

1160 

1^2 


Source: JSC 07700, Volume XIV, Rev. C, Julv 3, 19/4 










I 


i 


0.25°. Table 5.2. 3-2 shows that after 1.5 orbits the inertial 
attitude error of the Orbiter as related to its reference system 
is ±0.5°. Using the Orbiter star tracker continuously for atti- 
tude reference, the vehicle reference system error relative to the 
celestial reference can be held to v/ithin 0.25° indefinitely. 

In using the Orbiter IMU to point to an earth target, additional 
errors are introduced due to the Orbiter position and velocity 
o" uncertainties. In order to maintain a 0.5° error, the Orbiter 

IMU must be updated once in orbit. These errors are Orbiter 
reference axes errors relative to inertial or earth target refer- 
ences. For the purposes of payload pointing using the Orbiter IMU, 
an error source (>2°) can accrue due to vehicle flexu'^e, payload 
structure, and payload mounting misalignments. 

The Orbiter flight control system (FCS) is capable of providing 
stability (deadband) of ±0.1°/axis and a stability rate of ±0.01°/ 
sec/axis utilizing the Orbiter IMU and the vernier RCS thrusters, 
provided that all vernier thrusters are operational. When using 
the large RCS thrusters, the Orbiter FCS is capable of providing 
stability of ±0 . 1 °/sec/axi s . 

5. 2. 3. 5 Subsystem Description 

Comparison of the Orbiter pointing and stability capabilities and 
the ASF experiment requirements indicates a need for an accurate 
ASF attitude reference and pointing system for some of the instru- 
ments. The system defined in this section is an inertial system 
with optical updates. It consists of a three axis strap-down in- 
ertial system with star trackers and/or a sun sensor to provide 
updates to compensate for gyro drifts and computer errors. This 
is potentially the most versatile system and is equally effective 
for solar, stellar, and/or earth pointing missions. 




5,2. 3-9 


1 

i 


5. 2. 3. 5.1 Configuration 

The ASF pointing and control requirements defined in table 5. 2. 3-1 
dictate the need of a subsystem that can provide a pointing accu- 
racy of 60 arc seconds. This requirement is considered to be v/ithin 
the state-of-the-art of existing hardware and can be achieved 
using several techniques. Two approaches that can be used are: 

a. A central or master inertial reference system with optical 
updates . 

b. A distributed inertial reference system with optical updates. 

The master reference system can be placed on a separate gimbal 
system or incorporated with the instruments on a given pallet 
platform. This scheme introduces errors because of the required 
transfer of the reference error signals through the gimbals, how- 
ever, these are manageable by design and calibration. Because 
of mechanical errors between pallet segments, optical links between 
gimbal systems for aligniiient control are required. This subsystem 
approach is illustrated in figure 5.2. 3-2. 

The distributed attitude reference system places a star tracker 
assembly (STA) and a gyro reference assembly (GRA) on each 
pointing system in the payload bay. Table 5. 2. 3-4 discus-ses the 
merit of each approach. Both approaches can fulfill ASF pointing 
and stability needs. For this study, the initialized inertial 
reference system with optical updates (star and sun sensors) 
augmented with an optical alignment transfer device was chosen 
(reference figure 5. 2. 3-3). This system was selected because of 
the large range of pointing requirements (0.017® to 6.0°) and the 
non-severity of the pointing accuracy (0.017°) needed for ASF. 

The selected ASF PCSS configuration is shown in figure 5. 2. 3-4. 

It consists of the following hardware: 

a. Digital computer (part of CDMS). 


I 

I 


i 

I 

I 


f. 

■ i 


\ 




1 

I 

I 

i 

I 

i 

H 


5.2.3-10 


LL- 








TABLE 5. 2. 3-4 - CENTRAL VS DISTRIBUTED ATTITUDE REFERENCE SYSTEM 


Central 


Di stri buted 


Advantages 

• Only one star tracker and gyro 
reference assembly required. 

• Relative cost appears to be less 


• Weight requirement is less. 


Disadvantages 


• Computer requirements are higher, 


• Need optical links between gimbal 
systems . 

t Precision gimbal sets are 
required to minimize errors. 


Advantages 

• Experiments are more flexible and 
autonomous , 

9 Reduces computer involvement in 
the control loop. 

I Gimbal precision requirements are 
less stringent. 


Di sadvantages 

• Need one star tracker and gyro 
reference assembly for each 
gimbal . 

• Weight is increased. 


• Apparent increase in relative 
costs . 










EL- 


i# o 


tn 

ro 

OJ 


ORBITER 
GN & C 
COMPUTER 


ASF 

SUBSYSTEM 

COMPUTER 


OPTICAL 

TARGETS 


GIMBAL ANGLE COMMANDS 


GIMBAL POSITION READOUTS 


APS BASE 
MOUNT 
(PALLET 2)1 


LASER sounder! 
APS 1 A1M1A 


APS/PALLET #1 


MISALIGNMENT SIGNALS 


OPTICAL 


TRANSFER 

DEVICE 


GIMBAL ANGLE COMMANDS 


GYRO OUTPUTS 


APS GIMBAL POSITION READOUTS 


STAR TRACKER INPUTS 


STAR TRACKER COMMANDS 



APS CONTROL 
ELECTRONICS 

APS BASE 
MOUNT 
(PALLET 3) 

GRA 


STA 


GIMBAL POSITION COMMANDS 


GIMBAL POSITION READOUT 


TRACKING ERROR SIGNALS 


SOLAR INSTRUMENT 

APS 1 

AIM1B 

SUN SENSOR 


Figure 5. 2. 3-3. — Centralized AMS signal plan. 




MECHANICAL LINK 






ELECTRONICS 

APS 

COMPUTER 

AMS 


DISPLAY, CONTROL ANP 
MONITOR PANEL 


CONTROL 

AND 

MODE 

SELECTION 

K 


E 


Y 


B 


0 


A 


R 


D 

CRT 

DISPLAYS 




^ ORBtTER / 
GN&C COMPUTER^ 


DIGITAL 

COMPUTER 


A/D AND 
D/A 

CONVERTER 




GYRO TORQUING 
ELECTRONICS 


FLIGHT 

STA PACKAGE GRA 


STACIUZATIOM 
LOOP , 

”re¥alancing 1 

ELECTRONICS 


NUMBERS RELATE TO SUBSYSTEM OPERATION 
ISEE PARA 6.2.35.4) 


Figure 5. 2. 3-4. — Altitude reference & pointing control functions. 



b. Gyro reference assembly. 

c. Star tracker assembly. 

d. Processing electronics. 

e. APS. 

5. 2. 3. 5. 1.1 Digital Computer 

The digital computer is part of the CDMS described in paragraph 
5.1.4. The general pupose computer is used to support the ASF 
support subsystems. The com.puter is a major functional element 
of the PCSS providing coordinate transformations, gyro commands, 
gimbal commands, a star catalog, star identification processing, 
Orbiter GN&C inputs, and other related functions required for 
achieving the pointing, control and tracking necessary for the 
payload sensors operations. 

5. 2. 3. 5. 1.2 Gyro Reference Assembly 

The gyro reference assembly consists of three orthogonally 
mounted gyros on the AIM of the APS. The gyros provide stability 
error .signals to the support subsystem computer. Attitude data 
are obtained from readouts at the APS gimbals. 

Available gyro units have a drift rate of 0.01°/hour. The selec- 
tion of the gyro unit will be dependent on bias stability, power, 
weight, cost and gimbal drive rate requirements. Listed below 
are examples of current state-of-the-art gyros in the 0,01° to 
0.1°/hour drift range: 

Honeywell GG 248 Apollo 

A. C. Electronics IRIG 25 Apollo, Skylab, ASTP 

Kearfott 2519 Skylab 


5.2.3-15 



5. 2. 3. 5. 1.3 Star Tracker Assembly 

Star trackers are usually classified as gimbaled or strap-down 
with a variety of detectors ranging from solid state to photomul- 
tiplier tubes. State-of-the-art trackers are available with an 
accuracy of 10 to 30 arc seconds. There are star trackers adver- 
tised with an accuracy potential of 0.5 arc sec., such as the ITT 
or Nortronics trackers, but these are still in the development 
stages. Table 5. 2. 3-5 is a sample of state-of-the-art star track- 
ers that have been developed and have been qualified for the re- 
spective programs. 

Strap-down trackers are less complex from a hardware standpoint 
to implement. They can be mounted on the same platform as the 
GRA and payload instruments. In order to achieve the accuracy 
requirements desired, a narrow FOV is needed, thus requiring a 
lower star threshold and requiring the scanning of the entire 
gimbal system in order to conduct star searches. A minimum of 
two trackers are required in order to determine the attitude 
reference if large Orbiter maneuvering angles are to be avoided. 

The gimbaled tracker adds complexity to its design but can be 
operated independently of the main gimbal system when searching 
for stars. Star threshold levels for gimbaled systems are higher 

a 

because of the large area of celestial sphere available by the 
gimbal system. Usually, one star tracker is needed to determine 
the desired attitude reference. 

Either type of star tracker discussed has the accuracy needed 
for the ASF payload. The selection will be influenced by such 
factors as weight, cost, power, etc. 

5. 2. 3. 5. 1.4 Processing Electronics 

The processing electronics contains amplifiers, analog-to-digital 
(A/D) and di gi tal -to-anal og (D/A) converters, multiplexing equipment. 










U'l ^ 
1 

i-ri -i , 

ga 

h:* t> 

^2 
Ht tsi 

Sss 


TABLE 5. 2, 3-5. - TYPICAL STAR TRACKER PERFORMANCE CHARACTERISTICS 


Performance Data 

Kol 1 sman 
Oao 

ITT Federal 
labs 

Hughes** 

Tracker 

Kol 1 sman* 
KS-199 

Shuttle Tracker 
BBRC 

ATM Star 
Tracker 

Fie1d"Of-view 

r r 



1.2^^ diam 

10^ ^ 10^ 

V X r 

Star magnitude 
sensitivity 

+2 

+3 

+3 

+2.4 

+3 

+3 

Operational accuracy 

15 sec 

5 sec 

30 sec 

10 sec 

60 sec 

22 sec 

Weight - kg b) 

19.5 (43) 

4.3 (9.5) 

13.6 (30) 

13.6 (30) 

7.3 (16) 

32.6 (72) 

Power (watts) 

15 v/atts 

8 watts 

40 watts 

19 watts 

23 watts 

15 watts 

Gimbal freedom 

2 AXIS 

STRAP- DOWN 

2 AXIS 

2 AXIS 

Strap-dovjn 

2 AXIS 

Dimensions - meters (ft) 

i 

0,86-. 30^0. ^0 
(2. 8-1, 0^1. 3) 

0.12*0.27-^0.12 
(0.4*0. 9*0. 4) 

16"xl8">:12” 

0.43*0.30*0.30 
(1.4*1. 0*1.0) 

Unavailable 

0,40x0.28x0.10 
(1.3x0, 9x0, 3) 
(electronics) 
0,45x0.28x0.40 
(1.5x0. 9x1, 3) 
(Mechanical/ 
optics) 


^Developed for JSC under the cognizance of the Guidance and Control Division (EG). Unit has been qua! tested and 
performance verified by testing. 

**0eveloped during Apollo for JSC as a Lunar Optical and Rendezvous System and has been qual tested and performance 
verified by testing. 


s'" 







f 


processing electronics, switching logic, input and output signals, 
processing, routing and gyro rebalancing electronics for the 
operations of the PCSS. Included in the attitude measuring sub- 
system are the status indicator circuits for monitoring key system 
parameters for proper operation during the mission and ground 
self-testing. These include both hardware circuitry and software- 
aided programs. 

The pointing system control electronics as well as the torque 
motors, resolvers, etc., are considered part of the APS. 

5. 2. 3. 5. 1.5 AMPS Pointing System (APS) 

The type of pointing system selected and the system design 
features are discussed in Section 5.2. The general approach 
is to standardize the design for both systems. A particular 
gimbal arrangement or order is not necessary to meet pointing 
requirements. This is usually a function of mechanical 
obstruction in the desired FOV and gimbal range necessary to 
cover all desired targets. Each gimbal axis will have torque 
motors* and an angular readout device such as a resolver for 
torquing the platform to its desired position and providing 
position data. 

On some instruments, stability about the instrument L0*S is 
critical. Therefore, this may dictate the need of a three- 
axis gimbal system. 

5. 2. 3. 5. 2 PCSS Equipment Characteristics 

Excluding the APS and the digital computer which are described 
in other sections of this report, the equipment for the PCSS 
described in this section weighs approximately 100 kg and uses 
210 to 255 watts of electrical power. The estimated size, weight 
and power breakdown by subsystem equipment are shown in table 
5. 2. 3-6. 


5.2.3-18 


TABLE 5.2. 3-6. - SIZE, WEIGHT, AND POWER SUMMARY 
POINTING/CONTROL AND STABILIZATION SUBSYSTEM 


End Item 

Size 

WxHxD-nieters (ft) 

ght-kg (lb) 

Power (watts) 

AMPS Pointing System 

This data is provided in section 5.1.1 



Digital Computer 

The general purpose digital computer 
will be discussed in the data system 
section of this report 



Gyro Reference Assembly 

0.25 X 0.20 X 0.18 (0.8 x 0.7 x 0.6) 

30 (66) 

100 

Signal Processing Electronics 

0.40 X 0.35 X 0.15 (1.3 x 1 .1 x 0.5) 

20 (44) 

40 

Star Tracker Assembly 

FHST-0.60 X 0.21 x 0,21* 
(2.0 X 0.7 X 0.7) 

33 (73) 

75 


GST-0.58 X 0.45 x 0.42** 
(1.9 X 1.5 x 1.4) 

20 (44) 

30 

Sun Sensor 

0.30 X 0.20 X 0.15 (1.0 X 0.7 x 0.5) 

13 (29) 

10 

Optical A1 ignment Measuring 
Device 

0.25 X 0.10 X 0.10 (0.8 X 0.3 x 0.3) 

12 (26) 

30 

Total 


95*/ 108** 
(209/238) 



*GST - Gimbaled Star Tracker 

**FHST - Fixed Head Star Tracker 





























5. 2. 3. 5. 3 Interfaces 


Figure 5. 2. 3-5 describes the interface data flow within the ASF 
payload and to/from the Orbiter. The primary interface areas 
are power, data management, and Orbiter GN&C subsystem. The 
following paragraphs will describe in general the attitude 
pointing and control interfaces. 

a. PCSS-to-Orbi ter . The primary interface with the Orbiter 

is the GN&C subsystem. The Orbiter GN&C will provide to the 
payload initialization data such as vehicle attitude, timing, 
clock synchronization, etc., necessary for monitoring the 
attitude position of the Orbiter, The interface will 
allow for the transfer of pointing vector information to the 
Orbiter GN&C so that the payload attitude can be assessed 
by the Orbiter. Further, the interface requires the trans- 
mittal to the Orbiter GN&C of a pointing vector for reorien- 
ting he payload through use of the Orbiter RCS or vernier 
RCS. The Orbiter will have the capability through the inter- 
face to transmit override commands from the payload control 
panel to disable the payload attitude pointing system if 
requ i red . 

The requirement exists to interface the payload attitude 
sensor (star tracker) through the ASF support subsystem 
computer with the Orbiter GN&C computer so that the basic 
error between the payload attitude measuring system and the 
Orbiter reference system resulting from structural deformation 
can be established during flight. 

b. PCSS-to-Support Subsystems. The pointing and control disci- 
pline involves the management of several pointing systems. 
Figure 5.2. 3-6 shows typical inter-relationships of the various 
elements of the payload system and briefly describes the 
functions of the major interface subassemblies. The PCSS 

must interface with the support subsystem computer. This 


STAR TR^KBR STABLIZED PLATFORM 


i 




Figure 5. 2. 3-5 


Interface and Data Flow Dia 





















IGLOO (COMPUTER) 


/ 


-ET i 

UTY > 

I 

I 



I 

I 


Interface and Data Flow Diagram 


ORBITER 




! 



I 

i 


5 . 2 . 3-21 





J 
















-22 

















relationship and data and command flow is reflected in figure 
5. 2. 3-5. A summary of the command and data requirements is 
reflected in table 5. 2. 3-7. Power interfaces are discussed 
in paragraph 5.2.2. 

5. 2. 3. 5. 4 Operations 

In reference to figure 5. 2. 3-3 and 5. 2. 3-7 a and b, the PCSS 

operations are summarized below. Item letters below correspond to 

the item numbers contained in those two figures. 

a. The PCSS is activated by the PS. The PS has the option of 
operating the APS either automatically through the computer 
or manually through the fine pointing control lever at the 
PSS. 

b. The PS prepares the APS for alignment as follows: 

(1) The PS selects the proper program, mode, etc., and 
commands are transferred to the payload subsystem 
computer . 

(2) PCSS status and data are displayed on the CRT at the 
PSS during the alignment operation. 

c. The data link between the Arbiter GN&C and the payload support 

subsystem computer is activated. The ASF support subsystems 
computer receives the following information from the Orbiter 
GN&C for use in computation: Orbiter position (crosstrack, 

downtrack, altitude), velocity, attitude (3-axis), target 
coordinates, time reference, etc. The ASF support subsystems 
computer will update the Orbiter with the same data as 
required . 

d. Using Orbiter data, the payload is coarse aligned. The com- 
puter sends out commands via path k. The gimbal angles orien- 
tation is controlled by the computer and APS is positioned to an 


5.2.3-23 






TABLE 5. 2. 3-7. - SUMMARY OF POINTING AND CONTROL SUBSYSTEM 
PRELIMINARY DATA REQUIREMENTS 


Signal Name 

Source 

Signal Type 

Sample Rate 
(samples per 
second) 

APS 1 and 2 




APS Temperature 

Gimbal Platform Electronics 
Temp Sensor 

Housekeeping 

1 S/S 

Gimbal Resolver Axis 1 

Gimbal Position 

Output Data Word 

25 S/S 

Gimbal Resolver Axis 2 

‘ Gimbal Position 

Output Data V/ord 

25 S/S 

Gimbal Resol ver. Axis 3 

Gimbal Position 

Output Data Word 

25 S/S 

APS Power on Command 

Keyboard/Display 

Discrete 

1 S/S 

APS Mode Status 

Keyboard/D&C 

Discrete 

1 s/s 

APS Gimbal Slew Axis 1 

Subsystem Computer 

Input Data Word 

25 S/5 

APS Gimbal Slew Axis 2 

Subsystem Computer 

Input Data Vlord 

25 S/S ! 

APS Gimbal Slew 3 

Subsystem Computer 

Input Data Word 

25 S/S 

APS A/D Axis 1 Fan 

Gimbal Platform Electronics 

Housekeeping 

1 S/S 

APS A/D Axis 2 Fail 

Gimbal Platform Electronics 

Housekeeping 

1 S/S 

APS A/D Axis 3 Fail 

Gimbal Platform Electronics 

Housekeeping 

1 S/S 

Coolant Input Temp 

Pallet Sensors 

Housekeeping 

1 s/s 

Coolant Output Temp 

Pallet Sensors 

Housekeeping 

1 s/s 

APS Torquer Current Axis 1 

Gimbal Elect 

Housekeeping 

1 s/s 

APS Torquer Current Axis 2 

Gimbal Elect 

Housekeeping 

1 s/s 

APS Torquer Current Axis 3 

Gimbal Elect 

Housekeeping 

1 s/s 

Gyro Package Temp 

Temp Sensor 

Housekeeping 

1 S/S 

Gyro Electronics Temp 

Temp Sensor 

Housekeeping 

1 s/5 

Gyro Torque Command X 

Subsystem Computer 

Input Data Word 

25 S/S 

Gyro Torque Command Y 

Subsystem Computer 

Input Data Word 

25 S/S 

Gyro To roue Coiranand Z 

Subsystem Computer 

Input Data Word 

25 S/S 

Gyro Torque Rate X 

Subsystem Computer 

Input Data Word 

25 S/5 

Gyro Toroje Rate Y ! 

Subsystem Computer 

Input Data Word 

25 S/5 

Gyro Torque Rate Z 

Subsystem Computer 

Input Data Word 

25 S/S 

Gyro Warm-Up Time 

Keyboard/D&C 

Discrete 

1 S/S 

Gyro Power Present 

Keyboard/D&C 

Discrete 

1 5/S 

X Gyro Fail 

Gyro Package and Electronics 

Output/Housekeeping 

1 S/5 

Y Gyro Fail 

Gyro Package and Electronics 

Output/ Housekeeping 

1 S/S 

Z Gyro Fail 

Gyro Package and Electronics 

Output/Housekeeping 

1 S/S 


Oi- 




^ XT 


5.2.3-2A 



TABLE 5. 2.3-7. - SUMMARY OF POINTING AND CONTROL SUBSYSTEM 
PRELIMINARY DATA REQUIREMENTS - Concluded 


Signal Name 

Source 

Signal Type 

Sample Rate 

Star Tracker Temp 

S*** Temp Sensor 

Housekeeping 

1 S/S 

Star Tracker Elect Temp 

Electronics Temp Sensor 

Housekeeping 

1 S/S 

ST AZ Resolver 

ST Gimbal Position 

Output Data Word 

25 S/S 

ST EL Resolver 

ST Gimbal Position 

Output Data Word 

25 S/S 

ST AZ Slew 

Subsystem Computer 

Input Data Word 

25 5/S 

ST EL Slew 

Subsystem Computer 

Input Data Word 

25 S/S 

ST AZ Torquer Current 

ST Gimbal Electronics 

Housekeeping 

1 S/5 

ST EL Torquer Current 

ST Gimbal Electronics 

Housekeeping 

1 S/S 

ST Search Command 

Subsystem Computer 

Input/Command 

1 S/S 

Star Presence 

ST Selector Output 

Ou tpu t/ Hou se keep i ng 

1 S/S 

Star Tracker Engage 

Star Tracker 

Output/House keeping 

1 S/S 

Star Tracker Pov/er Present 

Keyboard/D&C 

Housekeeping 

1 s/s 

Star Magnitude 

ST Electronics 

Output/Housekeeping 

1 s/5 

ST A/D AZ Channel Fail 

ST Electronics 

Housekeeping 

1 S/S 

ST A/D EL Channel Fail 

ST Electronics 

Housekeeping 

1 s/s 

Bright Source Sensor 

ST Electronics 

Output Data Word 

1 5/s 

Optics Shutter Status 

Star Tracker 

Output Data Word ' 

1 s/s 


Note: Three fixed-head star trackers or one gimbaled star tracker required. Measurement 

list for each tracker is the same» 




a. PLATFORM ERECTION 


Figure 5. 2. 3-7. 


b ) ^1 COMPUTER 





b. ALIGNMENT AND UPDATING 


PCSS operational sequence. 


ifa w' 






! 


i 


i 

accuracy of Certain instruments such as the Optical 

Band Imager and Photometer system require manual fine point- 
ing using the "joystick". 

e. Signals are generated to torque the gimbals to a coarse align 
orientation. 

f. Resolver outputs which are proportional to the gimbal angles 
(position data) are provided to the payload computer (A/D 
conversion ) via k . 

g. Star position data (optical angles) are provided and trans- 
formed by the computer into inertial reference frame for 

positioning gyros. Sighting of at least two stars (non- | 

colinear) are required. These data are provided to computer 
via k. Status data are provided and can be displayed on 
the CRT if requested by the PS. 

•J 

h. The computer transforms optical measurements into inertial 
coordinates and compares desired coordinates with actual 
coordinates. The computer selects gyro(s) to be to'rqued and 
gates the required pulses through the gyro torquing 
electronics. 


p 

-or 



i. Each gyro is positioned as commanded by the computer until 
the PCSS is aligned. 

j. Stabilization loop is established. This loop holds the 
stabilized system inertially referenced as determined by star 
sensor and commanded by computer. Gyros generate error sig- 
nals (i) to indicate any change with respect to inertial 
space resulting in the gimbal torque motors being repositioned 
(e). 

k. PCSS data is routed through the system via: 

1. A/D data provided to computer as status and/or position 
indication. 

2. D/A commands provided by computer to perform required 
functions , 


i 







5.2.3-27 





Data to the CRT at the PSS are provided depicting the health of 
the system. 

5. 2. 3. 5. 4.1 Operational Modes 

The operational modes of the pointing and control subsystem can 
generally be classified into the following categories: initial 

alignment and updates, attitude determination, stabilization, 
and tracking/pointing. Each are described below. 

a. Initial Alignment and Updating. This mode utilizes the star 
tracker to establish the common reference frame for the pay- 
load experiments. In this mode, the gimbals can be slewed 

to zero or the Orbiter GN&C computer can transfer appropriate 
data to the ASF support subsystem computer for aligning 
the gyro system to a coarse reference frame. In order to 
perform the fine alignment, sightings on a minimum of two 
iion-col i near stars are required. The ASF support subsystem 
computer accepts the angular data received from these optical 
measurements along with star catalog data stored in memory 
and transforms it into an inertial reference frame for pre- 
cisely aligning the GRA. The same procedure is repeated to 
update the system to correct errors that are usually accrued 
from gyro drift. 

b. Attitude Determination. Outputs from the three gyros mounted 
on the gimbal system are used to maintain an updated attitude 
reference for the payload sensors and determination of the 
LOS with respect to the inertial reference frame established 
by star tracker sightings. The attitude data defining pay- 
load position is transferred to the Orbiter GN&C computer to 
maintain that the Orbiter spacecraft attitude is properly 
positioned during the payload operation. 

c. Stabilization. In the stabilization mode, the stabilized 
platform inertially referenced is isolated from spacecraft 


5.2.3-28 


motion. The three gyros generate error signals to indicate 
any change in orientation with respect to inertial space and 
these signals are supplied to the gimbal torque motors which 
reposition the APS. 

d. Tracking and Pointing. This mode allows the LOS to be pointed 
to a target and track the target in the presence of Orbiter 
motion. The Orbiter position and target position are trans- 
ferred into inertial coordinates and a command vector is 
determined. This command vector is then transformed into 
payload LOS coordinates and the gimbals are aligned to 
point the sensor{s) LOS to the desired pointing direction, 

5. 2.3. 5.4. 2 Operational Functions 

a. In-Flight Alignment. The in-flight alignment of the PCSS 
requires use of the Orbiter Gfi&C to maneuver the vehicle 
to an attitude where target visibility is obtained and to 
transmit star tracker pointing vectors to the support 
subsystem computer. The computer generates gimbal commands 
to point the star tracker along the star vector. Due to 
the potential misalignments between the star tracker and the 
Orbiter GN&C, it will be necessary to scan the star tracker 
LOS over a predetermined field to insure star acquisition. 

Prior to star acquisition, the gyro reference is initialized 
with an approximation to the desired inertial attitude for 
target tracking and placed in the inertial mode. Using 
this approximate alignment, simultaneous star tracker and 
APS gimbal angle readouts are taken by the subsystem computer 
for two non-colinear stars. 

Using this data the ASF support subsystem computer solves 
for the refined final pointing system gimbal angles and 
establishes a true inertial reference based on the desired 


5.2.3-29 




I 


I 


pointing vector and gyro reference assembly outputs. It is 
possible to update this alignment in a similar manner on a 
periodic basis, or, if desired, near-continuous updating may 
be performed utilizing optimal optional estimation techniques 
and continuous tracking of a single star. The sequence of 
operation discussed above is described in figure 5. 2. 3-8. 

Target Pointing and Tracking. This mode provides the capa- 
bility to track a target in the presence of Orbiter motion 
after initial alignment and/or acquisition has occurred. 

This operation requires that outputs of the GRA and the STA 
be combined by the subsystem computer to form an inertial 
frame in the APS. The present inertial attitude is compared 
to that desired, and appropriate gimbal torque commands are 
generated to position the platform(s) to maintain the desired 
inertial pointing vector. This vector may be fixed with 
respect to the earth; however, its position is. always 
referenced Instantaneously to an inertial frame and appro- 
priate bias rates are introdu'.ed by the ASF support sub- 
systems computer to enable tracking as desired. If it is 
desired to track a non- i nerti al ly fixed target, the support 
subsystems computer must be given the orbit ephemeris 
on a continuous basis to yield the desired tracking accuracy. 
Use of the star tracker in conjunction with the gyro refer- 
ence assembly will allow for periodic updating of gyro drifts 

In order to meet the pointing accuracy requirements utili- 
zing a centralized AMS, it will be necessary to establish and 
monitor relative base motion of the various mounts. This 
may be accomplished by optical transfer techniques. These 
alignment errors will be provided to the support subsystems 
computer to establish the relative location to the various 
mounts. The Laser Sounder (Instrument 213) pointing and 
tracking operation will require near continuous updating 
from the support subsystems computer. The following 


5 . 2 . 3-30 



-31 


BLOCK DIAGRAM 



Figure 5 . 2 . 3-8 . 


Inflight alignment sequence 





sequence of events is envisioned in the performance of 

target pointing and tracking. 

(1) The ASF support subsystems computer generates the 
desired pointing vector. Body fixed axis and inertial 
pointing is desired. 

(2) This vector is transferred to the Orbiter computer. 

(3) The Orbiter computer determines the necessary inputs to 
the Orbiter flight control system. 

(4) The RCS maneuvers the Orbiter to the desired orientation. 

(5) Once there, the Orbiter is placed in attitude hold with 
desired deadband, 

(6) The ASF support subsystems computer transforms the 
desired pointing vector into the payload LOS coordinates 
in terms of gimbal angles. The gimbal errors represent 
the rotation required to position the LOS to the desired 
di recti on . 

(7) The payload gimbals converge to the desired target 
using the payload attitude sensors. 

(8) During tracking, calculations for positioning the gimbals 
must be performed repeatedly to maintain the desired 
pointing direction while both the target and Orbiter 

are moving. 

Figure 5. 2. 3-9 illustrates a block diagram for perform- 
ing the pointing and tracking requirements. Tracking 
aids, such as a TV camera system, could be advan- 
tageous to the PS for those targets that require open 
loop fly-by tracking. This requires that the TV camera 
be boresighted to the flight package and slaved to the 
flight package gimbal electronics. 


5.2.3-32 



tl 





COMMANDS TO FLIGHT 
CONTROL SYSTEM RCS 


Figure 5. 2. 3-9. - Block diagram for pointing and tracking simplified 





c. Solar Pointing and Tracking. For solar pointing and tracking, 
sun sensors can be used to provide error signals to reposition 
the LOS. The AMS can be used along with the subsystems compu- 
ter to generate the desired pointing vector for the solar 
monitoring platform LOS. This can serve as a coarse alignment 
for the solar platform. Once the sun is in the solar tracker 
FOV, the error signal is supplied to the solar monitor gimbal 
system and the gimbals are torqued until the output error 
signal from the solar tracker is nulled. This is illustrated 
in figure 5.2.3-10. 

For those instruments that require scanning the sun disk or 
examining sections of the solar disk other than the center, 
offset signals can be introduced into the control loop. 

Another approach is to use optical wedge offset pointing. 

The fine sun sensor optical wedge subassembly i‘s rotated 
to vary the angle of the incoming sunlight and produce an 
offset of the experiment platform. This technique was 
utilized with a high degree of success during the Skylab 
missi on . 

Sun sensors are available, such as that used on Skylab, that 
have an accuracy capability of approximately four arc seconds. 

5. 2. 3. 5.4.3 Operations Management 

The ASF subsystem computer and the APS with its associated GRA 
and STA form the nucleus of the central reference system. Point- 
ing and control of any subsystem will involve management of the 
subsystem together with the APS. The outputs of the GRA and STA 
are combined by the ASF support subsystems computer to yield 
a constant APS inertial attitude. Gimbal drive commands for 
one or all payload subsystems are generated by the computer as 
required. Gimbal angle readouts together with other tracking 




Figure 5.2.3-10. —Solar monitoring pointing and control. 


5.2.3-35 




i 


I 


sensor outputs are accepted by the computer and used to generate 
a continuous update of pointing vector coordinates. These 
coordinates are compared to the desired coordinates (after 
suitable transformation) in the inertial frame defined by the 
APS and suitable gimbal torque commands are generated to null 
any existing tracking errors. 

Orbiter attitude must also be factored into the pointing and 
control tasks. Initial attitude vjhen alignment of the APS takes 
place v/ill require crev/ coordination. Thereafter, the support 
subsystems computer will monitor attitude through the APS. If 
particular experiments require Orbiter attitude changes for 
tracking or to prevent occultation, appropriate desired payload 
bay pointing vectors must be furnished to the GN&C computer or 
displayed to the crew. 

The pointing and control operations for the ASF payload involve 
two AMPS pointing systems plus a flight package that is hard 
mounted to the pallet. In general, the simultaneous operation 
of all pointing systems is not required; however, there will be 
occasions for the simultaneous operation of these systems. A 
typical example of this is the simultaneous operation of the Laser 
Sounder on module lA and the optical instruments on module IB. 

The experiment requirements will govern the need for multiple 
operati ons . 

The management of the APS will be performed by the support 
subsystems computer and the payload attitude reference system. 

In the previous paragraphs, two subsystem approaches were discussed 
for an AMS, central and distributed. Since the pointing accuracy 
requirements for the APS vary from .01° to 6°, a central attitude 
reference system with the capability of transferring alignment 
data for alignment control could be implemented for the operations 


5.2.3-36 


and management of the APS. Therefore, this approach is suggested 
for the pointing and control aspects of the ASF payload. Since 
the optical instruments module requires the most stringent 
stability and pointing accuracy, the AMS v/ill be placed on the 
pallet with this module. 

Pallets A-1 and A-3 contain the two ASF APS. 

a. Pallet A-1. Module lA on pallet A-1 contains the Laser 

Sounder (Instrument 213) and the Gas Release Module (Instrument 
532). Module IB contains optical instruments for investi- 
gating the atmosphere (Instrument 534), solar monitoring 
instruments (1002 and 1011) and particle beam diagnostic 
(Instrument 550). The pointing and stability requirements 
for the individual instruments within these flight modules 
are shown in table 5. 2. 3-1. 

The operational modes for the laser sounder system are point- 
ing and tracking. The pointing accuracy for this sytem is 
1°. Its reference base will be monitored and provided by 
the AMS located on pallet 3. The tracking mode provides the 
capability of maintaining lock on the target in the presence 
of Orbiter motion. The technique required is desc‘"ibed in 
the previous section. The Laser Sounder pointing control 
v/ill be provided by the computer. Resolvers provide position 
readouts for use by the computer in generating both the 
initial and update pointing commands. 

The minimum pointing and stability requirements for the solar 
flight package are 60 seconds of arc and 0.01“ per second, 
respectively. The solar instruments will be mounted together 
and boresighted to a common LOS. The solar pointing and 
tracking technique were discussed previously. A solar 
tracker is employed to provide the necessary tracking and 
maintaining lock on the solar disk during instrument operations 


5.2.3-37 


Pallet A-3. Pallet A-3 also contains two AMPS instrument 
modules. Module 3A is configured with the Airglow Spectro- 
graph (Instrument 116) and the Limb Scanning IR Radiometer 
(Instrument 118), while module 3B contains the UV-VIS-NIR 
Spectrometer (Instrument 122), the Fabry-Perot Interferometer 
(Instrument 124), and the Infrared Interferometer (Instrument 
126). The pointing and stability requirements for the indi- 
vidual instruments within these flight packages are shov/n 
in table 5. 2. 3-3. 

The primary function of this platform is to point and control 
the orientation of the common LOS of each flight module on 
the pallet. The minimum pointing and stability requirements 
dictated by the instruments on this platform are 0.1 degrees 
and 30 seconds of arc per second, respectively. An additional 
function of this platform is to provide the single ASF | 

attitude reference. Therefore, the present scheme is to 
mount the AMS on this pallet. Alignment data and control 
will be transferred through optical links to the other APS 
on pallet A-1 . To accomplish these two functions, the 
operational modes are: 

(1) initial alignment and update mode. 

(2) attitude determination and control mode. 

(3) target pointing and tracking mode. 

These three modes of operation are discussed in detail in 
previous sections. 

Pallet hard-mounted platform. This flight package consists 
of the particle accelerator instruments. Because of the 
gross pointing and stability requirements (i.e., 2® to 6° 
and l®/sec., respectively), there is no need for a gimbal mount 
These requirements can be achieved utilizing the Orbiter 
GN&C subsystem. Therefore, the Orbiter will be positioned 


5.2.3-38 


£>■ 


in order to point the LOS of this flight package to its 
desired target. 

d. Boom system. Instrument 550 (Faraday cup, Retarding Potential 
Analyzer, Cold Plasma Probe) is mounted on an extendable boom 
which is attached to AIM IB on pallet A-1 . The instrument 
must be extended generally above pallet A-4 during the 
accelerator operations. The instrument must raster scan an 
area covering the accelerator beam width. (See paragraph 
5.2.1). The scanning motion is provided by the APS for 

AIM IB. Instrument 536 (Triaxial Fluxgate) must be extended 
about 20 meters (66 feet) out of the payload bay. No special 
provisions other than holding to mechanical tolerances are 
required to meet the ±0.6° pointing accuracy for this 
i nstrument . 

e. Subsatellite. Some of the ASF experiments require the use 
of the PDS to obtain supporting data. When the subsatellite 
is deployed, the Orbiter is used to point the subsatellite 
in the proper direction for ejection. No other requirement 
has been identified for orientation of the Orbiter relative 

to the subsatellite except during recovery of the subsatellite. 
Subsatellite attitude and rates can be of significant impor- 
tance to ASF experiments and the compatibility of the AE 
satellite (used as baseline for the PDS) control ' system 
should be evaluated in the next study phase. 

5 . 2 . 3 . 6 Analyses 

Pointing error sources discussed in this section fall within two 
categories: (1) errors resulting from structural misalignments, 
and (2) errors related to the attitude pointing and control 
system (APCS). These two types of error directly affect the 
development of pointing techniques and system impel ementation . 


structural errors result from the multitude of structural 
interfaces separating the attitude reference sensors and exper- 
iments, structural assembly errors, thermal deflection, etc. 
Systems errors are a function of the attitude sensors, gyro 
drift, quantization of signals, noise, etc. 

In the candidate PCSS described in the previous paragraphs, the 
m"! sal i gnment between the STA and GRA and the transformation of 
error signals from the gyros through the gimbals are manageable 
by design and calibration techniques. However, the misalign- 
ment between gimbal systems can have a significant impact. This 
can be reduced by the arrangement of the gimbal platforms on the 
pallet. More sophisticated methods such as optical links for 
alignment control {reference figure 5.2.3*-2) or a gyro package 
for each gimbal system may be required to satisfy the pointing 
accuracy for the payload sensors. Until the design approach 
matures sufficiently to perform an error analysis, the pointing 
technique cannot be finalized. The selection of attitude sensors 
as well as the type of subsystem (central vs. distributed) is 
also dependent on the error analysis. 

The numerical values appearing in table 5. 2. 3-8 are typical errors 
of related sensors that were used on Apollo and Skylab programs. 

5.2.3.? Concl usi-^i^is and Recommendations 

5. 2. 3. 7.1 Conclusions 

A significant result of reviewing the Orbiter capability versus 
payload requirements for pointing and stability is that the 
uncertainties or errors in pointing knowledge of the Shuttle 
reference system will exceed the requirements of many payload 
sensors. Since the Orbiter GN&C cannot satisfy all of the ASF 
instrument pointing accuracy and stability requirements, it is 










concluded that one or more gyro stabilized platforms for stability 
and star trackers for pointing accuracy will be required to 
provide the pointing accuracy and stability desired by the pay- 
load instruments. 

A centralized reference system utilizing a gyro reference assembly 
and one or more star trackers can provide a common attitude 
reference frame for all pointing subsystems. However, mounting 
of individual gimbal systems, pallet segment flexures, and 
pallet segment misalignments may result in sufficiently 
large errors that the addition of optical links between tlie ^ 

individual gimbaled platforms and the reference system may be 
required for al ignment control . An alternate concept is to 
provide a separate gyro/star tracker attitude reference unit 
to serve each gimbaled system that requires a high degree of 
accuracy and stability. 

To summarize, the pointing/control and stabilization subsystem 
conclusions are as follows: 

a. The ASF pointing and stability requirements are more exacting 
than that provided by the Orbiter GN&C system; therefore, a 
payload attitude reference sensor and/or system is required. 

b. The error budget for the attitude measuring system demonstrates 

analytically that the ASF requirements can be met with state- 
of-the-art hardware consisting of a precision strap-down t 

gyro-reference assembly and a star tracker to provide 
attitude alignment and update. A solar sun sensor will be 
needed for the solar platform to maintain the stability and 
offset pointing requirements. 

c. Either a gimbal or strap-down star tracker can provide the 
necessary attitude reference for the payload. 

d. A central reference system can provide a common reference 
system for all gimbal systems but may require optical links 
for alignment control. 


5 . 2 . 3-42 



s. For open loop fly-by targets, the use of a TV system for 
monitoring the instrument LOS pointing could be useful. 

f. An interface between the payload star tracker and the Orbiter 
GN&C computer is mandatory so that an inflight calibration 
between the payload AMS and Orbiter reference system can be 
performed to determine the basic error resulting from 
structural deformation. 

5. 2. 3. 7. 2 Recommendations 

PCSS recommendations resulting from the study are as follows. 

a. A detailed error analysis must be completed early in the 
follow-on study so that the pointing techniques and attitude 
sensors selection can be solidified. 

b. Based on the error analysis, the type of subsystem, i.e., 
central or distributed, should be selected during the follow- 
on study. 

c. The sensors for the AMS can be selected after the follow- 
on study is completed. 

d. At the time this study was performed, studies for instru- 
ment pointing systems with an accuracy capability of 1 arc 
second were being conducted. During the fol 1 ow-on _study an 
assessment of these systems should be performed to determine 
applicability to the ASF missions. 


5.2.3-43 


5.2.4 COMMANU AND DATA MANAGEMENT SUBSYSTEM (CDMS) 

5. 2.4.1 Introduction 

The objective of this phase of the study was to determine the 
conceptual feasibility of acquiring, processing, displaying, 
storing and transmitting the scientific and engineering data 
generated by the ASF payload and to define a candidate CDMS. 

