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v1 ASTROPUSICAL PAM# ACCOMMODATION

N TE VACE STATION

RM MMIT U*X 124

(NASA - CF -178 56) ASIbOPHYSICAL PAYLOAD N86-13359

ACCOMMOCklICN Ch 11HE SEALE SIATICN Fiadl

H*port, 10 Occ. I584 - 15 Oct. 1985 (Allied

• beadAx Aerospdct) 164 p HL AC8/MF A01 U ac 1,1-,

CSCL 22b G3/18 154 1 ) r, 000

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Bendix Guidance Systems Division

AL

A WFN Eli
The

Allied

Corporation

Bendix Aerospace

Sector

Guidance

Systems

Division

Teterboro,

New Jersey 07608

Astrophysical Payload

Accommodation On The

Space Station

Final Report

15 October 1985

Prepared For:

George C. Marshall

Space Flight Center

Huntsville, Alabama

NASA contract No.

NAS8-36124

December 10, 1984

To

October :5, 1985

APPROVED BY: PREPARED BY:

4}E J. vinthal Bernard P. Woods

Engineering Manager Study Manager . Systems Design
FOREWORD

This final report is submitted in accordance with the requirements of Contract NAS8-

36124. This deliverable item is provided in direct response to Section H, Article H-3,

Item B, "Final Report" of Contract NAS8-36124. The study was performed at the Allied

Bendix Guidance System Division's facility in Huntsville, AL. Most of the effort in

support of this study was pei formed by Bernard P. Woo ]s, who prepared all sections of

this report. 'The study was directed from the Guidance Systems Division in Teterboro,

NJ. Mr. Joel Leviathal Is the engineering manager at this location. The guidance of Mr.

Max Nein and Mr. Joe T. Howell of MSFC during the course of this study is gratefully

acknowledged.

This study is one of a group of studies dealing with accommodation Cf astrophysics

payloads on the Space Station which were performed under the general direction and

guidance of Dr. David Gilman and Mr. Michael Kiya of NASA Headquarters, Code EZ.

The other studie- ;n the series were performed by the Jet Propulsion Laboratory, and the

GSFC. The subject of these other studies included pointing requirements and

technologies, contamination requiements, data requirements, and accommodation of the

Solar Optical Telescope on the Space Station. These studies were performed to support

the Space Station Definition and Preliminary Design activity which is presently being

performed by 14ASA and associated contractors.

I qlkl

i
ABSTRACT

Surveys of potential Space Station Astrophysics Payload requirements and existing

pointing mount design concepts were performed to Identify potential design approaches

for accommodating Astrophysics Instruments froin Space Station. Most existing

instrument pointing systems were designed for operation from the Space Shuttle and it Is

unlikely that they will sustain their performance requirements when exposed to the Space

Station disturbance environment. The study indicates that the technology exists or is

becoming available so that precision pointing can be provided from the Space Station

manned core. Development of a "disturbance insensitive" pointing mount is the key to

providing a generic system for Space Station. It is recommended that the MSFC

Suspended Experiment Mount concept be investigated for use as part of a generic

pointing mount for Space Station.

Availability of a shirtsleeve module for instrument change out, maintenance and repair is

desirable from the user's point of view. Addition of a shirtsleeve module or Space

Station would require a major program commitment.

ow
a

TABLE OF CONTENTS

T1 LE

INTRODUCTION .....................................

SUMMARY.........................................

ASTROPHYSICS PAYLOADS .. ........................ .

ASTROPHYSICS PAYLOAD REQUIREMENTS ................

THE SOLAR OPTICAL TELESCOPE .......................

THE PINHOLE OCCULTER FACILITY ......................

PRESENT POINTING MOUNT CONCEPTS ...................

SK YLAB ATM .......................................

SMALL INSTRUMENT POINTING SYSTEM (SIPS) ..............

MINIATURE POINTING MOUNT ............. ........... .

ANNULAR SUSPENSION AND POINTING SYSTEM (ASPS) ...... .

THE ADVANCED GIMBAL SYSTEM ... . ................... .

MODIFIED ATM STAR TRACKER .. , ..................... .

THE INSTRUMENT POINTING SYSTEM .....................

GIM BALFLEX .............. .............. — * ...****

SUSPENDED EXPERIMENT MOUNT .......................

TALON GOLD GIMBAL ................................

HUBBLE SPACE TELESCOPE ............................

2PACE STATION DESCRIPTION ..........................

SPACE STATION CONFIGURATION SUM"rIARY ..............

PAYLOAD ACCOMMODATION ...........................

SPACE STATION ORIENTATION ..........................

ATTITUDE CONTROL SYSTEM DESCRIPTION ...............

SPACE STATION DYNAMICS ............................

DISTURBANCE SOURCES ........... *is* , — * * * , , .... —,

ENVIRONMENTAL DISTURBANCE SOURCES ............... .

DISTURBANCES DUE TO OPERATIONS AND

INTERNAL SOURCES ..............,.................

EFFECT OF ASTRONAUT D[STU pPANCE AND

ORBITER ATTACHMENT . , ..... .................. .

POINTING MOUNTS FOR SPACE STATION ..................

INSTRUMENT FOCAL PLANE SHIRTSLEEVE ACCESS ..........

IOC SPACE STATION "SHIRTSLEEVE ENVIRONMMENT..........

IOC SPACE STATION SERVICING FACILITY .................

PRESSURIZED HANGAR DESIGN CONSIDERATIONS ..........

PAYLOAD DESIGN CONSIDERATIONS FOR

SHIRTSLEEVE SERVICING .............................

CONCLUSIONS AND RECOMMENDATIONS .................

.'APPENDIX 1

APPENDIX 2

ABBREV' ATIONS

REFERENCES

SECTION

1.0

1.1

2.0

L.1

2.1.1

2.1.2

3.0

3.1

3.2

3.3

3.4

3.5

3.6

3.7

3.8

3.9

3.10

3.11

4.0

4.1

4.2

4.3

4.4

4.5

4.6

4.6.1

4.6.2

4.6.2.1

5.0

6.0

6.1

6.2

6.3

6.4

7.0

FAGE

1

1

3

7

11

15

17

20

24

29

32

36

40

46

53

56

62

66

75

75

79

80

82

86

88

87

87

92

96

104

106

107

107

110

112

iii

0
TABLE OF CONTENTS (CONTINUED)

LIST OF ILLUSTRATIONS

FIGURE NO. TITLE PAGE

1 AGO ON THE SPACE STATION UPPER BOOM ........... 12

2 SOT-1 GENERAL CONFIGURATION .................. 14

3 PINHOLE OCCULTER FACILITY ....... , ............ 16

4 SKYLAB ORBITAL ASSEMBLY ...................... 21

5 ATM FINE POINTING CONTROL MECHANISM .......... 21

6 ECPS BLOCK DIAGRAM .......................... 23

7 SIPS CONFIGURATION ........................... 26

8 SIPS GIMBAL ARRANGEMENT ...................... 27

9 SIPS BLOCK DIAGRAM ........................... 28

10 MPM CONFIGURATION ... ........................ 30

11 MPM FUNCTIONAL BLOCK DIAGRAM ................ 31

12 ANNULAR SUSPENSION AND POINTING SYSTEM ........ 33

13 ASPS INTERFACES .............................. 35

14 AGS ELEMENTS ................................. 37

15 AGS CONTROL AND DATA INTERFACES .............. 39

16 AEPI PALLET MOUNTED EQUIPMENT ................ 41

17 AEPI RACK MOUNTED EQUIPMENT ................. 43

18 IPS ASSEMBLY ...... ........................... 47 

19 IPS ELECTRICAL AND D ATA INTERFACES ............ 50

20 IPS ATTITUDE CONTROL LOOP BLOCK DIAGRAM ...... 52

21 GIMBALFLEX -A VERNIER STAGE ................... 54

22 GIMBALFLEX MECHANICAL CONCEPT ............... 55

23 SEM CONCEPT AND PRINCIPLE .................... 57

24 SEM RETENSION AND SUSPENSION CONCEPT .......... 59

25 TALON GOLD SYSTEM ........................... 63

26 HST GENERAL CONFIGURATION ................... 67

27 HST EXPLODED VIEW ............................ 68

28 HST SUPPORT SYSTEM MODULE .................... 69

29 HST PCS FUNCTIONAL DIAGRAM ................... 71

30 HST PCS EQUIPMENT ............................

72

31 SPACE STATION REFERENCE CONFIGURATION ........ 76

32 RACETRACK MODULE CONFIGURATION 78

33 SPACE STATION ORIENTATION ..................... 81

34 AIRPLANE MODE ORIENTATION .................. , . 83

35 GN&C SUBSYSTEM BLOCK DIAGRAM ................ 85

36 NATURAL ENVIRONMENT DISTURBANCES 

TORQUE PROFILE ............................. 90

37 SWAY ABOUT THE X AXIS ......................... 91

38 EFFECT OF ASTRONAUT DISTURBANCES

AND ORBITER ATTACHMENT .................... 93

39 EFFECT OF CENTRALIZED CONTROL VERSUS

DISTRIBUTED CONTROL ON PAYLOAD POINTING ...... 95

iv
TABLE OF CONTENTS (CONTINUED)

LIST OF ILLUSTRATIONS (CONTINUED)

FIGURE NO. TITLE PAGE

40 ALPHA/BETA JOINTS FOR COARSE POINTING ......... 97

41 BETA JOINT FOR COARSE POINTING

ON TRANSVERSE BOOM ........................ 98

42 DISTURBANCE COMPENSATION METHODS ............ 101

43 A SIMPLE GIMBAL SYSTEM ........................ 103

44 SEM ADVANTAGES .............................. 105

45 ORBITER CARGO BAY ENVELOPE ................... 109

LIST OF TABLES

TABLE NO. TITLE PAGE

1 POTENTIAL ASTROPHYSICS PAYLOADS .................. 4

2 ASTROPHYSICS PAYLOADS POTENTIAL LOCATIONS ........ 6

3 ATTACHED ASTROPHYSICS PAYLOAD REQUIREMENTS ...... 8

4 NON—ATTACHED ASTROPHYSICS PAYLOAD REQUIREMENTS .. 8

5 SOT WEIGHT STATEMENT ............................. 9

6 POF WEIGHT STATEMENT ............................. 10

7 POINTING SYSTEMS SURVEYED ........................ 18

8 PRESENT POINTING SYSTEM CAPABILITIES ............... 19

9 POINTING REQUIREMENTS OF THE IPS ................... 48

10 SEM WEIGHT STATEMENT ............................. 61 

11 SPACE STATION ORBIT PARAMETERS ................... 80

12 DISTURBANCE SOURCES ............................. 88

13 POINTING CAPABILITY VERSUS

d

CUSTOMER SATISFACTION .......................... 99
1.0 INTRODUCTION

The purpose of this report is to present the findings and results of a :'udy entitled

"Accommodation of Astrophysical Instruments in the Space Station System" which was

performed by the Allied Bendix Aerospace, Bendix Guidance Systems Division (BGSD) for

the George C. Marshall Sp-ce Flight Center under contract NAS8-38124.

Although the term Astrophysics generally includes a study of the physical nature and

properties of the planets and stars, this study concentrated on the pointing requirements

of the solar and stellar instruments, which are potential candidates for Space Station,

since the earth looking instruments were being adequately covered by other NASA

activities.

The study had the following objectives and tasks:

A) To survey and organize astrophysical experiment pointing and control requirements

for Space Station and Space Platform missions, especially with respect to

anticipated pointing mount requirements;

B) To compare the experiment requirements to anticipate Space Station requirements

with respect to pointing mount design.

C) To consolidate the current concepts for pointing mount designs and innovative ideas

presently being pursued at NASA and contractor facilities.

D) To identify new pointing mount design approaches for accommodating future

experiments especially with respect to isolation mechanisms;

E) To identify potential techniques that will allow a "shirt sleeve" access to the

instrument focal plane structures.

1.1 SUMMARY

The Space Station is required to simultaneously accommodate instruments that have

Earth, anti-Earth, Solar and Stellar viewing requirements. Satisfaction of this

requirement from the Space Station flying in the baseline local vertical orientation will

i A
require a pointing system. The Space Station will provide some articulated attachment

points for externally mounted pointing Instruments. These coarse pointing/trarlking

gimbals will provide attitude cokitrol t--tween ±1 to ±5 degrees and attitude

determination to ±.01 degrees. Maximum angular rates would not exceed ±0.02

degrees/sec. In addition, tte Space Station will provide knowledge of the position of the

Instrument mounting surface relative to the navigation base to ±0.002 degrees. Users or

customers with Instruments that require higher pointing ace,. racy will be required to

provide their own pointing mounts.

The study surveyed the capabilities of eleven (11) past and present pointing concepts and

systems. Most of these systems were designed for a Shuttle application which has a more

benign disturbance environment than the Space Station environment is expected to be. It

Is unlikely that existing systems and cor. -pts will sustain their performance when

exposed to the Space Station disturbance environment. The study indicates that the

technology is either available or rapidly becming available to design a generic pointing

system for Space Station. Development of a disturbance insensitive pointing mount is

the key to providing a precise pointing mount for Space Station. With respect to pointing

mounts, the following conclusions and recommendations are made:

It is recommended that a trade study be performed to determine the cost effectiveness

of requiring the individual instruments to provide precise pointing versus the Space

Station providing a gerneric precise pointing capability.

An in±rtially reacting pointing system is less sensitive to base body disturbances than are

gimbal systems. It is recommended that the MSFC Suspended Experiment Mount (SEM)

be investigated for potential use as a pointing mount for Space Station.

Both the Spacelab Instrument Pointing System (IPS) and the Talon Gold Gimbal (TGG)

should be considered for use on Space Station because of the maturity of these systems

and to realize benefits from the investments previously made.

Availability of a module which will permit instrument change-out, maintenance and

repair in a shirt-sleeve environment is desirable from a customer's point of view.

Inclusion of a shirt-sleeve module would impact most Space Station subsystems. Addition

x of a facility of this type would require a major program commitment.

t

2
i^

It is recommended that a study be performed to determine the cost impact of providing a

"shirt-sleeve environment" maintenance and repair module versus the benefits gained by

the customer with respect to longer orbital lifetime, increased observing time and less

complicated EVA tasks.

2.0 ASTROPHYSICS PAYLOADS

Several sources were used to identify the potential Astrophysics Payloads and their

requirements. These are identified in references [ l] through ( 5). The major source of

information regarding potential payloads was the Langley Space Station Mission Data

Base, specifically, the Science and Applications Mission Data Base (SAAX). Both the

September 1984 and the May 1985 versions were accessed to identify potential payloads

and the payload requirements that would impact pointing mount design. Neither version

of the data base contains sufficient Information to perform detailed pointing mount

design; but, the information from the Langley Data Base (both versions) augmented by

other sources, such as the MSFC Pinhole Occulter Facility ( POF) Phase A Study Report

(61, does provide sufficient information to bound the problem. Table 1 identifies

potential astrophysics payloads and whether or not they were identified as such in

Reference ( 1) or in the two versions of the Langley Data Base. These payload lists are

intended to illustrate typical astrophysical payloads that the Space Station program may

have to satisfy. Some of the differences between the payloads identified in the two

versions of the Langley Data Base can be explained. The POF (SAAX0009) is included as

part of the Advanced Solar Observatory (SAAX0011). The X -Ray Timing Explorer

(SAAX0014), the Far UV Spectroscopy Explorer (SAAX0019), and the Solar Corona

Diagnostic Mission (SAAX0010) are the payloads for the Leased Platforms ( Explorer) 1

through 3 (SAAX0026, SAAX0027, SAAX0028).