Data rate and total data capacity requirements were derived from 
ASF ID's (see appendix B) and a conceptual CDMS was established 
using ESRO/ERNO designed equipment wfiere possible. Boundary 
conditions for data acquisition, processing, storage, and trans- 
mission were established and determined to be within existing 
ERNO equipment and Orbiter facility capabilities. 

Due to the ASF approach of providing complete onboard processing 
capability for scientific data and control of experiments, many 
areas of uncertainties in the data processing area exist. These 
areas have been identified for further study considerations. 

5. 2.4. 2 Requi rements 

I 

The CDMS performs executive functions for the entire payload | 

1 

system including the instruments and the support subsystems. 

The functional requirements for the ASF CDMS are to provide the 
fol lowing. 

a. Data acquisition. 

b. Data monitoring. 

c. Data formatting. 

d. Data processing which includes: 

(1) Instrument/subsystem checkout. 

(2) Sequencing and control of experiments and subsystems. 

(3) Data compression. 


5. 2. 4-1 


! 

(4) Filtering, averaging, hi stogrammi ng . 

(5) Computing. 

(6) Encoding and decoding. 

(7) Data display. 

(8) C&W display. 

(9) Data recording. 

C 

(10) Data transmission. 


5.Z.4.3 Guidelines and Assumptions I 

The CDMS, as defined for ASF, does not deviate from the ESRO base- 
line system. Through the use of the igloo and its command 
and data management components, and the extensive use of RAU's 
for controlling instruments and acquiring data, the CDMS is cap- 
able of performing the total ASF command and data management 
tasks as currently defined. 


Since certain details of the ESRO design are lacking, assumptions 
have been made regarding the CDMS baseline capabilities. These 
assumptions are listed as follows. 

a. The maximum number of RAD ' s per pallet segment is four. It 
is assumed that this figure is representative of each data 
bus; i.e., that four RAU ' s per pallet segment can be used 
for both the experiment bus and the subsystem bus, yielding 
a total of eight RAU ‘ s per pallet. 

b. The serial pulse code modulated (PCM) input to each RAU can 
be used simultaneously with the analog and discrete inputs. 
This is a critical assumption for ASF. 

c. Both experiment and subsystem RAU ' s can be mounted in the aft 
crew station. The ESRO documentation states that RAU ' s are 





5. 2.4-2 


mounted in the manned module for interfacing with the CDMS. 

In the pallet-only mode, these RAU ' s are required in the aft 
crev/ station. 

d. The softv/are resident in the CDMS mass memory may be altered 
during the course of the ASF mission by crew input. Changes 
may be in the form of different data processing routines on 
punched tape. These changes can be read into the mass memory 
as certain experiments are completed and their associated 
processing routines are no longer required. This mode of 
operation is required if the mass memory is unable to house 
in residence all required software for the 7-day mission. 
Software inputs to the mass memory may be uplinked from the 
ground as a secondary mode of operation. However, this will 
be done only if unexpected situations warrant such changes. 

5. 2. 4. 4 Additional General Assumptions 

In addition to the assumptions made based on preliminary ESRO 

descriptions, the following general assumptions are made. 

a. Data processing, to the maximum extent possible, will be per- 
formed onboard. 

b. Ground control over certain aspects of the ASF mission will 
be standard procedure if required. 

c. Adequate space will be available in the aft crew station to 
house two wideband analog recorders and associated electron- 
ics. The tape transports will be accessible in flight for 
tape changes. 

d. The primary communications link for PDS data and control will 
be with the Orbiter, although a direct link v/ith the STDN will 
be available to complement the primary link if required. Data 
from the PDS will be routed to the CDMS via the attached pay- 
load interface. The primary communications link with the SPS 
will be with the ground through TDRSS. These data may be up- 
linked to the Orbiter if required. 


5. 2. 4-3 


f 


t 


e. Communications links between the Orbiter and the ground, 
either through the TDRSS or STDN, and between the Orbiter 
and the PDS can be accomplished simultaneously. 

f. The subsystem computer will have adequate speed and computa- 
tional capacity to control the two APS required for ASF. 

5. 2. 4. 5 Capabilities and Constraints 

The ASF instruments and support subsystems will utilize the Or- 
biter avionics resources through the CDHS. The CDMS will share 
the use of the Orbiter C&W system to process and display safety 
critical data, the Orbiter PMS to process engineering data for 
both statusing and to back up the primary C&W system, the mission 
specialist station (MSS) PCM recorder for data storage, the 
Orbiter mass memory and general purpose computer (GPC) for constants 
and utility storage (for state vector, orbit ephemeris, and 
attitude data determination) and the FM and Ku band signal pro- 
cessors to process scientific data for downlink STDN or TDRSS 
transmission. 

The Orbiter payloads are limited to using lOK words of resident 
GPC memory and 35K words of mass memory storage capability. 

Hardline engineering data transmission is limited to five chan- 
nels and up to 64 kbps data rate and the hardline command rate 
for unmanned payloads is 2 kbps. 

The data rate from deployed payloads to the Orbiter is limited to 
16 kbps and the command rate to payloads is 2 kbps. 

Orbiter capability to handle scientific data is as follows. 

a. MSS PCM recorder - Analog, 2.0 MHz bandwidth 

- Digital, 1.024 Mbps rate 

b. S band FH dov/nlink - Analog, 4.0 MHz bnadwidth 

- Digital, 5.0 Mbps rate 


5. 2. 4-4 




c. Ku band downlink - Analog, 4.2 MHz bandv/idth 

- Digital, 50 Mbps rate 
uplink - 1 Mbps, max. 

The number of C&W annunciators at the forward crew station dedi- 
cated to payloads is limited to two at this time. The status 
panel at the MSS will accept up to five payload C&W parameters. 

5. 2. 4. 6 Subsystem Description 

A functional block diagram of the ASF COMS is shown in figure 
5. 2. 4-1, which depicts the total command and data flow, with 
all instruments, pallets, subsystems, and subsatellites defined. 

The CDMS consists of the following. 

a. Three computers: 

(1) Subsystem computer. 

(2) Experiment computer. 

(3) Backup computer (has capability to replace either exper- 
iment or subsystem computer, but not both simultaneously). 

b. Two I/O units: 

(1) Subsystem. 

(2) Experiment. 

c. Mass memory (shared by both computers). 

d. Keyboard. 

e. Data displays. 

f. C&W electronics. 

g. Wideband analog tape recorders. 

h. RAU's. 

i. CDU. 

j . A&A El ectronics. 


5. 2. 4-5 












































cdto 

#UATltl«l 













f 

b 

V 


The CDMS provides to the ASF payload all services associated with 
the command and control of each instrument, as well as data ac- 
quisition, preprocessing, compression and transmittal of all data 
generated during the conduct of each ASF mission. 

The command subsystem executes all ASF instrument command infor- 
mation in real time either by remote command from the ground, or 
from the Orbiter aft crew station, or by a stored program regu- 
lating the orbit operational schedule. The command subsystem has 
the capability to check validity of each command generated, 
regardless of its source. 

The command subsystem controls the operation of the full ASF in- 
strument payload, which includes the pallet, the APS, the sub- 
satellite, and the boom-mounted instruments. The proper 
sequences of turn-on, warmup, operate, standby, and turn-off for 
individual instruments or groups of instruments, consistent with 
the mission timeline requirements, are controlled through the 
command subsystem. 

Commands are sent to the pallet-mounted or boom-mounted instru- 
ments and igloo mounted subsystems through the 1 Mbps data bus. 
Again, these commands are initiated in real time by the flight 
crew, the ground controllers, or by preprogrammed command 
sequences in response to externally sensed conditions. 

Commands are sent to the subsatellite through the S band Phase 
Modulated (PM) link. These commands may be generated in the same 
manner as those generated for pallet and boom-mounted instruments. 

The data management subsystem provides acquisition capability for 
all data generated by ASF payloads. Data may be acquired from 
subsystems, pallet and boom-mounted instruments, and from 
subsatellites. All date, with the exception of subsatellite 
data, are managed through the 1 Mbps data bus, utilising RAU ' s , 


5. 2. 4-7 


or by wideband analog or TV data lines. Subsatellite data are 
received through the Orbiter S band PM communications system and 
routed to the CDMS as digital and/or analog inputs. 

The data management subsystem, after acquiring data from the var- 
ious data sources, formats the data for compatibility with the 
I/O units, thus allowing the onboard computers to perform pre- 
processing and data compression. This processing is dependent 
on the particular experiment ( s ) being conducted, the mission 
timeline, complexity of processing algorithms required, computer 
availability, etc. The data will be processed to the highest 
degree possible within the constraints imposed. Experiment end 
products are not defined to an extent which will allow details 
of processing to be defined at this time. 

The processed data will be stored on magnetic tape or downlinked 
in either real time or delayed depending on the experiment require- 
ments, detailed elsewhere in this report. 

5 . 2 . 4 . 6 . 1 Command Subsystem 

5. 2. 4. 6. 1.1 Command Generation 

Commands generated on the ground are generally in response to 
evaluations performed on downlinked data. Changes to the resident 
software for data processing may be uplinked through the Orbiter 
communications system. These changes are made and verified on 
the ground, and uplinked only on a programmed basis in accordance 
with the mission timeline. This technique of updating software 
will only be used if required, however, and is further described 
in section 5. 2. 4. 6. 4. The primary mode of software update/change 
will be by crew input. 

Commands generated by the crew are primarily entered through the 
keyboard input. These commands are limited to calling certain 
displays and diagnostic information to the monitors, and to 
initiating sequences for the conduct of certain experiments 


5. 2.4-8 


V 


consistent with the mission profile. The crew, through the key- 
boardi has the capability of overriding, preprogrammed sequences 
and of altering the resident software to a limited degree. The 
keyboard is the primary crew interface with the computers. Other 
commands generated by the crew consist of discrete and potentio- 
metric inputs for selection of operating modes and instrument/ 
subsystem tuning adjustments. The majority of these commands 
are routed to the igloo where they are converted to coded com- 
mands prior to insertion into the data bus. 

In addition to ground generated commands based on evaluation of 
downlinked data, ground controllers have the capability of con- 
trolling certain aspects of the ASF experiment, supplementing 
crew control, as required. Details regarding the crew-ground 
responsibilities are not treated in this report, and are greatly 
dependent on currently undefined aspects of the mission 
objectives . 

The majority of commands for ASF payload and subsystems operation 
are preprogrammed and stored in the CDMS mass memory. These 
software routines consist primarily of sequences of commands 
needed to conduct a specific experiment involving a number of 
instruments, pallets, stable platforms, etc. These routines are 
transferred from the mass memory into the appropriate computer 
by command from the crew or the ground. The computer then con- 
trols the conduct of the experiment or experiments, until 
completion, or until an override command is received. Following 
completion of the particular experiment, or experiments, the 
subject computer is reloaded with the next control program for 
subsequent experiments. 

5. 2. 4. 6. 1.2 Command Transmittal 

As previously stated, all commands, whether generated by the 
flight crew, the ground controller, or the flight computers, must 
interface with the total ASF system at the applicable I/O unit 


5. 2. 4-9 


within the igloo. Those commands affecting the TSMS, the 
pointing and control subsystem, the EPOS or the displays and 
controls (D&C) subsystem are routed to the subsystem I/O unit. 
Commands affecting instrument operation are routed to the exper- 
iment I/O unit. 

Within the applicable T/0 unit, the command, whether discrete or 
analog, is converted to a PCM code compatible with the RAU ' s 
and is routed through the 1 Mbps data bus to the RAU associated 
with the instrument/subsystem being commanded. This RAU converts 
the coded command into either discrete 0 to 5 Vdc outputs or 
serial bi-phase L PCM outputs, and routes the command to the 
instrument/subsystem. A full description of the RAU output capa- 
bilities and characteristics is provided in table 5. 2. 4-1. 

The subsatellite may be controlled by the payload or subsystem 
computer through the Orbiter S band PM link. These commands may 
be generated as preprogrammed sequences by the applicable com- 
puter or may be generated by the flight crew. Ground control of 
the subsatellites is yet to be assessed. In general, however, 
the subsatellite will operate in a continuous mode, and will have 
self-contained control sequences for such control functions as 
spin rate, stabilization, etc. Orbiter supplied commands to the 
subsatellite will primarily consist of initiating the prepro- 
grammed sequences and operational control overrides. 


5.2.4-10 


TABLE 5. 2, 4-1. - REMOTE ACQUISITION UNIT (RAU) 
DATA OUTPUT CHARACTERISTICS 


Discrete Outputs 

Number : 

16 

Type : 

Single-ended, positive with respect to 


0 Vdc RAU common 

Output Logic States : 

"1" -(on) +5 ± 1 .0 Vdc 


"O" -(off) 0 ± 0.5 Vdc 

Output Power: 

10 mA dc minimum at +4 Vdc 

Output Impedance: 

1 k ohms for "0" logic state 


2 k ohms for "1" logic state 


PCM Outputs 

Number : 

8 (data plus clock) 

Type: 

Manchester II bi-phase L code 

Logic States: 

“1 •' -(true) +5 ± 2 Vdc 


"0“ -(false) 0 ± 1 Vdc 







5. 2. 4. 6. 2 Data Management Subsystem 

5. 2. 4. 6. 2,1 Remote Acquisition Units (RAU*s) 

The majority of data generated by ASF experiments is handled 
through the RAU's. A maximum of 32 RAU ' s is available to accom- 
modate payload data. Each RAU has data acquisition character- 
istics as described in table 5. 2. 4-2. 

The maximum of four RAU's can be located on each pallet segment 
for payload data acquisition. Three RAU's can be located vnthin 
the igloo to manage subsystem data. Additional subsystem RAU's 
may be located in the payload bay. 

Each RAU is capable of managing data from 64 analog sources, 
sampling the sources at 1, 10, or 100 times per second, under 
preprogrammed computer control. The analog samples are converted 
to 8-bit digital words and are introduced into the 1 Mbps data 
bus, where they are routed to the applicable I/O unit for addi- 
tional processing, if required. The 64 analog inputs require 
data levels of 0 to 5.12 Vdc and are divided into 32 single- 
ended inputs and 32 differential inputs. 

In addition to the 64 analog inputs, each RAU can accommodate a 
single bi-phase L PCM serial input of up to 1 Mbps (see table 
5, 2. 4-2). This capability exists to accommodate those payloads 
which generate data not compatible with the low sampling rates 
of the analog inputs. 

Eight 8-bit discrete inputs are also available on each RAU to 
accommodate "mode" or "flag" data which is generated by certain 
instruments as housekeeping information. 

The mechanical configuration of a RAU is shown in figure 5. 2. 4-2. 


5.2.4-12 


TABLE 5. 2.4-2. - REMOTE ACQUISITION UNIT 
DATA INPUT CHARACTERISTICS 


Analog Inputs 

Number : 

64 

Type: 

32 0 to 5.12 Vdc single-ended, positive 
with respect to 0 Vdc RAU common 


32 0 to 5.12 Vdc differential 

Resolution: 

8 bits 

Source Impedance: 

100 ohms 

Input Impedance: 

500 k ohms with power on 
100 k ohms with power off 

Sampling Rate: 

Selectable - 1, 10, or 100 samples per 
second 


Discrete Inputs 

Number : 

64 

Logic States: 

"1“ -(true) -5 ± 1.0 Vdc 
"0" -(false) 0 ± 0.5 Vdc 

Type: 

Si ngl e-ended 


Digital PCM Inputs 

Number : 

1 

Source Code: 

Manchester II bi-phase L 

Logic States: 

"1” -(true) +2 to +5 Vdc 
"0" -(false) 0 to +1 Vdc 

Input Data Rate: 

1 Mbps (mean rate of all RAU ‘ s will be 
=: 300 kbps) 


5 .2.4-1 3 







Analog data which have bandwidths in excess of the analog data 
handling capabilities of the RAU are accommodated through the 
high frequency analog data lines. These lines, having one input 
terminal per pallet segment, can manage bandwidths to 6 MHz. 

The high frequency analog data lines are routed directly to the 
aft crew station, where the data are either recorded on the 6 MHz 
analog recorder (paragraph 5. 2. 4. 6. 2. 5) or interfaced with the 
Orbiter avionics via the FM signal processor for downlink transmission 

No significant data processing is performed onboard the vehicle 
with respect to high frequency analog data. 

5. 2. 4. 6. 2. 3 Television Data Lines 

Wideband analog video data are generated by certain of the ASF 
instruments utilizing TV cameras. These data are accommodated 
by the 4.5 MHz TV data lines. One input to this system is located 
on each pallet segment. 

These lines are routed directly to the aft crew station where 
they are interfaced with the Orbiter CCTV system. The video 
data may then be displayed on the TV monitors in real time, or 
may be recorded for later display. 

Downlinking in either real time or in a delayed mode is also 
available depending on downlink availability and the ASF mission 
timel i ne. 

Certain ASF instruments supply their own TV cameras. These in- 
struments v/ill utilize the TV input on their respective pallets. 
Multiple cameras required for the conduct' of certain experiments 
must be sequentially operated with output switching, as there is 
only one TV input available per pallet. 


5.2.4-15 




other ASF instruments do not supply TV cameras but require TV 
coverage of the phenomenon being observed. The Orbiter payload 
bay camera may be used in these cases, and the resulting video 
information displayed, recorded, and/or downlinked as required. 

5. 2. 4. 6. 2. 4 Data Processing 

Onboard data processing to the maximum degree possible consistent 
with ASF mission timeline and economic constraints is a primary 
goal of the ASF pallet-only mode CDMS. 

General purpose processing is supplied for experiment checkout, 
sequencing and control, data compression, data reduction, etc. 
Processing is accomplished in the igloo through the use of the 
experiment computer and the subsystem computer. The character- 
istics of these computers are shown in table 5. 2. 4-3. 

The basic software for execution and management of data process- 
ing is resident in the mass memory and is supplied by Spacelab. 
Application software for individual experiments is also resident 
in the mass memory, being supplied by the investigator. 

The purpose of onboard data processing is to deliver to the 
ground data dissemination center a product which can be rapidly 
reformatted into computer-compatible tapes and forwarded to each 
investigator for detailed analysis. The reformatting procedure 
would not include any manipulation, merging, curve-fitting or 
algorithm applications, as these functions would have been per- 
formed on the vehicle prior to delivery of the data to the 
dissemination center. 

While the detailed processing software will not be defined for 
some time due to the many external variables currently existent 
in the ASF mission profile, a conceptual description of the 
processing sequence is provided in paragraph 5. 2. 4. 6. 4. 


5.2.4-16 



TABLE 5. 2. 4-3 


COMPUTER CHARACTERISTICS 


Fonoats 




Operands: 6, 16» 32 and 24 ^0 (Floating points) bits 

Floating Point 32 Bits (24 f B) 


Instructions; 16 bits 

Add/Sub Direct 5 u sec 


Control Unit 


Indirect € u sec 


Hlcro-progr4En!ied control unit 

Hul/Div Direct 6 u sec 


Cycle time 30D NS 


Indirect 7 u see 


Micro-Interrupt capability 

Input/Cutput 


Micro-Instructions 4 K >fljrds of 16 or 20 bits 


Interrupts 


Instruction Set 


- Musrbcr of external 

B Levels 

■ Nuntber of InstrLCticns 123 


iluaber of internal 

5 Level? 

■ Fommt 16 bits 


- Nurber of software 

Program dependent 

Icmediote 3 bits 


Interrupt control 

Hicropragran + Software 

Direct 256 Bytes 


- Priority scheduler 

Software 

Indirect nectary double ivord 

• 

Data transfer rode 


Relative 512 bytes 


- Program controlled 


Based 2S6 bytes 


data rate 

60 yS/word 

Indexed 64 1C bytes 


no of addressable peri feral? 66 K 

■ Type 


- Direct fflcrory access 


Call and store 


data rate 

aOO'to "50 K word/sBC 

Logic and corTtparfson operations 


control 

direct 

Shift operations 

• 

Word length 

16 bits plus 1 parity 
+1 protection 

Fixed-to-floating and floating -to- fixed 
conversions 

« 

Discretes 

G Inputs and G outputs 

Conditional and unconditional Jur^ps 

■ 

Real tire work 

1 mS tc 2^^ HS 

Addressing Modes 

Meirory 



Itnedfate, direct* Indirect, 

• 

Type: la niil ferrits cores Z 1/2 D, configuration 

relati *; to a base. Indexed, relative 

to a prograni counter, half word, 1 

word, character, double word | 

t 

Capacity: 64 K 16-blt words (plus 1 parity bit and 1 protection 

bit) extendible to 512 K 16-bit words 

■ Addressing capability 

■ 

Modularity; 16 K words 


Byte* word, double word 

■ 

Cycle tine: 920 NS 


lluriber of addressable Registers 

t 

Appressing* 

Ouantun: Byte, word 


4 Speclallicd registers 

m 

Access tSir.e: 42D NS 


62 Oedlcaied registers 

« 

Ports: 2 


7 Base registers 





CDJspiitlng Speed 
Fixed Point 16 Bits 


Add/Sub Direct Z p set 

Indirect 3 u sec 

Mul/DIv Direct 4 p sec 

Indirect S a sec 

Fixed Point 32 Bits 


Add/Sub 

Direct 

5.5 

V sec 


Indirect 

6.6 

M sec 

Hul/Riy 

Direct 

3.3 

u sec 

’I 

Indirect 

9.3 

u sec 


* 


ORIGINAL PAGE IS 
OR POOR QUALITY 


5.2.4-17 



5. 2.4. 6. 2. 5 Data Recording 


There are two payload dedicated analog magnetic tape recorders 
in the ASF CDMS located in the aft crew station. Specifications 
for these recorders are shown in table 5. 2.4-4. 

The primary function of these recorders is to store high rate 
data during periods when the Orbiter data downlink capability 
is not available, or to store the data for delivery to the ground 
data dissemination center following the conclusion of the 
mission. The capability to store onboard the vehicle all PCM 
data from a 7-day mission exists. The capability to downlink 
data either in real time or in a delayed mode remains and can be 
accomplished as a complementary action. 

The recorders are designed to act as permanent or short term 
storage devices, or to be bypassed entirely if real time data 
downlink is available and the need to store data onboard the 
vehicle does not exist. 

These recorders must be located in the aft crew station (or other 
crew accessible areas). Tape changes will be required daily in 
order to provide a recording medium for raw and/or processed data 
for the 7-day mission. 

The MSS recorder is also available as a short term storage device 
for processed payload data. Its operation and interface with the 
data downlink system is very similar to that of the high rate 
recorders. 

5 . 2 . 4 . 6 . 2 . 6 Data Downlink 

The ASF payload is visualized as being totally dependent on the 
Orbiter communications system for transmittal of data to the 
ground either directly to the STDN or via the TDRSS. Details of 
the Orbiter communication system utilized by the ASF payload are 
shown in figure 5. 2. 4-1. 


TABLE 5. 2. 4-4. 


TAPE RECORDER CHARACTERISTICS 


Capaci ty ; 


Data tracks; 


Minimum bit rate; 


Maximum bit rate: 


Record/playback ratios 


Power: 


Weight; 


Dimensions : 


Height: 


Width: 


Depth : 


Vol ume : 


2.4 X 10^ Bits 


5.25 X 10’^ bps 


5.90 X 10® bps 


160:1 to 1:160 


15 - 30 Watts 


13.6 kg (30 lbs) 


Transport 


12.7 CM (5") 


El ectronics 


15.2 CM 16") 


33.0 CM (13") 33.0 CM (13") 


33.0 CM (13") 15.2 CM (6") 


0.045 M^ 0.025 M^ 

(0.489 cu. ft.) (0.271 cu. ft.) 


Mounting technique will include provision for stacking 
Recorders must be modified for reel change capability. 


5.2.4-19 























5. 2. 4. 6. 3 CDMS Equipment Characteristics 

Some of the characteristics of Interest for the CDMS equipment 
are listed in table 5. 2.4-5. 

5. 2. 4. 6. 4 Interfaces 

The CDMS comprises the system through which all experiments are 
commanded, controlled, and through which data are acquired. 

These interfaces are shown in figure 5. 2. 4-1. 

The CDMS interfaces with the pallet segments, where a maximum of 
four RAU ' s per data bus per pallet are located. These RAU's 
distribute command and control functions to various instruments 
and subsystems located on each pallet segment. They acquire data 
generated by these instruments, format the data, and route it to 
processing equipment located in the igloo. 

The heart of the CDMS is the igloo, where the experiment I/O unit 
and computer, the subsystem I/O unit and computer, the backup 
computer, and RAU's for monitoring subsystem performance are 
1 ocated . 

The CDMS interfaces with the subsatellite through the pallet 
A-2 RAU's prior to subsatellite deployment and via the Orbiter 
communications system after deployment. 

All ASF subsystems interface with the CDMS through subsystem 
RAU's. The subsystems include the TSMS, the PCSS, the EPOS and 
the D&C subsystem. 


5.2.4-20 




TABLE 5. 2. 4-4 


CDHS EQUIPMENT CHARACTERISTICS 






Power * 

Weight * | 

COMS Hardware 

Dimensions cm (in*)* 

Location 

Ave. 

Peak 

Suppl ier 

kg 

lUs) 

Experiment RAD 1 

23.1 X 12.2 X 8.9 
(9.1 X 4.8 X 3.5) 

Aft Crew Sta. 

30 

60 

E5R0 

2.7 

(6) 

Experiment RAU Z 

APS 1-A 

23.1 X 12,2 X B.9 
{9.1 X 4.0 X 3.5) 

Pallet A-1 

30 

60 

ESRO 

2.7 

(6) 

Experiment RAU 3 

23.1 X 12.2 X 8.9 
(9.1 X 4.8 X 3.5) 

Pallet A-1 

30 

60 

ESRO 

2.7 

(6) 

Experiment RAU 4 

APS 1-B 

23.1 X 12,2 X 8.9 
(9.1 X 4.8 X 3.5) 

Pallet A-1 

30 

60 

ESRO 

2.7 

(6) 

Experiment RAU 5 

23.1 X 12.2 X S.9 
(9.1 X 4.8 X 3,5) 

Pallet A-1 

30 

60 

ESRO 

2.7 

(6) 

Experiment RAU 6 

23.1 X 12.2 X B.9 
(9.1 X 4.8 X 3,5) 

Pallet A-2 

30 

60 

rcpn 

2.7 

(6) 

Experiment RAU 7 - APS III-A 

23.1 X 1Z.2 X 8.9 
(9.1 X 4,8 .-t 3.5) 

Pallet A-3 

30 

60 

ESRO 

2.7 

(6) 

Experiment RAU 8 

23.1 X 12.2 X 8.9 
(9.1 X 4,8 X 3.5) 

Pallet A-3 

30 

60 

ESRO 

2,7 

(6) 

Experiment RAU 9 - APS IH-B 

23.; X 12.2 X 8.9 
(9.1 X 4.8 X 3,5) 

Pallet A-3 

30 

60 

ESRO 

2.7 

(6) 

Experiment RAU 10 

23,1 X 12.2 X 8,9 
(9.1 X 4.0 X 3.5) 

Pallet A-3 

30 

60 

ESRO 

2,7 

(6J 

Experiment RAU 11 

23.1 X 12,2 X 8.9 
(9.1 X 4.8 X 3.5] 

Pallet A-4 

30 

60 

ESRO 

2.7 

(6) 

Experiment RAU 12 

23.1 X 12.2 X 8.9 
(9.1 X 4.8 X 3.5) 

Pallet A-4 

30 

60 

ESRO 

2.7 

(6) 

Experiment RAU 13 - 20m Boom 

23,1 X 12.2 X 8.9 
(9.1 X 4,0 X 3.5) • 

Pallet A-4 

30 

60 

ESRO 

2,7 

(6) 

Experiment RAU 14 

23.1 X 12.2 X 8,9 
(9.1 X 4.8 X 3,5J 

Pallet A-4 

30 

60 

ESRO 

2.7 

(6) 

Subsystem RAU 1 

23.1 X 12.2 X 8.9 
(9.1 X 4.8 X 3,5) 

Aft Crew Sta. 

30 

60 

ESRO 

2.7 

(6J 

Subsystem RAU 2 

23,1 X 12.2 X 0.9 
(9,1 X 4.8 X 3,5) 

Aft Crew Sta, 

30 

60 

ESRO 

2.7 

(6) 

Subsystem RAU 3 

23,1 X 12.2 X 8,9 
(S.l X 4.8 X 3.5) 

Igloo 

30 

60 

E'^RO* 

2.7 

(6) 

Subsystem RAU 4 

23.1 X 12,2 X 8.9 
(9.1 X 4,8 X 3.5) 

Igloo 

30 

60 

ESRO 

2.7 

(6) 

Subsystem RAU 5 

23.1 X 12.2 X 8,9 
(9,1 X 4.8 X 3.5) 

Igloo 

30 

60 

ESRO 

2,7 

(6) 

Subsystem RAU 6 

23.1 X 12,2 X 0.9 
(9.1 X 4.8 X 3.5) 

Pallet A-1 

30 

60 

ESRO 

2.7 

(6) 

Subsystem RAU 7 

23.1 X 12.2 X 8,9 
(9.1 X 4,8 X 3.5) 

Pallet A-1 

30 

60 

ESRO 

2,7 

(6) 

Subsystem RAU 8 

23,1 X 12.2 X 8,9 
(9.1 X 4.8 X 3,5) 

Pallet A-1 

30 

60 

ESRO 

2,7 

(6) 

Subsystem RAU 9 

23.1 X 12,2 X 8.9 
(9.1 X 4,8 X 3.5) 

Pallet A-1 

30 

60 

ESRO 

2,7 

16) 

Subsystem RAU 10 

23.1 X 12.2 X 8.9 
(9.1 X 4,8 X 3.5) 

Pallet A-2 

30 

60 

ESRO 

2,7 

(6) 

Subsystem RAU 11 

23,1 X 12.2 X 8.9 
(9.1 X 4,8 X 3.5) 

Pallet A-2 

30 

60 

ESRO 

2,7 

(6) 

Subsystem RAU 12 

23.1 X 12.2 X 8,9 
(9,1 X 4.8 X 3,5 

Pallet A-3 

30 

60 

ESRO 

2.7 

(6) 

Subsystem RAU 13 

23.1 X 12,2 X 8.9 
(9.1 X 4.0 X 3.5) 

Pallet A-3 

3D 

60 

ESRO 

2.7 

(6) 

Subsystem RAU 14 

23.1 X 12,2 X 8.9 
(9.1 X 4.8 X 3.5) 

Pallet A-4 

30 

60 

! ESRO 

2.7 

(6) 

Subsystem RAU 15 

23.1 X 12,2 X 3.9 
(9.1 X 4.8 X 3.5) 

Pallet A-4 

30 

60 

bSRO 

2.7 

(6) 

Computer, Experiment 

19.6 X 25.9 X 49.7 
(7.7 X 10.2 X 19.6) 

Igloo 

245 

350 

ESRO 

31.8 

(70) 

Computer, Subsystem 

19.6 X 25.9 X 49.7 
(7.7 X 10.2 X 19.6) 

Igloo 

245 

350 

ESRO 

31.8 

(70) 

Computer, Backup 

19.6 X 25.9 X 49,7 
(7.7 X 10,2 X 19.6) 

Igloo 

35 

350 

ESRO 

31.8 

(70) 


^Estimated 




5.2.4-21 



























































The CDMS interface with the Orbiter aft crew station is depicted 
in the functional block diagram, figure 5. 2. 4-1 . The diagram 
provides details regarding interfaces between various components 
comprising the CDMS and major D&C's required for their proper 
functioning. Many details regarding specific operational condi- 
tions are yet to be assessed, but should be compatible with 
this functional concept, 

The aft crew station also provides basic interfaces with the 
Orbiter communications sytem. These interfaces are depicted on 
the block diagram as being a part of the aft crew station. This 
is functionally correct; although the actual location of the 
Orbiter communications components may be elsewhere in the 
Orbiter. 

The primary CDMS components located in the aft crew station are 
subsystem dedicated D&C's, experiment dedicated D&C's, and the 
computer keyboard and CRT. 

All subsystem dedicated D&C's are routed to and from the igloo 
mounted subsystem I/O unit through subsystem RAU ' s located in 
the aft crew station. These RAU ' s convert all subsystem command 
and control fu ctions generated by the crew into a format compati 
ble with the subsystem data bus. The data bus then routes the 
command functions to the applicable subsystem where, through 
another RAU, the command is decoded and routed as a discrete 
word or as a PCM word to the subsystem. The D&C's required for 
proper control and monitoring of each subsystem are described 
in subsequent paragraphs. The display of required subsystem 
parameters is accomplished through monitor devices driven from 
the subsystem computer through the same RAU which receives the 
control inputs. 


5.2.4-23 




All experiment dedicated D&C's are accommodated in the identical 
manner described for subsystems. The experiment data bus, I/O 
unit, computer, and RAU's are used, however, in lieu of subsystem 
components . 

The computer keyboard is used for generating all instructions 
which do not require the use of manually-operated switches and 
tuning adjustments. The keyboard addresses both the subsystem 
and experiment RAU's located in the aft crew station. Thus, 
through the keyboard, both instruments and subsystems can be 
controlled. All of the various computer controlled experiment 
sequences which are initiated by the crew are entered through the 
keyboard. The data processing programs are transferred from m.ass 
memory into the experiment computer after keyboard instruction. 
Similarly, all computer driven displays are called through the 
keyboard . 

The aft crew station mounted subsystem and experiment RAU's inter- 
face with the keyboard and the computer driven display by the 
O&C unit - a device which has preprogrammed control recognition 
logic and display logic in residence. This unit is very similar 
to the Orbiter display electronics unit (DEU) and serves the same 
purpose . 

As stated previously, the aft crew station provides basic inter- 
faces between the CDMS and the Orbiter communications system. 

These interfaces are detailed in the functional block diagram, 
figure 5. 2. 4-1, which illustrates the various components within 
the Orbiter avionics system. A discussion of the total Orbiter 
communications system is not included in this report, but may be 
found in the reference documents. 


5. 2. 4. 6. 4. 2 


In the ASF pallet-only mode, the igloo houses the major portion 
of the CDMS. A description of the igloo is found in paragraph 
5.2.1 of this report; hence, this section is limited to the CDMS 
functions of the igloo. 

The igloo houses the mass memory, the subsystem and experiment com- 
puters plus a backup computer, two I/O units, subsystem RAU ' s , the 
A&A electronics, and the C&W electronics. The igloo is also used to 
route wideband analog and payload dedicated TV lines to the aft crew 
station. 

The mass memory houses the executive and application software 
for both the control of experiments and data acquisition and 
processing. The applicable software routines are transferred 
to the subsystem and experiment computers through their respective 
I/O units. The I/O units serve as the interface point between 
the 1 Mbps data busses and the computers. 

The subsystem RAU's located in the igloo are used to monitor 
housekeeping parameters from the EPOS power supplies and the 
active thermal control loop. These subsystems are also con- 
trolled through the subsystem RAU's. 

The A&A and C&W electronics units monitor both the subsystem and 
experiment I/O units for potential crew hazards and out-of-tol erance 
conditions existing within the total ASF system. Conditions 
recognized as being outside nominal tolerances are conditioned 
and forwarded to the aft crew station C&W and A&A panels. 

5. 2.4. 6.4. 3 Pallet Segments 

The CDMS has identical interfaces with each of the four pallets. 

Data busses for both experiments and subsystems are routed from the 
igloo-mounted I/O units to RAU's located in each pallet. The 
RAU locations within the individual pallets are flexible to 


5.2.4-25 



accommodate the various instruments located in each segment. 

Those pallets having instruments mounted to the pallet structure 
itself will have hard-mounted RAU's in close proximity to the 
instruments. Those pallets having pointing systems will have 
RAU's mounted on the platforms. 

Each pallet has inputs available for one wideband analog line 
(paragraph 5. 2. 4. 6. 2. 2) and one TV line (paragraph 5 . 2 . 4. 6 . 2 . 3 ) . 

5. 2. 4. 6. 4. 4 Pallet-Mounted Instruments 

The CDMS interfaces with the pallet-mounted instruments through 
the RAU's for control functions, and one or more of the three 
data acquisition components described in paragraph 5. 2.4. 6. 2. 

5. 2. 4. 6. 4. 5 Subsatel 1 i te 

The CDMS has a dual interface with the subsatellite mounted on 
Pallet A-2. For checkout of the various subsatellite sensors 
and subsystems prior to deployment, a RAU hard-mounted to Pallet 
A-2 is used. After deployment, the detached payload S band link 
is used. 

Subsatellite data composition is not yet defined, but will be 
generated to fit within the data bandwidth limitations imposed 
by the communications link available. 

5. 2. 4. 6. 4. 6 ASF Support Subsystems 

a. TSMS. The CDMS will control and monitor the TSMS through the 
subsystem computer, I/O unit, and RAU's. 

Since the ASF pallet-only mode has no life support require- 
ments, the function of the thermal subsystem is reduced to 
instrument cooling or heating as required. 


5.2.4-26 


The Freon loop is controlled via the subsystem RAU located 
in the igloo. The primary and backup Freon pumps will be 
turned on and off as required and the interloop heat ex- 
changer inlet and outlet temperatures will be monitored. 

The pump status will be continuously monitored, as will the 
inlet and outlet temperatures of the igloo heat exchanger. 

The inlet and outlet temperature of the cold plates will be 
monitored by pallet-mounted RAU's. 

The temperature and pressure of the internal environment of 
the igloo will be monitored, as will the status of the GNg 
f ans . 

Thermal conditions in the payload bay will be monitored as 
required. 

b. PCSS. The PCSS consists of three hardware groups with their 
associated electronics. These groups are: 

(1) Star tracker assembly. 

(2) Gyro reference assembly. 

(3) Pointing systems. 

Each of the platforms communicates with the CDMS subsystem 
computer through a dedicated RAU- In addition, the CDMS 
subsystem computer communicates directly with the Orbiter 
GN&C computer. 

The APS provides instrument positioning and tracking capabil- 
ities which exceed those of the Orbiter vehicle. The opera- 
tional modes of the subsystem are: (1) initial alignment 

and updates, (2) attitude determination, {3) stabilization, 
and (4) tracking. 


5.2.4-27 


During the initial alignment, a minimum of two non-colinear 
star sightings are made by the star tracker. This angular 
data is transferred to the subsystem computer, which has 
access to the star catalog stored in the Orbiter mass memory. 
The angular information is transformed to an inertial refer- 
ence frame and the resulting data is used to align the gyro 
reference assembly. The procedure is repeated for periodic 
updates . 

The outputs from the gyro are sent to the subsystem through 
the GRA RAU, v/here a determination of the LOS with respect 
to the inertial reference frame established by star tracker 
sightings is made. The data are then transferred to the 
Orbiter GN&C computer for Orbiter positioning. 

Stabilization is maintained through the monitoring of error 
signals by the subsystem computer. The error signals are 
generated by the GRA. Signals are then sent to the gimbal 
torque motors to reposition the instruments. 

During periods when the Orbiter is changing attitude but the 
APS's must remain in stable pointing modes, the Orbiter posi- 
tion and the target position are both sent to the subsystem 
computer. All information is transformed into inertial 
coordinates and command vectors are calculated. These data 
are transformed to target LOS coordinates, which are sent to 
the gimbal torque motors to align the APS. 

Because of the complexity of the CDMS/APS interfaces, the 
large interchange of information between the ASF subsystem 
computer and the Orbiter GN&C computer, and the operational 
computations required for proper APS operation, the practi- 
cality of the subsystem computer accommodating the total load 
is questionable. It is impossible to determine this factor 


5 . 2 . 4-28 


without performing a software analysis effort. Details 
are too preliminary at this time to initiate such an effort. 

An extensive follow-on to this report is needed to evaluate 
this situation. 

Should the computer be inadequate for the task, the use of 
APS dedicated microprocessors, which will perform operational 
calculations dedicated to each platform, will be investi- 
gated. These devices would receive Orbiter GN&C data through 
the subsystem data bus and perform the required computations. 

This technique would free the subsystem computer of the addi- I 

tional computation load. 

EPOS. The CDMS provides the basic controls over the switch- 
ing and monitoring of electrical power throughout the aft 
crew station, the igloo, and the payload bay. 

Monitoring of both voltage and load is accomplished through 
subsystem RAU's located at the power input point to the 
igloo, at the power converters, and throughout the payload 
bay at power distribution points. 

Control of power distribution is accomplished through the 
same subsystem RAU's. Remote circuit breakers and switches 
respond to commands generated by the crew or the computer and 
distributed by the subsystem data bus. 

Those instruments requiring capacitor banks for operation 
are automatically monitored by the subsystem computer to 
insure adequate capacitor charge prior to activating the 
discharge sequence. 

D&C subsystem. The CDMS interfaces with the D&C subsystem 
are described earlier in this section. A complete descrip- 
tion of the D&C subsystem is found in paragraph 5.2.5. 


5.2.4-29 


t 




( 


5 . 2 . 4 . 6 . 4 . 7 Instrument Interface Listin g 

The control, data and display interface requirements imposed on 
the pallet RAU's are listed in table 5. 2. 4-6 {for Pallet A-1 ) 
and table 5. 2.4-7 (for pallets A-3 and A-4). 

5. 2. 4. 6. 5 Operati ons 

The operation of the ASF CDMS involves continuous support of 
payload engineering (status, etc.) functions and on-demand 
support of the scientific instruments. 

As previously discussed, not all instruments operate simultane- 
ously. Instruments are divided into groupings of those which 
operate together. The controlling information and data handling 
requirements are preprogrammed to manage the operation of these 
sensor groupings. 

During the first day of the mission, following successful estab- 
lishment of the orbit, the instrument checkout sequence is per- 
formed. This command or series of commands may be generated by 
the crew or by ground controllers. During this operation, the 
executive routine and application routine for one group of instru- 
ments are transferred from the mass memory to the experiment and 
subsystem computers. The validity of the program is then verified. 
Following the validity check, each instrument is powered and the 
operational parameters are limit-checked to verify "in tolerance" 
conditions. 

Following this validation of software transfer and the checkout 
of all instruments and subsystems required to support the parti- 
cular sensor grouping, the entire sequence is repeated for the 
remaining sensor groups. This includes a verification of sub- 
satellite systems prior to deployment. The subsatellite is then 
deployed and stabilized. 




TABLE 5. 2. 4-6. - INSTRUMENT/CDMS INTERFACE LISTING 

(PALLET A-1 ) 


Instrument 



CDMS Interfaces 

Pallet A-1 

- 



213 

Control : 

RAU 2 

- Discrete Outputs 1 through 9 


Di spl ay : 

RAU 2 

- Discrete Input 1 

- Analog Input 1 

- Hardwire Displays 1 and 2 


Data : 

RAU 2 

- . Digi tal PCM Input 

- Analog Inputs 2 and 3 

- Discrete Inputs 2, 3, and 4 

532 

Control : 

RAU 3 

- Discrete Outputs 1 through 7 

- PCM Outputs 1, 2, and 3 


Data : 

RAU 3 

- Discrete Input 1 

- Digital PCM Input 

- Analog Inputs 1 through 24 

534 

Control : 

RAU 4 

- PCM Output 1 


Displ ay : 

RAU 4 

- Discrete Input 1 


Data ; 

RAU 4 
4 MHz 

- Digital PCM Input 

- Discrete Inputs 2 through 7 
TV Video Input 

1002 

Control : 

RAU 5 

- Discrete Outputs 1, 2, 3, and 4 


Data : 

RAU 5 

- Discrete Inputs 1, 2 , and 3 

- Analog Inputs 1, 2, 3, and 4 

1011 

Control : 

RAU 5 

- Discrete Outputs 5 through 10 

- PCM Output 1 


Data : 

RAU 5 

- Discrete Inputs 4 through 8 

- Analog Inputs 5 through 8 

550 

Control : 

Hardwire Controls 1 through 5 


Display: 

Hardwire Displays 3 through 7 
RAU 5 - Analog Inputs 9 through 28 


Data : 

RAU 5 

- Digital PCM Input 

- Analog Inputs 29 through 43 


5.2.4-31 










Instrument 


Pallet A-3 


116 


Pallet A-4 
303 


CDMS Interfaces 

Control 

RAU 

7 


Discrete Outputs 1, 2, 3, 

and 4 

Di spl ay 

RAU 

7 

- 

Discrete Input 1 


Data 

RAU 

7 

- 

Discrete Input 2 






Analog Input 1 


Control 

RAU 

7 

- 

Discrete Outputs 5 through 

10 




- 

PCM Output 1 


Data 

RAU 

7 

- 

Digital PCM Input 






Analog Inputs 2 through 6 


Control 

RAU 

8 

- 

Discrete Outputs 1 through 

9 




- 

PCM Output 1 


Data 

RAU 

8 

- 

Digital PCM Input 





- 

Discrete Input 1 





- 

Analog Inputs 1 and 2 


Control 

RAU 

9 

_ 

Discrete Outputs 1 through 

9 

Data 

RAU 

9 

- 

Digital PCM Input 





- 

Discrete Input 1 





- 

Analog Inputs 1, 2, 3, and 

4 

Control 

RAU 

10 

_ 

Discrete Outputs 1 through 

10 

Display 

RAU 

10 

- 

Analog Input 1 


Data 

RAU 

10 

- 

Digital PCM Input 





- 

Analog Inputs 2, 3, and 4 






Discrete Input 1 


Control : 

Hardwi re 

Controls 1 through 10 



RAU 

11 

- 

Discrete Outputs 1, 2, and 

3 

Display: 

Hardwi re 

Displays 1 through 5 


Da ta : 

RAU 

11 

- 

Analog Inputs 1 through 20 





“ 

Digital PCM Input 


Control ; 

Hardwire 

Controls 11 through 14 



RAU 

12 

- 

Discrete Output 1 


Display: 

Hardwire 

Displays 6 through 10 



RAU 

12 

- 

Analog Inputs 1 through 10 


Data : 

RAU 

12 

- 

Digital PCM Input 





** 

Analog Inputs 11 through 25 

Con trol : 

RAU 

13 

_ 

Discrete Output 1 





- 

PCM Output 1 


Data ; 

RAU 

13 

- 

Digital PCM Input 






Analog Inputs 1, 2, 3, and 

4 

Control ; 

RAU 

14 


Discrete Outputs 1 through 

5 

Data : 

RAU 

14 


Analog Inputs 1 and 2 



5 . 2 . 4-32 








stabilization of the subsatellite is automatic, activated by 
command from the flight crew following deployment. Deployment is 
accomplished by command the flight crew through the RAU 

located on the subsatel . i pallet (Pallet A-2). This command 
actuates the hold down c. vice, allowing the subsatellite to be 
ejected from the payload bay. Communication with the subsatellite 
is accomplished via the Orbiter S band PM communications link. 

The subsatellite range is monitored through the Orbiter tracking 
system until operational range (1 to 10 km) is obtained. By 
crew command, the spin stabilization sequence is initiated. All 
further stabilization adjustments to the subsatel 1 i *..e are auto- 
matic, being sensed and controlled by the internal stabilization 
system of the subsatellite. Crew override capability exists 
for repositioning, if required, and rendezvous following the 
completion of all experiments. 

The data from the subsatellite is complementary to that generated 
by the pallet-mounted instruments; i.e., the subsatellite dat' 
is used in real time or stored onboard the Orbiter for delayed 
use in processing and analyzing data from pallet-mounted instru- 
ments. The processing routines for this operation are entirely 
dependent on experiment definition, and will be addressed when 
mission requirements are finalized. The capability to transfer 
subsatellite data to the CDMS via the attached payload interface 
needs to be further assessed. 

Prior to the start of the first data acquisition sequence, the sub- 
system and experiment command and processing routines are transferred 
from the mass memory into the experiment and subsystem computers. 
Following validation of the software transfer, the preoperate 
sequence will be manually initiated by the crew. This operation 
will command all required subsystems and instruments to sequence 
through their preoperate modes; i.e,, to warm up, cool down, 
orient platforms, etc. Each stage of this operation is monitored 


5.2.4-33 


automatically for out-of-tol erance conditions, with the status 
being presented to the crew, if required, of each instrument or 
subsystem. 

Any instrument or subsystem which does not come within opera- 
tional tolerances within a preprogrammed time interval following 
initiation of the warm up sequence will be flagged to the crew 
automatically through the split-screen CRT. Diagnostic infor- 
mation on that instrument or subsystem may then be called up to 
determine the validity of the out-of -to! erance flag. Should the 
flag be false, the crew simply proceeds with the data acquisition, 
as the flag is only a monitor having no control over instruments or 
subsystem operation. Should the flag be true, i.e., the instru- 
ment or subsystem is malfunctioning, the capability exists to 
instruct the malfunctioning system to power down and to proceed 
with the experiment. This procedure, however, is entirely 
dependent on the system in question and the mission requirements. 

Any instrument or subsystem which demonstrates a malfunction 
which may affect crew or vehicle safety is immediately presented 
to the crew through the C&W system described in the preceding 
section . 

Assuming all systems accomplish the preoperate sequence success- 
fully, the total CDMS is then ready to initiate the operate 
sequence. This sequence is initiated by the experiment computer in 
response to a preprogrammed time label, which is recognized by 
the computer as it monitors the Greenwich Mean Time (GMT) and 
Mission Elapsed Time (MET) clocks. The sequence may also be 
initiated by the crew or by the ground, however, if required. 

The entire operation; i.e., subsystem operation, including point- 
ing and control, stabilization, thermal control, instrument 
operation, data acquisition, processing, and recording are con- 
ducted under the control of the CDMS. The operational sequence 


is ended, with all systems being commanded to "standby" or "off" 
by the CDMS automatically, provided the preprogrammed "stop time 
label exists. Again, the operate sequence may be terminated by 
crew or ground input. 

Certain instruments cannot be operated in a totally automatic 
mode, but require the man-in-the-1 oop capability to "fine tune" 
the instrument as the experiments progress. In these experi- 
ments, the command and processing routines transferred from mass 
memory to the computer include the automatic routing of observed 
phenomena to the various display devices required by the crew. 
The crew inputs are then routed to the instrument through the 
1 Mbps data bus, and the required adjustment is made. 


Following completion of one data acquisition sequence, the CDMS 
awaits the initiation of a second sequence by the crew. This is 
accomplished in the same manner as was the first sequence. 


The CDMS is able to conduct a number of experiments simultane- 
ously, being limited only by the speed and software capabilities 
of the computers. These limits are T3D and are dependent on the 
definition of experiment control and processing requirements, 
and on the definition of the mission profile. 

Data processing is accomplished in near real time. The limita- 
tion here is the complexity of the processing required. Where 
data rates exceed the capability of the computer for real time 
processing, the data will be stored on magnetic tape and played 
into the computer for processing during mission phases where the 
full capability of the computer is not required. 

As stated in paragraph 5.2.4. 3. d, if the mass memory is unable 
to house in residence all software required for a 7-day mission. 


5.2.4-35 


! 

1 



i 


updating of the mass memory during the course of the mission 
will be required. Thus, the mass memory becomes a system of 
temporary storage, and the flexibility required to assure on- 
board processing to the greatest degree possible is not re- 
stricted by the mass memory size. The data processing routines 
will be developed and verified on the ground prior to the start 
of the mission, and will be written on a punched tape format, 
taken on the flight, and loaded into the mass memory as required. 

5 . 2 . 4 . 7 Analyses 

Based on the overall ASF system timelines, data timelines for 
the instruments and the PDS were defined. Figure 5. 2. 4-3 shows 
the periods during the mission for revolutions 15 through 80 
for each instrument requiring CDMS support. Based on these time- 
lines and the data requirements defined in the ID's (in appendix B), 
data fait rate requirements during each minute of experiment oper- 
ations from revolution 16 through revolution 80 were established. 
These levels are illustrated in figure 5. 2. 4-4 through 5.2.4-12. 
These figures show that the maximum data rate for the ASF com- 
plement of instruments required is 123.192 kbps during revolu- 
tions 17 through 25 when Instrument 532 (Gas Release Module) 
with its requirement for over 77 kbps operates for 15 minutes 
each revolution. If these data were to be transmitted directly 
to the TORS or to the ground station, the data rate is well within 
the capability of the Orbiter S band FM downlink system and the 
Ku band TORS system. 

The total quantity of data required by the ASF instruments for 

the 7-day mission was determined by integrating the data rates 

over the total experimental operating period. The result is 

shown in figure 5.2.4-13, A total of 15.931 x 10^ bits of data 

will be processed by the onboard data management subsystem, 
g 

At 2.4 X 10 bits per reel, it would require seven reels to accommo- 
date onboard storage. The extent of data compression which will 


5.2.4-36 


-37 











DATA. KBPS 




DATA. KBPS 




DATA, KBPS 





130 r 








RE 

\ 

.VOLUTI 

ON 36 - 37 

120 

110 











100 

go 

8Q 

2 70 

§ 

CJl < 

• 

ro 

40 

30 

2D 

10 

0 











28.( 

>04 


45.692 



28.604 


37.260 

121,380 

! 

37.26 

0 



28.604 

















n 
























1 








10 20 30 40 I 50 60 70 80 90 


-« D AY NIGHT ► 


Figure 5. 2. 4-8. -ASF data rate requirements revolution 36-37. 


MINUTES 






















be performed, the real time versus stored data processing and 
transmission planned, and other related operations need to be 
further assessed to determine the full impact on onboard proc- 
essing and storage capability and the ground facilities. 

The 4 MHz TV requirement (Instrument 534) is not included in 
this analysis. 

This analysis does not address the recent developments regarding 
Instrument 126, the Infrared Interferometer. The original data 
rate was calculated at 1.2 kbps, and this figure was used in all 
analyses regarding data management. Recent developments place 
the data rate as high as 6 Mbps. Accommodation of this data rate 
with the currently defined CDMS can easily be made, although the 
RAU's and I/O units would be bypassed. This technique, while an 
inherent feature in the CDMS, is not discussed as part of this 
report. 

5. 2.4.8 Conclusions and Recommendations 
5. 2. 4.8.1 Conclusions 

The CDMS appears to be one of the most significant drivers in 
a Spacelab Manned Module versus pallet-only ASF approach. The 
capability to control a full complement of instruments by remote 
means and the requirement to perform a high degree of data 
processing onboard the vehicle present unique problems over the 
manned module approach. 

The study results indicate that, functionally, the ASF approach 
for onboard processing of scientific data is feasible with 
still-to-be-resolved issues being the computer executive, memory 
and throughput capacity and processing speeds, the application 
software sizing, and the possible need for advanced data compres- 
sion techniques. 


5.2.4-48 



5. 2.4.8. 2 Recommendations 

The study has resulted in the following recommendations. 

a. Establish firm data processing requirements. 

b. Define computer hardware and software required for the ex- 
tensive onboard processing required. 

Onboard data processing has been defined as one of the major 
goals of the ASF pallet-only mode. The degree of processing to 
be performed is dependent on many factors, most of which have not 
been adequately defined. Primary among these is the mission 
operating timeline. Since the requirement exists to operate 
one or more instruments continuously throughout the mission, the 
experiment and subsystems computers will be occupied in controlling 
instrument operation and will possibly be unavailable for data 
processing. The use of the backup computer as a data processing 
tool during these periods, i.e., when the primary computers are 
unavailable, should be investigated. This would relieve the prob- 
lem of having one computer perform both controlling and processing 
functions, and would assist greatly in the pallet-only mode ^ al 
of onboard processing to the maximum extent possible. 

Complete de^^i ni ti ons of experiments will not be available for 
some time, making sizing of software for applications routines 
undefinable at present. Processing routines cannot be defined 
until experiment end products are defined, making a determination 
of the adequacy of onboard computer capabilities for the total 
task impossible. The use of the backup computer and remotely 
located microprocessors must be considered. 

The ability of the subsystem computer to control and supply the 
required computational support for two APS's as well as control 
and monitor the electrical and thermal subsystems, cannot be 


5.2.4-49 


determined without performing a software design effort. There- 
fore, the use of APS dedicated microprocessors, capable of per- 
forming the high speed calculations and reference transformations 
needed for instrument pointing, thus reducing the computational 
load on the subsystem computer, is being explored. 


5 . 2 . 4-50 



5.2.5 AFT CREW STATION CONFIGURATION 
5. 2. 5,1 Introduction 


The primary objective of thi" phase of the study was to determine 
the feasibility of operating the ASF instruments from the Orbiter 
aft crew station. The purpose was to establish a baseline for 
future ASF and ArtPS pallet-only mode analyses. 

Requirements were established using inputs from the scientific 
id's and the support subsystem conceptual definitions. Guidelines 
and assumptions were established and Orbiter and other constraints 
were established. A full scale mockup of the Orbiter aft flight 
deck, constructed for this study, was configured to accommodate 
standup operation by a crewman in a zero-g erect position using 
only foot restraints. 

An operational D&C philosophy was developed from the results of 
a preliminary analysis which considered the following. 

a. ASF D&C requirements. 

b. Payload dedicated D&C area in aft crew station. 

c. PS workload for operating ASF instruments. 

d. One or two-man operating capability at the PSS. 

e. Support hardware (D&C units, recorders, etc.) stowage require- 
ments in the aft crew station area. 

f. The number of PS's required to perform an ASF mission. 

g. Support requirements for Orbiter supplied D&C (RMS, R&D 
CCTV, etc.) 

Study results indicate that some manual control is required but 
that most operations can be automatically controlled through the 
ASF payload CDMS. Sufficient space is available yjt the PSS console 
to accommodate ASF instrument D&C requirements. Stowage for support 

r 


5. 2. 5-1 


ASF hardware is required in the aft crew station. No attempt was 
made to assign a location to this hardware since Orbiter has not 
yet assigned dedicated aft crew station stowage volume to payloads 

It was determined that only one PS at a time can operate at the 
PSS. For 24 hour/day operations of the ASF, two PS’s will be 
required, each operating on a 12 hour/day shift. 

5, 2. 5. 2 Requirements 

5.2. 5.2.1 Instruments 

Instrument requirements were defined using the latest available 
information. These requirements are outlined in table 5. 2. 5-1. 

The table is organized fay instrument packaging per pallet and 
associated parameters. The parameters cover data collection, 
preparation time, controls, displays, forced A&A, C&W and other 
parameters such as provisions for data filming . 

5. 2. 5. 2. 2 Support Subsystems 

Functional D&C requirements for the support subsystems are minimal 
These are the following. 

a. Equipment power control. 

b. Boom extension and retraction, 

c. Platform latching and unlatching. 

d. Tape recorder control. 

e. Fine pointing control. 

5. 2. 5. 2. 3 Orbiter Support 

a. MSS Panels R-12 and R-T3 (figure 5. 2. 5-1 ) are dedicated to 
the mission station. Included in this area are the D&C for 


TABLE 5. 2. 5-1. - AMPS/ASF INSTRUMENT DISPLAY AND CONT 



V 

DATA COLLECTIflIt 







SEE ASF 

MISSION TIMELII 

ES (riOUBE 4,1 

.5-1) 

1 

FREPARATIOH TlHE 







SEE AS 

F Mission TIMEL 

!NES (FIGURE 4. 

K5-U 


CQIITROLS 

POWER 

(OWpOFF.STBY) 
POINTING 
TUNE A 
PULSE RATE 

DOORS 

(OPEN, CLOSE) 

POWER 

GAS BOTTLES 
(4) 

GAS PRESSURE 
(PSD 

GAS HDDE (Z 
srAi^E, SOX) 

GRATING 
(5 DIGITS) 
DDDRS 

HASS FILTER 
(5 DIGITS) 

PLftTFORfI 

POWER 
POINTING 
FILTER SELECT 
HOOE APERTURE 

2 CAMERAS 
(CCTV, UV) 
START* STOP 
RECORD 

DOORS 

n JEHISON REQ 

POWER 

DOOM- EXTEND, 
retract INCRE- 
MENTS 

JETTISON 

SCAN 

ODORS 

\nm\m - to q 

POWER 

DATA SEQUEHCE 

INITIATION 

CALIBRATION 

OOORS . 

'REITER 

POWER 

pointing 

EXPOSURE 

CONTROL 

DOORS 

POWER 
UNLATCH, 
EJECT. STOW, 
LATCH 

RCS CONTROL 
COVERS 

POWER 

CQNFIGURATIDH 

SELECT 

(SPECTRAL 

RANGES) 

POINTING 

EXPOSURE 

DOORS 

POWER 

POINTING-SCAN 

RATE 

FILTERS SCAN 
NODE 

DOORS 

POWER 

HDDE spectral 
(4 BANDS) 

SCAN RATE 
(GRATING) 

DDDRS 

PLATFORM S3 jq 

POWE 

ETAL 

PQIM 

FILT 

WAVE 

i 

DOOli 

j 

TIS( 

DISPLAYS 

WAVE LENGTH 
(VARIANCE) 

PULSE RATE 

RELATIVE 
POINTING ANGLE 

BEAH REFLECT 
MEDIA RANGE 

STATUS 

BOTTLES (4) 
GAS MODE (2) 

CHAMBER 

TEMPERATURE 

GAS PRESSURE 

CHAMBER 

PHOTODIODE 

(VOLTAGE) 

HAVELENGTH-X 

INTEHSITY-Y 

STATUS 

HASS COUNT-X 
IKTEHSITY-Y 

INDICATOR . 
LIGHTS 

TV HDHETDRS 
(2) 

VIDEO TAPE 
STATUS 

VARIABLE 
SAMPLING 
RATES (3) 

STrtTUS 

PULSE SHAPES 
[NEAR REALV 

\iint } 

STATUS 

■ ' ! 

■ 

INDICATOR 

LIGHTS 

STATUS 

SPECTRAL 

range scan 

film COUNTER 

STAR, SUN 

TRACKER 

INDICATORS 

STATUS 

TRACES 

EXPOSURE 

SPECTRAL 

RANGES 

TRACKER, . 
ACqUISITlOH 

RANGE , 

RANGE RATE 
INDICATORS 

ORIENTATION 
RUNNING LIGHTS 
ST- BE 
A . .IStTtOH 
LIGHT Dll, OFF 

STATUS 

INDICATORS 
FUH COUNTER 

RELATIVE 
POINTING ANGLE 

SPECTRDGRAH 
EXPOSURES 
START- STOP 

SCOPE TRACES 
(2 RANGES) 

STATUS 

TV HQMITQR 

SCOPE TRACES 
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VIEW LOOKING AFT 



Figure 5. 2. 5-1. - 


Orbiter aft flight deck panel location code. 












general payload subsystem operations. Panel R-11 will 
contain all unique D&C required by the MS to support ASF 
payload subsystem operation. The functions identified at 
this time are: 

(1) Subsatellite monitoring and control. 

(2) Experiment computer D&C. 

b. At the on-orbit station, Panel A2 is the R&D panel. The ASF 
support requirements are for retrieval of the subsatellite 
and include the following: 

(1) Range/range rate indicators. 

(2) Control and monitoring of subsatellite, if it plays 
active role during retrieval. 

Panel A2 is the CCTV panel. The ASF support requirements 
will include switching capability to allov/ the scientific TV 
cameras to utilize the onboard TV monitors. 

Panel A4 is the RMS panel. The ASF support requirements 
will include: 

(1) An indicator for stowed/unstowed status of the sub- 
satellite. 

(2) A control to activate the latch/unlatch mechanism. 

(3) An indicator for latch/unlatch status. 

Additional panel assignments at the on-orbit station are 
required to support ASF for the following* 

(1) APS jettison control. 

(2) Boom(s) jettison control. 

(3) Subsatellite jettison control. 

These will be controlled by guarded switches and will be 
required if the APS, booms, or subsatellite tiedown latching 
mechanism should fail. 

5, 2. 5-5 


5.2. 5. 3 Guidelines and Assumptions 

a. The ASF experiments will operate 24 hours/day. 

b. The inflight calibrations for the ASF instruments will be 
automatic. 

c. The Orbiter-provided RMS, CCTV, and R&D capabilities will 
be required. 

d. An AMPS/ASF dedicated computer will be provided. 

e. EVA will not be considered as a normal operational require- 
ment. 

f. The ASF will be considered the prime payload and the 
dedicated payload D&C area, lower console volume and allocated 
payload stowage space will be dedicated to it. 

g. There will be no audible alarms required for the operation 
of the ASF. An exception will be if the simultaneous crew 
sleep constraint is enforced. 

5 . 2 . 5 . 4 Capabilities and Constraints 

The ASF utilizes Orbiter facilities in accomplishing the required 
operations. The resources include space allocation at the PSS 
for p&C (figure 5. 2. 5-2) support equipment installation and stowage, 
and control, display, checkout, C&W, and other functions at the 
MSS and the on-orbit station. These stations are illustrated in 
figure 5. 2. 5-1. Figure 5.2. 5-T also shows the panel location 
codes which are as follows: 

A2 . R&D Panel 

A3 _ CCTV Panel 

A4 RMS Panel 

Rll , R12 & R13 MSS Panels 

LIO, 111 & L12 PSS Panels 


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PAYLOAD CONTROLS AND DISPLAYS 
ORBITER SUBSYSTEM CONTROLS AND DISPLAYS 


Figure 5. 2. 5-2. — Aft flight deck crew station 

















Rll at the MSS is dedicated to payload use. Following are the 

descriptions of on-orbit station, MSS, and the PSS. 

a. On-Orbit Station. This station is located at the aft flight 
deck wall and contains D&C for monitoring target range, 
range rate, LOS rates, and monitoring Orbiter attitude and 
attitude rate. It also contains a translation controller 
and rotation controller for controlling Orbiter maneuvers. 

The capability is provided for docking mechanisms control, 
lighting control, docking module system controls, and 
communications. A crewman optical alignment system (COAS) 
is located at the overhead window for visual alignment 

of target vehicles on-orbit. Two CCTV's are provided for 
monitoring manipulator operations and experiments. Addi- 
tionally, the on-orbit station contains D&C for manipulator 
operations, D&C for CCTV camera control (pan, tilt, zoom, 
focus), TV monitor switching, payload bay lighting, and 
payload bay door controls. Appropriate controls will be 
provided for operating with one or two manipulator arms. 

When required, controls will also be provided for manipulator 
jettisoning, payload latchl ng/unlatching, and payload/Orbiter 
umbilical connect and disconnect. 

b. Mission Specialist Station. The MSS on the starboard side 
contains D&C for checkout, monitor and control of the Orbiter/ 
payload subsystems interface. Command, control, and moni- 
toring, via rf, of deployed and detached payload support 
systems are also provided. A CRT display and keyboard panel 
are used to interface with the Orbiter systems management 
computer for onboard checkout and fault isolation of Orbiter 
subsystem malfunctions. A communications panel is provided 
for management of voice, TV and telemetry (TM) data to and 
from the ground. Onboard recording of Orbiter and payload 
subsystem data is provided via the PCM recorder. A standard 
48.26 cm (19-inch) v/ide by 53.34 cm (21-inch) high panel is 
provided for accommodating payload unique D&C. Mission and 


5. 2. 5-8 


event timers and lighting controls will also be located at this 
station. During launch and entry, some payload experiment and 
payload subsystem C&W and other monitor and control capability 
is provided on a panel located on the starboard side and facing 
aft. 

c. Payload Specialist Station. The PSS on the port side contains 
three standard panel spaces with required Orbi ter-to-payl oad 
standardized electrical power connectors for accommodating 
government-furnished equipment (GFE) and/or user provided 
unique modules for command, control , and checkout of experi- 
ment instruments. A junction box will be provided for routing 
wire from connectors located in the payload bay to the equip- 
ment modules at the aft crew station. Standard Orbiter audio 
panel and lighting controls will be provided at this station. 

5. 2. 5. 5 Aft Crew Station Configuration Description 

The ASF D&C consists of the following: 

a. An alpha-numeric keyboard (ANK)(part of CDMS). 

b. Two CRT’s. 

c. Two event timers. 

d. A fine pointing control panel. 

e. An Alarm/Advisory (A&A) display. 

f. An analog tape recorder control and status panel. 

g. A platform latch control and status panel. 

h. A boom control and status panel . 

i. Film status displays. 

j. Power control panels. 

k. A "scratch pad". 


5. 2. 5-9 


These D&C are located at the PSS on panels UlO, Lll and L12 
(see figure 5. 2. 5-1). Panels LIO and Lll comprise the "facility” 
D&C which are expected to support not only ASF but the full 
complement of AMPS missions. The D&C items located on LIO and 
Lll are shown in figures 5.2. 5-3 and 5.2. 5-4. The location of 
these D&C's is preliminary and will be updated as D&C requirements 
are better defined. Panel L12 (figure 5.2. 5-5) is dedicated to 
discrete instrument D&C's. These are not high usage items and 
are primarily status displays such as film status* etc. Other 
discretes would be time critical items such as manual power 
cutoff to instruments in case of critical malfunction requiring 
computer override. These items will change with the various 
ASF/AMPS mission requirements. The ASF discrete D&C items 
identified to date are: 

(1) Film status for instruments. 

(2) Computer override power cutoff for all instruments. 

(3) Latch/unl etch controls and status for pallets A-1 and A-3 
APS‘s. 

(4) Unlatch/1 atchj deploy/fetract controls and status for booms. 
5. 2. 5. 5.1 Cathode Ray Tube 

The CRT selected for ASF missions is the unit being developed 
for the Orbiter Multifunction CRT Display System (MCDS). The 
overall dimension is 26.06 cm x 18.75 cm (10.26" x 7.38") with a 
screen size of 17.8 cm x iz.7 cm (7" x 5"). The display has a 
split screen format capability. Character sizes can be either 
0.318 cm (0.125") or 0.381 cm (0.150") in height. The 0.318 cm 
(0.125") size allows 26 lines of 51 characters-per-line formats. 
The 0.381 cm (0.150") size allows 25 lines of 41 characters-per- 
line formats. Figures 5. 2. 5-6 and 5. 2. 5-7 show the Orbiter CRT 
selected for the ASF program. 



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Figure 5.2. 5-6.— CRT graphics display format 










Figure 5.2. 5-6 illustrates a sample split screen graphical display 
and figure 5. 2. 5-7 shows an alpha-numeric (A/N) page format. The 
examp.le shown in figure 5. 2. 5-7 illustrates a format for A&A 
information. The top two lines should be reserved for header 
information- The format for these two lines v/ill be the same 
for all pages. It includes the date, page heading, page number, 
orbit revolution number and the mission time. 

The alarm (third line) would be uni.que to this series of displays 
and the particular area would be brightened or flashing. For 
example, Instrument 532 on the alarm line would be flashing and 
the top line of 532 would be brightened across the entire CRT 
screen. Lines 4 through 23 (line 24 if the 26-line format is used) 
will display instrument data. The bottom two lines are reserved 
for keyboard communication and scratch pad use. 

5-2. 5.5.2 Alpha-Numeric Keyboard 

A typev/riter type of ANK was selected for the ASF program. 

Figure 5. 2. 5-8 shows the keyboard panel containing pushbutton 
indicators and the ANK subassembly. It is a self-contained unit- 
included in the back of the panel are control logic, gating shift 
register, and drivers. It is a standard 17.78 cm x 48.26 cm 
(7" X 19") panel, 33.02 cm (13") in depth excluding connectors 
with an internal power module. The power requirement for the ANK 
is 115 Vac, 60 Hz, single phase power. Lamp power (24 Vdc) is 
supplied from an external source. 

The numerics have been removed from the top key line of the 
typewriter format and grouped at the side for input convenience. 
Additional symbols have been added to the top keys. An alternate 
configuration would be to leave the numeric keys in the standard 
typev/riter location and reduce or move the symbols to other key 
locations. This decision will depend upon final instrument 
requirements. 


Figure 5.2,5 


“8. - Alphanumeric keyboard panel . 
























5 - 2 . 5 . 5 . 3 Fine Pointing Control 



The ASF mission requires the capability for fine pointing 
control of the APS and modules. There are tv^o directions of 
movements; (1) up-down, and (2) right-left. The tv/o APS 
can be rotated right-left in the "platform" mode. The two 
instrument modules on each platform can be individually rotated 
up-down and right-left in the "AIM" mode. The control is a 
two-axis displacement "joystick" with provision for spring 
return to center in each axis. Third axis control is not 
required at this time. 

The type of display to be provided to the crew in conjunction with 
the manual fine pointing control is subject to more detail study 
and will be considered in the follow-on AMPS study. 

5. 2. 5. 5. 4 Alpha-Numeric Dot Matrix Displays (time, film status) 
(TBD). 

5. 2. 5. 5. 5 Discrete Switches 

The type of switch selected for the ASF panels is the rack-mounted, 
plug-in, lighted pushbutton indicator (PBI) switch with full and 
split legend displays. Each contains four lamps with varied 
color capability. The lamps or legends can be removed from the 
panel front without tools. The majority of these PBI*s are 
alternate action switches and are 2.54 cm (1") high, 2.54 cm (1") 
wide, and 7.62 cm (3") deep. 

Using only one type of switch is more cost effective than mixing 
with toggles, thumbwheels, etc. Larger lots and simpler inventory 
requirements will result in low costs. Reasons for selecting this 
particular type of switch are as follows. 

a. Flexibility and growth is excellent. Functions can be changed 
by removing and replacing legends and making the desired 
connecti ons . 


5.2.5-18 



b. Visibility is excellent with a minimum scan time required 
for status information. 



5. 2. 5. 5. 6 Scratch Pad 

A scratch pad will be provided for recording remarks and notes. 
The design shown is conceptual. 

5. 2. 5. 5. 7 Support Hardware 


Support hardware for ASF requiring aft crew station lower console 
volume have been identified. No allocation of lov/er console 
volume to payload use has been made at this time. Therefore, the 
support hardware is identified as to requirements only, with no 
specific location assigned to them. ASF equipment located within 
the crew station envelope to support the ASF includes: 

a. One control and display unit (CDU). 

b. Three RAU’s. 

c. Two analog tape recorders. 

These items are part of the CDMS. 

The CDU is the interface between the keyboard and CRT’s. It 
generates the necessary symbology and data formats. The CDU 
will service the keyboard and two CRT's. The CDU will be located 
in tre PSS lower console a minimum distance from the keyboard 
CRT's. The unit is approximately 27.94 cm (11") in width, 

19.55 cm (7-1/2") in height and 22.99 cm (8.1/2") in depth. It 
weighs 18.2 Kg (40 Tbs), and requires 201 watts maximum power. 

The analog tape recorders are used primarily to store high frequency 
scientific data. The stored data can be transmitted to the ground 
station through the Orbiter rf links. Two analog recorders will 
be required to operate in series. To allow uninterrupted data 
recording they will require crew access for changing tapes. It is 
recommended that a crew-accessible location be allocated v/ith a 


5.2.5-19 


pull-out drawer in the lower portion of one of the aft crew 
station consoles. These recorders are each approximately 33.02 cm 
(13") wide by 33.02 cm (13") deep by 15.24 cm (6") high. 

Three RAU's will be required to interface pallet-located equipment 
with the D&C and Orbiter equipment. The MSS will require an 
experiment RAU and the PSS will require one experiment RAU and one 
subsystem RAU. RAU size is approximately 22.61 cm (8,9") in width, 
11.94 cm (4.7") in depth, and 8.64 cm (3.4") in height. 

5. 2.5. 5.8 Loose Equipment Stowage 

The loose equipment requiring stowage volume are: 

a. Tape recorder tapes (number and dimensions TBD), 

b. Backup keyboard. 

5. 2. 5. 5. 9 Interfaces 

The ASF D&C has three primary interfaces. These are with; 

(1) the payload, (2) the Orbiter equipment, and (3) the Orbiter 
crew. 

a. Payload. Interfaces with the ASF payload (instruments, 
subsatellite, and support systems) are made through the ASF 
CDMS RAU. These interface functions are discussed in 
section 5.2.4. 

b. Orbiter Equipment. Orbiter equipment which interface with 
the ASF D&C include: 

(1) C&W electronics and displays. 

(£'■ MSS. 

(3) On-orbit station. 

c. R and D Station. To retrieve the subsatellite, D&C capa- 
bilities at the R&D station are required. These are: 

(1) Orbi ter/subsatell i te relative range and range rate monitor. 


(2) Subsatellite attitude and control and monitoring (if 
subsatellite plays an active role during retrieval). 

d. RMS Station. The control and monitoring functions at the 
RMS station required during subsatellite retrieval are: 

(1) Indication to the operator that the subsatellite is in 
the stowed position. 

(2) Control to activate the latch/unlatch mechanism. 

(3) Indicator to show latch/unlatch status. 

In addition to the C&W status board, the MSS will provide 
control and monitor capability for the payload subsystem, 
payload computer, and the subsatellite. 

e. CCTV. The CCTV display will be used by the ASF system. 
Instrument 534 contains two cameras, one for low light tele- 
vision (LLTV) and the other for UV display. A capability 

to select either or both of these cameras simul-taneously 
for display on the CCTV is required. 

f. On-Orbit Station. In the event the pointing systems or the 
boom cannot be retracted into their stowed position in 
preparing for Orbiter return to earth, the following capa- 
bilities must be provided at the on-orbit station; 

(1) Jettison APS 1 and 3. 

(2) Jettison boom (Instrument 536). 

(3) Jettison boom (Instrument 550). 

5.2.5.5.10 Operations 

Utilization of keyboards, predefined CRT formats, operator inter- 
action, and computer/software interfaces must have an operating 
philosophy as a foundation for decisions. The following paragraphs 
discuss such a philosophy. 


Display design should be oriented toward ANK manipulation on a 
CRT. The CRT would present status, alarm message, attention 
coding, etc. Display request would be accomplished via ANK. 
Checklists and procedures will be incorporated into the ANK/CRT 
system (TBD). The following general characteristics would apply 
to most displ ays. 

a. Function Codes. The ANK function codes would be graphic 
descriptor or mnemonic labels, displayed on the CRT and linked 
through a software program or routine. These function codes 
would be displayed only if applicable to the particular format 
and picture being presented. An individual function code 
would occupy a common location on all displays; however, it 
may not appear on all displays, i.e., it would be suppressed 
when not applicable to a particular display. These function 
codes would be designated and activated by underlining with 
the ANK cursor either of the first two characters of the 
respective function code. The normal sequence of use would 
require the operator to designate an entity - such as a data 
item - and the operation to be performed on it, such as ERASE, 
ENTER, REVERSE, PAGE UP, PAGE DOWN, etc., and enter the request. 
The result would be an observable change in display content 

or format. 

b. Data Fields. Specified data fields would be intentionally 
protected and not subject to ANK manipulation. The data on 
some displays should be made impervious to change through the 
ANK to ensure the integrity of the display format designs. 

When the ANK cursor encounters such a field, it would 
automatically advance to the next eligible field. 

5. 2. 5. 5. 1 1 Crew Requirements 

ASF mission requirements dictate that several instruments operate 
simultaneously. Additional instruments may be activated as soon 
as some complete their data collecting cycles. Thus rapid and 


5.2.5-22 


continuous monitoring, experiment initiation and data analysis 
are required. The present assessment indicates that an excessive 
workload would be imposed on the operator if he were to have to 
change or correct incoming data, initiate new experiments, and 
monitor several sets of data. Therefore, the selected D&C 
approach is for the ASF pallet-only mode to limit the PS operations 
to the following. 

a. Initiate and interrupt preprogrammed sequences. 

b. Check initial conditions (modes, filters, etc,), 

c. Perform limited manual operations (pointing, TV monitoring, 
etc . ) . 

d. Act as a decision maker in off-nominal conditions. 

e. Perform real time updates and changes to sequences. 

In a very few instances the PS will analyze data, if it is the 
only way to ensure that a given instrument is performing 
correctly. 

Using the above as the PS's task, definition and based on the 
operating philosophy discussed in 5.2.5.5.10, one PS per shift 
will be able to operate the PSS for an ASF-configured mission. 

The ASF experiments will operate on a full time basis. It is 
recommended that two PS's, each working a 12-hour shift, be 
used. A clear definition of crew tasks is required to determine 
whether this payload will require a fifth crewman (chargeable 
to ASF) or whether the pilot or commander could serve as a PS 
on-orbit. 


For any ASF mission requiring detailed real time onboard data 
analysis, a preliminary study was performed to look at the 
possibility of providing a data monitoring station in the 

5.2.5-23 


mid-deck area v/here one or more principal investigators (Pi's) 
could monitor experiments and make real time inputs to the 
instruments through communication with the PS. Figures 5. 2. 5-9 
and 5.2.5-10 show the concept of a swing-out modular monitoring 
console in the mid-deck area. The console would be launched in 
position A. During on-orbit monitoring it would be moved to 
position B. The details of this concept will be supplied at a 
later time. Another possible location for this console is in the 
area of the airlock shown in positions A and B. This would require 
that the airlock be flown in the payload bay. Details of this 
concept will be the subject of further study. 

5.2.5.5.12 Crew Training Requirements 

Optimally, the crew training would be accomplished using a full 
scale, instrumented aft crew station simulator with all stations 
(PSS, MSS, and on-orbit) configured to the ASF mission. 

Since many of the scientific objectives .of the mission require 
the PSS and fiSS (and in some cases, the commander and pilot) to 
perform simultaneous or interdependent functions, the training 
should be done with the entire crew training to perform the 
scientific requirements. 

This simulator v/ould be dedicated to payload requirements and 
should be located in close geographical proximity to the Shuttle 
Mission Simulator. 

In addition to the Payload Mission Simulator, a flight config- 
ured mockup (non-i nstrumented ) of the aft crew station should 
be provided for crew familiarization of D&C, stowage, etc. 


5.2.5-24 


OPTIONAL MODULAR CONSOLE 



Figure 5. 2. 5-9. —Crew compartment, midsection plan view. 



original page is 

OE POOR QUALITY 







It is assumed that a part task trainer will be provided 
for RMS and R&D training, using out-the-wi ndow simulations. 

The training would generally consist of the following. 

a. Classroom briefings on scientific requirements. (The Pi's 
would be greatly involved with this portion of tii® training.) 

b. Mockup familiarization of D&C locations, stowage, velcro 
placements, etc. 

c. Simulator training for timelines, off-nominal situations, 
etc. A Payload Mission Simulator should be provided to 
accomplish the crew training with personnel from the various 
payload areas acting as briefing and training personnel. 

5. 2. 5. 6 Analyses and Trade Studies 

The functional requirements for D&C's defined by the instruments 
and support subsystems were evaluated. A key issue in establishing 
the ASF D&C approach was manual versus automatic (computer) control 
of the experiments, instruments, subsatellite, and the support 
subsystems. Using a full scale hard mockup, a preliminary layout 
evaluation was performed utilizing discrete switches, i.e., 
thumbwheels, windows, light indicators, pushbutton illuminators, 
etc., to ascertain the panel space required. 

Results indicated that slightly more than one bay 48.26 cm x 
53.34 cm (19" x 21") was required to accommodate these switches 
alone. This left less than two bays for CRT's, keyboards, timing 
devices, A/D recorder panels, etc., which was considered unaccept- 
able. 

Further analysis indicated that the panel space for D&C's could 
be reduced an order of magnitude with extensive automation 
utilizing keyboard interaction and display data formats for 
operator/computer communication. 


5.2.5-27 


The conceptual approach selected was to automatej to the maximum 
extent practical utilizing keyboard, CRT interaction and to use 
discrete switches or manual control elements where automatic 
jcontrol was unacceptable or impractical. 

Equipment selections were based on limited evaluations of avail- 
able candidates. The objective of these evaluations was primarily 
to verify that the selected equipment was compatible with the ASF 
D&C requirements and the Orbiter constraints. Evaluation of the 
keyboard concept went into somewhat more detail due to the 
importance of the man-machine interface and the impact of the 
keyboard concept on other segments of the D&C. 