The candidate payloads can be classified as: attached payloads, payloads located on coorbiting free flying platforms, and free flying satellites, which would be serviced by

Space Station. Table 2 identifies the performance location for each payload on the

composite list.

i

3
TABLE 1. POTENTIAL, ASTROPHYSICS PAYLOADS

REF. SEP. ' 84 MAY'85

CONF. LANGLEY LANGLEY

CODE PAYLOAD DOC. BASE BASE

SAAX0001 SPECTRA OF COSMIC RAY NUCLEI X X

SAAX0004 SHUTTLE INFRARED TELESCOPE FACILITY X X

SAAX0005 TRANSITION RADIATION do ION CALORIMETER X X

SAAX i^06 STARLAB X X X

8AAX0007 HIGH THROUGHPUT MISSION X X

SAAX0008 HIGH ENERGY ISOTOPE EXPERIMENT X X

SAAX0009 PINHOLE OCCULTER FACILITY X X

SAAX0010 SOLAR CORONA DIAGNOSTIC MISSION X

SAAX0011 ADVANCED SOLAR OBSERVATORY X X

SAAX0012 HUBBLE SPACE TELESCOPE X X X

SAAX0013 GAMMA RAY OBSERVATORY X X X

SAAX0014 X-RAY TIMING EXPLORER X X

SAAX0016 SMALL SOLAR PHYSICS ( NOTE 1) X X X

SAAX0017 ADVANCED X-RAY ASTROPHYSICS FACILITY X X X

SAAX0018 VERY LONG BASELINE INTERFER iMETER X

SAAX0019 FAR UV SPECTROSCOPY EXPLORER X X

SAAX0020 LARGE DEPLOYABLE REFLECTOR X X

SAAX0021 SUPER CONDUCTING MAGNET X

SAAX0022 SPACE STATION SPARTAN PLATFORM (NOTE 2) X X

SAAX0023 SPACE STATION SPARTAN PLATFORM 2 X

SAAX0024 SPACE STATION SPARTAN PLATFORM 3 X

SAAX0025 SPACE STATION SPARTAN PLATFORM 4 X

SAAX0026 LEASED PLATFORM (EXPLORER) 1 X X

SAAX0027 LEASED PLATFORM (EXPLORER) 2 X X

SAAX0028 LEASED PLATFORM (EXPLORER) 3 X X

SAAX0029 LEASED PLATFORM ( EXPLORER) 4 X X

SAAX0030 SPACE STATION HITCHHIKER 1 X

SAAX0031 SPACE STATION HITCHHIKER 2 X

SAAX0032 SPACE STATION HITCHHIKER 3 X

NOTE 1: IDENTIFIED AS SOLAR MAX MISSION IN SEPT. 1 84 LANGLEY DATA BASF.

NOTE 2: IDENTIFIED AS SOLAR SEISMOLCGY EXPERIMENT IN SEPT. 1 84 LANGLEY DATA

BASE.

r 

E' 

.^ "-va.»yr 
Attached payloads are those which operate a <.;-, fhed to the Space Station manned core,

but are external to the pressurized modules. Unmanned co-orbiting platforms will

accommodate one or more compatible Instruments. Co-orbit is used In the sense that the

orbital inclination is the same as the Space Station manned core. Free Flyers or Free

Flying Satellites refer to major facilities like the Hubble Space Telescope and other

satellites that may utilize the Space Station for maintenance and repair. The use of the

Orbital Maneuvering Vehicle (OMV) or the Orbital Transfer Vehicle (OTV) Is essential for

this purpose.

The Space Station Program Elements also include a polar platform. This unmanned

platform will accommodate those payloads which require a polar orbit. The payload

Identified for performance from the polar platforms support a strategy for obtaining

Integrated earth science measurements from low earth orbit. Polar Platforms and the

Instrument located on them will be serviced by the orbiter in an orbit that is within the

its capabilities.

Appendix 1 provides a summary description of each payload on tables 1 and 2. These

data were excerpted from the Langley Data Base.

ff

I

t,

5
TABLE 2. ASTROPHYSICS PAYLOADS POTENTIAL LOCATIONS

CO-ORBIT FREE

CODE PAYLOAD ATTACHED PLATFORM FLYER

SAAX0001 SPECTRA OF COSMIC RAY NUCLEI X

SAAX0004 SHUTTLE INFRARED TELESCOPE FACILITY X X

SAAX0005 TRANSITION RADIATION do ION CALORIMETER Y X

SAAX0006 STARLAB X X

SAAX0007 HIGH THROUGHPUT MISSION X X

SAAX0008 HIGH ENERGY ISOTOPE EXPERIMENT X X

SAAX0009 PINHOLE OCCULTER FACILITY X X

SAAX0010 SOLAR CORONA DIAGNOSTIC MISSION X

SAAX0011 ADVANCED SOLAR OBSERVATORY X X

SAAX0012 HUBBLE SPACE TELESCOPE X

SAAX0013 GAMMA RAY OBSERVATORY X

SAAX0014 X-RAY TIMING EXPLORER X

SAAX0016 SMALL SOLAR PHYSICS X

SAAX0017 ADVANCED X-RAY ASTROPHYSICS FACILITY X

SAAX0018 VERY LONG BASELINE INTERFEROMETER X X X

SAAX0019 FAR UV SPECTROSCOPY EXPLORER X

SAAX0020 LARGE DEPLOYABLE REFLECTOR X

SAAX0021 SUPER CONDUCTING MAGNET X

SAAX0022 SPACE STATION SPARTAN PLATFORM 1 X

SAAX0023 SPACE STATION SPARTAN PLATFORM 2 X

SAAX0024 SPACE STATION SPARTAN PLATFORM 3 X

SAAX0025 SPACE STATION SPARTAN PLATFORM 4 X

SAAX0026 LEASEI, PLATFORM (EXPLORER) 1 X

SAAX0027 LEASED PLATFORM (EXPLORER) 2 X

SAAX0028 LEASED PLATFORM (EXPLORER) 3 X

SAAX0029 LEASED PLATFOR.ti; (EXPLORER) 4 X

SAAX0030 SPACE STATION HITCHHIKER 1 X

SAAX0031 SPACE STATION H'TCHHIKER 2 X

SAAX0032 SPACE STATION HITCHHIKER 3 X

6
i

2.1 ASTROPHYSICS PAYLOAD REQUIREMENTS

Table 3 provides a listing of the key payload requirements for payloads which are

candidates to be performed attached to Space Station. Table 4 lists the requirement for

the payloads which are performed on co-orbiting platforms or are independent free

flyers. The listed requirements were derive + r-_.,. the Langley Data base, except in cases

where more definitive information was available. Some additional payloads, identified in

reference (5), were also included.

Caution should be taken when using summary descriptions like those contained in the

Langley Data Base. Often times pertinent characteristics or requirements are not

clearly identified in the Summary. For example, consider the Advanced Solar

Observatory (ASO), SAAX0011. The ASO consists of: the Solar Opt:?al Telescope (SOT);

the Solar Extreme U.V. Telescope Facility (SEUVTF); the Solar Soft X- Ray Telescope

Facility (SSXRTF)[7&8]; and the Pinhole Occulter Facility (POF). The SOT, SEUVTF and

the SSXRTF form the SOT cluster. Two pointing mounts are required for the ASO; one

for the SOT cluster, and the other for the POF. This requirement is not apparent from

the entries in the Data Base.

Tables 5 and 6 contain weight statements for the -)OT and the POF. These data were

excerpted from references [91,[5 ], and [6]. An examination of the SOT weight statement

in table 5 indicates that the SOT Observatory weight is between 4000 kg and 6600 kg.

Table 6 shows that the weight of the POF Instruments and mission peculiar equipment

would be 1 469 kg (3238 lb.). The Langley Data Base reports a total weight for the ASO

of 12,500 kg. The combined weight of the POF (1469 kg) and the SOT (4000 kg) is 5469

kg. This would leave 7031 kg for the SEUVTF and the SSXRTF. Requirements for tth9se

instruments have not been aefined, but a combined weight of 7031 kg for them appears

high. The data base does not indicate how the 12,500 kg for the ASO was arrived at.

Other examples can be cited. With respect to pointing requirements, the data base shows

that the ASO requires a pointing accuracy of 1.0 arc sec. and pointing stability of 0.1 are

sec. It is not clear if these requirements are those of the instrument or those of the

A

puinting system that the instrument is mounted on. In the case of the SOT, image

stability is provided for internally by precision control loops within the SOT in addition to

that provided by the pointing mount. This achieves a stability of .03 arc sec. The POF

pointing requirements, per reference [61, are a pointing accuracy of 30 arc sec. and a

stability of 0.1 arc sec.

7
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TABLE 5. SOT WEIGHT STATEMENT

(PER SP82 - MSFC-2583) Mass (Kg)

ITEM

SCIENCE INSTRUMENTS

Instruments and contingency/5046

Focal Plane

Soft X-Ray Telescope

XUV Telescope

INTEGRATION HARDWARE

Pointing System

Signal Interface Units (2)

Power Interface Units (2)

Freon Pump Package

S/S Cold Plate

Experiment Cold Plates (2)

Power and Signal Cooling

Pedestal

CARRIER

Pallet (or equivalent)

Berthing Adapter Assembly

2500

4100

426

20

80

63

14

44

7

80

741.

100

PAYLOAD TOTAL 8175

(PER GSFC SOT ACCOMMODATION STUDY)

ITEM

SOT OBSERVATORY

MODIFIED IPS

CMA

SUPPORT EQUIPMENT

TOTAL

WEIGHT

(Kg)

4000

1400

1200

2000

8600

k

9
TABLE 6. POP WEIGHT STATEMENT (LBS)

(per MSFC Phase A study)

EXPERIMENTS AND EQUIPMENT 2094

X-Ray Detectors 992

Ultraviolet Coronagraph 617

White Light Coronagraph 485

ESA and NASA Furnished Hardware 8239

Pallet + Igloo 3336

IPS incl P/L Clamp Assy (PCA) 2760

OAFD Equip 127

Fwd Utilities 1166

Extra F.C. Reactants (1-12 &02) 850

MISSION PECULIAR EQUIPMENT 1144

Expmt Platform STR. 500

Attach STR for Exp, canister, 120

and sensors

Payload Hold Down Mech. and STR 130

Integration Wiring 40

Thermal Insulation (MLI) 12

Alignment do Pointing Sensors 110

Must,Canister do Motor (SEPS Type) 120

Mask 112

Occulting Plane (A1) 17

Coded Aperature 24

Fourier Transform 3

Lightweight Shield 10

Attach. Str. to top 58

of Mask

TOTAL 11,477 LBS

10
For purposes of this study, it was assumed that the requirements shown in Table 3 are the

payload requirements which need to be accommodated by an instrument pointing system.

A generic pointing system would have to accommodate the combined worst case

requirements of all the attached payloads. The ASO represents the worst case payload

requirements shown on Table 3 with respect to mass, pointing accuracy and stability.

i

This is true even when the requirements for the SOT and the POE are examined

separately. As shown previously the mass of the SOT Cluster is estimated to be oetween

4000 Kg and 6600 Kg. The Transistien. Radiation and Ion Calorimeter has a mass of 5750

Kg, which is close to the high estimate for the SOT Cluster. A generic pointing mount

J

capable of satisfying the requirements of the attached payloads would have to point a

payload of approximately 7000 Kg with an accuracy of 1 are sec and provide a stability

of 0.1 are sec. This takes on added significance since the ASO has been recently approved

for attachment to the IOC SS in 1993. Figure 1 shows the ASO attached to the SS upper

boom.

a,

2.1.1 The Solar Optical Telescope

The core of the ASO is the SOT which will be a SOT-1 observatory with minimun

modifications to permit operation on the SS. SOT-1 will be the first of a series of Shuttle

missions of to 14 days duration carrying a 1.3 m solar telescope and science

instruments as a Spacelab attached payload. The primary goal of the SOT is to study

physical processes that occur on the sun using angular resolution and stability that

correspond to an atmosphere density scale height of 70 Km or 0.1 arc sec. In addition to

the resolution requirement the LOS stability must be maintained to approximately 0.03

arc sec F. MS and drift maintained to 0.06 arc sec for at least 60 sec. To meet the

science objectives the Orbiter will be placed in a solar inertial attitude and the Spacelab

Instrurr -.nt Pointing Subsystem !IPS) will provide coarse pointing of the SOT--1 to the sun.

Fine pointing will be performed within the SOT itself using a control system which moves

the primary and tertiary mirrors in tandem.

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Figure 2 illustrates the SOT-1 observatory in terms of overall dimensions and location of

components. The facility is approximately 4.4 m in diameter, 7.3 m long and weighs

approximately 4000 Kg. The observatory has two major functional components : the

! Telescope Facility and the Coordinated Instrument Package (CIP). The Telescope Facility

Is a folded, on-axis, Gregorian system with a 1.3 m diameter primary paraboliod mirror.

The primary mirror is mounted in an assembly called the articulated primary mirror

(APM) which is supported by six motor driven displacement actuators. These enable the

APM to be translated and rotated about the prime focus for offset pointing and rotated

about its vertex for LOS stabilization in six degrees of freedom. Additional, higher

bandwidth steering of the beam is achieved using the tertiary mirror. This is achieved

using three linear actuators which provide control in two degrees of freedom.

The CIP consists of four functional elements:

A dedicated experiment processor

An image processor

An electronic interface unit

A focal plane instrument package

3

The focal plane instrument package contains the optical and mechanical portions of the

s SOT instruments. They are a photometric filtergraph, a tuneable filtergraph, and a

combined ultraviolet and visible spectrograph.

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14
2.1,2 The Pinhole Occulter Facility

The POF Is designed to study the non thermal phenomena that lie at the heart of plasma

dynamics in the solar corona and will make possible high energy cosmic X-Ray

observations. The facility consists of a self deployed boom which separates an occulter

plane from a detector plane. The X- Ray detectors and coronal imaging optics mounted

on the detector plane are analogous to the Pc,^al plane instrumentation of an ordinary

telescope, except that the occulter Is used for providing a shadow pattern. The occulter

plane is passive and has no electrical interf9,e with the rest of the facility. Figure 3

shows the facility in a partially deployed configuration. MSFC performed a Phase A

engineering definition study of the POF [BI operated from the Space Shuttle. The study

considered a full instrument complement for the facility which included an X- Ray Coded

Aperature Pinhole, an X-Ray Fourier Transform Pinhole, a White Light Coronagraph

Imager, and a UV Spectrograph. The facility requires the use of the Spacelab IPS for fine

pointing. The primary technical challenge for performing the POF objectives is in the

area of pointing and control of the deployable boom. The boom is a continuous longeron

Astromast 0.4 m in diameter and deployed length of 32 m. The boom is quite flexible

having natural frequencies of 0.0948 Hz and 0.917 Hz ( first and second modes ) In Its

deployed condition.

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17

3.0 PRESENT POINTING MOUNT CONCEPTS

Information pertaining to past and present pointing mount concepts was gathered from

several sources to determine the basic design concepts and system capabilities. The

primary source was documentation and publications on the various systems. These

sources are identified in references 191 through 141). Verbal :nformation from telephone

contacts and attendance at presentations was also used. Much of the information

contained in the following sections describing these systems was summarized or exerpted

from references 191 through 1411. Table 7 lists the pointing systems considered, the

developer of the system, and the stage of system development. Various stages of

development are represented. The Suspended Experiment Mount is in concept

I development while the Instrument Pointing System flew on Spacelab 2. Three of the

systems t^ ^ /e flight experience; the Apollo Telescope Mount (ATM),the Modified ATM

Star Tracker (MAST), and the Spacelab IPS. The MAST is not a closed loop system but

was included to illustrate some of the techniques used to perform experiments on

Spacelab. The ATM was included to provided a historical perspective on what was

achieved during SKYLAB over a decade ago. The Hubble Space Telescope (HST), although

not a gimbal system, represents the present state of the art for precision pointing

systems. The basic capabilities of these systems are listed in Table 8. Most of the

systems were designed for Space Shuttle applications and the capabilities listed assume

the appropriate base body. Examination of Table 8 shows that none of the systems

surveyed will accomodate the " worst case " payload requirements as defined in section

4.1, i.e. accommodate a payload mass of 7000 Kg with a pointing accuracy of 1.0 are sec

and a stability of 0.1 arc sec. Appendix 2 provides a general discussion of pointing mount

design considerations, architecture, hardware, and terminology.
18

. , tt.-,-^,, , x.