5. 2. 5. 6.1 Cathode Ray Tube 

The instrument requirements dictate that the display unit must 
generate A/N's, symbols, vectors, and circles for display on a 
predetermined format of static and/or variable data. A double 
brightening and flashing of a character or group of characters 
is a desirable feature. In addition, expansion and contraction 
of displayed data as well as rotation are probably required. 

The viewing distance for the PS operator working in a zero-g 
erect position will range from 50.80 to 71.12 cm (20 to 28"). 

This will require a capability to present a character height 
of 0.318 cm (0.125") or more. This A/N and graphic capability 
is available in most CRT display units today, whether they use 
dot matrix, stroke writing, or plasma techniques. Recognition 
of some symbols in smaller size dot matrix configurations has 
been a problem in past studies so this type was not considered. 
Plasma displays have not been used operationally for a suffi- 
cient length of time so these display types were not investi- 
gated. The obvious candidate to investigate was the display 
I'liit being developed for the Orbiter MCDS. 


5.2.5-28 



The Orbiter CRT meets ASF display requirements. The symbols 
generated in the Orbiter display electronics unit are different 
from those required for the ASF system, as one might expect. 
However, this is a software change rather than one impacting 
hardware. 

The Orbiter MCDS format samples demonstrate that two separate 
graphs could be adequately displayed on the CRT screen (figure 
5. 2.5-6). For ASF mission applications, the split screen format 
is a significant advantage because of the large amount of data 
required to be displayed. Therefore, this feature is retained. 

5 . 2 . 5 . 6 . 2 Keyboard 

There are several keyboard devices available and table 5. 2. 5-2 
presents the types considered. 

a. Alpha-numeric Keyboard. The ANK is considered to be a slow 
input device, when compared to other input devices, because 

a number of separate operator actions are required to initiate 
a computer input. Evaluation of ANK usage indicates that 
picture modification (changing data on a display) of 10 item 
ch&nges will take approximately three minutes. Modification 
of three items containing 18 characters can be accomplished 
in approximately 32 seconds. The ANK is a highly versatile 
input device useful for all situations except those that are 
time critical. 

b. Page Overlay Keyboard (POK). The POK is designed to be the 
most rapid and useful of all keyboard input devices. Proper 
functional programming of the creative instructions for folio 
design can make this input design perform as desired. Soft- 
v/are for the folio and picture format is complex. Minimum 


Keyboard 
i Types 


Comparative' 

Factors 

Error 

Potential 


Operator 

Feedback 


Speed and 
Ease of 
Entry 


Training and 
Skill 


Addition and 
Modification 


Functional 

Grouping 


Operational 

Flexibility 


A/N 

Capability 


Discrete 
Capabf 1 f ty 


/ a"''?' 


Very high 


Requires 
extra dis- 
play area 


Slow and 
difficult 


Extensive 
training 
plus con- 
stant 
practice 


Easily modi- 
fied via 
software 


Does not 
exist 


Excel 1 ent. 

Poor in 

backup 

hardware 

logic 

mode 


Medium 


Very poor. 
Keys do not 
provide 
meaningful 
labels 


/ of 

/ 

/ 


Very low 


Minimum 

training 


Must add 
new 

hardware 






Very good 
(button, 
labels, 
lamps) 


High due 
to lack of 
groupi ng 


Fair 

(button, 

labels, 

lamps) 

Slow 

operator 
must search 
panel 


Moderate 

training 

constant 

use to 

retain 

skill 

Hay require 
new hard- 
wire 


Excellent Very poor 


Very good Good 
in backup 
logic 
mode 


Limited Limited None 


Medium Extensive Minimum 


Very good 
but 

redundant 


Very good 


/ <f <S* 

-S' 

1 ? 4 " j 


Very high. 
Inputs not 
user 

oriented 

Very low 

L . . . 

Very low 

Very poor 
unless dis- 
play is 
associated 

Very good 
(button , 

1 abel s, 
lamps ) 

Very good 
(button, 
labels , 
lamps ) 

Fast for 
single 
inputs . 

Sic -er for 
mul 1.1 - 
inputs 

Fast 

Fast 

Minimum 

training 

Minimum 

training 

Minimum 

training 

Easily 
modi fied 
via 

software 

Easily 
modtfi ed 
via 

software 

Easily 

modified 

via 

software 

Does not 
exist 

Excellent 
by page 
and folio 

Excellent 
by page 
and folio 

1 

Good 

Good 

Good 

None 

Limited 

difficult 

Some 

Minimum 

Minimum 

51 ight 
increase 
over POK 

None 

Excellent 

but 

redundant 

between 

pages 

Excellent 


original 
OR POOE QUALIFY 


5 . 2 . 5-30 



fixed functions (those that should be used on every page) 
should not be assigned to POK buttons. The major drawback of 
the POK is that its A/N capability is very limited. 

c. Fixed Function/Numeric Fixed Function Keyboard (FFK, NFFK). 

The FFK is a design which contains a limited number of 
possible inputs, thereby being the fastest input device 
available- The concept would be to assign fixed functions 
to this device which would be similar for all instruments 
or experimental packages. Its drawback (as in the case of 
the POK) is that its A/N capability is limited. 

d. Di spl ay/ Keyboard , Computer (DSKY). This keyboard has utility 
when extensive grov/th is not required. Normally considered 

to be a noun-verb-numeric (Apollo type) type, it has advantages 
when the same population will always be the users. Its 
primary disadvantage is its lack of functional grouping and A/N 
capability. It also requires more panel space for growth 
than the other candidates. 

The keyboard selected was the ANK because the A/N input 
capability is considered to be an important feature, the 
operational flexibility is highest of the candidates, 
functions and formats can be easily modified through 
software, and additional hardware is not required for 
future growth. Initial assessment indicates the speed of 
the ANK is adequate for the ASF application and that 
functional grouping is not necessary. 

The capabilities of the keyboard and CRT candidates selected 
can be greatly expanded without adding additional hardware. 

This concept requires the layout of formats and a command 
language interface easily handled and understood by the user. 
This type of keyboard/display/computer interface is presently 
being used in many operating systems and will meet the ASF 
requirements according to preliminary assessment. 


5 . 2 . 5-31 




1 


The one major disadvantage to this concept is its slow input 
capability even when meaningful abbreviations are used. 

Detailed task sequences and timelines which relate to the 
total instrument complement usage according to experiment 
objectives must be generated. It is only after these details 
are defined that a final assessment can be made to determine 
if the typewriter ANK can handle the data communication 
requirements within the time specified. 

Because of the total reliance on the keyboard (and CRT) for 
mission success, an additional ANK should be stowed onboard 
as a backup in case of failure. 

5. 2-5.7 Conclusions and Recommendations 

5. 2. 5. 7.1 Concl usions 

From the results of this study, the following conclusions have 

been made. 

a. Sufficient space is available at the Orbiter aft crew station 
port side console to accommodate the ASF D&C requirements. 

b. D&C panel area is available for future growth and for other 
ASF and AMPS configurations. 

c- Volume in the lower console bays of the aft crew station 

will be required for ASF supporting hardware (recorders, CDU(s), 
etc. ) . 

d. D&C space on the Orbiter C&W, R&D, RMS, MSS, and CCTV panels 
will be required to support the ASF missions. 

e. Creative display format design and keyboard interaction will 
be requi red . 

f. The operator-computer interaction will require complex soft- 
ware development. 

g. Instrument scientific data collection must be accomplished fay 
preprogrammed experiments initiated by the operator but 
executed by the computer. 


5.2.5-32 


h. The PS's primary functions are to operate the payload 
Instruments and monitor the scientific data collection. 

1.^ The ASF mission requires 24 hr/day operation. One operator 
per shift Is adequate to operate the ASF Instruments. 

V. . 2 . 5 . 2 . 7 Recommendations 

Recommendations are as follov/s. 

a. Lower console volume be assigned for dedicated payload use. 

b. Dedicated panel space on the RMS, CCTV, and R&D panels be 
assigned for payload use. 

c. ASF data display formats be designed and keyboard interactio 
defined for submittal to software analysis and computer 
storage requirements. 

d. ASF task sequence/timelines be further detailed to complete 
the D&C arrangement/layout definition. 

e. Complete AMPS requirements be defined to establish total D&C 
requirements. 

f. A follow-on study be performed to define the details of the 
data monitoring station In the mid-deck area for scientific 
real time onboard data analysis by experiment Pi's. 


5.2.6 PARTICLE DETECTOR SUBSATELLITE (PDS) 
5. 2. 6.1 Introduction 


In defining the ASF pallet-only concept and feasibility, it was 
determined that a subsatellite was required to carry out the 
experiments previously described. The instruments required were 
similiar to those on the AE satellite and hence, this satellite 
was the baseline for the subsatellite described herein. 

5. 2. 6. 2 Requirements 

The experiments AS-4, AS-5, AS-9, AS-10, AS-11, AS-12, and AS-13, 
described in section 4.1.3 and appendix A, require information 
that cannot easily be obtained using instruments mounted on the 
Orbiter vehicle. Other instruments are required on the vehicle 
and in the near vicinity (10 km) of the Orbiter vehicle. These 
two factors alone dictated a subsatellite and, since most of the 
instruments were particle detectors (ions, electrons, etc.), the 
name "particle detector subsatellite" was selected. 

The instruments which were determined to be best suited for the 
subsatellite require the following engineering support functions 
onboard the subsatellite. 

a. Provide for attitude control, maneuvering, and stabilization 
of the subsatellite. 

b. Provide velocity change capability. 

c. Provide attitude determination capability. 

d. Provide downlink communication capability for scientific 
and engineering data. 

e. Provide for data processing and formatting. 

f. Provide for uplink command capability. 

g. Provide for onboard experiment programming and control. 


5 . 2 . 6-1 


h. Provide pov/er and power control for instruments and support 
systems . 

i. Provide instrument and support system health status and 
diagnostic data- 

5. 2. 6.3 Guidelines and Assumptions 

The following guidelines and assumptions were made for the study. 

a. The basic AE satellite instruments and support systems will be 
used to the maximum extent possible in their existing 
locations. 

b. The AE satellite will be in production during the ASF time 
frame {beyond 1 981 ) . 

c. The subsatellite will operate at or near the same orbital 
altitude as the ASF payload at a distance of about 10 km. 

d. The subsatellite will be passive, cooperative (except for 
attitude control) for the rendezvous and retrieval operations. 

e. Control of the subsatellite attitude and velocity v/ill be 
provided from the Orbiter Mission Specialist Station (MSS). 

f. Scientific and engineering data link will be primarily with 
the Orbiter. 

5. 2. 6. 4 Capabilities and Constraints 

The capabilities and constraints of the basic AE satellite are 
shown in table 5. 2. 6-1. 

5. 2. 6. 5 System Description 

The subsatellite will be used as the platform on which the particle 
detection instruments will be mounted. The instruments will 
provide the necessary particle data in support of the experiments 
being conducted by the ASF. Thi s subsatel 1 i te wi 1 1 be of the AE 
type v/ith the configuration shown in figure 5. 2. 6-1. 


TABLE 5.2,6-K- AE CAPABILITIES SUMMARY 



Parameter 

Value 

Spacecraft Weight (less payload) 

560 kg 

Payload Weight (typical) 

100 kg 

Projected Area 

1 ,5 m^ 

Experiment Footprint Available 

0.8 m^ 

Experiment Volume Available 

0.2 m^ 

Energy Available to Experiments 
(orbit average) 

4000 watt minutes 

Regulated Voltage 

-24.5 V +2% 

Temperature Range 
(upper baseplate) 

10“ C to 15° C 

Temperature Range 
(lower baseplate) 

10“ C to 28“ C 

Attitude Determination Accuracy 

0.5“ 

Attitude Control Accuracy 

1.0“ 

Spin Rates Available 

1 revolution/orbit; 1 to 10 rpm 

Minimum Operating Altitude 

i 

120 km 150 km (depends on 

stabilization mode and apogee 
altitude) 

Orbit Adjust Capability 

fijei 0 m/ sec 

Maximum AV per Burn 

7.6 m/sec 

Memory Capacity 

2 X 32 kilobits 

Memory Delay Time (maximum) 

72 hours maximum 

Command Op-Codes Available to 
Experiments 

260 

Recorder Capability 

2 X (1/2 X 10^ bits) 
2x2 hours record time 

Maximum Playback Data Rate 

•vl 30 ki 1 obits/sec 

Maximum Communication Rates 

1 6 transmi tted 


8 received 


“A plan for the use of the Basic Atmosphere Explorer 
Spacecraft System as a Subsatellite of the Shuttle". Goddard 
Space Flight Center* May, 1973. 











MIRROR 


CEP PROBE 


TOP PANEL COVER 


HAT-UPPER . 


APERTURE-EXPERIMENT - 


DRIVE UNIT 
MOMENTUM WHEEL 


YAW THRUSTER 


AV THRUSTERS 


SIDE PANELS 


EXPERIMENTS AND 
EQUIPMENT 


SHEAR PLATE 


BELT ANTENNA - 


HAT-LOWER 


PROPELLANT TANKS (6) 


CENTRAL COLUMN 
MOMENTUM WHEEL 


EXPERIMENTS AND 
EQUIPMENT 

BASEPLATE-UPPER 


BASEPLATE-LOWER 


NUTATION DAMPER ^ 
MOMENTUM TORQUER 
ASSEMBLY 




5 . 2 . 6 . 5 . 1 General Description 


In general configuration, the AE subsatellite is a 16-sided 
polyhedron, 136 cm (53.5 inches) in outside diameter and 114 cm 
(45 inches) high, weighing 678 kg (1494 lbs). The subsatellite 
contains a 3-axis attitude control system utilizing a momentum 
wheel to provide roll-yaw stiffening and pitch orientation, and 
magnetic torque coils to maintain momentum axis orientation in 
inertial space. A combination of fluid-filled loops and a cage- 
tuned pendulum is used for nutation damping. A thruster and 
monopropell ant hydrazine fuel supply is used to provide orbital 
adjust capability. An active thermal control system maintains 
subsatellite temperatures within operating limits. Command and 
communication systems are compatible with the Orbiter communica- 
tion system. Power is obtained from a skin-mounted solar cell 
array and a battery pack. The subsatellite is designed to be 
launched by a system of pressurized gas thrusters to impart the 
small velocity increment necessary to achieve the desired sepa- 
ration between the subsatellite and the Orbiter. Recovery of the 
subsatellite will be accomplished using the remote manipulator 
system. 

5. 2. 6. 5. 2 Internal Configuration 

The internal configuration consists of two baseplates separated 
by a central column and connected by six shear ties- Spacecraft 
components and experiments are mounted to one side of each 
baseplate. Six cono-spheri cal shaped propellant tanks, carrying 
a total of 169 kg (373 lbs) of hydrazine propellant, are grouped 
symmetrically about the central column, sandwiched between the 
two baseplates. The momentum wheel assembly for spacecraft 
stabilization is mounted on one end of the central column within 
the spacecraft, with its attitude sensing mirror assembly 
projecting through a hole in the center of the upper surface. 


5.2. 6-5 


5. 2. 6.5. 3 Subsatellite Coordinate System 

The subsatellite coordinate system is a right hand system as shown 
in figure 5. 2. 6-2. The origin is taken at the geometric center 
of the subsatellite, on the axial centerline midway between the 
end surfaces. The coordinate system is body fixedj the orbital 
directions shown are for reference only. 

The coordinates are defined as follows. 

a. Subsatellite top: -Z axis 

b. Subsatellite bottom: +Z axis 

c. Pitch: -Z axis 

d . Rol 1 : +X axi s 

e. Yaw: +Y axis 

5. 2. 6. 5. 4 Instruments 

The ASF experiment support instruments on the subsatellite are as 
follows; 

a. Cylindrical Electrostatic Probe (CEP) 

b. Low Energy Electron Probe (LEE) 

c. Airglow Photometer (VAE) 

d. Photoelectron Spectrometer (PES) 

e. Triaxial Fluxgate Magnetometer (MAG) 

f. Planar Ion Trap (RPA) 

g. Neutral Atmospheric Composition (NACE) 

h. Neutral Atmospheric Temperature (NATE) 

i. Cold Cathode Ion Gauge (CCIG) 

j. Low Energy Ion Detector (LEID) 

k. High Energy Particle Detector (HEPD). 




NUMERICAL AXIS 


2 

1 

-3 

RIGHT HAND 
ABOUT -3 

ALONG +2 

ALONG +1 


ALONG +3 














The first nine instruments are existing AE devices. The last two 
(LEID and HEPD) are additional instruments required to meet ASF 
requirements . 

The instruments providing the particle data and ancillary informa- 
tion are mounted along the outer edge of the baseplates as shown in 
figure 5. 2, 6-3. The particle detector instruments are oriented so 
that the primary axis of measurement is at a 67.5° angle to the spin 
axis of the subsatellite. Each of the electron and ion detectors is 
repeated. Each pair of instruments is placed so that the primary 
axes are in opposition. Thus, as the subsatellite spins, measure- 
ment of the particle flux is made in both directions along two lines 
near the orbit trajectory. Due to the spin, these lines sweep out 
conjugate cones and each instrument repeatedly scans above and 
below and to each side of the local magnetic field lines. 

5. 2. 6. 5. 5 Support Subsystem 

Except for the addition of three nickel -cadmi urn rechargeable 
batteries, and the deletion of the tape recorders, the support 
systems provided by the basic AE satellite remain basically 
unchanged for the ASF application. 

5. 2. 6. 6 Anal ys i s 

9 3 

The AE * was used as the starting point for the ASF subsatellite. 
Comparing the basic AE scientific and engineering capabilities 
with the ASF support requirements led to changes. These changes 
include the following. 


"Atmosphere Explorer (AE Spacecraft System Description)"; RCA 
Government and Commerical Systems, Astro Electronics Division; 

AED R-3816F; March 30, 1972; Attachment B (updated August 8, 1974). 

^GSFC Specification for Atmosphere Explorer (AE-C, D and E)"; 
Goddard Space Flight Center, S-e^O-P-l; September, 1973. 


UPPER BASE PLATE 


LOWER BASE plate 



cep-cylindrical electrostatic probe 

LEIO-LOW ENERGY ION DETECTOR 
HEPD-HIGH ENERGY FARTClE DETECTOR 
LEE -LOW ENERGY ELECTRON DETECTOR 

vAE-AiRGLOw Photometer 
pes-photoelectron spectrometer 
MAG - magnetometer 

R PA - PLANAR ION TRAP 

NACE -neutral HASS SPECTROMETER 
NATE -neutral ATMOSPtHERiC TEMPERATURE 

ccig-cold cathode ion gauge 

BAT - BATTERY, ADDED 


HOTi: 

tMAOCO AREAS DENOTE E(iSTin« At SATIlliTE 
SURRORT StSTEHS. 


Figure 5. 2. 6-3 — Subsatellite Instrument Layout 




Scientific. Of the 17 scientific instruments used on the 
combined AE-C, D and E satellites, eight will not be required 
for ASF experiments and will be removed. Two new instruments, 
the Low Energy Ion Detector (LEID) an/d the High Energy Particle 
Detector (HERD) will be added to respectively provide , 

He and 0 ion detection in the 0 to 10 keV range and to cover 
the energy range of from 25 keV to 10 MeV for electrons and 
protons . 

Engineering. The total power required by the instruments and 
support systems for the ASF subsatellite remains at about the 
same level as that required by the AE satellite. However, 
three additional batteries v/ill be added to increase the 
experiment duty cycle capability about 50 percent. The tape 
recorders are not required since the subsatellite will 
communicate continuously with the Orbiter, and data time 
compression is not required. The accuracy (about 1 km) 
achievable with ground tracking using pseudorandom noise 
techniques is not adequate for primary mission state vector 
determination due to Orbiter crew and vehicle safety con- 
siderations. Therefore, the range and range rate determina- 
tions will be made using the Orbiter rendezvous radar 
system. However, since the tracking system is a small part 
of the total communication transponder, these circuits will 
be left in the system. Provisions must be made to disable 
these circuits when Orbi ter/subsatel 1 ite communication is 
required. For verification purposes, the ground tracking 
facility will then be able to use the subsatellite track- 
ing system to determine approximate subsatellite position 
and velocity. 


5 . 2 . 6-10 



5. 2. 6. 7 Conclusions and Recommendations 




5. 2. 6. 7.1 Conclusions 

The conclusion resulting from the study is that the AE satellite 

is an acceptable candidate for use as the ASF remote subsatellite. 

Some moderate changes in the instrument complement and minor 

changes in the support systems will be required, 

5. 2. 6. 7. 2 Recommendations 

The following recommendations apply to the PDS definition: 

a. Establish a subsatellite experiment and instrument timeline 
which is correlated with the overall ASF mission timelines. 

b. Using the subsatellite timeline, determine power and data duty 
cycle requirements to be used in further defining support 
system needs. 

c. Develop a detailed operational sequence including that for 
the rendezvous and retrieval phase. Establish the safest and 
most effective way of capturing and retrieving the subsatellite 


5.2.6-11 


5.3 GROUND SYSTEM 


\ 


* 

The ASF ground systems are comprised of two major facilities: 

(T) one v/hich processes the flight data and disseminates the 
results to the scientific community, and (2) the ground support, 
test, and checkout facilities which include both mechanical and 
electrical GSE. 

5.3.1 GROUND PROCESSING OF FLIGHT DATA 

5.3.1 .1 Introduction 

During an ASF mission, electronic data will be delivered to the 
ground data reduction complex in one or more of four forms, all 
of which are technically feasible within the concepts explored 
during this JSC study. 

a. Unprocessed data transmitted from Orbiter to ground. 

b. Unprocessed data on magnetic tape delivered to ground at 
completion of mission. 

c. Processed data transmitted from Orbiter to ground. 

d. Processed data on magnetic tape delivered to ground at 
completion of mission. 

5. 3. 1.2 Considerations 

The processed data as defined in (c) and (d) above represent the 
most desirable forms, and as such, are the types of data the CDMS 
is designed to produce. Because these data will have been pro- 
cessed onboard the Orbiter prior to transmittal or storage on 
tape, the ground-based operation is greatly simplified. As the 
transmitted data is received, it will be stored on ground-based 
tape recorders. These tapes will then be reformatted to computer 
compatible tapes, screened, and forwarded to the scientists. 
Similarly, processed data delivered on magnetic tape at the end 
of the mission will be handled in like manner. There are numer- 
ous trade-offs to be considered regarding which of (c) and (d) 


5.3-1 


represents the greatest advantage and is yet cost effective. A 
combination of these two data forms is recommended. 

Data forms (a) and (b) represent the concept utilized in previous 
manned spacecraft experiments. Form (b) represents the worst 
case and is considered undesirable, although it is technically 
feasible. Analyses have indicated that all unprocessed digital 
PCM data generated during a 7-day ASF mission can be stored on 
seven reels of magnetic tape. This figure does not include wide- 
band analog video from instruments utilizing TV cameras. The 
degree of difficulty in transforming these seven reels of tape 
data into usable products for the scientific community is much 
greater than that applied to data forms (c) and (d). 

It is anticipated that even though data will be processed on- 
board the vehicle, certain unprocessed data will be trans- 
mitted in real time to allow some recovery ability in the case of 
processing errors or malfunctions. Unprocessed data will also 
be required on the ground in near real time to allow diagnostic 
analyses of certain instrument operations. This data, however, 
would not be intended for dissemination to the data user. 


5.3.2 GROUND SUPPORT, TEST AND CHECKOUT SUBSYSTEM 

5. 3. 2. 1 General 

The objective for this phase of the study is to define the 
conceptual design and requirements for an overall ground-based 
hardware support system considered both feasible and practical 
for the AMPS/ASF pallet-only payload concept. The system must 
accomodate all levels of preflight and postflight payload hard- 
ware testing plus considerations of major transportation, sto- 
rage, installation, and logistical requirements. 

5. 3. 2. 2 Requirements 

The AMPS/ASF ground support, test, and checkout subsystem must 
satisfy the following functional, hardware/software , and data 
requi rements . 

a. Provide verification that scientific and engineering para- 
metric requirements are met. 

b. Provide diagnostic and evaluation capabilities beyond the 
performance verifications of a. above. 

c. Incorporate optimized test flexibility, mechanical and 
structural support, and mobility concepts wherever possible 
within existing constraints (safety, reliability, quality 
assurance, programmatic, etc.). 

d. Encompass test and/or checkout at all assembly levels of 
AMPS/ASF flight equipment. This includes both preflight and 
appropriate postflight calibrations and tests. 

e. The ground support functions must encompass applicable trans- 
portation, storage, and logistical requirements. 

5. 3. 2. 3 Guidelines and Assumptions 

The guidelines and assumptions in paragraph 2.3.4 were used 
where applicable to this portion of the study. The major items 
pertained to standardization of equipment and utilization of 
Space! ab facilities wherever practical. 


5.3-3 


5 . 3 . 2 . 4 Subsystem Description 

The ground support, test and checkout subsystem is segregated 
into three general categories: (1) electrical GSE (EGSE), 

(2) mechanical GSE (MGSE), and (3) logistics. 

5. 3. 2. 4.1 Electrical Ground Support Equipment (EGSE) 

5. 3. 2. 4. 1-1 General 

The EGSE provides functional test and checkout of all physical 
parameters of the ASF i nstruments/experiments pertaining to 
electric, electronic, magnetic, electrostatic, and optical 
functions. It is made up of various types of test instruments, 
many varied readout and display devices, and numerous recording 
devices . 

The EGSE design is based on the use of computer-controlled auto- 
matic test equipment augmented by simulators. Figures 5. 3.2-1, 

5. 3. 2-2, and 5. 3. 2-3 depict the equipment comprising the EGSE. 

It is designed to support the ASF pallet-only mode during the 
integration, prelaunch, launch, postflight, maintenance, and 
refurbishment phases. The primary purpose is to assure that the 
ASF instruments and subsystems are operating within their design 
1 imi ts . 

The EGSE simulators support instrument integration at user sites 
and at the payload integration site. At the payload integration 
site the two-computer configuration shown in figure 5. 3. 2-2 allows 
for checkout of the ASF instrument pallets supported by payload 
subsystem simulators and Orbiter signal simulators. 

The EGSE is used for payload final checkout at the launch site 
by interfacing with the ASF CDMS via hard lines and telemetry 
(see figure 5. 3. 2-3). 

Overall test control is implemented via the EGSE GPC and check- 
out software and CDMS data acquisition capabilities. EGSE 




GROUND AIR COOLING UNIT 
SUPPLIED BY ECS 


ELECTRONIC TEST AND SERVICE 
EQUIPMENT 


• GENERAL PURPOSE TEST EQUIPMENT 

• CABLE KITS 

• SPARE PARTS 


GROUND FLUID COOLING UNIT 
SUPPLIED BY ECS 


ORBITER INTERFACE ADAPTER 


• TM DATA ACQUISITION 

• ORBITER SIGNAL SIMULATORS 

• UMBILICAL INTERFACE 


GROUND POWER SUPPLY 


POWER SOURCE 

CONTROL AND DISTRIBUTION 


• A/D CONVERTERS 

• D/A CONVERTERS 

• MULTIPLEXER 

• DEMULTIPLEXER 

• TRANSMITTER/RECEIVER 


OPERATOR CONSOLE 


• ATE. CRT AND KEYBOARD 

• SUBSYSTEM COMPUTER CRT AND KEYBOARD 

• EXPERIMENT COMPUTER CRT AND KEYBOARD 

• CAUTION AND WARNING DISPLAY 

• CCTV MONITOR 

• INTERCOM 

• QUICK LOOK DATA DISPLAY 
« STATUS DISPLAYS 

• TIME DISPLAYS 

• DEDICATED CONTROLS AND DISPLAYS 

• ALARMS 


EGSE COMPUTERS 


• PROCESSOR (21 

• MEMORY (2) 

• I/O (2> 

• RAUS 


EGSE COMPUTER PERIPHERALS 


CARD READER 
TELETYPEWRITER 
LINE PRINTER 
MAGNETIC TAPES (21 
DUAL DISK STORAGE 
PAPER TAPE READER/PUNCH 


INTERFACE UNIT 


RECORDING AND TIMING UNIT 


• HIGH SPEED A/D 

• PCM OECOMMUTATOR AND 
D/A CONVERTERS 

• SWITCH MATRIX 

• PATCH PANEL 

• PCM SIMULATOR 


• PCM TAPE RECORDERS (21 

• TIME BASE 

• TIME CODE GENERATOR 

• ANALOG TAPE RECORDERS (2) 

• TAPE SEARCH UNIT 

• STRIP CHART RECORDER 


♦LOCATED IN A SEPARATE CONTROL ROOM 


AMPS EGSE assemblies 















O o 

tog 

F Q 
a © 


CDMinOL CONSDLE 


TV MDNUfQH 
LNTtnCDM 

SYSTEM POWER 
ANP mope 

SWUCHES/iriDICATOnS 


TO PALLET CAMERA ^ 
TO PAU.ET 


ELECTRONIC TEST EOUn-MEHT 

• OINERAl. PUnPflSE lOSClLLOSCOPES. VOWS. ETCi 

• CABLE XITS 

• SPARE PARTS (CDWPDNEJjTS, MODULESi 


onBtten interface APfineR 

ORBITER SIGRAL 
SIMULATORS 

» ORBITER DPS BUSES IZi 
• CPUS TM BUSES tZ) 

* . ECSE POSES <» 


PAYtOAP SUBSYSTEM 
SIMULATORS 

* PBS 

* CCS 

* STRUCT UR AL/MECICAUiCAL 

* PCSS 


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RECDflOtRS ^71 
PCM RECOnPERS IZI 
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TtME COPE PENERATOn 
TAPE SEARCM UHtT ^ 
ginip CHART REcanpER 
CONTnOLLiR 

iNTERFAtE UNIT 


PCM SIMULATOR 
PATCH PANEL 
EVJ1TCH matrix 

itrcHSPEEP A/p 
and COMPUTCn 
DATA FORMATTER 


, 4:2 MRi ANALOG 


CRPUNP 

— 

POWER 


SUPPLY 




UKREO PC 
REG DC 
400 H» AC 
Cfl Mr AC 
«D Ht AC 


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cpmputeii 



.CRT DIEPLAVS 1 ARO 2 


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PAYLOAD 
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1 


1 

insthument 


CmWEnCIAttV 


EQUIPMENT 



RAU 


FLiGHT EQUIPMENT 


Figure 5.3. 2~3. — AMPS EGSE (launch site) 




















control, measurement, stimuli, recording and processing capabili- 
ties allow detailed testing, fault isolation and tasks such as 
data reduction and test result display and printout. The 
control console controls the EGSE 6PC and accesses the onboard 
payload and subsystem computers via CRT keyboard terminals 
connected to CDMS I/O units. Manual controls are also provided. 

The EGSE Ground Subsystem Computer at the launch site will be 
used as a backup computer to the 6PC, It will also perform off- 
line duties such as processing of TM tape dum.ps. 

Additional EGSE equipment simulate ASF CDMS interfaces to facili- 
tate payload preparation prior to integration into the Orbiter. 
Integration of instruments into the pallets is supported by the 
Spacelab Simulator for Experiments (optional, interface verifica- 
tion testing only) and the Core Segment Simulator (CSS). These 
items are discussed in later paragraphs. 

5. 3. 2. 4. 1.2 Utilization of EGSE 

Utilization of the ASF EGSE for four stages of instrument check- 
out is depicted in figure 5. 3. 2-4. 

a. Early Interface Verification. Verification of instrument 
interfaces will be supported at any site by the portable 

ASF simulator for instruments (option), which allows for early 
interface verification. 

b. Instrument Integration or Pallets. Integration and test of 
complex instruments will be supported by the CSS at any site. 
The CSS is a modular portion of the ASF EGSE and it simulates 
the flight CDMS and EPOS services provided for the payload. 

The CSS includes a payload computer and its associated peri- 
pherals. 

c. Payload Integration for All Pallets. Integration and test 
of all pallets is supported at the payload integration site 
by the full complement of ASF EGSE, excluding the TM data 





cn 

CaI 

I 

V£> 


C. PAYLOAD I.MTEGRATION SITE 

CENTRAL INTEGRATION FOR ALL PALLETS 





B. ANY USER SITE 

EXPERIMENT INTEGRATION ON PALLETS 



D, LAUNCH SITE 

PRELAUNCH, LAUNCH AND POST-FLIGHT CHECKOUT SUPPORT 



E6SE utilization 













acquisition equipment. Flight-packaged instruments and 
RAU's are tested with actual flight software at this site. 

d. Launch Site Integration. At the launch site, the ASF EGSE 
is used with the TM data acquisition equipment. During the 
prelaunch and launch phases, the EGSE is used for ASF pay- 
load and subsystem verifications, interface verifications, 
and final checkout. The ASF EGSE interfaces only with the 
ASF CDMS via TM and hard lines. 

5. 3. 2. 4. 1.3 Interfaces and Functions 

a. Simulator for Instruments. The ASF Simulator for instruments 
is depicted in figure 5. 3. 2-5. The interface characteristics 
are as follows: 

(1) Facility Interfaces: 

Wei ght : 

Dimensions : 

Power : 

Type of Connectors: 

£nvi ronment : 

(Operating) : 

(2) Instrument Interfaces: 

(a) EPOS Interface Simulation: 

Unregulated dc - 26 to 32 Vdc - (TBD) kW 

Regulated dc - 28 Vdc ± 2 % - (TBD) kW 

400 Hz ±1% ac - 115 Vac ±5% - (TBD) kW 

60 Hz ±1% ac - 115 Vac ±5% - (TBD) kW 

50 Hz ±1% ac - 220 Vac ±5% - (TBD) kW 

(b) RAU Interface Simulation: 

RAU interface technical requirements are identical 
to flight RAU. 

b. CSS. The CSS interfaces are not fully determined. Essentially, 
the CSS will provide limited EPOS and CDMS simulation. 

Figure 5. 3. 2-4 (B) depicts the use of the CSS at any user site. 
Figure 5. 3. 2-6 shows the functional units of the CSS. The 
CSS would comprise several racks of equipment whose functional 


(TBD) 

(TBD) 

115 V,.60 Hz, (TBD) kW 
(TBD) 

22 ±5*’C 

40 to 80% relative humidity 


220 V 
115 V 


50 Hz 


1 


INTERNAL 

POWER 

SUPPLY 


OPERATOR 

INTERFACE 

• CONTROLS 

• INDICATORS 

• METERS 

• TEST POINTS 


POWER CONTROL 
AND MONITORING 


ADDRESSING/ 

SAMPLING 


EPDS 

INTERFACE 

SIMULATION 


DC 

REG 


AP 

400 Hz 


AC 

60 Hz 


RAU 

INTERFACE 

SIMULATION 

• ANALOG 

• DISCRETES 

• PCM 


RECORDER 

INTERFACE 

SIMULATION 

• 6 MHz ANALOG 

• 50 MBPS PCM 


OUTPUTS 


TO INSTRUMENT 


INPUTS 


Figure 5. 3. 2-5. 


AMPS simulator for experiments. 




CONTROL CONSOLE 


'I 


I-., 


AUTOMATIC 

TEST 

EQUIPMENT 


r 





PAYLOAD (EXPERIMENT) 
SEGMENT OR PALLET 


Figure 5. 3. 2-6. -Core Segment Simulator (CSS). 



i 


units are identical to those named in the full complement EGSE, 
and are described tn the next section, 

(1) Control Console. Provides operator control of automated 
checkout, contains C&W, TV monitor, recorder control, and 
i ntercom . 

(2) Ground Payload Computer. 

(3) Ground Payload I/O Unit. 

(4) Mass Memory. 

(5) Recorders. 

(6) EPOS Simulator. 

c. Payload Integration Site EGSE. A full complement of EGSE is 
required to integrate and test all of the ASF experiments 
(see figure 5, 3. 2-2). All equipment shown to the left of the 
vertical dotted line is commercially packaged. 

Two modes of checkout will be used. The first is the single 
instrument checkout. A single flight-packaged instrument will 
be tested before it is mounted on a pallet by using a commer- 
cially packaged RAU, and the rest of the EGSE. The second 
mode is checkout of one or more pallets with flight packaged 
RAU ' s and instruments. Commands to the pallet RAU’s are sent 
in PCM format at a 1 Mbps maximum rate. The RAU's send output 
discretes and PCM signals to control the attached instruments. 
Housekeeping and scientific data are sent from the instrument 
to the RAU's in PCM, discretes, or analog form. The RAU con- 
verts the data into PCM for return to the Ground Payload I/O 
Unit. The Ground Payload I/O Unit formats the data, buffers 
it, and sends it to the GPC for processing. Additional wide- 
band scientific data from some instruments are recorded on 
analog and PCM tape recorders. 

The operator uses the items in the control console to auto- 
matically perform extensive instrument checkouts using both 


5.3-13 


computers, I/O units, mass memory, computer peripherals, 
recording and timing equipment and the interface unit. Use 
of computer-controlled instrument calibration equipment can 
help shorten instrument checkout time. 

Recommended EGSE equipment: 

(1) Ground Payload Computer, 

(2) Ground Payload I/O Unit. 

(3) Ground Subsystem Computer. 

(4) Ground Subsystem I/O Unit. 

(5) Mass memory, 

(6) Peripheral switch. 

(7) Disk storage. 

(8) Magnetic tapes. 

(9) Line printer, 

(10) Paper tape reader/punch . 

(11) Tel etype. 

(12) Card reader. 

(13) Instrument calibration equipment. 

(14) Control console. 

(15) Recording and timing equipment. 

(16) Interface equipment. 

(17) Orbiter interface adapter. 

(18) PSS. 

(19) Electronic test equipment. 

d. Launch Site EGSE. The ASF EGSE required for the launch site 
is shown in figure 5. 3. 2-3. The GPC and checkout software 
communicates with the Orbiter AMPS CDMS via payload and I/O 


5.3-14 


i 



GSE buses through the umbilical interface and via TM. In 
this configuration, the GPC requests the ASF CDMS computers 
to perform checkout of the instrument pallets and send the 
checkout results down to the EGSE for real time and post test 
analysis . 

The Ground Subsystem Computer and Ground Subsystem I/O Unit 
may serve as backups. The computer will be used for processing 
TM tape dumps. 

The same equipment shown in figure 5. 3. 2-3 has been previously 
specified in the paragraphs on the Payload Site EGSE, with the 
following exceptions: 

(1) Orbiter Interface Adapter (Launch Site). 

(2) Computer Test Equipment. 

5. 3. 2. 4. 1,4 Software 

a. Ground Checkout Software. The ASF ground checkout software is 
used with the EGSE computers at user sites, the Payload Inte- 
gration Site, and at the Launch Site. The software will be 
additional application software added to the onboard CDMS 
subsystem and payload software such that certain additional 
standard routines may be utilized. 

The software will provide for the ability to operate in both 
an automatic checkout sequence as well as manual mode. The 
onboard computer routines for formatting data, display, 
sequencing and initialization will be used to the maximum 
extent possible. 

b. Support Software. The support software will provide the 
capability to develop, verify and maintain the ground checkout 
softv/are. This software will be developed to operate on 

IBM 360/370 and/or CDC 6500 computers to assure its useability 


5.3-15 


at various installations. The development process requires 
that all software must be coded in the GPC assembler or com- 
piler language. The software is then assembled on an IBM 
360/370 or CDC 6500 computer and verified with a simulator 
resident in the same computer. The simulator-verified soft- 
ware is thei used with both computers in the EGSE to integrate 
and checkout the ASF instruments. 


Provision will be made for an interpretive simulation for 
the onboard CDMS computers and functional simulations of 
external devices to provide a source of data for the software 
test functions. Main items in the support software package 
are : 

(1) Program generation and maintenance. 

(2) Simulation and program validation software. 

(3) Utility software. 

5. 3. 2. 4. 1.5 Packaging and Modularity 

The ASF EGSE is designed to use commercially packaged equipment 
in a modular arrangement to reduce equipment and maintenance 
costs . 


5. 3. 2. 4. 1.6 Quantitative Requirements 


Equipment Quantity Required 


ASF Simulator for Experiments (TBD) 
Core Segment Simulator (TBD) 
Payload Integration Site EGSE 1 Set 
Launch Site EGSE 1 Set 
Payload Crew Training Simulator 1 Set 


5.3-16 


I 


I 






5. 3. 2. 4. 1.7 Payload Mission Simulator (PMS) 

The PMS 1s comprised of two major subassemblies as follows: 

a. A full scale, Instrumented operating aft crew station co.sole 
simulator capable of utilizing flight (mission) software. 

b. A programmable electronic payload simulator and operator's 
consol e . 

The PMS will be designed to the applicable EGSE requ Irements . 
Software will be provided to program the PMS for ASF requirements. 
The PMS will be designed to be integrated into the Shuttle mission 
simul ator . 


5.3. 2.4.2 Mechanical Ground Support Equipment (MGSE) 

5. 3. 2. 4. 2.1 General 

The ASF MGSE consists of operational equipment for the handling, 
transportation, servicing, measuring, aligning, and protection 
of ASF payload hardware instruments and experiment assemblies. 

The ASF payload is modular in its major subassemblies and therefore 
requires a highly flexible MGSE capability to acommodate all of 
the alternative configurations. The flexibility is provided by 
MGSE which includes the following. 

a. Support for each instrument to enable bench handling and local 
transportation (see figure 5. 3. 2-7). 

b. Support, with local mobility, for major subassemblies, experi- 
ments, and pallets. These support assemblies will be capable 
of interlocking for test and checkout purposes (see 

fi gure 5.3 .2-8) . 

c. Instrument protective/shipping containers. 

d. Servicing equipment to fill, leak check, and drain coolant 
loops of the thermal control system, provide adapters for 
nitrogen purge, and to provide experiment gas for instru- 
ments 304, 532 and 549. 


5.3-17 



6L- 


0 



Figure 5. 3. 2-8 


Concept of AMPS pallet handling frame 


t 


e. A matched rail assembly system with local mobility features 
which is capable of orienting experiments/pallets in any 
desired longitudinal pattern for checkout and/cr vehicle 
integration (reference ESTEC SLP/2104 - design coordination 
with ESRO in this regard is highly desirable). 