TABLE 7. POINTING SYSTEMS SURVEYED

SYSTEM

SKYLAB ATM

SMALL INSTRUMENT

POINTING MOUNT (SIPS)

MINIATURE POINTING

MOUNT (MPM)

ANNULAR SUSPENSION AND

POINTING SYSTEM (ASPS)

ADVANCED GIMBAL SYSTEM

(ACS)

MODIFIED ATM

STAR TRACKER (MAST)

INSTRUMENT POINTING

POINTING SYSTEM (IPS)

GIMBALFLEX

SUSPENDED EXPERIMENT

MOUNT (SEM)

TALON GOLD GIMBAL

HUBBLE SPACE TELESCOPE

DEVELOPMENT STAGE

FLIGHT HARDWARE

DESIGN

DESIGN

GROUND TEST HARDWARE

GROUND TEST HARDWARE

FLIGHT HARDWARE

FLIGHT HARDWARE

GROUND TEST HARDWARE

CONCEPTUAL DESIGN

GROUND TEST HARDWARE

FLIGHT HARDWARE IN

ASSEMBLY AND TEST

MSFC

BALL BROTHERS

FOR GSFC

MSFC

SPERRY FOR

LANGLEY

RESEARCH CENTER

SPERRY FOR

MSFC

MSFC

DORNIER FOR

ESA

MARTIN

MARIETTA

MSFC

LMSC FOR

DOD

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19
3.1 SKYLAB ATM

Skylab was launched on May 14, 1973 and the mission ended on February 8, 1974, with the

undock.ng of the Skylab 4 Command and Service Module. The Skylab Orbital Assembly

and its associated control axes are shown in Figure 4. The Skylab Attitude and Pointing

Control System (APCS) provided three axis attitude stabilization and maneuvering

capability throughout the mission and pointing control of tPhR Skylab experiment package

during experimentation periods.

The A?CS was comprised of three control systems. The CMG Control Subsystem

p rovided three axis attitude stabilization and maneuvering capability for the vehicle.

The Experiment Pointing and Control (EPC) Subsystem was used for maintaining attitude

control and fine pointing of the solar Instrument package during solar experimentation.

While the EPC subsystem was fine pointing the solar instrument package, the CMG and

TACS control systems maintained attitude control of the vehicle. The EPC Subsystem

also provided the astronaut with the capability of offset pointing the solar instrument

package to targets of opportunity oo the solar disk or its outer perimeter. The TACS

provided additional attitude control capability to the vehicle.

The solar instruments were mounted on a cruciform spar within 'ie ATM canister, which

was located within the basic structure of the ATM. The instruments which were mounted

on the spar required accurate and stable pointing control.

The requirements were: offset pointing from the sunline of t24 arc-min., with an

accuracy of t2.5 arc-sec.; a stability of t2.5 arc-sec. with a jitter of 1 arc-sec/sec for a

15 minute period; and roll about the sunline of 1120 degrees with an accuracy of t10 arcmin., a stability of t7.5 arc-min. for a 15-minute period, and a jitter of 3 arc-min for a

1-second period.

To achieve the required pointing accuracy and stability, the instrument canister was

decoupled from the main portion of ATM structure as much as possible by the use of

frictionless flex pivot bearings. Figure 5 illustrates the ATM fine pointing control

mechanism. The canister was attached to the structure by a three-degree-of-freedom

mechanism actuator assembly. The outer ring of this assembly was attached to the sour

observatory structure by four roller bearings, and was driven by an electric motor to roll

the canister. The inner ring was attached to the outer ring by two frictionless flex-pivot

20
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mm to

FIGURE 5 ATM FINE POINTING CONTROL MECHANISM

21
bearings with ±2 degrees freedom and was driven about the up-down axis by redundant

actuators. The canister was attached to the inner ring by two frictionless flex-pivot

bearings with ±2 degrees freedom and was driven about the left-right axis by redundant

actuators. Orbital locks constrained the rings and the canister while the solar

observatory was not being used. A roll stop limited the roll to ±120 degrees, to avoid

damage to electrical cables. Launch locks protected the assembly from damage during

launch. A roll resolver provided experiment roll position information to the digital

computer.

The EPC subsyst?m provided automatic stabilization about the canister's Yaw (left/right)

and Pitch (up/down) axis and open loop positioning was provided for the Z (Roll) axis.

Figure 6 is a block diagram of the EPC subsystem.

The Fine Sun Sensor (FSS) and Rate Gyro (RG) sensors, mounted on the lanister, supplied

position and rate feedback to the Experiment Pointing Electronics Assem,11y (EPEA).

The EPEA contained the electronic functions to command the gimbal actuators, closing

the control loop. The Z-axis stabilization was provided by the ACS.

Manual positioning vbout all three axes was provided for offset/roll pointing of the

experiment canister. X and Y- Axis offset pointing was achieved by means of rotating

optical wedge mechanisms within the FSS. Offset commands could be issued from the

Manual Pointing Controller (MPC) or from the ATMDC. The Roll Position Mechanism

(RPM) was activated by command switches located on the Control and Dispiay Panel and

on the ATM EVA Rotation Control Panel. All offset/roll commands were processed by

the EPEA. A line-of-sight pointing capability during roll operations was provided by the

ATMDC, utilizing the ATMDC wedge drive capability.

The EPEA also provided an interface between the Star Tracker and the MPC for manual

positioning of the Star Tracker gimbals. Other EPCS related functions were canister

caging control, experiment alighnment calibration, and experiment and FSS door control.
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A number of new concepts were proven by the performance of the Skylab APCS, which

are applicable to Space Station. Skylab was the first manned spacecraft to utilize large

Control Moment Gyros (CMG) for momentum storage and attitude control, the first to

utilize vehicle maneuvers for CMG momentum desaturation, the first to utilize a fully

digital control system with in-orbit reprogramming capability and extensive automatic

redundancy management, and the first to util ;ze an attitude reference system based on a

four oara ► neter strapdown computation.

The Skylab APCS successfully met or exceeded all design specifications. In particular,

the EPC Subsystem provided pointing accuracy and stability well within the Skylab

instrument requirements.

3.2 SMALL INSTRUMENT POINTING SYSTEM (SIPS)

The SIPS was a precision fine pointing system designed by Ball Brothers Research for

GSFC in the middle 1970's. The pointing system was designed to accommodate two

instruments, each approximately 1 x 1 x 3 meters, from a Spacel.ab pallet. Individual

pointing and control was provided for each instrument. The SIPS configurat i on is shown

in Figure 7. The SIPS consisted of four major sections: a Pallet Interface Structure; a

Deploy/Retract Pedestal; a Pointing Section; and Instrument Cannisters.

The Pallet Interface Structure provided load-carrying attachments for mating with the

hard points on a standard Spacelab pallet. The deploy/retract pedestal provided the

suport structure and mechanisms for raising the pointing section to a position above the

shuttle bay, to obtain an unobstructed hemispherical field of view for the instruments.

The pointing section supported two individually pointed instruments and provided

independent fine pointing for Each instrument. The Instrument Canisters were modular,

to accommodate'nstr+lments of various lengths and provided complete instrument

environmental protection.

24
25

y

iq

The design allowed instruments to be integrated and tested in the instrument canisters

before being mated to the SIPS gimbals. The SIPS was a center of mass mounted system

to decouple shuttle translational motions from payload pointing. The -nit, without

instruments, weighed approximately 764 Kg. The overall gimbal arrangement is shown in

Figure 8. The fine pointing gimbals were normal to the line of sight eliminating the need

for coordinate transforinations and some independence of spacecraft orientation. Figure

8 identifies the gimbal capability with respect to degrees of freedom. Azimuth is motion

about the shuttle "Z" axis. An optional roll gimbal was designed to fit within the SIPS

canister gimbal frame for use with instruments requiring roll about the LOS. Depending

on instrument size, the SIPS could provide payload angular acceleration from 1.5 to 25

deg/sec 2 and a typical slew rate of 2 deg/sec.

Both the SIPS and the mounted instruments depended on Spacelab for electrical and

command and data services. In many ways, the SIPS acted as an "extension" providing

the user with easy access to Spacelab systems. Figure 9 is a block diagram of the SIPS

Pointing System.

The SIPS was nearly autonomous in that all control loop closures were made through the

SIPS control electronics. Sequencing and supervisory control of pointing was

accomplished by command inputs from the Spacelab experiment computer by means of

the data bus and a Remote Acquisition Unit (RAU).