5. 3. 2. 4. 2. 2 Instrument/Instrument Cluster Support and Handling 
Equi pment 

Each instrument requires handling and servicing for test and 
checkout; therefore, a simple handling/interface assembly which 
permits positioning, alignment, functional test, and servicing of 
the instrument is required. The design of these handling fixtures 
provides that after the instrument is fabricated (or reworked), 
it will have minimal contact with human hands to minimize contami- 
nation during instrument/experiment test, checkout, and instrument 
cluster integration. 

5. 3. 2. 4. 2. 3 Instrument/Pallet Handling Equipment 

The MGSE for pallet level test and checkout is made up of matched- 
rail handling and servicing equipment. This enables connecting 
the pallets or integrated instrument assemblies to each other in 
any order or sequence required, and connecting to the E6SE check- 
out complex for integration tests (see figure 5. 3. 2-8). The 
matched-rail pallet handling dolly has steerable, removable wheels 
and is provided with hoisting/sling loops and slots for fork-lift 
handling from either side. Manual handling of a pallet/cart must 
be limited to movement over the floor with the handling dolly 
wheels installed. 

Pallet interface points are connected to and supported by MGSE 
support and lifting assemblies capable of supporting and orienting 
the pallet In both horizontal and vertical positions. 


Orientation of a cluster in horizontal and vertical positions 
is accomplished v/ith slings, hoists, and hydraulic positioners 
(see figures 5. 3. 2-9 and 5.3.2-10), The ASF payload is stabilized 
and supported during integration, Orbiter loading, and unloading 
operations. This is done with the hoist/sling assembly. Each 
segment (pallet) of the payload is spanned by interconnected struc- 
ture which is assembled or connected for support of any ASF pallet 
or payload configuration (figure 5.3.2-11). Note that during 
Orbiter/pal 1 et integration, the instrument cluster yoke will be 
retracted and the clusters will be in the launch positions. 

The MGSE design provides for complete integration of the instrument/ 
instrument clusters with the pallet(s) plus ASF payload integra- 
tion prior to installation and integration with the Orbiter. 

The instrument handling fixtures are designed of tubular support 
struts with angular interface frames to optimize the rigidity and 
strength of the assembly. This essentially single plane design 
provides, with the instrument attached, complete manual, sling/ 
hoist and laboratory cart handling capability for an instrument. 
Further, the design provides, sling/hoist attach points (rings), 
round end pieces (handles) which have soft molded plastic or 
rubber covers for manual usage, instrument interface attach po,.,ts 
compatible with instrument mechanical interface design, plus 
interface points on the bottom of the frame to mate with certain 
standard laboratory wheeled carts. See figure 5. 3. 2-7. This 
design is adaptive to adjustable configurations through the use 
of extension members. The adjustable design embodies handling 
capability through instrument cluster integration with the pallet 
instrument cluster support/positioning yoke. 


5.3-21 










5. 3. 2. 4. 2. 4 Vertical Payload Installatlon/Removal (Contingency 
Operation) 

Payload installation and removal from the Orbiter in the vertical 
position is a contingency operation. Hov/everj an ESRO concept 
has been reviewed and is considered acceptable, v/ith minor 
changes, for ASF use. See figures 5.3.2-12, 13, 14 and 15. 

5.3.2. 5 Logistics and Transportation 

Appendix E is a conceptual treatment of one set of fabrication and 
checkout flow requirements compatible with the section 5.3.2 
ground support and test philosophy which in turn was derived for 
the ASF pallet-only mode payload. These requirements are not only 
a baseline for ASF hardware logistics but a driver for the defini- 
tion of transportation requirements. 


5.3-25 



OBBITER IN 
VERTICAL POSITION 


ENVIRONMENTAL ENCLOSURE 
OVERHEAD SUPPORT RAIL 


PAYLOAD VERTICAL INSTALLER 


MANIPULATOR (DEPLOY OUTBOARD 
FOR PAYLOAD CLEARANCE) 


PAYLOAD INSTALLATION 
FIXTURE (GSE) 


FLOOR INSTALLER SUPPORT RAIL SYSTEM 


Figure 5.3.2-12. -AMP/ASF payload vertical handling concept 


5.3-26 






5 . 3-28 








r 

i 


5.4 SUPPORT SYSTEMS 

In addition to the ASF unique flight and ground systems described 
in paragraphs 5.2 and 5.3, respectively, the accomplishment of 
each mission depends on support from other systems not dedicated 
to the ASF program but in operation as part of the overall 
national space program. These include the Orbiter and ground 
facilities within the STS inventory, the TDRSS and a SPS 
which are being planned for operational status in the late 
197C‘S to early 1980's time period. 

5.4.1 ORBITER 

In the pallet-only mode, the ASF missions depend extensively on 
a number of Orbiter support facilities in the areas of structural/ 
mechanical, thermal control, avionics, and electrical power. In 
addition, Orbiter crew and vehicle operations support are required. 
The support functions required by the ASF payload are discussed 
in greater detail in paragraph 5.2 of this report. Details of tne 
Orbiter vehicle facilities and operational capabilities are 
provided in the "Space Shuttle Systems Payload Accommodations," 

JSC 07700, Volume XIV, Rev. C, July 3, 1974, and the "Space Shuttle 
Flight and Ground System Specification," JSC 07700, Volume X, 

Rev. A, January 2, 1974. 

5 . 4 . 1 . 1 Payload Placement 

The first ASF launch will be from KSC in a 28.8® inclination, 

400 to 500 km orbit. Polar orbit missions at the same altitude 
are also scheduled from the western launch facility. 

Placement accuracy required for ASF missions are expected to be 
well within baseline Orbiter capabilities. 


> 


5. 4. 1.2 Orbit Changes 

The PDS will be deployed at a distance of 1 to 10 km in front 
of the Orbiter at approximately the same altitude. The subsatel- 
lite will be ejected at 20 cm/sec and will continue to separate 
for the duration of the mission. At the end of tne mission, 
either the Orbiter or the subsatellite will provide the AV re- 
quired to rendezvous and retrieve the subsatellite. Both vehicles 
have AV capability. 

5. 4. 1.3 Attitude Control, Maneuvering and Pointing 

The Orbiter will be maneuvered to some preselected attitude and 
will maintain payload attitude within an accuracy of 1 or 2 
degrees. In this mode, the Orbiter will be operating at minimum 
deadband (0.1°) and at lowest rate (0.01°/sec). The ASF pointing 
system will decouple the instrument clusters installed in the AIM's 
from Orbiter motion and provide the precision accuracy required. 

During the mission, a number of maneuvers each revolution may be 
required to reorient the vehicle for the different experiments. 

5. 4. 1.4 Communi cati ons 

a. Direct with STDN. The maximum real time downlink data rate 
for unprocessed data (worst case condition) is approximately 
123 kbps. This condition exists only during the 15 minute 
bursts when Instrument 532 operates. During these periods, 
the Orbiter S band PM direct downlink system will be inade- 
quate, forcing use of the S band FM link, if the direct link 
is required. 

b. Relayed through TDRSS. The TDRSS S band PM downlink capa- 
bility is limited to 96 kbps. The ASF missions will use the 
TDRSS Ku band PM link which has the capability of transmitting 
up to 50 Mbps. The Orbiter baseline provides a Ku band 
antenna on one side of the fuselage. A second payload charge- 


5.4-2 


able antenna (available in kit form) will be required on the 
other side of the vehicle to provide near 4‘fr steradian 
coverage, 

c. PDS to Orbiter. The only communication link betv^een the 
Orbiter and deployed payload is the S band PM link. The 
capacity of this link is 16 kbps. 

5 . 4 . 1 . 5 T racki nq / 

a. Orbiter state vector (position and velocity). The ASF payload 
will utilize the baseline Orbiter one-way doppler tracking 
capability to establish Orbi ter/payl oad state vector. 

b. PDS tracking. Tracking of the PDS, v/hich will be deployed 
approximately 10 km from the Orbiter, v/ill be performed 
using the baseline Orbiter rendezvous radar. This microwave 
radar is capable of detecting and tracking a passive target 
at ranges between 10 km (5.4 n.mi.) and 30 m (IQO ft.) 

5. 4. 1.6 Data/Command Interface and Processing 

a. Engineering data to Orbiter performance monitoring system. 
Engineering data from the payload will utilize the 64 kbps, 
5“Channel interface provided by the baseline Orbiter. 

b. Scientific data for downlink transmission. Scientific data 
will utilize the 5.0 Mbps digital interface with the Orbiter 
FM signal processor for data being transmitted direct to 

the ground stations. The 50 Mbps digital interface with the 
Orbiter Ku band signal processor will be used for data being 
relayed through TDRSS. The interface with the MSS recorder 
is optional. 

c. Commands and data from the Orbiter. The ASF payload will 
interface with the 8 kbps Orbiter payload signal processor 
output for commands from the ground station (through the 
Orbiter rf link). The ASF payload will receive commands 
from the MSS and GN&C data through the Orbiter general- 
purpose computer and the multiplexer-demultiplexer. 


d. C&W, The payload C&W interface is with the Orbiter baseline 
C&W electronics (primary) and with the PMS (backup), 

e. Video. The ASF payload will utilize the Orbiter bay and 

RMS arm TV cameras proyided in the Shuttle baseline configura- 
tion. The two TV monitors with their associated camera and 
monitor controls at the on-orbit station will be used during 
experiment preparation, instrument pointing, accelerator 
operation, and subsatellite operations. 

f. Time codes. The mission elapsed time and GMT codes will be 
used by the ASF experiment and support subsystem computers 
and the timing displays at the PSS. 

g. ASF unique interface. The PSS interconnects with the ASF 
payload through the patch board. The patch board is located 
in the Orbiter junction box which provides the interface among 
the MSS, PSS and the payload through the station 14,630.4 
(576 in) bulkhead electrical connector panels. The ASF PSS 
will use the Orbiter audio system for voice communication with 
the rest of the Orbiter crew and with mission control. 

5.4. 1.7 Displays and Controls 

The ASF mission will utilize the Orbiter C&W annunciator at the 
forward station and the status board at the MSS. The CRT displays 
and keyboard control at the MSS which are part of the Orbiter CDMS 
will be used to monitor ASF instrument and support subsystems for 
health status, to perform failure diagnosis, and to control the 
deployed subsatellite. 

5 . 4 . 1 . 8 Remote Manipulator System 

The RMS will be used to retrieve the PDS at the end of the mission. 


5. 4. 1.9 Electrical Power 

As discussed in paragraph 5.2.2, the ASF payload will use the 
Orbiter dedicated fuel cell as its primary power source and the 
shared fuel cell as the secondary source in case of dedicated 
fuel cell failure. Two energy kits will also be required, 

5.4.1.10 Structural /Mechanical 

a. Payload attachment. The ASF program will utilize four of the 
ERNO designed, ESRO furnished standard equipment pallets to 
install instruments and support system equipment in the payload 
bay, (see paragraph 5.2.1 for details of pallet configuration 
and equipment installation). Each of the four pallets required 
for the ASF payload will be attached to the payload bay 
structure using the standard Orbiter primary payload structural 
attachment points as discussed in paragraph 5.2.1. No special 
provisions for these installations, or the vernier bridge 
fittings, are required. 

b. Payload bay cabling and fluid line accommodations. The ASF 
payload will require electrical cable break-out points for 
signal, command and power outlets from the wiring tray to the 
four pallets and the igloo. The Orbiter wire tray will provide 
harness routing from the pallets and the igloo to the forward 
bulkhead station Xg 14,630.4 (576 in), the aft bulkhead 
station Xg 33,197.8 (1307 in) and to the prelaunch (T-4) 
umbilical. 

Fluid lines are required from the launch (T-0) umbilicals 
through the aft bulkhead at station Xg 33,197,8 (1307 in) for 
the following functions: 

(1) Cryogen (LNe or LN 2 ) fill, vent, dump and relief. 

(2) Energy kit (LOg and LHg) fill, vent, dump and relief 
(part of Orbiter baseline). 


ff 


c. Payload service panels. The ASF payload will require access 

to the following service panels: 

(1) Station Xq 14j 63C (576 in) for signal, data, and 

command interface with aft crew station facilities. 

(2) Station Xg 33,197.8 (1307 in) for signal, data and 
command interfaces and fluids with the launch (T-0) 
umbilical . 

(3) Station Xg 17,653 (695 in) for primary and secondary 
electrical power interfaces. 

(4) Preflight (T-4) umbilical station Xg 20,688.3 (814.5 in) 
for functions (TBD). 

(5) Launch (T-0) umbilical, left and right side, station Xg 
26,444 (1435 in) to interface with the launch control 
complex for signal, data, commands, electrical power, 
ground cooling, and LHg and LO 2 fill, drain, and dump. 

Fill and drain for the active thermal control freon and 
fill, vent and drain interface provisions for the GNg 
for the subsatellite ejection system are (TBD). 

5.4.1.11 Thermal Control 

The ASF payload requires the use of the Orbiter ATCS to dissipate 
up to 24,000 Btu's of thermal energy per hour generated by the 
use of about 6.9 kW of electrical power over prolonged periods 
of time (e.g., from orbit revolutions 43 through 47), 

The baseline Orbiter ATCS coolant loop will provide the required 
support. Two additional heat radiator panels provided In a kit 
will be required. 


5.4-6 


5.4.2 TRACKING AND DATA RELAY SATELLITE SYSTEM 

The ASF mission will utilize the forv;ard and return link communi- 
cation services of the TDRSS. The single access (SA) return 
link is planned for the scientific data since the multiple 
access (MA) link capability (50 kbps) is inadequate to handle 
the 123.192 kbps ASF data rate requirement. However, due to 
possible priority scheduling problems, a combination of multiple 
and single access system usage may be necessary with selective 
transmission of continuously acquired data being provided by the 
low data rate MA system. The forward link can be on either MA 
or SA services. 

The standard tracking service will be utilized by the Orbiter to 
determine its position and velocity. The TDRSS v/ill not be used 
for the primary tracking of the PDS since this function v/ill be 
accomplished using the Orbiter microwave rendezvous radar. The 
TDRSS tracking capability could be utilized in a back-up mode if 
the TDRSS compatible transponder is incorporated into the subsatel- 
lite communication system. 

5.4.3 SOLAR PHYSICS SATELLITE (SPS) 

A SPS is scheduled for operational status in the late 1970 to 
early 1980 time period. The solar physics program is part of 
the overall scientific program for the NASA to investigate short 
and long term solar phenomena. The ASF missions which include 
solar measurements require data from the solar satellite to support 
experiments. 

ASF optical instruments would be used to calibrate the satellite's 
optical instruments. The most logical arrangement is for the 
Orbiter to have optical instruments which duplicate those on the 
satellite. Both sets of instruments would have the identical 
spectral sensitivity and would eliminate one source of error in 
the calibration. This calibration by the Orbiter instruments is 


5.4-7 


necessary because the responsivity of the optical instruments on 
the satellite would drift due to long term UV radiation, con- 
tamination, aging, and gamma rays affecting the optical surfaces. 

The instrument planned for this purpose in the baseline system 
is the Pyrhel iometer/Spectrometer, described as Instrument 1002 
in appendix B. The solar calibration instrument complement in the 
ASF payload cannot be truly defined until the instrumentation that 
will be used on the SPS has been specified. However, the physical 
size, weight, pointing requ i remen ts , and data sampling rate of 
the Pyrhel i ometer/Spectrometer make the instrument representative 
of the Orbiter instrumentation that will ultimately be selected. 

The Orbiter and the satellite need not be at the same altitude for 
optical calibration. If they are not, then atmospheric corrections 
must be made to the data before applying them to the calibration. 

It is sufficient for the satellite optical instruments that the 
orbit be sun-synchronous at a sufficiently high altitude to require 
little correction, if any, for atmospheric absorption (such as may 
occur for H and 0 atoms). Hov/ever, the particle detectors may 
need to be outside the earth's magnetic field, 25 or 30 earth radii 
from the earth in order to measure the particle flux and energy 
from the sun. The final altitude at which the satellite will 
operate is yet to be determined. 

The data required by the ASF program will be provided by the solar 
physics program. To support the ASF experiments, the sampling rate 
from the satellite is not critical. The data rate depends upon the 
number of wavelengths and energy intervals sampled. The particle 
detection instruments may sample more often, possibly as rapidly 
as once every four seconds. However, the optimum sampling rate 
can be found only by examining experimental data to determine how 
rapidly the changes occur. The data rate of the instruments on 
the Orbiter should be at least as great as that on the satellite. 


5.4.4 SPACE TRANSPORTATION SYSTEM (STS) GROUND FACILITIES 

The STS ground operational facilities include the OPF, the VAB, 
and the LCC at KSC and the v/estern launch site; mission controls 
at JSC; data monitoring and handling facilities at JSC, MSFC and 
GSFC; and the landing area facilities. Developmental ground 
facilities utilized by the ASF program will include the Shuttle 
Avionics Integration Laboratory (SAIL), the Mockup and 
Integration Laboratory (MAIL), the payload integration 
facilities, and instrument development and integration 
facilities. 

The unique GSE required for ASF payload checkout and test are 
discussed in paragraph 5.S 

The four ASF pallets and the igloo will be installed into the 
Orbiter in the OPF. The ASF ATCS will be installed and connected 
to the Orbiter payload heat exchanger. The coolant loop will be 
filled with freon and checked for leakage. The pumps and valves 
will be checked for operation. Fluid connections between the 
payload and the station Xq 33,197.8 (1307 in) bulkhead service 
panel will be made and electrical connection at stations Xq 

14.630.4 (576 in), 17,653 (695 in), 33,197.8 (1307 in) and at 
other harness breakout points will be made. Mated checkout 
requirements for the ASF payload at the OPF are yet to be assessed. 
The ASF unique GSE required to interface with the STS facilities 

is discussed in paragraph 5.2. 

At the VAB, after the Orbiter has been erected and mated to the 
external tank, ASF payload test and checkout will be performed. 

The details of the ASF payload checkout operations at the VAB have 
not been established. However, the same type of ASF unique GSE 
used for OPF checkout operations would be available to interface 
with the VAB STS facilities, if required. 


1 


The prelaunch operations support required by the ASF payload 
through the LCC includes verification of the engineering level 
operation of ASF support subsystems and as much of the instrument 
operations which might be practical at this time. After the 
launch readiness checks are completed, the subsatellite ejection 
system gaseous nitrogen tank is filled to a pressure of 2.41 x 10^ 
N/m^ (3500 psig). After the countdown process is initiated at T-2 
hours, the cryogenic coolant tank is filled with cryogen (LNe or 
LN 2 )- The electrical energy kit reactant tanks are filled with 
LO 2 and LH 2 as part of the Orbiter fuel cell reactant loading 
operations. During the countdown and through liftoff, the ASF 
payload data monitoring and operation checks are performed by the 
ASF unique GSE and the baseline LCC complex. 

The flight operations approach selected for ASF missions requires 
most of the processing and experiment operations to be performed 
onboard the Orbiter. Data not processed by the payload will be 
processed by the ASF ground facility. Most of the processing 
and controls will be performed automatically using the experiment 
and support subsystem computers. 

Using this approach, the support role for the MCC for the ASF 
payload is primarily to provide backup capabilities in the 
areas of experiment control and data processing analyses. It 
is expected that, as for most payloads, ASF unique stations 
will be required at the control center where engineering and 
scientific data monitoring will be conducted and non-safety 
related contingency decisions vnll be made. The stations will 
require support from the mission control complex for communications; 
data processing, stripping, storage, and routing; command and 
mission operational data (GN&C, attitude update maneuvering 
parameters, etc.) generation and other standard payload services. 

The ASF payload will not require special operations during the 
period immediately after Orbiter landing except for normal safing 


5 . 4-10 




1 


operations. Tapes and films from the recorders and camera will 
be removed and transported to the data handling facilities within 
a short time of landing. 

During the maintenance, refurbishment and repair cycle, the ASF 
payload will be removed from the Arbiter payload bay at the OFF 
and serviced in preparation for the next mission. 

5.5 CONTAMINATION 

5.5.1 INTRODUCTION 

This section addresses the STS environments and provides a treat- 
ment of salient environmental susceptibilities and radiative 
characteristics of ASF/AMPS instruments. Appendix C (4 parts) 
presents results from various supporting tasks undertaken during 
the study of this most important subject. 

5.5.2 STUDY APPROACH 

The contamination portion of this study started with a summary 
definition of the STS environment (appendix C-1 ) as derived from 
authoritative documents listed in paragraph 2. 3. 3. a. An environ- 
mental analysis (appendix C-2) was performed which treated the 
EMI and dust/gas/particulate contamination characteristics of the 
ASF/AMPS instruments. The analyses encompassed EMI controls and 
effects of contamination upon the payload instruments. Specific 
interference problems for ASF/AMPS instruments were identified 
and presented in apprndix C-3. The results of this study are in 
context with a June 1975 correspondence originated by the Space 
Shuttle Program Office (appendix C-4) which quotes an expected 
magnetic flux density of three orders-of-magni tude greater than 
the maximums specified for AMPS instruments. 


I 




5.5.3 CONCLUSIONS 

The complexities of the environments and preliminary nature of 
information relating to characteristics of both the advanced 
scientific instruments and the Arbiter itself, result in obvious 
major problems requiring extensive, in depth study and analysis 
before either quantitative definitions or solutions can be expec- 
ted. These same problems, in most cases, will confront pallet- 
mounted instruments comprising most scientific payloads presently 
conceived for the Shuttle era. The potential impact to Orbiter 
mounted scientific payloads presents important trade-off consider- 
ations relating to "cleaning up" the Orbiter design versus long 
range resource requirements to fortify the design of each instru- 
ment and payload subsystem to live with the Orbiter payload bay 
environment. Impact estimates, trade-off, and follow-on actions 
related to these environment problems will be addressed in para- 
graphs 8. 1.1,1 (Problems), 8.1. 1,2 (Impacts), 8.2 (Trade-Off 
Considerations), and 9.0 (Recommendations). 

5.5.4 RECOMMENDATIONS 

5. 5.4.1 EMI Contamination 

A number o" areas related to electromagnetic and electrostatic 
compatibility remain to be evaluated in depth after the ASF 
instruments and the supporting subsystems are further defined. 

Some of these are the following. 

a. Further definition of instrument susceptibility character- 
istics. 

b. Further definition of Orbiter, support subsystem and instru- 
ment EMI generation characteristics. 

c. Determination of the electrostatic potentials expected 
between the Orbiter and the surrounding plasma during particle 


5.5-2 



discharge. Determination of the impact on experiments of 
these potentials and evaluation of possible solutions, if 
requi red . 

d. Evaluation of the effect on Instrument electromagnetic 
contamination of the Orbiter multipoint structural dc power 
return system. 

e. Generation of a preliminary electromagnetic control plan. 

5. 5. 4. 2 Dust, Gas, Particulate Contamination 

The following information may be considered as potential 
approaches toward improvements in the overall Orbiter contamina- 
tion problem. 

a. The raising of lens temperatures a few degrees to inhibit 
condensation. 

b. The use of lens covers which are only opened shortly before 
sensor activation. 

c. The use of "shades" to limit the angle from which particulates 
may impinge on the lens surfaces. 

d. The use of clean inert gas to purge various areas periodically 
or create a slight positive pressure inside cameras, 

optical systems, etc. 

If contamination of radiative coolers by condensibles cannot be 
avoided, the contaminants can be removed by periodic cleaning 
through evaporation. Heaters for this purpose may be incorporated 
into the cooler design and operated on ground command. This 
approach has been used effectively on the Surface Composition 
Happing Radiometer (SCMR) on Nimbus E. 


Contamination control /avoidance on sensor optics may be achieved 
by a number of preventive designs and operational countermeasures 
The utilization of an Orbiter work shroud at the OPF during all 


I 




open cargo bay operations and the maintenance of rigid clean 
room standards will help greatly to reduce the major source of 
contamination. Materials selection will obviously be controlled 
by the Shuttle Project Office, and this will also help. 

For flight operations, there are a number of approaches which 
the Orbiter can use to limit contaminants. The most obvious is 
the control of Orbiter RCS/OMS thrusters, especially during periods 
when lens covers are opened. Without a high duty cycle use of the 
RCS, tight attitude control cannot be maintained by the Orbiter, and 
there might be some image smearing; however, this problem is not 
unique to the pallet-only mode configuration. 

Another critical operational method to avoid contamination is the 
avoidance and/or control of venting. Since it would be impractical 
to prohibit venting, appropriate control measures are required. 
Quantities can be made low and vent ports should be located away 
from critical areas and designed to provide high velocity, short- 
duration, directional flow. Designing tankage to provide minimum 
duty cycle is also desirable. 

5.6 STANDARDIZATION 

Appendix D is a detailed treatment of the feasibility of standard- 
dized electronics for the AMPS/ASF pallet-only mode payload. 

5.6.1 SUMMARY 

Using the current AMPS/ASF ID's, a typical system was defined. 

Using this definition, a study was made to determine if it was 
feasible to remove from the individual instruments the support- 
ing electronics and combine those for several instruments into 
a single instrument electronics package (lEP). The study also 
considered the cost savings that might be realized by standardiz- 
ing and modularizing the sensor support electronics within the 
lEP. It further considered the possibility of utilizing existing 
packaging methods such as NIM-CAMAC, Navy QED, Navy SHP and ATR. 


5.6-1 


This study made a special application of this centralizing, 

standardizing, modularizing concept to ASF. 

5.6.2 CONCLUSIONS 

a. It is feasible to apply the centralized instrument electronics 
concept to AMPS/ASF, 

b. Considerable cost savings can be realized in the areas of 
design, fabrication, and test. 

c. Standardizing and modularizing the lEP offer a number of 
advantages in the areas of repl aceabi 1 i ty , maintainability, 
system expansion, and preflight checkout. 

d. Although the concepts are good, the use of existing standard- 
ized electronics such as NIM-CAMAC, Navy QED, ATR, and others 
are not acceptable v/ithin the existing constraints for space- 
flight type experiment hardware. This is largely because of 
the constraints on weight and volume and also the existing 
packaging technique is not suitable for spaceflight use. 


/i 



6.0 MISSION OPERATIONS 


6.1 INTRODUCTION 

This section describes the ground and flight operations required 
to support ASF missions. These missions include both low in- 
clination and polar flights to provide global coverage and to 
provide flight paths both parallel and normal to the earth's 
magnetic field. 

The first launch will be into a 28.8° inclination orbit from KSC 
ii! 1981. Succeeding launches from the eastern launch facility 
(KSC) will be at orbital inclinations from 28.8° to 67°, depend- 
ing upon individual mission requirements. Launches from the 
western launch facility sometime after it becomes operational 
will inlude 90° inclination (polar) orbits. Orbits will be at 
altitudes between 200 and 500 km. 

The mission operations will be supported by the prelaunch check- 
out facilities at the launch site (both ASF dedicated and part 
of the STS), the launch pad checkout facilities, the Orhiter 
vehicle and crew, the STDN and TDRSS, the MCC at JSC, the ASF 
dedicated ground data handling and processing facility, the 
landing site facilities, and the ASF payload refurbishment and 
modification facilities. In addition, supporting data from a 
PDS deployed from the Orbiter and a SPS already in orbit will be 
required to accomplish the ASF missions. The in-flight mission 
system support facilities and the fundamental interfaces re- 
quired are illustrated in figure 6.1-1. 

6.2 GUIDELINES AND ASSUMPTIONS 

The following guidelines and assumptions were utilized for the 
ASF study. 

a. Normal Orbiter operation will be nose up and nose down 
Z/velocity vector (ZVV) and Y/perpendicul ar to Orbiter 
Plane (Y-POP) for least fuel consumption (figure 6.2-1). 


<®IGIWAL PAGE IS 
OP POOR QUAIOT 







•f 


b. For each ASF mission, the Orbiter vehicle will be dedicated 
to the ASF payload. 

c. ASF instrument pointing will be accomplished primarily from 
the Orbiter attitude shown in figure 6.2-2. Maneuvers from 
this position v/ill be required to accomplish some experiment 
objectives. 

d. SPS instrument data will be available at the time the ASF 
missions are ^lown, and the solar instruments on the Orbiter 
will be used )r calibration only. 

e. The identical instrument/experiment complement will be flown 

in polar as well as 28.8° inclination orbits. Time in attitude 
hold angle will be limited for Beta angles greater than 60°. 

6.3 OPERATIONS DESCRIPTION 

For each mission, the operations required will include the 

fol 1 owi ng . 

6.3.1 PREFLIGHT OPERATIONS 

a. Pallet level integration and test. 

b. Payload level integration and test (SAIL). 

c. Vehicle integration, test and launch preparation (launch 

base facilities). 

6.3.2 FLIGHT OPERATIONS 

a. Launch and mission orbit injection. 

b. Orbi ter/payl oad systems verification and experiment prepara- 
tion. 

c. Deploy subsatellite. 

d. Perform experiments. 

e. Retrieve subsatellite. " 

f. Re-entry preparation. 



6-4 




g. Re-entry, descent ana landing. 

6.3.3 P0STFLI6HT OPERATIONS 

a. Onboard stored data recovery, 

b. Data processing. 

c. Distribution of data to users, 

d. Payload refurbishment, change and preparation for storage 
or subsequent flight. 

e. Logistic support. 

6.4 PREFIIGHT OPERATIONS 

6.4.1 PALLET LEVEL INTEGRATION AND TEST 

Subsequent to delivery of the instruments, PDS, and the sub- 
system equipment to the ^allet integration facility, the instru- 
ments and the PDS will be installed into handling fixtures. 

These handling fixtures (see paragraph 5,3.2 for further details) 
will allow instrument and PDS positioning, alignment, functional 
test and servicing with minimal manual contacts, thereby re- 
ducing possible contamination. 

The instruments (or PDS) and support equipment will be installed 
directly onto the pallet in the case of pallets A-2 and A-4. 

For pallets A-1 and A-3, the instruments will be installed into 
the AIM's and the AIM's will be mounted on the APS. Support 
equipment for pallets A-1 and A-3 will be installed directly to 
the pallets or, in the case of the star tracker and sun sensor 
assemblies, on the AIM. Mechanical alignments will be checked 
and adjusted as necessary using optical alignment tools for the 
critical alignments. 

One or more pallets with the complement of instruments (or PDS) 
and support subsystems (pallet packages) will be installed in 


the matched-rail handling and service fixture (see paragraph 
5.3.2), The pallet packages will be mated with the EGSE de- 
scribed in paragraph 5.3.2 and comprehensive test and checkout 
will be performed using simulators to provide compatible support 
subsystem and Orbiter Interfaces. The tests will be automati- 
cally controlled through the EGSE computer and test software. 
Power, signal, and data interfaces will be verified; comple- 
mentary operation of instruments will be tested for compati- 
bility; and the operation of the integrated pallet package will 
be checked for compatibility. 

The precision of many of the instruments and the measurement 
thresholds or the sensitivity of instrument operations are such 
that under ground level environments, verification of instrument 
accuracy or operational capability may not be possible. The 
operation of these instruments will be checked, to the accuracy 
level possible, for functional compatibility and to verify 
that gross malfunctions have not occurred. The support sub- 
system equipment on the pallet including the APS will be 
checked for both functional operation and to verify in-limit 
performance. 

After the pallet level tests are completed, the pallet (or 
pallets) together with the matched rail handling fixture will 
be installed into an enclosed transporter. The transporter 
will then be placed in temporary storage or shipped to the 
payload integration facility. 

6.4.2 PAYLOAD LEVEL INTEGRATION AND TEST 

If pallets are shipped individually, the four pallets comprising 
the ASF payload will be assembled at the payload integration 
and test site. The SAIL will be utilized for electrical and 
electronic integration verification. The MAIL will be used 
to verify the mechanical interfaces. The full complement 
of ASF EGSE will be available to 


6-6 


be used at either integration laboratory, although the SAIL 
will require little additional ASF unique GSE. 

Tests and checkout at the payload integration facility will be 
performed to verify: (1) compatibility among the different 
elements of the integrated payload, and (2) thermal, fluid, 
mechanical, electrical and electronic compatibility of the 
integrated payload with the Orbiter, As in the case of the 
pallet level test and checkout, precise verification of in- 
strument accuracy or operational capability may not be possible 
at the integrated payload level. The same test approach 
utilized at the pallet level will apply at the integrated pay- 
load level. 

After integration and tests are completed at the payload inte- 
gration site, the payload will be loaded into the same type 
of enclosed transporter used for the individual (or combined) 
pallets and the transporter will either be placed into storage 
or will be shipped to the launch site. 

6.4.3 VEHICLE INTEGRATION, TEST AND LAUNCH PREPARATION 

A^ the launch site, the transporter will be delivered to the 
OPF. When the Orbiter is ready for payload installation, the 
pallets (and the handling dolly) will be removed from the trans- 
porter and inspected for possible damage. An integrated payload 
test will be performed using the ASF dedicated EGSE to verify 
operational integrity of the instruments and the support equip- 
ment. After tests are completed, the pallets will individually 
be installed into the Orbiter payload bay using the slings, 
hoists and hydraulic positioners discussed in paragraph 5. -3. 2. 

The pallets will be attached to the payload bay structure using 
the standard Orbiter attachment provisions (see paragraph 5.2). 
Mechanical alignment checks will be made, electrical connections 
will be made and verified. Fluid and gas lines will be connected 



and pressure checks will be made using helium or other inert 
gases to verify pressure integrity of the lines, valves and 
containers. After the lines are purged and cleaned, the ATCS 
coolant loop will be filled with freon and checked for leakage. 
The pump and valve operations wilT be verified. The pyrotechnics 
for latch actuation and boom and platform jettison will be 
installed and made safe using safe plugs or electrical safeing 
circuits. 

Using the ASF EGSE, a payload functional test similar to those 
conducted at the pallet and integrated pay! oad . 1 evel s will be 
performed. 

The Orbiter, with the ASF payload installed, will be transported 
to the VAB and will be erected and mated to the external tank 
and the solid rocket motors. T[«e integrated Shuttle/payl oad 
will then be transported to the launch pad. At tlie launch pad, 
with the Shuttle in position, electrical and fluid interfaces 
with the launch complex will be made at the T-0 and T-4 umbili- 
cal connections. 

A Shuttle system launch readiness test wilT be conducted to 
verify all Shuttle onboard and ground interfaces using command 
and test controls from the LCC. All payload functional 
electrical interfaces with the Orbiter will be verified and 
the payload and support subsystem computers and mass memory will 
be loaded with the final flight programs. The programs, as 
loaded into the computers, and the ma|s memory will be verified 
through memory dumps and a simul ated fl i ght sequence will be 
performed using the onboard computers and the mass memory. 

After these 1 aunch/ readi ness checks are compl eted, the, pyro- 
technic safe plugs will be removed {after the circuits are 
reverified to be in the safe conditions). The subsatellite 


ejection system gaseous nitrogen tank will then be filled. The 
Orbiter payload bay doors will be closed at 1-2 hours and the 
final countdown process will then be initiated. During the final 
countdown phase, the cryogenic coolant tanks will be filled 
with cyrogen and the electrical energy kit reactant tanks v/ilT - 
be filled with L02 and LH2 as part of the Orbiter fuel cell 
reactant loading operations. Power to the Orbiter will then 
be switched from the external ground support source to the 
i nternal ‘ fuel cell s . 

6.5 FLIGHT OPERATIONS 

6.5.1 LAUNCH AND MISSION ORBIT INJECTION 

During launch and until the Orbiter is in its operational orbit, 
the payload will basically be passive although the ATCS will be 
operational and the C&W parameters will be displayed onboard 
the Orbiter at the MCC. 

6.5.2 ON-ORBIT OPERATIONS 

The on-orbit operations will be separated into 5 phases. These 
will be: (1) payload preparation; (2) subsatellite deployment; 

(3) perform experiments; (4) subsatellite retrieval; and (5) re- 
entry preparation. 

The ASF program approach for control of payload operations and 
for data processing will be to provide as much of these opera- 
tions as possible through the payload or Orbiter systems. De- 
pendence on ground stations will be minimized. Therefore, 
throughout the. mission, the role of the MCC for payload opera- 
tions control will be primarily one of backup. Also, data 
processing will be limited to that required to display the 
downlinked data at the MCC stations since an ASF dedicated 
ground data handling facility is currently planned. 


6-9 


Although tfi'e mission of the SPS will be independent of the ASF 
mission, coordination between the JSC and GSFC for the operation 
of these two systems in orbit will be required. The data 
obtained from the ASF payload and the deployed ASF PDS will 
be used to calibrate the instruments onboard the SPS and the 
data obtained from that sateTl ite. These coordination factors 
will be established during subsequent ASF studies. 

The on-orbit operations are shown in the ASF mission timelines 
presented previously in figure 4. 1.5-1. 

6 . 5 . 2 . 1 Payload Preparation (Revolutions 1 through 10) 

After the Orbiter has achieved orbit insertion, the payload 
bay doors will be opened and the Orbiter system will be pre- 
pared to support the mission. 


The ASF support subsystems, which have been powered through 
the launch and ascent phase, will be verified for mission 
readiness and safety. Power to the APS gimbal torque motors, 
the instruments and the subsatel Tite wil 1 be applied with the 
instruments in the standby mode. The cryogenic coolant 
systems for instruments 118 and 126 will be activated. After 
a short warm-up period (5 to 10 minutes), the APS and all in- 
struments except instruments 118 and 126 will be checked to 
verify readi ness -status . After the temperature of the "detectors 
for instruments 118 and 126 have stabilized (approximately _ 

10 hours after coolant system is activated) the operational 
status of these two instruments will be verified. 

By orbit revolution 10, all verification checks will be com- 
pleted and the payloaLd^wil 1 be ready for operations. 


PDS Deployment (Revolutions 11 through 15 


Between Orbit revolutions 10 and 15, the Orbiter will adjust 
its orbit, if required. The Orbiter will be maneuvered to the 
desired attitude for deployment and inertially stabilized to 
that attitude, during the PDS deployment operations. 

On or about revolution 15, the PDS deployment sequence will 
be initiated by the MS. Subsequent deployment operations 
will be controlled by the subsystem support computer. 

The ejection system will be armed, preparing the GN 2 system for 
ejection operations. The command to eject the PDS will be manual 
and will actuate a solenoid pilot valve which will introduce 
the gas into a cylinder bore containing a piston. Under the 
action of the gas, the piston will move and the movement will 
allow the PDS holding mechanisms (collets) to unlatch the PDS, 

The piston will continue its movement until a striker fixture 
attached to the piston rod impacts the PDS and imparts to it 
a separation velocity of about 20 cm/sec. Teflon guide rails 
will be used to assure liftoff in the desired direction, 

Details of this ejection system are provided in paragraph 5,2 .1 . 

Subsystem support will be provided before, during and after PDS 
separation. Prior to ejection, the PDS commu’^ication system 
will be checked through hardline connections with the Orbiter. 
Operational data will be processed through the experiment RAU's 
on Pallet A-2 and the experiment computer in the igloo before 
the data is displayed at the aft crew station. Commands will 
be programmed by the subsystem computer in the igloo and trans- 
mitted to the PDS ejection system through the subsystem RAU 
located on Pallet A-2. Electrical power will be provided to 
the PDS from shortly after Orbiter insertion into orbit to 
PDS separation. 






The PDS will be a modified AE satellite as discussed in paragraph 
5.2,6. After separation from the Orbiter, the PDS thrusters will ^ , 

provide it with a spin rate of one revolution per minute (rpm) 
for spin stabilization. During this period the PDS communication 
link with the Orbiter will be verified. After about 14 hours the 
PDS will be at a separation distance of 10 km from the Orbiter 
and thrusters will be actuated automatically by the PDS control 
system to reduce its velocity relative to the Orbiter. The PDS * 

will be reoriented to the desired attitude and will go into a 
stationkeeping mode at this distance from the Orbiter. 

All operations on the PDS subsequent to separation (except sta- 
tionkeeping velocity changes) will be autonomous although 
control can be exercised either from the ground stations or the 
Orbiter through the respective rf communication links. Orbiter 
control capability, if required, will be provided at the MSS. 


After the PDS has been stabilized in the stationkeeping mode, | 

it will begin to transmit data from its instruments (and support 

systems) to the Orbiter. | 

i 

Payload attitude initialization and update will be performed -j 

during the latter stages of this phase. Commands v/ill be given I 

to the star tracker assemblies to start the star search. Since \ 

the tracker reference axes orientation will be known to within 
^2°j the star search and recognition processes, and alignment 
of the APS, will be completed within a few minutes. The star 
angle data will be processed by the payload subsystem computer 
to align the APS inertial reference system. 

Subsequent to this attitude initialization, updates will be 
required at least every 1-1/2 revolutions during the conduct 
of the experiments. 


6.5. 2.3 Experiment Operations (Revolutions 16 through 8Q) 

The ASF experiments will be conducted during a period of about 
4 days (64 revolutions). During this entire period, the 
support subsystems including the APS, the ATCS, and the cryo- 
genic cooling system will be operational. 

6 . 5 . 2 . 3 . 1 Revolution 16 

Instrument 116 (Airglow Spectrograph) will be operated inter- 
mittently to study upper atmospheric emissions and absorptions 
and Instrument 1002 (Pyrhel iometer) will operate for about 
15 minutes during mid-daylight to measure the solar constant, 
the solar spectral irradiance, and to determine possible varia- 
tions of total and spectral flux associated with changes in sun 
radiation. Instrument 1002 is used to calibrate solar instru- 
ments on the SPS . 

6. 5. 2. 3. 1.1 Instrument 116 

The frequency of operation of Instrument 116 will depend on the 
occurrence of discrete phenomena such as aurorae, by the exist- 
ence of observable conditions such as noctilucem: clouds, and 
the frequency of data required for normal day-glow and night- 
glow studies. The spectrogram exposure times will range from 
1 second to 1000 seconds. 

The subsystem support required will include the following. 

a. Pov/er, power control for standby and operate modes. 

b. Instrument pointing using the APS to within 0.5°. 

c. Computer controls to shift instrument collimating mirrors 
into and out of the energy path. 

d. Exposure control to control the start and duration of the 
spectrogram exposure. 


e. Displays for indication of the optical configuration of the 
instrument, relative pointing angle, indication of spectro- 
gram exposure completion and indication of the exhaustion 
of the film supply. 

f. Operational status monitor displays. Inflight calibration 
of Instrument 116 is not planned at this time. 

6. 5. 2. 3. 1.2 Instrument 1002 

This instrument will be pointed at the sun using the APS, The 
instrument has a door covering the opening for the optical input 
The door will be opened before the data take and closed after 
the measurements are made. Scan frequency will be 2 or 3 times 
during this revolution and scan time will be 10 minutes, A 
light source will be used during flight to calibrate the in- 
s trument. 

Subsystem support will include the following. 

a. Power and power control . 

b. Door opening and closure control. 

c. Data sequence control. 

e. Displays indicating control activation. 

f. Data processing for data taken at one sample per minute 
for the Pyrheliometer and 270 pairs of samples per minute 
for the Spectrophotometer. 

g. Operational status monitor displays. 

h. APS pointing to within j^2. 5° of the sun line. 


6 . 5 . 2 . 3 . 2 Revolutions 17 through 31 


During revolutions 17 through 31 the following instruments are 

operated. 

a. . 118 “ continuously scanning with scan periods between 40 and 

66 seconds to measure trace gas concentrations in the 
spectral range of 3 to 40 urn. 

b. 122 - continuously measuring atmospheric and ionospheric 
gaseous spectral emissions and absorptions in the range of 
1100 to 10,000 A. 

c. 1 24 “ continuously cycling at 2 minute intervals for dura- 
tions up to 100 seconds. The instrument will be used to 
measure spectral or photo emissions in the range of 0.2 ym 
to 1 0 ym in the stratosphere, mesosphere and thermospheraf. 

d. 126 - continuously acquiring data in the 1 to 150 ym spectral 
region. 

e. 213 - continuously cycling at a rate of one pulse per 
second. The laser will operate over the spectral range of 

o o 

1000 A to 30000 A and will be used with other instruments 
to study the composition, structure and dynamics of the 
atmosphere through backscatteri ng and absorption of the 
laser beam. 

f. 532 " once for 15 minutes near niid-dayl i ght during revolu- 
tions 17 through 25. The instrument will be used to release 
gases and to monitor orbital and solar effects on these 
gases . 

g. 1011 - twenty or more exposures during 20 minute scans of 
earth limb during solar occultation. The instrument mea- 
sures solar energy absorption by certain molecules and 
free radicals at different altitudes above the earth. 


6^15 



6 , 5 . 2 . 3 . 2 , 1 Instrument 118 

The Limb Scanning IR Radiometer will scan the atmosphere 
vertically from the horizon upwards to 120 km and normal to 
the vertical up to 10° on either side of the nominal position 
using the APS. 

The scanning operation will be provided by a preprogrammed 
routine in the subsystem computer which will drive the APS, 
Either the computer program or the crew manual control will 
provide the alignment of this instrument which will be on AIM 
3A with Instrument 124, on AIM 3B with Instrument 213, and 
on AIM lA when co-alignment operations are required, 

a. Controls will include: 

(1 ) Power control , 

(2) Selection of scan rate and scan angle, 

(3) Selection of scan mode (sawtooth, zig-zag, sinusoidal) 

(4) Inflight calibration using internal black body source 
and space background, 

(5) Control of instrument temperature using cryogenic 
cooling system. 

b. Displays required will include: 

(1) Scientific data from 4 to 12 spectral channels. 

(2) Verification of selected scan rate and angle. 

(3) Verification of scan mode. 

(4) Relative pointing angle of radiometer. 

(5) Detector temperature and bias voltage. 

(6) Telescope temperature. 

(7) Status of cr^ogen supply, 

(8) Instrument operational status monitor. 



6, 5. 2. 3. 2. 2 Instrument 122 


The Ultraviolet-Visible-Near Infrared Spectrometer will be 
operated in both a fixed orientation mode and in an earth 
limb scanning mode. Tracking and scanning control will be 
provided by the APS, Covers will be used to protect the in- 
strument optics from contamination when possible. 

a. Controls will include: 

( 1 ) Power control . 

(2) Selection of spectrometer mode. 

(3) Selection of grating scan rate. 

(4) Selection of wavelength to be measured. 

(5) Instrument pointing using the APS to within ±0,017“ 
of the Instrument 1011 reference axes. 

b. Displays will include: 

(1) Verification of scan ratej scan mode and wavelength. 

(2) Detector counts as function of integrated time, 

(3) Relative pointing angle of instrument. 

(4) Instrument operational status monitors. 

6. 5. 2. 3. 2. 3 Instrument 124 

The Fabry-Perot Interferometer will operate in two modes; limb 
scanning (during even-numbered revolutions) and nadir scanning 
(during odd-numbered revolutions). During limb scanning opera- 
tions, the instrument will scan the earth from side to side 
from the tangent point to a depression angle of about 20“ below 
the tangent point. The scanning is provided internally to the 
instrument through a scan-dri ven pi anar mi rror . However, the 
initial pointing is provided by the APS. Nadir scanning will 
occur when operating in conjunction with the Laser Sounder 
(Instrument 213) to measure resonance backscatter energy. For 


this operatiorij the instrument must be co-aligned with Instru- 
ment 213 to within 1 milliradian using the respective APS on 
pallets A-1 and A-3 and the optical transfer system between 
the two pallets to transfer the attitude reference. 

Inflight calibration of this instrument will be performed 
throughout the mission using integral spectral /radi ance calibra 
ti on 1 amps . 

a. Controls will include: 

( 1 ) Power control , 

(2) Selection of operating modes (interferometer, photo- 
meter or infrared photometer). 

(3) Selection of scan rate. 

(4) Instrument pointing using APS to wi thi n +0 . 06° of the 
Instrument 213 reference axes. 

b. Displays will include: 

(1) Scientific data (interferometer diagnostics and 
parameters for plotting intensity versus wavelength). 

(2) Verification of operating mode and scan rate. 

(3) Photomultiplier tube power supply voltage and dark 
current calibration data. 

(4) Detector temperature. 

(5) Relative pointing angle of instrument. 

(6) Integration time. 

(7) Instrument operational status monitors. 

6 . 5 . 2 . 3 . 2 . 4 Instrument 126 

The Infrared Interferometer will operate in two modes; limb 
scanning and nadir scanning. The scanning function will be 


provided by the APS. During operation with the Laser Sounder 
{Instrument 213), the two instruments must be co-aligned to 
within ±0.1°. Since Instrument 126 will be in AIM 3B and 
Instrument 213 will be in AIM lA, the optical transfer between 
pallets A"1 and A-3 will be used to determine the attitude 
alignment. Instrument 126 viill require cryogenic cooling of 
the detector to 4 K. A common set of storage tanks will supply 
make-up cryogen to both instruments 118 and 126. 

Internally provided black body sources and the space background 
will be used for in-flight calibration. 

a. Controls will include: 

(1 ) Power control . 


(2) 

Selection of scan rate. 


(3) 

Selection of scan angle. 


(4) 

Duration of data take. 


(5) 

Inflight calibration. 


(6) 

Control of instrument temperature using 
cooling system. 

cryogeni c 

(7) 

Initial pointing of instrument using APS 

to within 


±0.1° of instruments 213 and 118 reference axes. j 

b. Displays will include: 

(1) Spectrogram of observed data. 

(2) Verification of spectral range. 

(3) Verification of scan rate and angle. 

(4) Relative pointing angle of instrument. 

(5) Detector temperature and bias voltage. 

(6) Instrument internal temperature. 

(7) White light in terferograms (for calibration). 

(8) Instrument operational status monitors. | 


6-19 


6. 5. 2. 3. 2. 5 Instrument 21 3 


The Laser Sounder will operate similar to a radar in which the 
laser beam will be directed towards the atmospheric mass under 
observation, generally toward the nadir using the APS. The 
receiver section will measure the reflected ( bac kscatter ) 
energy. The laser will be operated on both the dark side and 
the daylight side of the earth in conjunction with instru- 
ments 118, 124 and 126. Some question remains as to the effec- 
tiveness of the laser operation during the daylight, and this 
will be further assessed during the next phase of study. 

a. Controls will include: 

(1) Power control. 

(2) Selection of wavelength to be emitted. 

(3) Selection of pulse width and repetition rate. 

(4) Instrument pointing using the APS to within ±0.1° of 
the reference axes of instruments 118, 124 and 126. 

b. Displays will include: 

(1) Indication of receipt of backscatter energy. 

(2) Indication of pulse height and duration. 

(3) Laser head temperature. 

(4) Relative pointing angle of instrument. 

(5) Instrument operational status monitor. 

6. 5, 2.3. 2. 6 Instrument 532 

The Gas Release Module will admit gases into the excitation 
chamber, and the gases will be elevated to an excited state by 
exposure to the solar flux introduced into the chamber. The 
excited gases will be observed by a monochromator. The ions 
produced in the chamber will be analyzed by a mass spectro- 
meter. Gases will also be released into space and analyzed by 
the monochromator. 


The LOS of the sun sensor on the excitation chamber will be 
pointed to within ±1° of the sun line such that sun sensor 
will be able to acquire the sun and control the reflection- 
of the solar radiation into the chamber. 

The monochromator will be calibrated in flight using a special 
light source attached to the system. 

a. Controls will include: 

(1 ) Power control . 

(2) Selection of gas. 

(3) Selection of gas release mode {chamber or space re- 
1 ease) . 

(4) Gas rel ease . 

(5) Control of gas pressure./ 

(6) Monochromator grating control. 

(7) Mass filter control. 

(8) Instrument sun sensor pointing to within ±2.0° of sun 
line using APS . 

b. Displays will include: 

(1) Verification of sun acquisition. 

(2) Veri -Pi cation of selected gas and gas mode. 

(3) Gas system pressure. 

(4) Chamber pressure. 

(5) Chamber photodiode signal . 

(6) Chamber temperature. 

(7) Monochromator intensity versus, wavelength. 

(8) Mass filter ion intensity versus mass count. 

(9) Mass filter rf voltage. 

(10) Instrument operational status moni tors . 



6. 5, Z, 3. 2, 7 Instrument 1011 


The Ultraviolet Occultation Spectrograph will he pointed at 
the sun through the earth's atmosphere. The initial LOS will 
be at altitudes about 100 km or more above the point of tangency 
with the earth. The altitude will decrease as the Orbiter 
makes its revolution. A sun tracker will be used integral with 
this instrument to provide the control signal for sun tracking 
by the APS. Exposures will start just before detectable ab- 
sorption. Ten or more one-second exposures will be made until 
the data is rendered useless by the reduced tangency altitude 
or the excessive absorption. 

Infl i ght cal i brati on wi 1 1 be achi eved peri odi cally using a 
calibrating source in front of the small telescope. 

a. Controls will include: 

(1 ) Power control . 

(2) Exposure control and sequencing. 

(3) Opening and closing of protective door. 

(4) Fi 1 m advance . 

(5) Calibration control. 

(6) Initial pointing or instrument to allow sun sensor to 
acquire the sun. 

b. Di spl ays. wi 1 1 include: 

(1) Verification of door position. 

(2) Sun acquisition. 

(3) Verification of calibration source position and verifi- 
cation that it is on. 

(4) Film frame count. 

(5) Exposure timing. 

(6) Instrument operational status monitors. ' . , I 


6. 5. 2. 3. 3 Revolutions 32 through 47 

During this span, the payload will operate as follows. 

6. 5. 2. 3. 3.1 Instrument 116 

This instrument will function in the same manner as it had during 
revolution 16 except that it will be operating in support of 
Instrument 303 (revolutions 32 to 42) and Instrument 304 (revolu- 
tions 43 to 47). The instrument will observe the effects of the 
accelerated particles on spectral emissions from the elements 
of the Upper atmosphere. Subsystem support, control of the 
instrument and the displays required will be the same as that 
required during revolution 16 operations. 

6. 5. 2. 3. 3. 2 Instruments 118, 126, 213 

These instruments will operate as they did during revolutions 
17 through 31 except that operations will be at standby for about 
50 minutes on the dark side of the earth. This period will 
begin about 17 to 18 minutes before the accelerators start to 
operate and will continue for 17 to 18 minutes after the ac- 
celerators are turned off. 

The support subsystem operation will continue as before except 
processing and display of scientific data will not be required. 
Computer controls required will be the same as for revolutions 
17 and 31 except that switching the instrument to and from the 
standby, ^mode will be required. 

6 . 5 . 2 . *3 . 3 . 3 Instrument 303 

The Electron Accelerator will operate for 15 minutes each 
revolution, during revolutions 31 through 42, while on the 
dark side of the earth. At all other times, the instrument 
will be on standby status . 


6-23 


The accelerator win operate in a continuous dc, pulsed or 
modulated mode. When pulsed, the repetition rate and pulse 
duration will vary such that the duty cycle remains at 5 per- 
cent at maximum power. In the modulated mode, the amplitude 
of beam energy will vary from 0 to 100 percent at a frequency 
of 0 to 10 MHz with a .5 percent duty cycle at maximum power. 

When energized, the accelerator will provide a beam of electrons 
with energies between 1 keV and 30 keV which will be used in 
conjunction with instruments 116 and 534 to study the excita- 
tion of the upper atmosphere and ionosphere elements, to map 
the magnetic field lines of the earth, to determine ionospheric 
electric field magnitude and direction, and to study plasma 
wave excitation in the ionosphere. 

The instrument pointing requirement (< 2° error) will be pro- 
vided by the Orbiter. attitude control system. 

a. Controls will include: 

(1) System power control. 

(2) Accelerating voltage control, 

(3) Control grid voltage and frequency control,. 

(4) Diverging and converging lens voltages control. 

(5) X-Z and Y-Z sweep coil voltages control. 

(6) Control interlock with Triaxial Fluxgate (Instrument 

536) to prevent accelerator operation when direction 
of earth's magnetic field could cause beam return. 

b. Displays will include: 

(1) Power unit output voltage and current amplitude and 
wave shape. 

(2) Acceleration voltage and current amplitude and wave 
shape . .. . 


6-24 


(3) Grid current amplitude and wave shape. 

(4) Accelerator operational status monitors. 

6.5.2. 3. 3.4 Instrument 304 

The MPD Arc will operate for 15 minutes each revolution while 
on the dark side of the earth during revolutions 43 through 47. 
Simultaneous operation of instruments 303 and 304 are not 
planned because of the thermal dissipation constraint of the 
Orbiter ATCS ... 

The MPD Arc will operate in a pulsed mode with the pulse dura- 
tion and rate controlled to. keep the. power drai n on the Orbiter 
supply below 10 kW. 

Instrument 304 is a plasma accelerator which will discharge 

5 

currents up to 2 x 10 amperes. It will be used to study the 
exeitation of the upper atmospheric and ionospheric elements, 
to trace and map the earth's magnetic field lines, to modify 
ionospheric conductivity in certain regions, and to generate 
plasma waves in the very low to extreme low frequency regimes. 

The pointing requirement for this instrument (< 2“ error) will 
be provided by the Orbiter attitude control system. 

a. Controls will include: 

(1) System power control. 

(2) MPD Arc pi enum pressure control . 

(3) Solid state switch (for high voltage) control. 

(4) Interlock control with Magnetometer (Instrument 536) to 
prevent beam return due to direction of earth's magnetic 
fiel d 1 ines. 


b. Displays will include; 

(1) Discharge current and voltage pulse amplitude and 
waveforms. 

(2) MPD Arc operational status monitors. 

6. 5. 2. 3. 3. 5 Instrument 532 

The operation of the Gas Release Module during this span will 
be concurrent with the operation of the accelerators (in- 
struments 303 and 304). The mode.vn’ll be used to release the 
gases into space and this will occur during the dark phase of 
the orbit rather than during the daylight side. The mode for 
gas release into the excitation chamber will not be used during 
this phase and therefore the sun sensor and the spectrometer 
will not be required. All other operations will.be the same 
as those described during revolutions 17 through 25. Controls 
and displays will not include those associated with the gas 
release into the excitation chamber, 

6.5. 2.3.3. 6 ’Instrument 534 

The OBIPS will operate for 30 minutes during each revolution 
from revolutions 32 through 47. The 30-minute span will start 
7 or 8 minutes prior to the operation of the accelerator 
(instruments 303 and 304) and end about 7 or 8 minutes after 
accelerator operation terminates. 

The OBIPS will be used to obtain images of faint, transient 
atmospheri c energy phenomena such as artificial or induced 
aurorae and glows produced by chemical tracers. The electron 
and MPD Arc accelerator beams will be used to provide the 
energetic particles required for the excited states. The 
images produced by the beams will also be picked up by the 
Orbiter TV cameras. 




1 


1 

] 

] 

] 


i 


1 

] 

) 

1 

1 

1 

i 


] 

i 


\ 







6-26 


The orientation of the OBIPS reference must be known to within 
0.02* of the Orbiter reference. Initial pointing and subsequent 
target tracking will be provided by the APS. 

A cover will be necessary to prevent contamination of the optics 
This will be provided by a door located in front of the lens 
which will be opened Just before data measurements are made 
and closed when the instrument is on standby status. The in- 
strument can be damaged by direct sunlight and operational con- 
trols will be provided to prevent direct solar incidence onto 
the photometers through the lens. 

An inflight calibration will be performed. The calibration 
source will be selected during the next study phase. 

a. Controls will include: 

(1 ) Power control . 

(2) Opening and closing of door. 

(3) Aperture control. 

(4) Selection of filter (if turret is used). 

(5) Image processing gain control. 

(6) TV pointing and controls mode. 

(7) Calibrator source position and light control. 

(8) Instrument pointing control using APS. 

b. Displays will include: 

(1) TV monitor. 

(2) Door position verification. 

(3) Filter selection verification (if turret is used). 

(.4) TV camera direction indicator. 

(5) Calibrator position and light indicators. 

(6) Instrument operati on al status monitor. 


6-27 


6. 5. 2. 3. 3. 7 Instrument 536 


The Triaxial Fluxgate will be deployed on a boom during (or 
before) revolution 32. The instrument will be used to measure 
the direction and amplitude of the earth's local magnetic 
field. 

Data from the instrument will be used to provide an interlock 
for accelerator operations if the earth's magnetic field 
direction is such that beam return might occur. 

There are no special pointing requirements for this instru- 
ment but the knowledge of the reference axes orientation must 
be accurate to within 0.5°. 

The boom will be retracted at the end of this phase. If the 
retraction mechanism malfunctions, the boom and instrument 
will be jettisoned. 

a. Controls will include; 

(1) Power control. 

(2) Boom extension and retraction control. 

(3) Boom jettison control . 

b. Displays will include: 

(1) Magnetic field lines direction relative to Orbiter 
reference axes and field strength. 

(2) Verification of boom extension and retraction. 

(3) Instrument operational status monitors. 

6. 5.2. 3.3.8 Instrument 549 

The Gas Plume Release instrument is a diagnostic tool used in 
conjunction with the Electron Accelerator (Instrument 303). 

It will be used to determine accelerator-^produced electron beam 


6 -28 


flux density and emergence angles. This function will be per- 
formed by the release of gas into the electron beam which will 
allow visual observation of the beam profile OBIPS, or the 
Orbiter TV camera will be used to pick up the beam profile 
image. 

The instrument will operate concurrently with Instrument 303 
only during revolutions 32 through 35 since it is not expected 
that the beam characteristics will subsequently change from 
that initially observed. 

a. Controls will include: 

( 1 ) Power control . 

(2) TV camera angle and mode control. 

(3) Gas release sequence control (synchronized with In- 
strument 303 operation). 

b. Displays will include; 

(1) TV display of images. 

(2) TV camera angle relative to Orbiter reference 
axes. 

{'%) Instrument operational status monitors. 

6. 5. 2. 3 .3. 9 Instrument 550 

The Faraday Cup Retarding Potential Analyzer Cold Plasma Probe 
is a di agnosti c: instrument whi ch wil T be used to supplement the 
operation of Instrument 549 in determining the electron beam 
characteristics of Instrument 303. It will also be used to 
determine the exhaust potential of the Instrument 304 plasma. 
The instrument will be mounted on a boom installed in the AIM 
IB. The instrument will scan the beam fields in a raster scan 
using the APS which will be controlled through a software pro- 
gram resident in the subsystem computer or 1 oaded from mass 
memory . 


There are no special pointing requirements other than that 
imposed on the boom system (< 0.5° knowledge of instrument 
reference axes). The boom will be retracted at the end of this 
phase. If the boom retraction mechanism mal f uncti ons , the 
boom and instrument will be jettisoned. 

a. Controls will include: 

(1 ) Power control . 

(2) Faraday cup inner and outer grid potential control. 

(3) Retarding potential analyzer outer, retarding and 
suppressor grid potential control. 

(4) Boom extension and retraction control. 

(5) Instrument scan sequence control. 

(6) Boom jettison control. 

b. Displays will include: 

(1) Faraday cup collector current amplitude versus in- 
strument position. 

(2) Retarding potential analyzer collector current and 
retarding potenti aT versus instrument position. 

(3) Gold plasma probe current and floating potential versus 
instrument position. 

(4) Boom position relative to Orbiter reference. 

(5) Boom extension and retraction verification. 

(6) Instrument operational status monitors. 

6. 5. 2. 3. 4 Revolutions 48 through 80 

During this phase of the mission, the operations of instru- 
ments 118, 122, 124, 126 and 1011 will be i dentical to those 
conducted during revolutions 17 through 31. The controls and 
displays will also be the same. 



6-30 


Instrument 1002 will operate once at mid-daylight during 
revolution 80. Its operations will be identical to those 
conducted during revolution 16. The controls and displays 
will also be the same. 

Instruments 116, 303, 304, 532, 534, 536, 549 and 550 will 
be on standby or powered down. The support subsystems will 
be operational during this period. 

At the end of this phase, all experiments will have been com- 
pleted. 

6.5.2. 3.5 Crew Operations Timeline (Revolutions 1 through 81 ) 

A PS and a MS will operate and monitor the payload and payload 
subsystems from their respective consoles in the aft crew 
station. Since ASF requires 24-hour operation, two teams each 
consisting of a PS and a MS will operate on 12-hour shifts 
with the last hour of each 12-hour shift serving as a shift- 
over period for the new team. 

The control of the instruments wil 1 be accomplished by grouping 
them into operating sequences based on the instrument operating 
timeline (table 6. 5. 2-1). These sequences will be preprogrammed 
in the payload computer and will be initiated by the PS. The 
initiation can be accomplished by inserting a "At" to initiate a 
sequence into the computer via the keyboard, thus allowing the 
PS^and MS to check system status, set up parameters well in 
advance of the sequence initiate time, and allow for other 
setups and monitoring tasks.. 

Table 6. 5. 2-2 shows a preliminary crew task timeline for opera- 
ting the ASF mission. 


6-31 


TABLE 6.5. 2-1 . - ASF 


Sequence 

1 . 

2 . 

3 . a . 

b. 


4. 

5. 

6 . 


Sequence Duration 

Rev . 1 6 to Rev . 81 

Rev .16 to Rev .17 

Rev , 16 to Rev . 16-1/2 
Rev. 80 to Rev. 80-1/2 

Rev . 1 7 to Rev . 81 

Rev. 17 to Rev. 32-1/2 

Rev. 32-1/2 to Rev. 48-1/2 


7. 

8 . 


Rev. 36 to Rev. 36-1/2 
Rev.' 48 to Rev. 81 


OPERATING TIMELINE 


Instruments 

Subsatellite 

116 

1002 

122, 124 
118, 126, 213 

116, 118, 126, 213, 303, 304, 534, 
536, 549 

550 

118, 126, 213 



TABLE 6.5. 2-2. -CREWMAN TASK TIMELINE 


Rev . 

Mission Specialist {MS) 

Payload Specialist (PS) 

1-12 

1« Set up payload subsystems: 

a. electrical 

b. thermal 

c* data management 

d, computers 

e. subsatellite 

1 . Set up payload D&C 

2. Check out payload in- 
struments 

3. Set up APS 

■ 

1* Deploy subsatellite 

2. Monitor subsatellite data 
and payload subsystems 

1 * Set up for Seq , 2 , 
Seq* 4 and Seq* 5 

■ 

K Monitor subsatellite data 
and payload subsystems 

1 * Initiate Seq * 2 and 
Seq. 3 

2. Monitor Seq* 2 and 
Seq. 3 

17- 

32-1/2 

1. Monitor subsatellite data 
and payload subsystems 

1 . Ini ti ate Seq . 4 and 
Seq* 5 

2* Monitor Seq. 4 and 
Seq. 5 

3. Set up for Seq. 6 . 

32-1/2 

-36 

1. Monitor subsatellite data 
and payload subsystems 

1 . Initiate Seq* 6 

2 . Moni tor Seq* 4 and 
Seq. 6 

36- 

36-1/2 

1. Monitor subsatellite data 
and payload subsystems 

1 . Ini ti a te Seq . 7 

2 , Moni tor Seq * 4, Seq* 6 
and Seq. 7 

36-1/2 

-4S 

K Monitor s ubsa tel 1 1 i,e data 
and payload subsystems 

1 . Moni tor Seq * 4 and 
Seq . 6 

2* Set up for Seq. 8 

48-80 

1* Monitor subsatellite data 
and payload subsystems 

1, Initiate Seq. 8 

2, Monitor Seq* 4 and 
Seq. 8 

3, Set up for Seq. 3 

80-81 

T. Monitor subsatellite data 
and payload subsystems 

1 * Ini ti ate Seq* 3 

2. Monitor Seq. 3, Seq. 4, 
and Seq. 8 

81 

1. Retrieve subsatellite 



NOTE; Seq ^ Sequence 


PE lOOE QiOALiK 


6-33 



This preliminary timeline presents no craw overload periods for 
setting up and monitoring the instrument sequences. 

6 . 5 . 2 . 3 . 6 PDS Retrieval (Revolutions 81 through 95) 

At the beginning of this phase, power to the instruments will be 
turned off and the APS will be retracted and stowed in place. 

If either of the APS fails to return to its stowed position or 
if more than two of the redundant latching mechanisms fail , the 
entire APS will be jettisoned from the payload bay. Individual 
microswitches on the latches indicate the positive latch condi- 
tion. After the APS are safely stowed, the Orbiter will ren- 
dezvous with the deployed PDS. Rendezvous can be automatic 
for the initial approach and manual for the last stage, or 
the entire rendezvous operation can be under manual crew con- 
trol if visibility of the PDS is unimpaired. Relative range 
and range rate will be displayed at the on-orbit station. 

When the final approach is completed and the Orbiter is about 
15 meters (50 feet) from the PDS, the RMS will be deployed. 

The PDS grab collar will be grasped by the RMS attach mechanism. 
Upon verification of positive capture, the PDS will be re- 
tracted into the payload bay and reseated onto the PDS re- 
tention structure on Pallet A-2. The tapered mount cone assists 
in guiding the PDS to the proper location on the structure. 

When the PDS is fully seated on the retention structure, the 
collet piston will be automatically actuated, setting the 
latches and locking the PDS to the structure. Microswitches 
will be used to indicate positive seating and locking of the 
PDS. If the PDS fails to seat or the locking mechanisms mal- 
function, the PDS will be lifted out of the payload bay and 
left clear of the Orbiter. • 

After the PDS is safely stowed and latched into place, the RMS 
will be disconnected from the PDS grab collar. The RMS will 


then be returned to its stowed position. The baseline Orbiter 
configuration has the provisions to jettison the RMS if safe 
stowage cannot be achieved. 

6 . 5 . 2 . 3 . 7 Preparation for Re-Entry (Revolutions 96 through 1T2) 

During this last on-orbit phase... high pressure gases and cryogens 
win be dumped from the payload system.^ using the dump lines 
provided by the Orbiter. Power to all payload systems except 
for those required to verify safety will be removed. The pay- 
load bay doors will be closed an°d the Orbiter systems readiness 
checks will be performed. 

The Orbiter will be maneuvered to its retro-attitude and the 
Orbiter maneuvering system engines will be fired to provide the 
delta velocity required. 

6.6 RE-ENTRY, DESCENT AND LANDING 

During this phase, the payload will be passive. 

6. 7 POST-LANDING 

After landing the Orbiter will be transported to the OPF to 
allow ground crew access, the payload bay doors will be opened 
and the films will be removed from the cameras. The tape re- 
corders in the aft crew station will also be removed. The 
films and tapes will be transported to the ASF dedicated data 
handling facility. 

The films and tapes will be catalogued, 
processed, reformatted, re-recorded and 
to the responsible scientific centers. 

The ASF payload will be removed from the Orbiter payload bay and 


stripped, data reduced, 
stored or^ transferred 


loaded into the payload transporter. The transporter will then 
be sealed and shipped to the payload integration facility. 

At the payload integration facility^ full functional and perfor- 
mance tests will be performed to determine health statu'^ of the 
support subsystems and components. The instruments will, be 
tested to the extent possible at this facility or will be re- 
moved and returned to the supplier for detail tests. 

Damaged or marginal performance elements will be replaced and 
the payload will be reconfigured for the next mission. 

The logistic aspects of operations are yet to be defined. 
Maintenances repair, spares and inventory management, trans- 
portation and handling, and packaging requirements and approaches 
will be further defined in the next study phase. 


6-36 


7.0 ASF SYSTEM DEVELOPMENT STATUS AND REqUIREMENTS 
7 . 1 INTRODUCTION 

One objective of this study was to assess the potential of a 1981 
ASF mission. Accordingly, three schedule hard point requirements 
relative to delivery of a flight-ready ASF payload were assigned 
for this pHase of the study: (1) the launch date of July 1981, 

(2) the payload hardware delivered to KSC at T-6 months, and (3)' the 
payload hardware to integration site at T-9 months. 

These requirements are depicted in tabl e 7.1 -1 . This firm require- 
ment precipitated a comprehensive look at the design status and 
development lead times for all entities comprising the conceptual 
ASF payjoad system. The required program functions may be cate- 
gorized as follows: 

a. Development, test, and acceptance of individual instrument and 
subsystem bl ocks . 

b. Development, test, and acceptance of software packages. 

c. Assembly and test of a qualification ASF system [4 different 
pallet configurations). 

d. Assembly and checkout of a flight model of the ASF system. 

e. Initial installation/checkout of the ASF payload system with 
Orbi ter , 

f. Prelaunch activities. This category contains those functions, 
which are spanned by the above T-9 months delivery requirement. 
Thus, category (a) through (f) functions must be accomplished 
between authority-to-proceed and T-9 months. The category (a)- 
functions must be completed prior to start of category (b) or 

f c) functi ons , i . e . , engineering models must be successfully 
tested prior to fabrication of the qualification of flight 
model hardware blocks. For realistic planning it is assumed 
that production of the qualification arid flight Units will not 
be sequential , and the flight . model production will follow the 










qualification model. Another assumption made during this study 
was that the initial installation, checkout or integration of 
the ASF payload with the Orbiter will be performed with quali- 
fication units of the payload hardware (4- pallets) after comple- 
tion of a qualification test program at the payload system 
level (4 pallet configurations). 

With these assumptions, an estimate of nine months schedule 
block is programmed between delivery of the qualification model 
payload hardware and the T-9 months milestone. The nine months 
time block will optimistically provide time for the first 
assembly and qualification testing of four different pallet- 
mounted payload configurations. The block will also provide 
time for the initial instalTatlon, fit/functional test and 
checkout of the ASF payload (four different pallet configura- 
tions) in the Orbiter^ This nine months will not be a serial- 
ized function; but rather, it is a planning provision for 
defining lead-time requirements to. develop an overall ASF pay- 
load system, This nine months schedule requirement is referred 
to as a first-article integration time. 

7.2 INSTRUMENTS 
Z.2.1 PRIME 

A prime instrument is one which has been described by the scientist 
for a particular experiment or group of experiments. 

The prime instruments are all treated in detail in section 2.0 
and 4.0 and appendix B. It is sufficient at this point to reem- 
phasize two points. 

a. The ID‘s (appendix B) and ED's (appendix A) generated as part 
. of thist.study are preliminary and require more refinement 

fay the NASA AMPS SDWG. 

b. Many of the preliminary ID's contain specifications which are 
beyond current state-of-art instrumentation technology. 


7. 2. 1.1 Technical Considerations 

This section summarizes the more obvious technical considerations 
(for each prime instrument) influencing development time required 
to produce instruments for the ASF payload. 

a. Instrument 116 (Airglow Spectrograph). Similar instruments 
have been flown and used successfully on sounding rockets. 

o 

Some development difficulty may be expected in achieving 300 A 
with the normal incidence grating and with the focusing magnet 
coil required. Changing direction of view with the collimating 
mirror complicates pointing operation. Extreme care must be 
exercised in integrating into the ASF payload due to the instru 
ments susceptibility to stray magnetic fields and EMI. 

Technical risks are :ow. 

b. Instrument 118 (Limb Scanning IR Radiometer). A smaller non- 
cryogenic radiometer with lesser capability will operate on 
the Nimbus "F" spacecraft. Unmanned satellites in recent years 
have carried, as payloads, radiometers which are somewhat simi- 
lar to the instrument described although they were of a lesser 
degree of sophistication. The cooling of the optics required 
by the instrument described is often quite risky. Problems 
are anticipated iri protecting against off-axis interference in 
achieving suitable spectral rejection. Minimizing heat loss, 
cryogenic Teaks, and maintaining proper cryogenic temperature 
pose difficult engineering problems. Technical risks are rated 
high. 

c. Instrument 122 (UV-VIS-NIR Spectrometer). Several similar 
instruments have successfully flown on sounding rockets and a 
satellite version of the proposed instrument is scheduled to 
be flown on the Naval Research Laboratory (NRL) SOLRAD II 
Satellite in November 1975. The small grating in this instru- 
ment may make it difficult to obtain the desired dynamic range. 

■ Technical risks are low. 






d . . Instrument 124 (Fabry-Perot Interferometer). Fabry-Perot 

Interferometers with significantly less sophisticated compon- 
ents have flown successfully in rockets. These instruments 
used ruggedized Piezo-Electric scanning etalons of the type 
considered for this instrument but smaller in diameter. The 
large etalons required for this unit will be difficult to pro- 
duce and keep in proper adjustment through a launch environment. 
To achieve the required degree of optical flatness over the 
large etalon diameters requires advancing the state-of-the-art 
in optical component fabrication. The necessity to maintain 
extreme optical flatness while under thermal and mechanical 
stress may also require significant advances in optical mater- 
ial, Technical risks are high. 

e. Instrument 126 (IR Interferometer). Laboratory models of con- 
ceptually similar spectrometers have been developed and others 

with significantly reduced technical specifications have been 

* 

developed for aircraft operation. However, significant devel- 
opment effort remains to be done on this instrument to achieve 
the full range of specifications and provide the cryogenic 
cooling required. Technical risks are medium. 

f. Instrument 213 (Laser Sounder). Fixed wavelength lasers have 
been employed in both airborne and ground installations to de- 
tect and profile various atmospheric constituents. A reasonably 
high-powered tunable dye laser has been used in a mobile van 
for profiling sodium atoms. There has not been a forerunner 
instrument that has accomplished all the capabilities desired 
for this application. Significant advances must be made in 
energy output capability , laser efficiency, and operational 
lifetime of dye materials. Useful measurements can be achieved, 
however, of at least some of the constituents by using different 
laser heads and wavelengths for different applications.. Techni- 
cal risks are high. 

g. Instrument 303 (Electron Accelerator). Similar devices with 
significantly Tower capability have been flown on sounding 


7-5 


rockets. Current devices have a maximum output energy capa- 
bility of approximately 5,000 Joules v/hich is about a factor 
of 20 less than that envisioned for the proposed instrument. 

To :hieve the desired output energy, voltage levels and current 
lev s wilt present some difficult engineering problems in the 
design of the capacitor storage bank and the output switching 
circuitry. Technical risks are medium. 

h. Instrument 304 (Magnetopl asmadynami c (MPD) Arc). A plasma 
accelerator somewhat similar to the one proposed but with sig- 
nificantly lower output capability has been flown on unmanned 
rockets. This device employs the same energy storage capaci- 
tor bank as that used for Instrument 303. Significant develop- 
ment problems other than those associated with the development 
of the capacitor bank and output switching circuitry are not 
anticipated. Technical risks are medium. 

i. Instrument 532 (Gas Release Module). Development of this instru 
ment is essentially a combining of subsystems that have flown 
successfully in space before. No particularly difficult devel- 
opment or integration problems are anticipated. Technical 
risks are low. 

J. Instrument 534 (Optical Band Imager and Photometer System 

(OBIPS)), All major components of this instrument have been 
developed and employed in either spacecraft, aircraft, or field 
applications. The major significant problem remaining is the 
design of a suitable baffle and the integration of the various 
items into a unified assembly capable of meeting the pointing 
requirements. Technical risks are low. 

k. Instrument 536 (Triaxial Fluxgate Magnetometer). Several flux- 
gates have been flown, however, further instrument development 
is required, to achieve the desired sensitivity. Technical 
risks are low. 

l. Instrument 550 (Particle Accelerator System Level II Diagnos- 
tics). This dia:gnostic instrument package comprises a Faraday 


7-6 


cup, a Retarding Potential Analyzer, and a Cold Plasma Probe- 
Retarding potential analyzers have flown on several satellites. 
The Faraday cup and the Cold Plasma Probe are passive sensors 
- used extensively in ground based ion and plasma studies. 
Significant engineering design effort will be required to 
increase the high voltage capability of the Retarding Poten- 
tial Analyzer and to integrate the three units into a suitable 
packaging configuration- Technical risks are low. 

m. Instrument 1002 { Pyrhel i ometer/Spectrometer ) . This instrument 
is currently under development. Major components have been 
built and are in use. No significant development problems are 
anticipated. Technical risks are low. 

n. Instrument 1011 (Ultraviolet Occultation Spectrograph). The 
basic spectrograph has been designed and breadboarded . Sub- 
systems used have all been developed and employed in previous 
applications. No uniquely difficult development problems are 
visible at this time. Technical risks are low. 

7.2.1 .2 Development Schedule Requirements 

Table 7. 2. 1-1 depicts the estimated lead times required to pro- 
duce each of the prime ASF i nstrunients . The production of instru- 
ments encompasses all the necessary functions between project 
approval and delivery of a qualified and acceptance-tested flight 
instrument to a pallet integration facility. The following typical 
functions are performed in the production of a flight instrument. 

a. Program start upon authori zation-to-proceed . 

b. Procurement cycle and preliminary design studies. 

c. Procurement cycle and final design study. 

d. Procurement cycle and prototype model development and test. 

e. Qualification and flight model development and acceptance 
test . ■ 

f. Qualification testing. 


ORIGINAL PAGE IS 
OF POOR QUALirf 



✓ 


f 




g. Flight model acceptance testing, 

h. Government acceptance. 

Table 7. 2. 1-1 shows a spread of from 24 to 66 months estimated for 
production of all ASF prime instruments. This is due to the diverse 
design schedules. One category is those instruments of current 
design, i.e., similar instruments have been produced for other 
uses. These instruments require some study to define minor modifi- 
cations necessary to adapt them for ASF use. The second category 
is those instruments whose basic design is not current; they must 
be developed, and/or modified, and then fully qualified. These 
instruments are extremely complex and sophisticated. In many cases 
the design is pressing current state-of-the-art technology but can 
be developed with adequate funding. The third category encompasses 
those (futuristic) instrument concepts which could have major 
impact to flight schedules and financial resources. 

Table 7, 2. 1-1 includes the nine months first-article integration 
time block which must be considered in planning the overall system 
development program. 

7 ♦2.1.3 Conclusions 

The follov/ing schedule ground rules were directed for this study. 

a. Assume project approval on January 1, 1977. 

b. Deliver an assembled and checked-out ASF payload (4 individually 
configured pallet assemblies) to an integration site by 
October 1 , 1 980 . 

Table 7. 2,1-2 depicted the results when payload development and 
qualification lead time requirements are superimposed on the 
January 1, 1977, and October 1, 1980, schedule hard points. They 
are grossly incompatible. This incompatibility necessitated con- 
sideration of substitute instruments and other trade-offs treated 
in the following sections of this report. 



INAL PAGE IS 
OOJR QUALllY 















7.2.2 SUBSTITUTE INSTRUMENT CONSIDERATIONS 
7. 2. 2.1 Introduction 

The ASF experiments have been reviewed for suitability of existing 
instruments that may be considered for use in lieu of prime instru- 
ments for program schedule and/or economic trade-off considerations. 
Such instruments are referred to as substitute instruments in this 
report. They are defined as instruments tuat are fully developed 
and have been used to performed similar observations. While they 
may not provide the full degree- of scientific fulfillment that is 
anticipated from those instruments described by scientists of the 
AMPS SDWG, they will nevertheless provide valuable and useful data. 

Economics and time constraints require that the sensor portion of 
the instrument be complemented with the most cost effective off- 
the-shelf subsystems. Since weight is not a critical factor for 
the ASF pallet-only mode payload, considerable flexibility exists 
in the selection of support subsystem hardware which are already 
developed. In addition, equipment installation in the pressurized 
crew compartment and the igloo allow consideration of orbiter or 
even aircraft types of hardware. These factors should result 
in considerable reduction in the cost of support subsystems. 


The Orbiter will have EMI, dust and gas environmental contamination 
which may be manifested as background noise. Therefore, the advanced 
state-of-the-^art sensitivities desired of many of the prime instru- 
ments may not yield usable data that could not be acquired with 
sensors that are already developed. Instruments in existence which 
are less sensitive than those specified by the AMPS SWDG but which 
are adequate for measuring data above the Orbiter background noise and 
are readily available to fill many observational requirements. 