The payload specialist could monitor and control the operation of the SIPS from the

keyboard and CRT that are associated with the experiment computer. Normal operation

was largely pre-programmed, but the payload specialist would have full capability to

conduct such operations as boresighting, guide star acquisition, and sensor calibrations.
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28
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3.3 MINIATURE POINTING MOUNT (MPM)

The Miniaturized Pointing Mount (MPM) was a modified ATM Star Tracker intended to

accommodate relatively small payloads on the space!ab pallet that was studied at MSFC

during the middle 1970's.

1

i The configuration of the MPM is shown in Figure 10. The star tracker was modified by

t. removal of the optical telescope assembly; addition of a roll gimbal torquer and resolver;

fabrication of a base mount pedestal, isolators, and clamping mechanism; and the

addition of an experiment mounting structure. Payloads were end-mounted to the

experiment base plate. The gimbal system nea enclature and freedom were as follows:

The outer or elevation gimbal ±180 degrees; the middle or cross elevation gimbal ±50

degrees; the inner or roll gimbal ±180 degrees.

This gimbal arrangement allowed a conical field of view of 100 degrees included angle,

and when the outer gimbal freedom was along the orbital trajectory, permited horizon to

horizon track along me orbital path or fixed point earth tracking. The roll gimbal being

the innermost, allowed roll about the payload LOS to be achieved with a one-axis

command.

The estimated weight of the MPM system was 110 Kg. The MPM torque capability was

0.6 N-M per axis adequate for solar and stellar fine pointing, but limiting slowing and

scanning operations. Slewing and scanning operations are a function of payload inertia

and payload center of mass offset f pom the center of rotation. End mounting the payload

increased the payload inertia about the center of rotation, placing additional

requirements on the torque motor than would be seen by a e.g. mounted system.

The soft mount isolators on which the MPM was mounted were provided to attenuate the

effects of external disturbances to the mount (i.e., those prc ;uced by crew motion or

Orbiter RCS firings) and to pass only frequencies within the controller bandwidth. Figure

11 is a functional block diagram of the MPM.

The 6iPM interface with the Spacelab Electrical Power Distribution System (EPDS) at a

standard bulkhead connector. The MPM required both 28 VDC and 115 VAC, 400 Hz, 30

for on- orbit operation. The 28 VDC was utilized to power the elements of the MPM.

while the 115 VAC was taken across the gimbal for payload use. Crossing the gimbal at

29
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31
high voltage levels minimized the wire diameter across the gimbal, thus yielding leas

cable torque. The AC voltage was converted to 28 VDC for those payloads requiring DC

excitation. The prime controlling element for the MPM was a self-contained

miniprocessor. This computer controled the MPM, utilizing error signals either from the

MPM sensors or from a payload supplied source.

The MPM attitude sensors consisted of three rate gyros, a fine sun sensor and a fixed star

tracker..

End mounting the payloads on the MPM allowed the experiment canister to be designed

Independently. The basic interface requirement on the canister was that it mate with

the MPM base plate, both mechanically and electrically. Due to the variety of payloads

requirements, no "standard" canister design was established as in the case of the SIDS.

3.4 ANNULAR SUSPENSION AND POINTING SYSTEM (ASPS)

The ASPS I,9 a modular payload pointing system which was intended for use on Spacelab.

The system was designed by Sperry Flight Systems Division under contract to Langley

Research Center. The system, shown in figure 12, was comprised of: a Vernier Pointing

Assembity, two Coarse Gimbel Assemblies, a Mounting and Jettison Assembly, a Control

Electronics Assembly, and assorted connecting hardware.

The complete ASPS was expected to provide payload pointing stability better than 1 arcsec and, when coupled with a quality sensor, the pointing accuracy should have

approached that of the sensor. The unique arrangement of gimbal and vernier subsystem

permitted the utilization of ASPS for stellar, solar, and terrestial observations.

Actuators in the ASPS were sized to accept payloads up to 600 Kg in mass with center of

mass offsets up to 1.5 maters; however, larger instrL vents could be used if reduced

acceleration was acceptable.

Two identical Coarse Gimbal Assemblies (CGAs) were stacked to form an elevation and a

lateral gimbal pair. Design of the Gimbal Mounting Bracket and the Gimbal Mounting

Structure was such as to provide a mechanically limited travel of ±100 degrees (from

vertical) along the lower elevation gimbal axis, and ±60 degrees about the upper lateral

gimbal axis. The Vernier Pointing Assembly (VPA) contained the roll axis drive which

provided unlimited rotation about the payload longitudinal axis and a vernier rotation of

32
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33
1.75 degrees about any axis in the plane normal to the payload roll axis. The Payload

Mounting Plate formed a removable base plate for mounting and aligning experiments

prior to installation on the ASPS. For missions not requiring roll freedom or very high

pointing stability, the Payload Mounting Plate could be attached directly to the upper

Gimbal Mounting BrAcket. ASPS electronics were contained within a separate

electronics package, which was mounted to a pallet cold plate near the bottom of the

Mounting and Jettison Assembly.

Magnetic Suspension was utilized in the VPA to provide noncontacting isolation and

vernier positioning of the payload. Three-axial Magnetic Bearing Assemblies (MBAs)

provided axial translation and vernier pointing about the transverse axes. Proximeters

associated with each MBA were used to linearize the force/c.;.splacement characteristics

of each actuator. The proximeter outputs were al o combined electronically to compute

the axial displacement and the tilt of the payload about the two axes normal to the

payload line-of-sight.

The ASPS interfaced with the Spacelab Electrical Power Distribution system for power

to the ASPS components and for the payload. Twent,. five nickel-cafthim cells were

mounted below the rotating top plate to supply 300 watt hours of electrical energy to the

payload. The selected cells were NASA Standard and provided a nominal 28 Vdc with a

peak power output of 300 watts. A recharge circuit with a one- hour duty cycle was

included. Rechanging the batteries required the VPA to be caged, allowing a solenoid

operated brush block assembly to connect power required to recharge the batteries.

Figure 13 shows •r y e interfaces of the ASPS with the Orbiter/Spacelab, pallet and

payload. The experimenter began control of the system via the Spacelab Digital Display

Unit (DDU) keyboard. He provided servo position commands to the ASPS. These

commands were routed through the ASPS resident software in the experiment computer,

through the CDMS bus to an RAU, and eventually arrived at the Control Electronic

Assembly (CEA). The CEA could then command the ASPS to the desired direction, using

feedback from the ASPS resolvers and proximeters. An optical coupler, which connected

the Spacelab Experiment Data Bus to a second RAU located on the payload plate,

allowed the experimenter to communicate with the payload.

Pointing error data generated by experiment sensors could be sent to the CEA through

the CDMS for use in the pointing control servos.

34
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3.5 THE ADVANCED G'INIBAL SYSTEM

The AGS is a three-axis precision pointing system that was designed by Sperry Flight

Systems under contract to NASA :Marshall Space Flight Center, for use on Spacelab

Missions. Its initial design was based on the ASPS with the magnetic system removed. It

was comprised of a three-axis gimbal system, which accommodated end mounted

payloads through the use of a payload mounted structure (PMS). Total weight of the AGS

was approximately 1025 Kg. The AGS could accommodate payloads from 60 to 7200 Kg,

which Included all hardware mounted above the PMS plane. The dimensions of the

payload were restricted only by the allowable movement of each of the three gimbals, by

the space in the pallet available to the payload and by other payloads resident in the

Orbiter payload bay. Figure 14 illustrates the major hardware elements that made up the

AGS. Two Identical gimbal assemblies formed an elevation and a lateral gimbal pair.

The lower (elevation) gimbal provides an angular range of + 100, - 60 degrees (from

vertical), and the upper (lateral) gimbal provides cross-axis positioning of ±60 degrees.

The gimbals were mechanically limited by an adjustable stop arrangement to prevent

contact with the Shuttle. A roll gimbal permited rotation (±180 degrees) about the

payload line-of-sight and was mounted above the lateral gimbal. The gimbals were

connected to the payload through a Payload Mounting Structure (?MS). The flat endmount configuration of the AGS permited overhanging the PMS to accommodate payloads

of various sizes and shapes. A Mounting and Jettison Assembly (MJA) wss used to

connect the 6lmbals with the underlying mount:ng structure. During launch and landing,

a separation device within the MJA physically disconnected the mounting base from :he

AGS gimbals, which were independently mounted with the experiment to the support

structure.

The MJA also contains a pyrotechnically actuated jettison system which provided for

jettisoning of the experiment/AGS gimbals in the event of a multiple failure in orbit. An

accelerometer unit mounted on the AGS structure or the pallet sensed Orbiter motion for

use in a feed forward decoupling control law.
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The AGS angular error sensors (NASA standard fine sun sensor and star tracker) were

mounted on top of the payload mounting structure with the experiments. The DRIRU-I1

rate integrating gyro package was mounted inside the thermal shroud on the bottom of

the payload mounting structure, along with the Payload Electronics Assembly (PEA), and

up to three (RAUs).

Figure 15 identifies the AGS control and data interfaces. Interfaces with the attitude

sensors were provided by the PEA which is connected to the Digital Electronics Assembly

(DEA) through a serial data link. The PEA is a general purpose microprocessorcontrolled interface unit, which provided a Remote Interface Unit (RIU) compatible

serial interface with the angular error sensors, and a special purpose interface with the

DRIRU-11. Experiment control and data was handled by the Experiment RAUs, wnich

were connected to the Spacelab Command and Data Management System (CDMS).

Additional lines were also provided to the experiments for power, control, and Spacelab

High Rate Multiplexer (HRM) input data. A Power Distribution Unit (PDU) supplied

primary power through the ACS gimbals to an Experiment Power Distribution Unit

(EPDU) mounted within the thermal shroud under the PMS. The L?DU could be

"patched" to supply different power feeds on a mission-dependent basis.

Other AGS electronics were a Gimbal Electronics Assembly (GEA), Power Conditioning

Electronics (PCE), Backup Electronics Assembly (BEA), Jettison Electronics Assembly

(Ji A), Digital Electronics Assembly (DEA), and NASA Standard Spacecraft Computer

(NS°-i!i, all of which were mounted to Spacelab coldplates attached to the AGS Support

Pedestal. The GEA contained analog electronics for modulation and demodulation of the

gimba! angle resolvers and for power drive of the two-phase, brushless do torquers used

to position the gimbals. The GEA also contains electronics which operated the payload

clamping system, separation mechanism, and retractable gimbal stop. A passive Wiring

Junction Unit (WJU) mounted adjacent to the coldplate mounted electronics and provided

for flexibility in mission- dependent harnessing.

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Power switching for all of the AGS electronics (sensors, GEA, BEA, PEA, DEA, and

NSSC-11) and for the AGS heaters was performed in the Power Conditioning Electronics

(PCE). The PCE also provided EMC filtering of the primary power utilized by the GEA

and BEA.

3.6 MODIFIED ATM STAR TRACKER (MAST)

The MAST is a two-degree-of-angular-freedom, manually controlled, gimballed system

that was used on STS-9 (Spacelab 1) as part of the "Atmcspheric Emissions Photometric

Experiment" (AEPI). The pointing mount (a modified ATM Star Tracker) and its

associated electronics was designed by MSFC to accommodate the AEPI instrument and

could be adapted to support instruments with similar requirements and mass properties.

A general description of the MAST includes the description of several of the AEPI

components and functions since these were required by the pointing system for power,

command and c(.ntrol interfaces.

The AEPI experiment was designed to (1) observe both the natural and induced aurora. (2)

observe air glow and other atmospheric phenomena, and (3) observe and measure

contamination around the Shuttle/Spacelab. The major hardware components of the

AEPI included Spacelab pallet mounted equipment and rack mounted equipment within

the Spacelab pressurized module.

The Pallet mounted equipment, shown in Figure 16, was c:;mprised of:

- MAST

- MAST Control Electronics Package

- A Load Isolator

- Linear Actuator

- Detector Assembly

AEPI Support Pedestal

- Cradle/Locks Assembly

- Lens Covers

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41
The experiment detector assembly is affixed to the MAST pointing gimbal system. 'the

lens covers (which also contained the experiment calibration source) are attached to the

pedestal. A cradle provides for mounting of the locking mechanisms near the cg of the

detector assembly. The MAST electronics box is located beneath the pedestal.

The rack mounted equipment, shown in Figure 17, located in the Spacelab pressurized

module include:

- a Dedicated Experiment Processor (DEP)

- a Video Data Encoder (VDE)

- a Mount Manual Control (MMC) Panel

The MAST is a two-degree-of-angular-freedom gimbal system. Each gimbal has two

pivots. Each pivot contains a duplex bearing pair. One pivot contains a multi-speed

resolver, which permits accurate control of the angular position of the gimbal with

respect to the MAST mounting plane. The other pivot contains a d.c. torquer, a d.c.

tachometer- generator, and a system of flexible wires to carry electrical signals and

power to the MAST components and the load isolator. The torquer is driven by servoelectronics to provide pointing control torques. The tachometer permits compensation

for control stability, and is also the rate reference. Each pivot also contains a

thermostat, a heater, and a thermistor for thermal control of the pivot environment.

The MAST Control Electronics Package is used for pointing control and for power

distribution for the experiment. Relays in the box switch power to heaters, the detector 

assembly, and the pointing controls. They also control the shutter, load isolator and

launch lock linear actuators. The analog servo system for the MAST gimbals are part of

the box, as are the digital interfaces with the computer (DEP).

The load isolator is a device which mechanically connects the detector assembly package

to the gimbal system. it contains a screw mechanism driven by a d.c. torquer. Driving

the torquer in the extend direction decouples the detector lssembly package from the

gimbal system which is required to prevent damage to the gimbal system during launch.

42
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43
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Driving the torquer in the retract direction mechanically couples the detector assembly

package to the gimbal system which is required for pointing in orbit. The load Isolator is

launched in the extend (decoupled) position. Once in orbit, it is retracted (coupled), and

remains this way until the end of the mission.

The detector assembly contains two cameras and three lens systems and associated

control electronics. One of the cameras is a low light level television camera (LLLTV)

and can operate through either of two lens systems, 5 or 20 degrees field of view. The

other camera is a 10 x 10 diode array which operates through a separate 5 degree field of

view.

The linear actuator is an electric motor provided to open or close the TV lens covers and

to actuate the lock mechanism (close or open).

The lens cover assembly is mounted on the open end of the TV lenses. The covers in this

assembl y cover the input aperture of the TV lenses during periods when the camera is not

in operation and when the cameras are being calibrated. The cover assembly is released

to the open or closed position by a linear actuator within the assembly.

The detector assembly is held in position during boost phase and re-entry by launch

locks. The locks are contained in a structural "U"-shaped cradle that engages the

detector assembly at the two sides and bottom surface. The lock mechanisms are

actuated by three linear actuators.

The support pedestal is constructed of fiberglass material. Fiberglass structure is used

becalse of its low thermal conductivity. The modified ATM star tracker, the MAST

control electronics package, the detector assembly experiment, the stowage device

(cradle), lens covers, and the "launch locks" locking mechanism are all mounted on the

pedestal. Thermal insulation blankets cover all external surfaces of the pedestal and

equipment with the exception of the star tracker gimbals and cradle/lock assembly,

which will be covered to the maximum extent possible.

The DEP is an off-the-shelf minicomputer which is used to compute pointing commands

and to cycle the detector assembly for its specific functional objective.

44
The VDE is the control panel for the detector assembly. It provides power to the

detector assembly and interfaces the camera and video display. It facilitates operator

Interaction with the experiment when unde)• computer control. In addition, it also

provides an emergency lock switch, which automatically stows and secures the detector

aasaa Ibly,

The MMC panel has the basic tasks of controlling experiment mechanisms Involving the

pointing system, shutter System, locks system, and manual stowing of the experiment

system should compute controlled stowing fail. The mount manual control system

receives primary power (28 Vdc) from the VDE and essential power (28 Vdc) from the

Spacelab. The MMC provides manual control of the launch locks, load isolator, lens

covers, calibration source, and servo-electronics. In the event of a major system failure,

the MMC provides switch capability to the servo-electronics from the primary power

mode to the secondary system (essential power). The MMC also provides for manual

operation of the gimbal system to stow the experiment and activate the launch locks.