7-11 


7. 2. 2. 2 Technical Considerations 


The substitute instruments listed in table 7. 2. 2-1 for ASF are a 
collection of instruments from many other programs wherein weight 
v/as critical. As a result of the Orbiter's larger payload volume 
and Weight capability, they can now be used simultaneously on the 
same mission. Key spacecraft are AE, Nimbus, ISEE, GEOS, OGO, OSO, 
ISIS, and ATS. Tabl e 7 . 2 . 2-1 1 i sts the prime instrument complement 
for the ASF payload and indicates whether a potential substitute 
has been identified by the study team. A candidate substitute has 
been identified for seven of the 15 prime instruments. More may be 
in existence or in development in the academic and/or industrial 
communities. A continuing search is recommended as follow-up action 
to this study. 


The following pages present a technical comparison of each of 
the seven prime instruments for which a candidate substitute has 
been identified. 

7 • 2 . 2 , 3 Development Requirements 

Table 7.2. 2-2 shows a comparison of devel opment requirements for 
the ASF prime instruments and the potential substitutes identified 
to date. It must be emphasized that the scientific suitability 
of these candidate substitutes has not been assessed. Also the 
development lead times are estimates based on the preliminary 
information currently available. The purpose of this comparison 
is to indicate availability of possible substitutes and the impact 
that the use of these substitutes might have on schedules and 
costs. Therefore, these estimates are carried forward in subse- 
quent sections as potential trada-off factors . 


Instrument 


Substitute 


n 6 

No candidate substitute 

118 

118X Nine-channel radiometer 

122 

No candidate substitute 

124 

1 24X (Several possible candidates) 

126 

126X Instruments from Nimbus satellites 

213 

No candidate substitute 

303 

303X Accelerators flown on sounding rockets 

304 

No candidate substitute 

532 

No candidate substitute 

534 

534X Photometers flown on ISIS, DMSP satellites 

536 

536X Commercially developed magnetometers 

549 

No candidate substitute 

550 

No candidate substitute 

1002 

1 002X Instrument from Nimbus satellites 

1011 

No candidate substitute 






COMPARISON OF PRIME/SUBSTITUTE INSTRUMENTS 


Instrument - 
Configuration - 


Prime 

n 8-Limb Scanning Infrared Radiometer 

Cryogenically cooled instrument 
in dewar construction; 60 cm to 
100 cm clear aperture; detectors 
are copper-doped or gold-doped 
germanium; 12 channels. 


Substitute 

Lower Atmosphere Composition and 
Temperature Experiment (LACATE) 

Cryogenically cooled instrument 
in dewar construction; 20 cm 
clear aperture; detectors are 
Hg:Cd-Te; 10 channels. 


Specifications - 

Physical measurements; 
Resolution: 
Sensitivity: 
Field-pf-view: 

Power;. 

Physical dimensions; 
Size; 

Weight; 

Other: 

Constraints - 


3 urn to 40 vm 6 ym to 18 ym 

(TBD) 0.5 rar (spatial) (TBD) (spectral) 

1 to 5 X 1 0 ^ ^ W cm ^ sr ^ ym ^ (Not known) 

0.02° desirable; 0.08° acceptable 0.04° x 0.11° 

15 W (standby) 100 W (operating) 34 W 


1 to 5 X 10“^^ W cm ^ 


4.52 cu m 
115 kg g 

Dynamic range=10 , off-axis rejection= 
(TBD 3 10-5); nutates and scans 10° each 
side; dewar operates at 28°K or 77°K; 
detector at 4°K 

Operation must be completed 
before cryogen is exhausted. 


0.18 cu m 
77 kg 

Methane/ammonia cooler, operates at 
80°K; detector temp=80°K; cooler is 
mechanical. ■ 

(Not known) 


Procurement - 
Design status: 
Delivery time: 
Relative cost: 


Conceptual only 
36 months 
100 percent 


Design has been flown in space 
18 months 
12.5 percent 

Detectors and filters v/ould require 
change; nutating scanning system would 
be added. System has been flov/n in 
balloons. 



COMPARISON OF PRIME/SUBSTITUTE INSTRUMENTS 



Prime 

Substitute 

Instrument - 

1 24-Fabry-Perot Interferometer 

Beamont Fabry-Perot Interferometer • 

Configuration - 

Combination of 23 cm Fabry-Perot 
interferometer, photometer with 
variable frequency filter and an 
Infrared photometer. 

(TBD) 

Specifications - 



Physical measurements: 

2000 A to 1 0 urn 

Selected wavelengths 

Resolution: 

1 A (spectral) 3 km (spatial) 
25 detector photons 5“^ Raleigh"' . 

0.015 A 

Sensitivity: 

1“ circle 

(Mode I) 

Field-of-view: 

2 rarad (Mode I) ranging to 50 ym (Mode 

50 W 

Power; 

Physical dimensions: 

II } 

1 4 W 

50 W 

Size: 

0.86 cu m 

0.012cum1 

Weight: 

45 kg 

1 Q j^g / excluding telescope 

Other: 



Constraints - 

No specified constraints 

Not known 

Procurement - 

1 ■ 


Design status : 

Conceptual only; concept proven 

Versions have flown on OGO-6 

Delivery time ; 

24 months 

1 2 months 

Relative cost: 

100 percent 

28 percent 

Remarks - 

25 cm etalon 

Requires attachment to larger 
telescope and adding selective 
filters for wavelengths of inter- 
est. Would only measure pre- 
selected discrete lines. 




COMPARISON OF PRIME/SUBSTITUTE INSTRUMENTS 


Prime 


Substitute 


Instrument - 


126-Infrared Interferometer 


(Unnamed) Michel son Interferometer 


Configuration - 


Specifications - 

Physical measurements 
Resolution: 
Sensitivity: 
Field-of-view: 

Power : 

Physical dimensions: 
Size: 

Weight: 

Other : 


Constraints - 


Michelson configuration inter- Double-pass interferometer; has 
ferometer; encased in dewar solar tracker with 0.25° accuracy 

housing; cryogenical ly cooled; 60 for absorption measurements; di- 
em telescope; detectors; Hg:Cd-Te gitally stepped or continuous 
up to 50 pm; InSb above 50 pm; movement; can be used for emis- 
instrument has four ranges. sion measurements . 


1 pm to 150 pm, in four ranges 
0.05 cm"^ 

10”^^ W cm'^ sr'^ pm~^ 


0.45 cu m 
114 kg 

5 

Dynamic range = 10 ; off-axis 
rejection = >10-6; signal-to- 
noise ratio ratio = 100:1 

Protect against contamination; 
use before cryogen supply has 
been exhausted. 


1 pm to 8 pm 
0.25 cm”^ 

(Not specified) 
(Not specified) 


10 W (standby); 25 W (operating) 'v30 W 


0.3 cu m 
1 00 kg 


Procurement - 
Design status: 
Del i very time : 
Relative cost: 

Remarks - 


Design concept only 
36 months 
100 percent 


Has been flown on Nimbus 4 
9 months 
2.5 percent 


Dewar and cryogen designed to maintain Non-cryogenic; designed for nadir obser- 
instrument at internal temperature of vations from satellite; ruggedized for 

77 K or 28 K depending on cryogen; de- aircraft vibration; has been flown on 

tector operates at 4 K. Concorde. Does not have sufficient reso- 

lution for upper atmosphere wind speed 
measurements. 



COMPARISON OF- PRIME/SUBSTITUTE INSTRUMENTS 


Prime 


Substitute 


303-El ectron Accelerator 


Electron Echo Experiment 


DC instrument; pulsed or modu- 
lated heated cathode electron gun 
with magnetic beam steering capa- 
bility; beam modulation capabili- 
ty to 10 MHz . 


Battery powered accelerator with 
ten el ectron’ guns ; deployable 
col 1 ector screen to prevent 
build-up of charge on carrier 
vehicle. 


Instrument - 

Configuration - 

Specifications - 

Physical measurements: 

Resolution; 

Sensitivity: 

Fiel d-of-vi ew: 

Power: 

^ Physical dimensions: 
Size; 

Weight: 

Other: 

Constraints - 

Procurement - 
Design status : 

Del ivery time: 

Relative cost: 

Remarks - 


«-■> . rv. i,^ m -■■■ i.... -'i-* . 


1 keV to 30 keV 
0.1 (max) 

0-7 Amperes 
±5° (max) 

400 W (standby); 5 kW (avg); 10 kW (max) 

1 5 . 75 cu m 
740 kg 


9.5 keV 
(Not known) 

0.5 Amperes 
(Not known) 

'^■5 kW (average) 

2.4 cu m (Collector screen folded) 
'^*300 kg 


Shares some of its volume and 
weight with other accelerators, 
if they .are flown. 


Cathode may be contamination sensitive. 
Pointing with respect to magnetic field 
restricted because of beam return to 
spacecraft. Operate above 200 km. 


(Similar to those of Instrument 
303) 


This ins 
48 months 
100 percent 


trument not designed 


Instrument has flown on Aerobee 
24 months 
28 percent 


Problems involving charge build- 
up and high voltage discharge and 
corona will require study. 


(Same as Instrument 303) 



8L- 



COMPARISON OF PRIME/SUBSTITUTE INSTRUMENTS 



Prime 

Substitute 

Instrument - 

534-Optical Band Imaging Photometer 

Multifilter TV Camera 

Configuration - 

System has 2 TV camera-UV System, 
and Visible-NIR; 2 monochromatic 
radiometers, all on the same 
1 ine-of-sight . Large light 
shields are used, 

Wide angle (150°) TV camera with beam 
compression: filter wheel with 4 tilt- 
ing filters: SEC time integrating TV 
tube: minicomputer for exposure cycling, 
image processing: digital image proces- 
sing: B&W and color monitors. 

Specifications - 

Physical measurements: 
Resolution : 

Monochromatic images are presented and 
spot monochromatic measurements made. 
0.02“: (spatial ) 

(Not specified) 
(Not specified) 

Sensitivity: 

1 0~^footcandl es at TV faceplate 

(Not specified) 

Field-of-view: 

16“ 

150°, with beam compression 

Power: 

Physical dimensions: 

50 W 

550 W 

Size: 

2,5 cu m 

2.8 cu m 

Weight: 

TOO kg 

318 kg 

Other: 



Constraints - 

Protect instrument from high 
light levels in field of view* 

(Similar to those of prime 
instrument) 

Procurement - 
Design status: 
Delivery time: 

Conceptual only 
24 to 36 months 

Ground use; proposed for 
aircraft use 
9 months 

Relative cost: 

1 00 percent 

50 percent 

Remarks - 

UV capability important in other 
AMPS missions, desirable in 
acclerator experiments. 

Requirement for TV camera coverage of 
accelerator operation will be met by 
closed circuit camera installed as part 
of Orbiter baseline configuration. 






■7 


IT 



COMPARISON OF PRIME/SUBSTITUTE INSTRUMENTS 



Prime 

Substitute 

Instrument - 

536-Triaxial Fluxgate 

Triaxial Fluxgate Magnetometer 

Configuration - 

Three sets of excitation and pick- 
off windings on high permeability 
core forms . 

Essentially identical to prime 
instrument. 

Specifications - 


■ • 

Physical measurements: 

Magnetic field vector 

Magnetic field vector 

Resolution: 

N/A 

N/A 

Sensitivity: 

±2 degree, ilO"^ gauss 

±2 degrees, ±10“^ gauss 

Field -of- view: 

4 Ti sr 

4 IT sr 

Power; 

Physical, dimensions: 

(TBD) 

(TBD) 

Size: 

(TBD) 

0.002 cu m 

Weight; 

(TBD) 


Other: 



Constraints - 

Will only operate efficiently in 
EMI below 3 X 10“7 gauss RMS. 


Procurement - 
Design status: 
Delivery time: 

(TBD) 

18 months 

Instruments are commercially 
avail a bl e 
6 months 

Rel ati ve cost : 

100. percent 

5 percent 

Remarks - 

Instruments have flown; some 
idevelopment may be required to 
achieve desired sensitivity. 
Requires space qual ification. 

Have flown on many spacecraft 

i 








COMPARISON OF PRIME/SUBSTITUTE INSTRUMENTS 




Prime 

Substi tute 

Instrument - 


1 002-Pyrhel i ometer/Spectrophotometer 

Pyrhel i ometer/Spectrophotometer 

Configuration - 

Pryhel i ometer and a spectrometer 
with parallel containers in a 
common package. 

Pyrhel iometer flown on NIMBUS R06 
combined with quartz prism 
spectrometer . 

Spec i fi cat ions - 

Physical measurements: 
Resolution: 

Sensi ti vi ty : 
Field-of-view: 

0.25 urn to 4 pm (spectrophotometer) 

0.2 pm to 5 pm (pyrhel i ometer) 

X/AA 'V 100 

0.5 percent ( pyrhel i ometer ) 
Spectrophotometer accuracy 2-5 percent 
5° 

0.2 pm (min) to (TBD) 
(TBD) 

(TBD) 

5° 

Power : 


10 W 

10 W 

Physical dimensions: 

30 X 30 X 10 cm 

30 X 30 X 10 cu 

Size: 


9x10^ cm^ 

9x10^ cm^ 

Weight: 


<1 0 kg 

<1 0 kg 

Other : 




Constraints 


Protection against contamination 
of calibrating radiation source 
is critical. 

Same as 1002 

Procurement - 
Design status: 
Delivery time: 

Under development-models built 
18 months 

Pyrhel iometer proven in space- 
nominal; development required. 
9 months 

Rel ati ve cost : 



Remarks - 



Will cover 99% of solar radiant 
energy . 





Prime Instrument/Stibstitute Instrument 


6 Airglovr Spectrograph 
No substitute identified 
8 Limb Scanning Infrared Radiometer 

8X Lower Atmosphere Composition S 
Temperature Experiment [LACATE) 

.2 UV-VIS-NIR Spectrometer/Photometer 
No substitute identified 
4 Fabrir-Perot Interferometer 
:4X Blamont Fabry-Perot Interferometer 
6. Michel son Infrared Radiometer 
16X Michel son Interferometer 
3 Laser Sounder 

Ho substitute identified 
13 Electron Accelerator 
I3X Electron Echo Experiment 
f4 Magnetoplasmadynami c Arc 
Ho substitute identified 
12 Gas Release Module 

No substitute identified 
^4 Optical Band imager & Photometer System 
14X Multifilter TV Camera 
16 Triaxial Fluxgate 
16X Triaxial Fluxgate Magnetometer 
^9 Gas Plume Release 

No substitute identified 
iO Level II Beam Diagnostics Group 
No substitute identified 
)2 Pyrhel iometer/Spectrophotome ter 
)2X Pyrhel tometer/Spectrophotameter 
1 Ultraviolet Occultation Spectrograph 
No substitute identified 


Design 

Status 

Devel opment 
Time (MO) 
(see notes) 

Relative 
Cost iZ) 

10.0 

57 

100 

5.0 

80 

100 

10.0 

30 

12.5 

5.0 

18 

100 

0.0 

24 

100 

5.0 

21 

28 

2.5 

36 

100 

10.0 

18 

2.5 

0.0 

36 

100 

0.0 

48 

100 

10.0 

33 

38 

0.0 

48 

100 

5 . 0 

24 

100 

5.0 

.42 

100 

7.5 

21 

50 

TBO 

18 

100 

10.0 

15 

5 

5.0 

- 

100 

5.0 

36 

100 

o 

18 

100 

5.0 


30 

2*5 

42 

100 

— 

■ -- ■■■. ■ 



NOTE: Design status and relative cost ratings are assigned as follows: 

Concept state only 0 

Laboratory breadboard exists 2.5 

Operational componenets exist 5.0 

Fully developed* not space operated 7.5 
Operationally proven in space 10.0 

Prime instrument costs are rated 100 percent. Substitute instrument costs 
are relative to prime costs. 


i-S'mal page is 

-“'OCR QUALOT ” 




1 .Z.IA Cone 1 us i ons 


Estimated development times for the potential substitute and prime 
instruments (where no substitutes have been identified) are plotted 
in table 7. 2. 2-3 against the program schedule hard points. Incom- 
patibility still exists for six of the ASF instruments. The table, 
v/hen compared to 7. 2. 1-2, illustrates a trade-off potential signi- 
ficant enough to warrant a more detailed, follow-on investigation 
of possible substitute instruments; especially in view of a forth- 
coming refinement of ASF ED's and ID's from the NASA AMPS SDWG, 

The significance of this preliminary treatment is described in 
sections 8.0 and 9.0. 

7.2.3 SUPPORT SUBSYSTEMS 

The basic approach to developing the subsystem concept for the ASF 
pallet-only mode study was to utilize equipment planned for the 
Spacelab and Orbiter systems to the maximum extent possible. 

The support subsystem equipment selection and development status 
are shown in table 7. 2. 3-1. Of the 33 major groups of elements 
listed, 13 are being developed by ERNO for the Spacelab program 
and seven are being developed by the NASA for the Orbiter. Five 
other items have been (or are being) used on various space vehicles 
and satellites. Only 8 of the major subsystem items reqi. ^ re f ul 1 
scale development at this time. The major new development elements 
required for ASF are the following: 

a. PCSS signal processing electronics. 

b. A&A electronics, 

c. PSS ASF unique C&D panels. 

d. APS including gimbals, torques, and extendable column. 

e. AIM or modules holding the instrument clusters. 

f. Subsatellite retention/ejection mechanism. 

/■■■ 7-22 


-23 






























TABLE 7. 2. 3-1. -SUPPORT SUBSYSTEM EQUIPMENT DEVELOPMENT STATUS 


Ec|ui]iniQnt 

Program 

Suppl ier 

Development Status 

PoiRtinq, Control and StabH 





K 

Oyro Reference Assembly 

Apollo, Skylab, 

Ifoneywel 1 , 

Delco, Kearfott 

Existing, spaceflight proven 

2. 

3tar Tracker Assembly (3) 

Shuttle 

Ball Bros* , 

Roneywel 1 

[n development 

3. 

Sun Sensor 




Existing, spaceflight proven 

A. 

Optical Alignment Measuring Devices 


ITT, Perkin 

“Elmer 

Existing 

S. 

Signal Processing Electronics 

- 


- 

New 

1 Command and Data Hanaaacient 





1 . 

Computer (3) 

Spacelab 

ESRO 


In development 

2. 

Input/Output Unf ts (2) 

Spacelab 

ESRO 


In development 

3. 

Mass Memory (1) 

Spacelab 

£5RC 


In development 

A. 

Remote Acquisition Unit (29) 

Spacelab 

£SRC 


In development 

5. 

Caution and Warning Electronics Unit 

Spacelab 

ESRO 


In development 

6, 

Alarm and Advisory Electronics Unit 

- 


- 

hew 

7. 

Analog Recorder (2) 

Shuttle 

Ode tics 


In development 

3, 

fleyhoard 

Shuttle 

ION 


In development 

9. 

CRT 

Shuttle 

ISM 


In development 

10. 

P55 CiD Panel LIO 

■ - 


- ■ 

flcyi (existing switches, etc.) 

11, 

PSS CiD Panel Ll 1 



- 

flew (existing switches, etc.) 

12. 

PSS C&D Panel L12 

- ■ 


- 

flew (existing svritches, etc.) 

13. 

Control and Display Unit 

Modified 
Shuttle Display 
Electronic 
Unit (DEU) 

IBM 


DEU in development 

1 Electrical Patfer and Distribution 





1. 

Emergency Battery 

Spacelab 

ESRO 


In development 

Z. 

OC/AC Inverter 

Spacelab 

ESRO 


In development 

3- 

Power Control Box 

Spacelab 

ESRO 


In development 

A. 

Secondary Power Distribution Box 

Spacelab 

ESRO 


In development 

5, 

Pallet Power DistHbiition Box (4) 

Spacelab 

ESRO. 


In development 

6. 

Energy Kit (2) 

Shuttle 

Beech Aircraft 

In development 

I Thormal , Structural . Mechanical 





1. 

Pallet (4) 

Spacel ab 

ESRO 


In development 

2. 

APS (2) 


Ball Bros. 


Goddard study conducted . 

3. 

AIM Structure (4) 


Ball Bros. 


Goddard study conducted 

4. 

Boom and Actuator (2) 

Thor, Agcna, 
Delta, Titan 
and others 

Astro Research Corp, 

Existing, spaceflight proven 

5v 

Igloo Container 

Spacelab 

ESRO 


In development 

6. 

Coolant Pumps, Heat Exchanger, 
Capacitors, Plates 

Spacelab 

ESRO 


: In developpiERt 

7. 

Subsatellite Retention/Ejectlon 
Mechanism 

" “ ■ ■ 



New 

a. 

Cryogenic Coolan': Tank, Valve, etc. 

Apollo 



Existing, spaceflight proven 

9. 

Heat Radiator Ki t 

Shuttle 



In development 




7-24 



Some of these items, Including the APS and AIM which are currently 
under study by GSFC, could be in development by 1981. 

Table 7. 2. 3-2 shows the support subsystem development schedule 
assuming a contract start in January 1977. The schedule shows 
development completion including payload integration with the 
Orbiter in early 1981. This assumes that the scientific instru- 
ments will be available for the pallet level development and inte- 
gration tests in early 1979. 

Since the assessment of the instrument development schedule shows 
many of the prime instruments will not be available until late 
1980 and early 1981, it is apparent that the critical development 
paths are those related to prime instrument rather than the support 
subsystem development. 

7.2.4 GROUND SUPPORT AND TEST EQUIPMENT 

The philosophy employed during conceptual definition of paragraph 
5.3.2 requirements was that of utilizing equipment developed and/or 
planned for the Spacelab, Orbiter, or other sources to the maximum 
extent . 

The only visible potential problem area related to status or 
development lead times is that of possible unique experiment and 
system test sets. It is unrealistic to address the potential 
impact at this time due to the conceptual status of ail levels of 
instrument, subsystems, and system designs. 














TABLE 7. 2.3-2 


SUPPORT SUBSYSTEM DEVELOPMENT SCHEDULE (Concluded) 


VAUTltORIZATION TO PROCEED (ATP) 


3. Subsystem 

a. Development and 
Integration 

4. Pallet 

a* Structural 

b. Development and 
Integration 

c. Qualification 

5, Payload System 

a. Development and 
integration 

b. Qualification 

6, Urblter 

a. Development and 
integral ion 


1977 

1978 

1979 

1980 

1901 

19B2 

19B3 

- 

z 

■ 

- 

1 






7»2,5 GROUND DATA HANDLING AND PROCESSING 

The conceptual description of the types of ASF data expected for 
ground data processing is contained in paragraph 5,3 of this report. 
There are no identifiable handling or processing requirements that 
cannot be implemented with existing JSC equipment. 


8.0 CONCLUSIONS 


8.1 INTRODUCTION 

Many worthwhile conclusions may be drawn from results of this 
study which was oriented towards assessing the potential of a 
1981 ASF pallet-only mode STS mission. In this case, assessing 
the potential involved much more than a go-no-go determination 
of scientific and/or technical feasibility of the concept. The 
mission-level treatment, as opposed to only a flight package 
evaluation, precipitated many tangential studies into facility 
level interspaces exposing technical, scientific and programmatic 
factors of significant impact to STS missions planning. They 
extend beyond ASF or even AMPS missions to perhaps planning fac- 
tors applicable to all STS missions dedicated to the use of the 
STS as a scientific platform. 

The scope of this ASF study is depicted in figure 8-1 which 
illustrates the major spacecraft and facility interfaces. 

The study was initialized with a preliminary set of IFRD's devel- 
oped by the AMPS SDWG from which ED's and ID's were derived and 
are contained in appendix A and B, respectively. The prime in- 
struments were then packaged into one of four pallets in a phy- 
sical and functional manner compatible with an Orbiter 7-day 
mission timeline (section 5). In section 6, operational com- 
patibility was verified between the Orbiter/payload and supporting 
facilities (PDS, SPS, TDRSS, STDN, Mission Control and Ground 
Data Processing facilities). 

In the course of the study some problems were encountered, most 
of which were resolved. Some potential problems remain and will 
be summarized in the following trade-offs and recommendations 
sections. In general, however, the feasibility conclusions can 
be summarized as technical, scientific, and programmatic. 






8.1.1 TECHNICALLY FEASIBLE 

It is technically feasible to conduct an ASF mission in the 
pal-let-only mode aboard the STS. Only one factor remains which 
could conceivably negate technical feasibility. The Orbiter EMC 
environment as currently defined is not compatible with the full 
AMPS experiments and will have adverse effects on even the ASF 
missions. The EMC problem and a possible solution are addressed 
in detail in appendix C. Many areas require further indepth study. 
They are addressed in the technical feasibility section below. 

8.1.2 SCIENTIFICALLY FEASIBLE 

A pallet-only ASF mission aboard the STS is scientifically feasible. 
Much refinement is required in the area of scientific requirements. 
This subject is treated in detail in the scientific feasibility 
section below. 

8.1.3 PROGRAMMATICALLY FEASIBLE 

The programmatic feasibility factors (cost, schedule, etc.) 
for the ASF pallet-only mission are addressed in section 8.2. 

Many trade-off options have been identified; many others will 
be available from follow-on studies. 

8.2 FEASIBILITY ASSESSMENT 

8.2.1 SCIENTIFIC OBJECTIVES FULLFILLMENT 

Results of the study show that all scientific objectives estab- 
lished for atmospheric science Orbiter missions can be met with 
a pallet-only payload. The added instruments made possible by 
the additional payload volume enables acquisition of much more 
scientific information in each mission than would otherwise be 
possible. This will reduce the necessity to time-phase operations 
over several missions, as would likely often be required with a 


8-3 




smaller complement of Instruments. This will permit more accurate 
determination of interacting time-variant parameters. 

The baseline system established in the study is not fixed; however, 
it shows that a large complement of highly sophisticated instru- 
ments can be appv apri ately poi nted, powered, and thermally 
controlled. Also, very importantly, the study shows the atten- 
dant data can be adequately handled and the instruments controlled 
and their outputs displayed in the limited volume of the aft crew 
station. 

This system is not final. It is based on preliminary and sketchy 
information from the AMPS SDWG. Their requirements, conceptuali- 
zations, and experiment prioritization are still in progress and 
may result in some changes in the instrument complement. The 
modular approach taken, for packaging and pointing, readily 
accommodates changes in instruments and their configurations that 
may evolve as the AMPS SDVJG continues its study. 

There is also room for growth of instrument numbers or size. 

Pallets A-2 and A-4 have substantial space available that can 
accommodate additional instrumentation. Small instrumentation 
or electronics packages can also be located on pallets A-1 
and A-3 if needed. 

In summary, it appears highly probable that as scientific require- 
ments are further defined, this baseline system will evolve into 
a complete ASF that need not be reconfigured for successive 
atmospheric science missions. It is also anticipated that only 
replacement of certain pallets and software packages v/ill be 
required to convert to payloads suitable for other AMPS experiments. 


8. 2. 1.1 Problems 


The foremost cause for uncertainties related to scientific 
feasibility stems from the preliminary nature of the instrument 
and experiment descriptions. Although changes to these documents 
and additional information are forthcoming from the AMPS SDWG, 
only the preliminary versions of these documents were available 
during the study. As a result, the study had to contend with 
futuristic performance requirements for experiment objectives 
and procedures which in some cases were not fully defined or 
understood. Maximum emphasis is placed on the necessity for the 
forthcoming upgrading of the IFRD's and ED's. With a better 
definition of the instruments, the follow-on studies will result 
in improvements in cost effectiveness of the ASF pallet-only 
mission. The improvements will be manifested in both scientific 
return and cost/schedule factors. 

Problems related to scientific feasibility of the ASF pallet-only 
mode concept can be categorized as follows. 

a. Those scientific objectives requiring the operation of instru- 
ments in a low EMI/particle contamination environment may 

not be compatible with the contamination environment presently 
postulated for the Orbiter payload bay. However, this problem 
is not unique to pallet-only mode, but is of concern for all 
payload configuration on STS. The EMI portion of this poten- 
tial problem is detailed in appendix C. 

b. Many of the proposed prime instruments cannot be developed 
in time for the baseline 1981 launch. Also, there are ques- 
tions as to v/hether or not they can be developed at all within 
realistic budget limits. This fact again points to the neces- 
sity for an upgraded set of instrument and experiment defini- 
tions against which the use of substitute instruments may be 
v/eighed. The following information presents more detailed 
insight into this type of problem. 



Of the prime instruments treated in section 7.0, at least four will 
probably not be developed in time for a 1981 launch. Of the four, 
two substitute instruments have been identified that could be 
used to yield a good percentage of the desired scientific infor- 
mation. The two instruments that may not be ready are Instrument 
304 (Magnetoplasmadynamic Arc) and Instrument 1011 (UV Occulation 
Spectrograph). Unless either substitute instruments or other means 
of acquiring the desired information are found, experiments using 
these instruments may have to be postponed for later missions. 

The three instruments that represent high technical risks present 
problems that could adversely affect scientific fulfillment. 

These are: Instrument 124 (Fabry-Perot Interferometer), Instru- 

ment 118 (Limb Scanning IR Radiometer), and Instrument 213 (Laser 
Sounder). The technical difficulties delineated in the previous 
section may necessitate relaxing critical specifications in order 
to achieve a realistic unit. Substitutes which can be used in 
lieu of instruments 124 and 118 are identified. However, none 
is available for Instrument 213. This instrument presents oper- 
ational problems in addition to technical problems that must be 
resolved before it can be used as it is currently envisioned, 
and for some of the purposes for which it is intended. Specifi- 
cally, it is intended to measure the intensity and temperature 
of various atmospheric constituents by a laser fluorescence 
technique. Some of these measurements require energy and power 
levels far in excess of those considered possible by the early 
1980's. Increasing the energy and power output to these levels 
exceeds safe allowable limits for ground personnel by at least 
three orders of magnitude. 

•j 

As an example of the problem, an experiment was reported which 
required accumulation of returns from 250 pulses, 0.5 joule each, 

5 seconds apart, from a ground-based experiment to accumulate 

^"Composition, Structure, and Dynamics of the Atmosphere," 
Sandford and Gibson, FATP, ^ 1423 (1970) 

' 8 - 6 ' . . 


statistics on the sodium layer. The Orbiter would travel 
10,000 km while accumulating these data. To accumulate similar 
statistics at a range of 300 km above the Na layer would require 
approximately 1 kilo joule per second. The safe level from 
ground observe eye damage standpoint is 1 joule or about three 
orders of magnitude less than that required. This problem needs 
careful study to define operational usage and realistic design 
specifications for Instrument 213 to support any payload confi- 
guration, i.e., pallet-only, Spacelab, etc. 

While there is no instrument that covers the broad range of wave- 
lengths, power and energy requirements stipulated for Instrument 
213, it is likely that further evaluation of requirements and 
instrument availability may reveal that operational lasers 
already developed for other applications can be modified to 
meet a majority of scientific requirements. 

8. 2.1.2 Impacts 

As inferred above, there are areas of uncertainty which remain 
in assessing the potential of the pallet-only mode ASF mission. 

The potential involves not only scientific merit but also cost/ 
schedule factors. The more obvious trade-offs will be summarized 
i n paragraph 8.3. 

8.2.2 SUPPORT SUBSYSTEMS 
a . 2 . 2 . T Cone! usions 

In general, the conclusion in the support subsystem area is that 
by using Spacelab, Orbiter and other proven equipment and approaches, 
the ASF subsystem concepts selected are compatible with the ASF 
requirements and constraints. Feasibility at the conceptual level 
has been assessed in each area and no major functional feasibility 
problem is anticipated. However, a number of areas require fur- 
ther definiti'on as to sizing and capacity. These have been iden- 
tified and are suggested for further study. 


The particularly critical subsystem for pallet-only operations 
proved to be the CDMS. The conceptual designs for the CDMS, as 
developed during this study, were found to be adequate to support 
the additional instruments afforded by the pallet-only mode. The 
command and control functions intricately involve the crew and 
the ASF experiment and subsystem computer capabilities. Although 
adequacy of the control and command techniques herein developed 
may appear marginal, the question of feasibility only involves 
sizing (i.e., memory and processor size) which will be established 
during the follow-on studies. 

The specific conclusions resulting from the study were as follows. 

a. The Orbiter ATCS capability of 29,500 Btu/hr with the radiator 
kit is adequate for the ASF payload requirement of 24,000 
Btu/hr. 

b. Open loop cryogenic cooling appears to be the only practical 
approach to cooling instruments 118 and 126 to less than 4K. 
However, a detailed heat load analysis is required after the 
instrument designs are better defined. 

c. The APS approach selected provides instrument pointing 
accuracy capability (0.007° 1 sigma) with adequate margins 
over requirements (0.017° 1 sigma). 

d. The selected boom and boom deployment approach meets the ASF 
accuracy requirements of 0.6°. 

e. The subsatellite retention/ejection mechanism meets both the 
Orbiter launch and landing static load requirements of 9 g 
with a capability of 17 g and provides a simple means of 
ejecting the subsatellite at the required 20 cm/sec separation 
rate. 

f. The pallet loads and center-of-gravity locations are well 
within specified constraints. Considerable growth potential ^ 


8-8 




exists with the pallet-only mode since equipment and instru- 
ments can be relocated from pallet to pallet. 

Two Orbiter energy kits with a capability of over 1700 kWh 
will supply, with sufficient margin, the 897.3 kWh of energy 
required by the ASF payload. The Orbiter provides for the 
addition of four energy kits. Therefore the growth potential 
is significant. 

The maximum power level available from the Orbiter, 12 kW, 
is adequate for the 9 kW peak required by the ASF payload. 

The Orbiter thermal constraint reflected by the 815 kW 
maximum average power capability is sufficient to handle the 
6.9 kW average required over an extended (>1 Orbit) period 
of time . 

An independent instrument pointing and attitude measuring 
system is required since the ±2.0° accuracy predicted for 
the Orbiter for pointing payload is inadequate to meet the 
0.017° minimum instrument req irement. 

The attitude measurement approach selected (gyro reference 
with star tracker update) provides an accuracy capability 
will within requirements (0.007° capability versus 0.017° 
required) . 

Onboard computer control of instruments, subsatellite and 
subsystem operations poses no fundamental issues of functional 
feasibility since the types of inputs, outputs, equations, 
and algorithms used will be similar to those currently being 
used on many commercial and space applications. The question 
of memory, executive, I/O and software capacity and timing 
remains to be resolved. These comments regarding onboard 
computer control apply also to onboard data processing. 

The data transmission requirements of 16 kbps and 123.192 kbps 
from the deployed subsatellite and from the ASF instruments, 
respectively, to the Orbiter are within the capability of the 
Orbiter. Since the Orbiter capability of processing ASF 
instrument data located in the payload bay is 5.0 Mbps for 



S band and 50 Mbps for Ku band data links, considerable growth 
margin exists for these data. Since the subsatellite to 
Orbiter data rate capability Is 16 kbps, some data compres- 
sion may be necessary or individual instrument data rates 
may have to be reduced if additional data grovJth occurs. 

n. Available space at the aft crew station limits the number of 
PS's to one at any given time. 

0 , Due to the experiment timelines which require 24 hr/day oper- 
ations, two crew members will be required to man the PSS; 
each on a 12 hour/shift basis. 

р. The AE satel 1 i te appears to be an ideal carrier for the particle 
measurement support instruments required by the ASF experi- 
ments. Nine of the 17 existing instruments will be used and 
two new instruments will be added. 

8. 2. 2. 2 Assessment 

Conceptual feasibility was established for the major support sub- 
system areas. The issues involved and the study results are dis- 
cussed in some detail in each of the subsystem sections (paragraphs 

5.2.1 through 5.2.6). The results are summarized in this section. 

8. 2. 2. 2.1 Thermal, Structural and Mechanical Subsystems (TSMS) 

The major issues in the TSMS were as follows. 

a. Installing all 15 ASF instruments and the support equipment 
on the ESRO furnished equipment pallets within the require-^ 
ments and constraints of the ASF missions. 

b. Providing, accurate instrument pointing and tracking for massive 
(up to 691 kg) clusters of instruments. 

с. Maintaining better than 5° reference axes accuracy relative 
to payload reference, axes of instruments extended on long 
(20 m) booms under environmental conditions, Orbiter limit 
cycle operations, and boom scanning operations. 


8-10 


d. i'lai ntaini ng retention integrity of subsatellite installation 
under launch and landing dynamic environments while providing 
a simple and effective means of ejecting the subsatellite 

at the desired separation rate. 

e. Providing the required cryogenic cooling of instruments 
118 and 126. 

Table 8. 2.2-1 summarizes the TSMS issues, approaches taken, and 
compares capabilities with requirements. The table indicates 
the growth potential of the ASF payload configuration in terms 
of weight, volume, and pallet-mounting space. 

8. 2. 2. 2. 2 Electrical Power and Distribution Subsystem (EPPS) 

The major feasibility issues in the EPOS area were the following. 

a. The compatibility of the ASF power needs with Orbiter 
capabi 1 i ty . 

b. The ability of the Orbiter to provide the required energy 
levels.' 

c. The compatibility of the ASF thermal energy dissipation 
expected relative to the Orbiter ATCS capability. 

d. The selection of a practical high voltage electrical pow^r 
source for instruments 213, 303, and 304. 

Table S.2.2-2 summarizes the EPOS issues, approaches taken, and 
compares the capabilities of each approach to the ASF requirements 
or constrai nts . 

8 . 2 . 2 . 2 .3 Pointing, Control and Stabi 1 i zati on Subsystem ( PCS S ) 

The major feasibility issues in the PCSS area were the folTcwing. 

a. The conceptual approach to be taken for instrument pointing 
and tracking . 

8-11 


Table 8. 2. 2-1. -TSMS ISSUES 



00 


Item 

Approach 

Capability 

Requirement 

la. 

Installation space and volume 

Install instruments 
and support equipment 
on 3 pallets, sub- 
satellite on one 
pallet 

2 

Space: 4 x 17 m 

Volume: 4 x 33 m^ 

Unused Available 
Space Volume 

Pallet 1 0% 0% 
Pallet 2 50% 50% 
Pallets 0% 0% 
Pallet 4 40% 40% 

lb. 

Installation weight {not 
including pallet). 

Distribute total pay- 
load weight among 
four pallets 

Maximum weight: 

3,000 kg/pallet 
with igloo 

3,500 kg/pallet 
without igloo 

Pallet 1: 2449 fcg (with igloo) 

Pallet 2: 721 kg 

Pallet 3: 1852 kg 

Pallet 4: 998 kg 

2. 

Instrument pointing (APS) 

Modified Ball 
Brothers SIPS 

1 arc sec 

Accuracy: <1 arc minimum 

3, 

Instrument pointing (boom) 

BI-STEM concept 

Accuracy: < 0.6° 

0.5° 

4. 

Subsatellite retention/ 
ejection 

Collet/GNg activation 

17 g 

2 cm/ sec 

Launch and landing 
Loads: 9 g 

Ejection AV: 2 cm/sec 

5, 

Cryogenic cooling i 

(Instruments 118 and 126) 

Open loop joules - 
Thompson Expansion 
Ne or N 2 

4K 

(TBD depends on 
instrument design) 

Detector: <+4K 
Housing: <+77K 


J. ^ 

V- : i -k* 





TABLE 8. 2. 2-2. - 


Item 

Approach 

1. Power level 

Use Orbiter fuel 
cells 

2. Energy level 

Use 2 Orbiter 
energy kits 

3. Thermal energy 
dissipation 

Use Orbiter heat 
radiator kit 

4. High voltage, 
high power 
source 

Use capacitor bank 
{0.8 joules 
capacity and power 
converters ) 




} 

v..> 


8-13 


ISSUES 


Requirement 

Capabil ity 

9 kW peak 

Dedicated fuel 
cell -12 kW 

897.3 kWh 

1730 kWh 

24,000 Btu/hr 

29,500 Btu/hr 

Up to 30 kV 

>30 kV at a 
>70% efficiency 


i 


I 











The main feasibility issues in the CDMS area were the following. 

a. The ability to perform most of the ASF experiment operations 
automatically with the onboard computer. 

b. The capability of performing most of the data processing 
automatically through the ASF computers. 

c. The compatibility of ASF data transmission requirements with 
Orbiter downlink capability. 

Table 8. 2. 2-4 summarizes the CDMS issues, the selected approaches, 
and the comparison between the capabilities of the selected 
approaches and the requirements. 

8. 2, 2. 2. 5 Aft Crew Station 

The main feasibility issues in the aft crew station support area 
were the following. 

a. Crew control versus automatic computer control of experiments 
and instruments. 

b. Adequacy of PSS space allocation. 

c. Workload versus crew capability 

If' 

Table 8. 2. 2-5 summarizes these issues, describes the selected 
approaches, and compares the capabilities of the selected approaches 
with the requirements. 



TABLE 8. 2. 2-3. - PCSS ISSUES 


Item 

Approach 

Requirement 

Capability 

1. Instrument pointing 

ASF independent 
pointing system 

<0.017'=’ 

Orbiter: +0.4° ASF 

system (including 
AMS): <0.007° 

2. Attitude measurement 

Three axes gyro 
referenced with 
star tracker up- 
dates 

<0.017“ 

<0,007° 

3. Centralized versus 
distributed AMS 

L.. 

Central using 
optical attitude 
reference transfer 

Provide reference 
.for 2 pointing 
systems with 
0,017° accuracy 

Provide attitude 
transfer accuracy 
of better than 
0,001° 



TABLE 8. 2. 2-4. - CDMS ISSUES 


Item 

Approach 

Requirement 

Capability 

1. Onboard computer 
control versus crew 
control of experiments 

Maximize onboard 
computer control 

Control 15 
experiments using 
15 instruments 

Conceptual feasibility 
of automatic experiment 
operations not an issue. 
Computer, I/O and soft- 
ware sizing and timing 
impact yet to be 
assessed. 

2. Onboard computer data 
processing versus 
ground data processing 

Onboard computer 
processing for most 
data 

Processing of data 
from 15 instruments 
and 5 subsystems 

Conceptual feasibility 
of extensive data pro- 
cessing onboard not an 
issue. Computer, I/O 
and software sizing 
and timing yet to be 
assessed. 

3. Data transmission 
downlink (STDN or 
TDRSS) 

Use Orbiter baseline 
S band EM with STDN 
and Ku band with 
TDRSS 

123, 192 kbps data 
rate 

S band FM: 

Analog 4.0 MHz 
bandwidth. 
Digital 5.0 Mbps 
Ku band; 

Analog-4.2 MHz 
bandwidth. 
Digital -50 Mbps 

4. Data transmission from 
subsatellite to 
Orbiter, 

Use Orbiter baseline 
S band PM link 

16 kbps data rate 

1 6 kbps 







TABLE 8. 2. 2-5 - AFT CREW STATION SUPPORT ISSUES 


Item 

Approach 

Capabil ity 

Requirement 

1. Crew control versus automatiCj 
computer control of experiments 
and instruments 

Automatic, computer 
control of most 
instrument. 

Feasibility of controlling 
15 or more experiments and 
15 or more instruments 
through computer program 
is not an issue. The 
impact on computers, I/O. 
and software sizing and 
timing is yet to be 
evaluated. 

Control operations of 
15 experiments and 
15 instruments. 

2. PSS space allocation 

Minimum manual , 
direct control , 
minimum dedicated 
displays. Use key- 
board/CRT for 
control and display. 

Using the computerized 
control and data 
processing approach, the 
standard panel allocations 
are adequate. 

Three standard 
48.26 X 53.34 cm 
{19 X 21 in) panels 
for control and 
displ ay. 

3. Workload versus crevy 
capability 

Tvm payload 
specialists 
each working a 12- 
hour shift. Use of 
a second specialist 
or possibility of 
pi 1 ot or commander 
sharing PS duties 
yet to be assessed. 

Using the computerized 
control and data 
processing approach, two 
crew members, each 
working separate 12-hour 
shifts should be 
adequate. 

24 hour/day mission 
coverage. Control 15 
experiments and 15 
instruments. Monitor 
instrument data. 



8. 2, 2. 2. 6 Particle Detector Subsatellite (PDS) 


The main issue in the PDS area was the question of adequacy of 
the AE satellite for ASF mission support. Results indicate that 
the AE satellite, with a few replacements and deletions of instru 
ments, will meet ASF support requirements. Of the 15 AE instru- 
ments, eight are not applicable to ASF missions and two different 
instruments must be added. The AE support systems appear to be 
fully compatible with ASF mission requirements. However, the 
electrical energy storage capacity will be increased by 100 per 
cent to provide additional margin for instrument operation duty 
cycles . 

8. 2. 2. 2. 7 Other Key Issues 

Other major feasibility issues may exist which are not unique to 
the ASF pallet-only mode of operation, but should nevertheless 
be mentioned since they may be more fundamental to the question 
of feasibility than those related to ASF unique areas. These 
other issues include: 

a. The impact of Orbiter background EMI on the practicality 
of performing experiments. 

b. The impact of the electrostatic charge on the Orbiter and 
payload on accelerated electrons and ions. 

c. Effect of Orbiter background contamination (e.g., water and 
other vapors) on experiments operations. 

d. The capability of ground data handling facilities to store 
and segregate billions of bits of data. 

These areas were not fully addressed during the study since they 
are common to all AMPS or ASF types of missions. Assessments 
must be made to determine if the ASF missions can be undesirably 
impacted by these issues and to determine if practical solutions 
exist. 


8.2.3 PROGRAMMATIC FACTORS 


8. 2. 3.1 Schedule 

As described in section 7.0, there are no major constraints iden- 
tified to date in meeting the program schedule milestones for 
delivery of a flight-ready ASF payload system on October 1, 1980, 
except in the case of prime instruments. An alternative approach 
involving consideration of existing substitute instruments offers 
potential relief to the overall instrument problem; but not a 
total solution. This prime versus substitute instruments approach 
is addressed in detail in paragraphs 8.3, 8.4, and 8.5 

8 . 2 . 3 . 2 Costs 

Detailed cost analyses are not included in this technical report. 
They are contained in the Executive Summary. 

The major cost items for the ASF pallet-only mode are centered 
around development of the advanced state-of-art prime instruments 
and the costs associated with end-to-end testing (development, 
qua! if ica’ti on , and flight acceptance) required for a scientific 
payload of this size and complexity. 

8 . 3 TRADE-OFF CONSIDERATIONS 

8.3.1 SCIENTIFIC 

Scientific requirements are presently quite preliminary, hence 
hard instrument specifications are not feasible in many instances. 
Many of the prime instruments are very sophisticated and some 
require technological advances that cannot be accurately timed. 

The result is that lengthy development times may preclude their 
inclusion on earlier Orbiter flights, as shown in table 7. 2. 1-2. 

In such cases, it will be necessary to: (1) Use a different 



technique to derive the desired scientific information; (2) 
postpone experiments that require, the instrument; or (3) use a 
substitute. instrument. 

Consideration of different techniques that might provide requisite 
information requires scientific investigation that is beyond the 
scope of this study. The advisability of postponement of desired 
experiments is also beyond the scope of this study, however, it 
seems obvious that such a choice would only be made if there 
were no other alternative. The possibility of using substitute 
instruments on early ASF missions, however, is a likely choice. 

Summarized development schedules and relative costs of prime 
versus candidate substitutes, depicted in figure 7. 2. 2-1, show 
that the schedule problem could be alleviated and substantial 
cost savings realized if substitute instruments could be used. 

Throughout this study, it was assumed that those instruments 
described by the AMPS SDWG in the IFRD's are the best choice 
and will provide the highest yield of scientific information. 

Based on that assumption, using substitute instruments will, to 
some extent, reduce the degree of scientific fulfillment that 
might otherwise be obtained. The quantification of this reduc- 
tion is hard to derive, because the full number and range of 
physical quantities to be measured is not known. They will not 
be known until they have been measured with instruments that have 
bandwidths and dynamic ranges greater than the quantities to be 
measured. . 

Nevertheless, useful and valuable information concerning atmos- 
spheric dynamics can be derived from spatial and temporal 
measurements made with instruments with less than the ultimate 
capabilities. Their use could preclude detection of some obscure 
phenomena or subtle effects; however, the global coverage of 
simultaneous measurements of interacting parameters afforded by 


the Orbiter should yield information that will be invaluable in 
characterizing and modeling the principal dynamic processes of 
the atmosphere. 

Trade-off parameters that might influence whether or not a sub- 
stitute instrument is selected include size, weight, power 
requirements, thermal requirements and operational constraints. 
There is insufficient information about both prime and substitute 
operational constraints to consider the latter further as a trade 
off parameter. Regarding size, weight, power, and thermal re- 
quirements, the Orbiter can accommodate any but the most gross 
increases in these parameters. The brief comparative descrip- 
tions available at this time do not reveal any difference of 
such magnitudes, except for Instrument 534 versus Instrument 534X 
power requirements. In that case, the power required by the 
substitute instrument is about 500 watts greater than that 
estimated for the prime. The timeline developed during this 
study shows this instrument will be used only for a very small 
fraction of the mission time. Furthermore, the ASF payload 
power requirements are far below the full capacity afforded by 
the ASF configuration. The remaining trade-off factors, namely 
schedule and relative cost, are shown in tables 7. 2. 1-1 and 
7. 2. 1-2. The assumed cost of each prime instrument is expressed 
as 100 percent and the cost of the corresponding substitute is 
expressed as a percentage of that cost. Generally, as would be 
expected, substitute instruments cost substantially less than 
the closest corresponding prime instruments. Many examples of 
potential ly significant cost and schedul e options can be derived 
from tables 7. 2. 1-1 and 7.2. 1-2. 

8.3.2 TECHNICAL 

8.3.2. 1 Approach 

The primary approach to concept and design selection for this 
study was to make use of the ERNO designed Spacelab and the 


8-21 


Orbiter designs wherever possible. The purpose was to show 
conceptual feasibility and not necessarily the most cost effec- 
tive or optimized performance or design approach. The issues 
identified and discussed are not related only to the pallet-only 
mode but are considerations that must be addressed for any payload 
configuration. Future studies should consider the impact of 
alternative approaches to cost, risk, schedules, performance, 
capability or capacity margins, reliability, weight, size, power, 
and other trade-off parameters considered to be of prime significance. 

Table 3. 2, 4-1 summarizes the alternatives to the selected ASF 
approaches which should be considered in future studies. 


8 . 3 . 2 . 2 Structural | 

Installation of large structural elements; mounting the APS or | 

the subsatellite installation structure directly to the standard J 

payload attach points provided by the Orbiter can result in | 

significant weight savings since pallets A-1 , A-2 and A-3, each .. j 

weighing 428 Kg, would not be used. Additional attach structures j 

would be required which would reduce the weight savings to some | 

degree. | 

3 

1 

The pallets have great flexibility for installation of different •; 

sized equipment in different locations with standard provisions 

for active thermal control , if required. These capabilities would ’ 

require considerable development effort if individual installation 

provisions were to be provided. ^ 

- ’ ■ 1 

8 . 3.2.3 Thermal Control System ^ 

Current evaluations indicate that there is about a 14 percent 

margin between payload heat dissipation requirements and the 

Orbiter ATCS capability. If greater margin is required, additional 

heat dissipation capabilities can be provided through payload " I 

unique radiators. Considerable effort would be required for I 


development of these radiators and they v/oiild result in additional 
payload weight. The a’’ternative would be to constrain instrument 
operations to reduce • .rage power consumption, which might 
reduce the effectivene , of the experiments. 

8. 3. 2. 4 Remote vs. Direct Access Circuit Breakers 

Direct access circuit breakers would be located at the aft crew 
station. They would provide direct means for manually control- 
ling primary power to each instrument and equipment if individual 
power control was lost. Also, direct access provides a reliable 
way of resetting the circuit breaker switches. Remote circuit 
breakers reduce the weight of power lines since Targe wires (e.g. 
4/0 gauge) carrying primary currents need not be routed to the 
aft crew station and back. 

8. 3. 2. 5 High Current Transmission Media 

Large cross sectional area copper busses interconnecting the 
pallets with the central power distribution point would be the 
most efficient way of providing the high current capability 
required and would allow greater flexibility for reduction of 
common impedances. However, this approach will probably result 
in greater levels of magnetic field generation since the enclosed 
area of the total current loop will be increased. Tv!/q busses 
(pov/er and return) adjacent to each other will not provide the 
same level of field cancellation as a tv/o-wire twisted pair. 

8. 3. 2.6 AMS 

The distributed system is more accurate than the centralized 
system since optical transfer of attitude reference from one 
user to the next is not required and operations become more 
complex. However, distributed systems require more hardware 
and software . 


8-23 


8. 3. 2.7 Payload Specialist Work Station 

An additional mid-deck data monitoring station will provide in- 
creased space such that additional displays could be provided. 

With the added space, real time onboard data analyses of the 
experiments could be provided. The added equipment and crew 
member will increase cost and weight. 

8. 3. 2. 8 Instrument Sequence Initiation 

Control of instrument initiation from the ground station would 
reduce crew workload. It would, however, complicate ground 
operations. 

8 . 3 . 2 . 9 Data Processing 

Processing the scientific and engineering data at ground facili- 
ties would reduce the burden on the onboard computer and result 
in smaller, less complex hardware and software. However, the 
downlink data transmission requirements could be increased 
significantly and the ground facility software complexity would 
increase. 

8.3.2.10 Mass Memory Operational Programs 

Providing full end-to-end mission operational programming capa- 
bility resident in the onboard mass memory reduces the uplink 
communication load and the dependence on timely ground support. 
However, a much larger mass memory capability is required. 

8.3.2.11 Data Compression 

High density data compression techniques can result in significant 
reduction of downlink communication data rate and quantity required 
However, high compression systems with capability of reducing 
data quantities by a factor of 10 or more have not yet been 
developed to the operational stage. Considerable development 


8-24 


effort would be required before such a system could be used on 
the ASF program. 

8.3.2.12 Computer, Processor 

The centralized computer results in a more efficient utilization 
of the machine since the executive, central processor, memory, 

I/O and power supplies can be time shared by the different users. 

Dedicated distributed microprocessors are more flexible since 
any change in software programs or computational requirements 
affect only the processor directly Involved. Problems of priority 
are reduced and faster processing is possible without the exten- 
sive time sharing required. 

8.3.2.13 Subsatellite Retrieval 

Recovery of the subsatellite and returning it for Refurbishment 
to new mission requirements and reusing it is obviously a more 
economical approach than to provide a different subsatellite for 
each mission. The possibility of continued use of a single sub- 
satellite with the capability of supporting many missions left 
in orbit after thefirst mission should be further explored. 
Spacecraft retrieval of the subsatellite increases Orbiter opera- 
tional complexity. 

8.3.2.14 Orbiter and Payload EMI Environment 

The baseline Orbiter using structure as the return for electrical 
current will present a high background EMI environment to payloads. 

If the levels are such as to affect the validity of the experiments 
the trade-off considerations are: (1) provide extensive electro- 

static and magnetic shields to protect the instruments, (2) operate 
the instruments on extended booms, (3) operate the instruments 


on deployed subsatelHtes, and C4) change Orbiter structure 
return to a two-wire system. 

Extensive shielding around the instruments may not be adequate 
to reduce the EMI fields to acceptable levels. Booms 'omplicate 
operations and reduce pointing accuracy. Two of the ASF instru- 
ments are already on booms, although only one, the Triaxial Flux- 
gate, is deployed because of the EMI effect of the Orbiter and 
payload. Operating instruments on deployed subsatellites is an 
expensive approach. Also, accurate co-alignment of instruments 
on the subsatellite with instruments onboard the Orbiter, if 
required, is more difficult to achieve. Changing the Orbiter 
structure return to a two-wire system will increase Orbiter 
wiring weight by about 317 Kg (7D0 lbs) but is probably the most 
cost effective approach for the ASF program. 

8.3.2.15 Support Subsystem Equipment Trade-Off 

In each subsystem area there are a number of alternative equipment 
approaches currently available. Others are almost certain to be 
available by 1981. These equipment include star trackers, gyro 
reference assemblies, computers, mass memories, tape recorders, 
high voltage supplies, power inverters, remote circuit breakers, 
and CRT' s . 

Each equipment area will have a number of trade-off considerations 
which must be assessed. The primary considerations will be those 
associated with direct support of the instruments, such as accuracy 
and data rates. Total power usage is also an important consider- 
ation due to the limitation of the Orbiter ATCS. Program cost 
is a prime consideration in any area. Weight, size and volume 
will probably not be critical factors although they may be the 
deciding factors if all else among the options are equal . 


8-26 


8.3.3 PROGRAMMATICS 

Many major trade-off considerations of a programmatic nature are 
evident from this ASF pallet-only mode Orbiter mission. However, 
th'ese considerations are not unique to the pallet-only mode and 
must be addressed for any ASF mission configuration. Information 
is available now for some; additional information is required for 
many others. 

i 

1 

Perhaps the most important trade-off to be considered at this 
time is one related to a timely modification to the Orbiter elec- 
trical wiring design to reduce the EMI contamination to something 
more compatible with ASF and AMPS instrument requirements. 

Appendix U-4 addresses the problem and an improvement technique 
(for a price) which should reduce the Orbiter's ac magnetic field 
by about three orders of magnitude. This study has addressed 
the problem and derived a requirement for the use of an AE type 
subsatellite to remove the more critical instruments away from 
the Orbiter's EMI contamination environment. This concept also 
has a price, and there still remains a question of the remaining 
ASF instruments being operable in the cargo bay without an in- 
tolerable degradation of scientific return. 

This problem will be magnified many times during planning for the 
MPS experiments because of the design and operation characteristics 
of the MPS particle instruments. 

Candidate trade-offs appear productive in the following areas. 

a. The price of rewiring power cables in the Orbiter cargo bay 
and the cost of analytical models of the resulting wideband 
EMI environment in which all future payload bay-mounted 
scientific instruments must operate. 

b. The cost of analytically modeling the non-modif ied , wideband 
EMI environment throughout the bay; information which is re- 
quired before the feasibility of future payloads can be 
established. 


8-27 


c- The projected costs of individually shieldingj as applicable, 
each scientific and/or subsystem to be flown on future STS 
flights. 

d. An overall assessment of the applicability of the STS as a 
platform for scientific payloads. 

A major trade-off treatable at this time is that of the 1981 
launch. Many major, prime instruments will not be available. 

One trade-off factor is a start date of January 1976 instead of 
1977. However, this will not provide a satisfactory probability 
margin that all prime instruments will be ready. Quite obvious 
then, is the trade-off factor of substitute instruments. How- 
ever, there are no adequate substitutes for some of the long 
lead, major instruments. The third factor in this area is the 
inadequacy of a 28° inclination orbit of the 1981 launch in 
satisfying the scientific requirement for global coverage. The 
final trade-off consideration is that of a 1 983-85 launch v/hich 
would achieve two goals: (1) higher probability of all prime 

instruments, and (2) a polar orbit which would result in one 
mission to achieve all ASF scientific objectives. 


8.4 TECHNICAL FOLLOW-UP REQUIREMENTS 


Although study results indicate functional feasibility of the 
conceptual ASF payload design, more accurate capacity and sizing 
definitions are required in most areas (e.g., the quantity of 
cryogen required to cool instruments to 4.0 K and the memory^ 
executives I/O and ^software capaci ty and timing capability re- 
quired to perform extensive onboard data processing). In order 
for capacity and sizing to be further defined, many details of 
the design and operation of the various instruments are required 
(e.g., detector and housing design for cryo-cooled instruments, 
and total payload data characteristics and timelines affecting 
data processing). 


Therefore, the first priority of follovz-up efforts should be to 
define in greater detail a comprehensive set of requirements for 
experiments, instruments, subsatellite and support subsystems. 
This effort should include the generation of a more detailed 
mission timeline for experiment, instrument and subsystem opera- 
tions then that developed to date. 

The second priority for follow-on efforts is to provide better 
and more comprehensive design and operational definitions of the 
instruments and subsystems. For example, the detector holding 
structure and instrument housing for the cryogenical ly-cool ed 
instruments should be defined in some detail and the complemen- 
tary operations of two or more instruments, and the operational 
constraints of each instrument, should be defined. 

The third priority is to perform various analyses and trade-off 
studies to verify the preliminary selections or to update the 
design and operations with more optimum approaches. 

The fourth priority for the follow-on efforts i s to generate 
preliminary design and operational specifications which will be 
used as a basis for downstream development. 

The last priority is to develop programmatic factors such as 
estimates of total program development, production, and opera- 
tional costs; funding plans including expenditures by phases, 
allocation of resources, funding constraints and optional expen- 
diture approaches; development, production and operational sche- 
dules including expected critical paths and availability of 
non-ASF support such as the Orbiter, the SPS, the TORS system, 
etc.; development, production and operational plans for each 
major program element (e.g. , flight hardware, flight software, 
ground support facilities and ground support software); and an 
analysis of the technical, cost and schedule risks involved with 
full scale development. 


8.b UNRESOLVED f^lAJOR ISSUES 


During the course of this study, initial concepts and approaches 
were selected in the development of a pallet-only mode ASF mission 
utilizing the STS which required modification at a later stage of 
the study. Some were major in scope and others minor. A prelim- 
inary mission timeline resulting from limited definition of the 
experiment and instrument requirements was developed and subse- 
quently updated. As appreciation of the STS contamination envir- 
onment developed, a PDS and a boom-mounted equipment design were 
implemented. This evolution did result in a conceptual functional 
design considered technically feasible, but with certain qualifi- 
cations stemming from key assumptions developed along the way. 

The validity of some of the assumptions could not be fully veri- 
fied. As a result, several potentially significant issues remain 
which warrant identification at this time, and require future 
investigation. These issues are not unique to the pallet-only 
mode and must be resolved for any ASF mission configuration. 

a. Upon receipt of the forthcoming upgraded set of AMPS/ASF 
experiment/! nstrument requi rements from the SDWG , revised 
mission timelines will be required to establish operational 
boundaries. These boundary timelines will then be used to 
complete the task of sizing the ASF system, followed by a 
reassessment of the ASF design concepts relative to the new 
timeline. Particular emphasis should be given to the aft 
crew station, command and data management, power, and thermal 
subsystems for probable impacts. 

b. There is need to operate the particle detector instruments 
at a relatively short distance away from the Orbiter contam- 
ination environment. The AE satellite was chosen for two 
primary reasons. It is presently operational and the 
instrument complement requires minimal change. There are 


obviously many unresolved problems associated with this 
approach : 

(1 ) What Is the overall Impact to operational and safety 
aspects of the Orbiter during release, deployment, and 
retrieval of this suhsatel 1 ite? 

(2) How do the above impacts compare with those r a tethered 
satellite? 

(3) Would it be feasible to modify the proposed subsatellite 
to remain in orbit and. possibly be used for other 

s c i e n t i f i c tn i s s i 0 n s ? 

(4) How practical is the boom concept to implement in view 
of the requirement for STS attitude changes? Potential 
boom dynamics problems warrant further investigations 
related to safety and scientific, as well as the opera- 
tional factors. 

With the above instruments deployed away from the payload bay, 
a valid question exists as to the operation of the ASF instruments 
remaining in the cargo bay being compatible with the EMI and 
contamination environments. A comprehensive analysis is re- 
quired for an answer to the question. If the answer is neg- 
ative then this subject becomes a problem of the highest 
priority because there is a limit to how much of a sophisti- 
cated payload, such as the AMPS, can be deployed away from 
the Orbiter. 

Several assumptions were made during this study related to the 
data management philosophy. The resulting subsystem is 
technically feasible but it does approach a marginal capability 
and flexibility. A detailed Took at the philosophy provides 
insight to the situation which could very easily become a 
problem if even a slight increased demand were made of its 
function. However, this problem is not unique to the pallet- 
only mode and must be addressed for any ASF mission configuration. 



The operational philosophy of data management for AMPS(ASF) 
pallet-only mode places a larger demand on automation of 
mission conduct than does the pressurized module approach. 

This automation is accomplished through the use of the experi- 
ment and subsystem computers located in the igloo. The 
practicality of providing computers of adequate capacity to 
accomplish the total task must be further assessed. In 
addition a mid deck monitoring station could be provided 
that would reduce the required level of automation approaching 
that of the pressurized module approach. 


9.0 RECOMMENDATIONS 


9.1 ASF PAYLOAD SYSTEM DESIGN 

Results of this study warrant the follovnng design recommenda- 
tions for an ASF pallet-only mode payload system. These recom- 
mendations incorporate an extensive use of Spacelab and Orbiter 
equipment and approaches. Although follow-on efforts are re- 
quired to better refine the design concept, the recommended con- 
figuration establishes a feasible baseline from which to initiate 
a preliminary system design study. These design recommendations 
are presented by subsystem. 

9.1.1 TDRS 

a. Use ESRO furnished pallets for instrument and support equip- 
ment installation. 

b. Cluster instruments by pointing requirements. 

c. Use independent instrument pointing systems to achieve 
desired accuracy. 

d. Use open loop cryogenic cooling because of excessive power 
required by closed loop cooling systems. 

e. Use Orbiter ATCS with the addition of the heat radiator kit 
and ESRO approach for active thermal control. 

f. Use the BI-STEM configuration for the deployable booms. 

g. Use the collet/cold gas velocity separation mechanization for 
the retention and ejection of the subsatellite. 

9.1.2 EPOS 

a. Use Orbiter dedicated fuel cell for primary power source. 

b. Use two Orbiter energy sets to meet ASF requirements of 

897.3 kWh. 


9-T 


c. Use ESRO furnished dc/ac inverters power control box, primary 
and secondary power distribution boxes. 

d. Use remotely controlled circuit breakers except where safety 
may be involved. 

e. Use a high capacity capacitor bank and power converters to 
provide the 30 kV voltage required. 

f. Use conventional heavy gauge wires for power and return lines 

9.1.3 PCS 

a. Use an independent ASF AMS consisting of a gyro reference 
and siar tracker update. 

b. Use the centralized reference approach with optical transfer 
of attitude from Pallet A-3 to Pallet A-1 . 

c. Use Orbiter for only coarse instrument pointing. 

9.1.4 CDMS 

a. Use ESRO furnished computers, I/O's, mass memory, C&W elec- 
tronics and RAU's. Use Orbiter designed tape recorder, key- 
board and CRT display, and modified display electronics 
unit. Use Orbiter furnished TV cameras and monitors. Use 
additional ASF supplied, Orbiter designed TV cameras, if 

required. Use bi-phase L Manchester coded PCM data bus ' | 

approach. 

b. Perform as much of the data processing with the onboard 
computers as is practical. 

c. Operate the instruments, subsatellite and subsystems with the 
onboard computers to the maximum extent possible. 

d. Use Orbiter baseline S band FM link for communication with 
STDN and Orbiter Ku band for communication with TDRSS. 


9-2 


e. Use Orbiter baseline payload S band PM link for communication 
with subsatellite. 

f. Use mass memory for temporary storage of operational 
sequences with real time ‘reloading prior to next set of oper- 
ational routines or programs. 

9.1.5 AFT CREW STATION SUPPORT 

a. Use all three standard panels at PSS to support ASF mission. 

b. Use one PS at any given time. Support of two crew members 
will be required to man the station 24 hr/day. 

c. Limit the PS functions primarily to initiating and interrupt- 
ing programmed sequences, checking initial conditions, per- 
forming limited manual operations, analyzing and making de- 
cisions for off nominal conditions, and performing real time 
updates and changes to sequences. 

d. Provide a limited number of manual controls at the PSS in- 
cluding a manual fine-pointing capability to point instru- 
ments and control power to equipment and instruments. 

9.1 .6 PDS 

Use AE satellite as baseline. Delete eight of the 17 AE instru- 
ments and add two new ones (low energy ion detector and high 
energy particle detector). Add three rechargeable batteries 
to improve duty cycle capability. Delete the tape recorders. 

9-2 FOLLOW-ON STUDY 

The need for foTlow-on study efforts was addressed in paragraphs 
8.4 and 8.5. Two major study areas were identified and rationale 
presented for follow-on study efforts. The first contained 
several unresolved major issues which must be addressed because 
they not only constrain technical effectiveness of this concep- 
tual payload but they also involve major cost and schedule 
impacts to an ASF pallet-only mission. 


The second category of recommended study efforts relate to 
certain facets of this conceptual design. These design features 
need more detailed definition since they too offer promise of 
increased technical efficiency and cost/schedule effectiveness 
in an ASF payload system. 

9.2.1 UNRESOLVED MAJOR ISSUES 

a. The preliminar;/ nature of the ASF ID's and ED's used as a 
baseline for uh's study created an unresolved issue neces- 
sitating the following studies: 

(1) Using the upgraded ED's forthcoming from the AMPS SDWGj 
develop upgraded ASF mission timelines. The new time- 
linesj utilizing the new ED's and revised ID's, should 
be analytically exercised by the conceptual payload sys- 
tem to verify continuing feasibility of the payload con- 
cept with a more realistic ASF pallet-only mode mission. 

(2) Choice of instruments. Because of the unavailability of 
some ASF instruments for a mid-1981 launch date, it is 
recommended that a study be conducted with the following 
objectives. 

• Search for availability of instruments that can be 
used in lieu of those prime instruments presently 
described that cannot meet launch date and for which 
substitutes are not identified. Such instruments 
could be currently under development by either Govern- 
ment or industry, and could be completed in time to 
meet the scheduled launch date. Assess the impact 

to scientific value from the use of substitute and/or 
alternate instruments. 

• Explore alternate means of acquiring desired scien- 
tific information without the use of those instruments 
that cannot meet launch date and for which there are 
no substi tutes . 







t 


9 



• Assess scientific and cost impacts of flying certain 
eKperiments during 1981 and deferring others until 
requisite instruments are available. 

b. EMI assessment. Some of the ASF instruments are extremely 
sensitive to EMI (see paragraph 5.4). In order to assess 
the total impact of this problem, the levels of EMI expected 
to be generated by the Orbiter and the payload should be 
established. In parallel, the susceptibility levels (using 
conventional EMC design practices to reduce susceptibility) 
of instruments should be established. The effect on instru- 
ment measurements of the expected EMI levels should then be 
evaluated. If the effects are not acceptable, evaluation 
should be performed on the practicality of incorporating 
methods of reducing both EMI generation and susceptibility, 
including possible changes to existing Orbiter systems. 

c. Electrostatic charge assessment. The amount of electrostatic 
charge expected on the surface of the Orbiter vehicle and the 
payload structures should be established. In parallel, the 
maximum charge acceptable for particle accelerator operations 
should be established. If the two are not compatible, vari- 
ous possible means of reducing the charge potential of the 
vehicle and payload should be evaluated. Reference is made 
to appendix C. 

d. Particle contamination evaluation. The expected particle 
contamination from the Orbiter and from the payload should 
be established. Water vapor, cryogenic coolant gases, leak- 
ing fluids, subliming solids and other outgassing products 
can affect the validity of certain experiments. In a par- 
allel effort, the susceptibility levels of instruments 
sensitive to each of the expected contaminants should be 
established. Depending on the seriousness of the problem, 
changes to materials, methods of reducing the production rate 
of contaminants, and, the possible time sequencing of 


I 


I 

i 


i 




I 




A 

■I 

I 




9-5 


experiments, instruments and support equipment operation to 
minimize the impact of the contaminants should be assessed. 

e. Study the overall issue of the use of booms, subsatellites, 
tethered satellites, or other concepts to cope with problems 
posed by the operation of AMPS particles instruments. This 
study should encompass the following factors: 

(1) All Orbiter interfaces {physical, operational, etc.) 

(2) Gross cost factors 

(3) Scientific merit 

(4) Program schedules 

(5) Boom structural analyses. 

9.2.2 ASF PAYLOAD SYSTEM DESIGN 

For each of the support subsystems, those areas requiring follow- 
on primary emphasis were identified in section 8.0. The following 
specific studies are recommended as follow-up efforts. 

9 . 2 . 2 . 1 TSMS 

a. Cryogenic cooling system requirements definition. In order 
to define the cryogenic cooling system, requirements should 
be further defined. The designs of instruments 118 and 126 
in the areas of detector installation, housing and associated 
structure should be defined in enough detail that a meaning- 
ful heat load analysis can be performed. 

b. Cryogenic cooling system trade-offs. After the heat load 
analysis is completed, the open and closed loop cooling sys- 
tem design parameters (flow rates, quantity of cryogen, power 
required, system weight, size, etc.) should be established, 
environmental impacts such as contamination should be eval- 
uated, and development risk and other programmatics should 
be. assessed. Assuming both approaches meet basic ASF ... 

■■ ' 9 - 6 ''''" 


criteria, a trade-off study should be conducted to determine 
the most cost effective approach. 

c; Payload static and dynamic loads analysis. The structural 
and mechanical provisions for instrument, APS equipment and 
pallet installation should be further defined. Preliminary 
static and dynamic loads analyses should be performed to 
determine design margins available. The results of ESRO/ERNO 
analyses should be utilized, as applicable. 

d. ATCS definition. The thermal energy dissipation expected 
from instruments and support equipment should be further 
established. A thermal analysis should be performed to 
determine heat transfer characteristics and to determine the 
heat loads expected at the active thermal control interface. 

Thermal capacitors and cold plate requirements should be 
established and coolant loop characteristics including choice 
of fluid and flow rates should be established. The number 
of pumps, valves and heat exchangers, the routing of the 
coolant, the need for flexible conduits, and other design 
features should be defined at the preliminary design stage. 

e. Boom dynamics analyses. The capability of the boom, and the 
instrument mounted at the end of the boom, to withstand 
vehicle dynamics and the Instrument 550 scanning operation 
on instrument alignment accuracy should be analyzed. If 
necessary, constraints on Orbiter maneuvering and limit 
cycle acceleration should be established. 

9.2.2.Z EPPS 

a. High voltage, high power source definition. There are many 

issues involved with the use of high voltage (up to 30 kV), * -I 

high power (5 to 10 kW) sources. Each one of these should | 

be evaluated in some detaiT since this capability is funda- I 

mental to three of the ASF instruments (213, 303, and 304). I 

Trade-offs should be conducted to determine the most effec- | 

tive approach for the generation of the required voltages. | 


Design features such as best transmission media for the high 
current (200 amps), optimum location of the high voltage 
sources relative to the instrument, adequacy of available 
insulation techniques to eliminate breakdown and corona 
effects, and adequacy of EMC techniques to minimize effects 
of radiated and conducted EMI generated by the high voltages 
and current, should be fully evaluated. 

b. Inverter trade-offs. The need for a centralized 400 Hz 

inverter should be evaluated. The effectiveness of a cen- 
tralized system compared to individual inverters provided 
by the using instrument or equipment should be assessed. 

9. 2.2.3 PCSS 

a. AMS trade-offs. Trade-offs should be performed on the 
effectiveness of a centralized versus distributed AMS. An 
AMS located on Pallet A-1 using optical media to transfer 
attitude reference to Pallet A-3 simplifies the system. In- 
dividual star trackers and gyro packages on each AIM dr APS 
improve attitude accuracies. 

b. Equipment trade-offs. Trade-off studies should be conducted 
on most applicable gyro packages, on selection of strapped 
down or gimballed star trackers, and the use of the Orbiter 
GN&C system to calibrate and improve the accuracy Capability 
of the ASF system . 

9. 2. 2. 4 CDMS 

a. Data processing requirements definition. The instrument and 
support subsystem data processing requirements should be 
further defined . 

b. Experiment, instrument and subsystem operations definition. 
The payload operational functions and timelines should be 
defined in greater detail. 


9-8 



IV 


t 


4S 


c. Processing system definition. Based on better definition 

of the data processing and operations requirements » computer, 
I/O, and software sizing and timing analyses should be per- 
formed. The analyses results would be the basis for estab- 
lishing computer and I/O design and selection criteria. 

Based on these criteria, trade-off studies should be con- 
ducted to compare effectiveness of centralized versus dis- 
tributed processors, and to select the processors most com- 
patible with ASF program requirements. 

d. Mass memory utilization assessment. The issue of using the 
ASF mass memory for temporary processing routine storage or 
for permanent storage with full mission operational capa- 
bility should be resolved. Based on the payload operational 
requirements previously established, an assessment should be 
made to determine mass memory storage capacity required to 
provide the full mission operational program capability with- 
out recourse to crew update. The memory size required for 
this approach should be assessed against the experiment, 
instruF ..nt, and support subsystem mission timelines and the 
possible constraints imposed by dependence on ground opera- 
tions and availability of Orbiter or ground facilities for 
communications at the required time. 

e. Data transmission compatibility assessment. Based on the 
data processing requirements established previously (includ- 
ing that for the fixed payload and the deployed subsatellite) 
a determination should be made as to whether the margins 
between the requirements and capabilities are adequate. If 
data compression is required, trade studies should be con- 
ducted to determine the most cost effective approach! 


I 


-j jiiot 


9-9 


9. 2, 2, 5 Aft Crew Station Support 


a. C&D requirements definition. The C&D support required by the 
individual instruments for the subsatellite and the support 
subsystems should be further defined. 

b. PS operations definition. The role of the PS in controlling 
the experiments and other payload operations and in monitor- 
ing and assessing the displayed data should be further 
evaluated with scientific community participation. The need 
for a second (mid-deck) work station to expand the aft flight 
deck capability should be examined. Based on the updated 
definition of the PS functions and responsibilities, the 
need for additional specialist/crew member support should be 
defined . 

9. 2. 2. 6 PDS 

An assessment of the economic impact on the total ASF program 
should be made if the subsatellite was left in orbit rather than 
retrieved subsequent to each mission. Subsatellite retrieval 
significantly complicates mission operations and crew training 
and more importantly, increases the possibility of jeopardizing 
the safety of the crew and the vehicle. The total ASF (and AMPS) 
mission traffic should be analyzed to determine if a single 
deployed subsatellite might be able to support a number of sep- 
arate ASF launches over an extended period of time. 

9.2.3 CONCEPTS OF STANDARDIZING 

The concepts of centralizing and standardizing described in 
appendix D of this report will be applied to the ASF configura- 
tion to determine the savings offered in areas of reliability, 
schedule, cost, etc. 


9-10 


10.0 BIBLIOGRAPHY 


10.1 AUTHORITATIVE DOCUMENTS 

1. Space Shuttle System Payload Accomodations. JSC 07700, 

Volume XIV, Revision C, July 1974. 

2. Space lab Payload Accomodations Handbook. ESTEG SLP/2104, 

May 1975. 

3. Orbiter 102. Feb 1975 PDR, Payload Interfaces Team 

Documentation, Mission Kits, JSC 09325, February 1975. 

4. Orbiter 102, Feb 1975 PDR, Payload Interfaces Team 

Documentation, Avionics Baseline, JSC 09320, February 1975 

5. Orbiter 102. Feb 1975 PDR, Payload Interfaces Team 

Documentation, Aft Flight Deck Payload Accomodations, 

JSC 09321 , February 1975. 

6. Crew Station Specifications. MSC 07387, October 1972. 

7. Payload Interfaces Team Documentation. Payload Interface 

Drawings, JSC 09326, February 1975. 

8. Electromagnetic Compatibility Requirements, Systems for the 

Space' Shuttle Program. JSC Specification SL-E-OOOl , 

June 1973. 

9. Electromagnetic Interference Characteristics, Requirements 

for Equipment for the Space Shuttle Program. JSC 
Specification SL-E-0002, June 1973. 

10. Space Shuttle Lightning Protection Criteria. JSC 07636, 

September 1 973 . 

11. Space Shuttle Flight and Ground System Specification. JSC 

07700, Volume X, Revision A, January 1974. 


lO.Z REFERENCE DOCUMENTS 

1. Phase "A" Conceptual Design Study of the Atniospheiric , 

Hagnetospheric , and plasmas in Space (AMPS) Payload. 

NASA TM-Xt^ 64895, November 1 974. 

2. An Assessment of the Instrument Pointing Subsystem (IPS) 

Requirements for Spacelab Missions. NASA TM X-64896j 
November 1974. 

3. Skylab Operations Handbook, MSC 01549, Volume IV, October 

1 972. 

4. Skylab Pointing Techniques. MSC 07220, October 1972, 

5. Preliminary Concepts from Woods Hole Atmospheric and Space 

Physics, An Informal In-House Study. October 1973. 

6. Shuttle Orbiter IMU and Star Tracker. Rockwell International 

Report Number Sk-72-SH-Ol 01 -2. 

7. Apollo Guidance and Navigation Systems Manual. AC Electronics, 

September 1 968. 

8. Preliminary Design Study for an Atmospheric Facility. 

Martin-Marietta Aerospace Report MCR-72-322, Contract 
NAS 9-12255, December 1972. 

9. Ramos, Daniel 0.: Collet Release Mechanism. General Electric 

Company (Aerospace Mechanism, Volume I, Part A. General 
Applications, pp. 163-.70. Distributed by J. W. Stacy, 

Inc., a Subsidiary of Bro-Dart Industries). 

10. Shapiro, H.: Visual Appearance of Polymeric Contamination. 

NASA Technical Note TN D-5839, 1970. 

11. Scialdone, J. J. : Predicting Spacecraft Self-Contamination 

in Space and in a Rest Chamber. NASA Special Publication 
SP-298, Space Simulation, 1972, pp. 349-360. 

12. Nimbus 3 User's Guide. National Space Science Data Center, 

GSFC, Greenbelt, Md., 1969. 

13- Nimbus 4 User's Guide. National Space Science Data Center, 
GSFC, Greenbelt, Md., 1970. 

14. Human Engineering Guide to Equipment Design. Joint Army- 
Navy-Air Force Steering Committee, American Institute 
for Research, Superintendent of Documents, Washington, 

D.C., Revised Edition, 1972. 


10-2 


15. Atmosphere Explorer (AE) Spacecraft System Description. 

RCA Government and Commercial Systems? Astro Electronics 
Division Report AED Rtv 3816F, March 1972? Attachment B 
(Updated August 1974), Contract NAS 5-23003. 

16. GSFC Specification for Atmosphere Explorer (AE-C, D, and 

E). GSFC Specification S-620-P-1 , September 1973. 


10.3 INFORMATION DOCUMENTS 

1. Crawford, R, F.; Strength and Efficiency of Deployable 

Booms for Space Application. Astro Research Corporation, 

April 1971. 

2. Heaney, J. B.'. Results from the ATS-3 Refl ectometer 

Experiment, Thermophsics, Applications to Thermal 
Design of Spacecraft. PIAA, Vol . 23, pp. 249-274, 

J. Bevans, Ed., Academic Press, New York, 1970. 

3. McKeown, D.; and W. F. Corbin; Space Measurements of the 

Contamination of Surfaces by (G)-6 Outgassing and Their 
Cleaning by Sputtering and Desorbtion. NBS Special 
Publication 336 , Space Simulation, J. C. Riclimond, ed., 
pp. 113-127, 1970. 

4. Heath, D. F.; and Sacher, P. A.,; Effects of a Simulated High 

Energy Space Environment on the UV Transmittance of Optical 

O O 

Materials Between 1050 A and 3000 A. Applied Optics 5, 

937, 1966. 

5. Heath, D. F. i and McElaney, J. H.: Effects of High Energy 

Particle Environment on the Quantum Efficiency of Spec- 
trally Selective Photocathodes for the Middle and 
Vacuum UV. Applied Optics 7, 2049, 1968. 

6. Hass, G.; and Hunter, W. R.: Laboratory Experiments to Study 

Surface Contamination and Degradation of Optical Coatings 
and Materials in Simulated Space Environments. Applied 
Optics 9, 2101 , 1970. 

7. Shapiro; H., and Hanyok, J.; Monomol ecular Contamination 

of Optical Surfaces. Vacuum 18, 587, 1968. 

8. Bauer, S. J. ; Brace, L. H.; Grimes, D. W.; and Spencer, N. W.; < 

A Plan for the Use of the Atmosphere Explorer Spacecraft | 

System as a Subsatellite of the Space Shuttle. Goddard 
Space Flight Center, Greenbelt, Md., May 1973. 


9. Cothran, C. A.; McCa.rgoj M-; and Greenberg, S. A.: A Survey 

of Contamination of Spacecraft Surfaces. AIAA Paper No. 
71-157, presented at the 6th Thermophysics Conference, 
Tullahoma, Tenn, April 1971. 



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