45 
3.7 THE INSTRUMENT POINTING SYSTEM

The Instrument Pointing System (IPS) was developed at Dornier Systems under contract

from the European Space Agency (ESA) as a multipurpose, precision pointing system for

Spacelab missions. Figure 18 contains an assembly view and an exploded view of the

IPS. The IPS Is a three-axes gimbal pointing system, which points an inertially stabilized

mounting plate on which an end-mounted payload Is attached.

The total operational range of the gimbals Is ±180 degrees about the roll axis and 30

degrees to 60 degrees, 1/2 cone angle, about the two lateral axes. The range of the

lateral gimbals Is limited by adjustable mechanical stops. For safing any uncontrolled

motion of the gimbal structure during an orbit operation, the Gimbal Support Assembly

has a bumper device which is adjustable to different angles before the flight and prevents

any damage to the Orbiter, Spacelab or Payload. The payload may be a single or a

clustered scientific array of instrument payloads having varied mass and dimensions that

observe inertially fixed or moving targets.

The IPS design permits pointing of payloads with mass properties up to 7,000 kg,

moments of inertia about their C.G. up to 30,000 kg-m 2 about the lateral axes and 12,000

kg-m 2 about the roll axis and C.G. offsets from the payload Interface plane up to 3 m.

Payload masses above 3,000 Kg will require modifications to the Payload Clamp

Assembly (PCA), which clamps anu structurally holds the payload during launch, ascent,

reentry and landing phase cf the flight. Payload dimensions are restricted by the width

and available length of the Spacelab pallet and diameter restrictions associated with the

PCA. Wiring for payload power, for three RAUs and for direct data transmission to

Spacelab/orbiter, is provided across the gimbal at the payload mounting plate.

Attitude control of the payload is based on the inertial reference of a three-axis gyro

package mounted on the payload mounting plate that is updated by star trackers or sun

sensors mounted on the payload. A Digital Control Unit (DCU) also mounted on the

payload mounting plate, governs all slew maneuvers, payload stabilization and tracking

operations by issuing commands to the three-axis gimbal drive assembly. A three-axis

accelerometer package, mounted to the IPS support structure, measures accelerations

caused by external disturbances. Outputs from the accelerometers are fed .forward to

the primary feedback control loop to allow for compensation of external disturbances.

The pointing performance requirement for the IPS are listed in Table 9. The disturbance

46
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TABLE 9. POINTING REQUIREMENTS OF THE IPS

IN THE 2 AXES

PERPENDICULAR

TO THE

PERFORMANCE EXPERIMENT

REQUIREMENTS LINE-OF-SIGHT ABOUT THE ROLL AXIS

Pointing accuracy of the 2 are sec 20 arc sec

experiment line-of-sight

Quiescent stabijl ± y error 1.2 are see 3 are sec

Man-motion disturbance 4 arc sec 15 are see

Stability rate 60 are sec/sec 130 are sec/sec

response requirements listed apply for a 2,000 Kg payload with the IPS center of rotation

(COR) located on a "Z" axis through the Orbiter center of gravity (e.g.). For a 200 Kg

payload, the disturbance error is required to be less than 1.7 times that of the 2,000 Kg

payload.

The IPS has seven major assemblies, which are:

- Gimbal Structure Assembly

- Attitude. Measurement Assembly

- Power Electronics Unit

is - Data Electronics Assembly

- Thermal Control Assembly

- Payload Clamp Assembly

- Harnesses

The total mass of the IPS, excluding the PCA, is approximately 1,176 Kg.

F

r The Gimbal Structure Assembly incorporates all static and moving mechanical

structures, drive units, stow/unstow mechanisms, and all special devices mounted onto it.

48
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The Attitude Measurement Assembly consists of a three-axis gyro package and an optical

sensor package that provide inertial and optical references for IPS attitude control.

The Power Electronics Unit distributes power to all other IPS elements and to the

Payload. It also provides contingency control functions to enable manually controlled

contl; ,rncy retraction or jettisoning of the system.

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The Data Electronics Unit contains the circuitry which processes the signals required to

maintain the IPS attitude control. It also contains the three-axis accelerometer package,

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which enables feed forward control of disturbances introduti^,,d by external sources.

t The Thermal Control Assembly protects critic:i, q onents from the extreme

temperature variations of space.

The Payload Clamp Assembly (PCA) is a mechanical structure designed to secure the

payload whenever it is separated from the IPS. It clamps and structurally holds the IPS

payload and payload attachment ring during launch, ascent, reentry and landing phases of

a flight. The PCA holds the payload in a statically determinate manner by three attach

flanges lying in a plane through the payload center of mass.

The IPS wiring harness provides all the necessary interconnection of the electronic units

and assemblies. It is subdivided into Pallet Harness, the Gimbal Support Assembly

Harness and the Payload Interface Harness.

The IPS electrical configuration envolves extensive interfaces with the Spacelab

Command and Data Management System (CDMS), and the Electrical Power Distribution

System (EPDS). A functional Block diagram of IPS electrical and data in' , --faces is

snown in Figure 19 .

Pointing, tracking, and stabilization control of the IPS is maintained by a digital control

loop operated within the IPS DCU and the CDMS subsystem computer. Data flow within

the control loop is seperated into two computation loops: fast loop computations

operating at 25 Hz between the DCU and IPS attitude components, and slow loop

computations operating at 1 Hz between the DCU and subsystem computer.

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The IPS also contain all software necessary for monitoring and exercising ground and

mission operations. A modular software structure is used to facilitate modifications and

to minimize the demand on the Spacelab CDMS.

The IPS Attitude Control loop is shown in Figure 20 .

The fast loop computations are based on the inertial reference furnished by a three-axis

rate-integrating gyro package mounted on the equipment platform that sends 100 Hz

signals to the DCU. The DCU averages the data over every four samples, integrates by

using quaternion notation, and compares it with the desired attitude for generation of the

a

attitude errors to be used in the subsequent control arithmetics. When the inertial

reference is disturbed by Orbiter and/or IPS axis deviations, the DCU issues control

signals to the PEU to apply control torque to the drive unit torquers. In this manner, any

attitude discrepancy is compensated for by the IPS automatically moving to a corrected

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attitude.

In addition, a three-axis accelerometer package mounted onto the supporting framework

measures accelerations caused by Orbiter perturbations and sends a feed-forward control

signal to the DCU. The DCU adds this signal to the fast loup computations to

compensate for the accelerations.

To compensate for mechanically-induced gyro drifting, slow loop computations are used

to update the inertial reference of the fast loop by taking optical measurements supplied

by the optical sersor package (OSP). 'The OSP consists of three fixed head star trackers

mounted on a precision base. For stellar missions, a boresight and two skewed trackers

provide IPS line-of-sight and roll attitude references respectively. For solar missions,

the boresight tracker is reconfigured as a solar sensor with the two skewed trackers

maintained for roll references. The OSP signals are sampled once a second by the

subsystem computer and processed by a Kalman filter to generate the attitude and rate

updates. The OSP is mounted on the payload to assure optimum viewing and to minimize

misalignments between the IPS line of sight and the experiment sensor line of sight.

Besides using the OSP for pointing references, attitude offset commands generated by

the experiment computer or experiment sensor signals can be used to provide additional

pointing references.

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3.8 GIMBALFLEX

Gimbalflex is a fine pointing system, which has been under development by the Martin

Marietta Aerospace Co. The systen-, is meant to serve as a vernier stage on top of

another gimbal system such as 1PS )r AGS, as shown in Figure 21 . The system combines

a linear isolation system with rotation axis control to achieve a fine pointing system with

vibration isolation. Figure 22 illustrates the basic mechanical configuration. Vibration

isolation is made possible by the use of flex hinges and an intermediate gimbal. As shown

in Figure 22, the outer linkage set allows motion along the Y-axis between the payload

and the gimbal. The inner linkage allows motion along the X-axis between the gimbal

and the base body or base pointing system. This provides complete isolation in the X and

Y planes. Angular flexures between the linkages and the gimbals allow payload pointing

about the Z-axis. Direct drive brushless torquer motors are used to apply the pointing

and stabilization t6-ques. This linear isolation system provides a mechanically stiff path

back to the base body along all rotational axes. Linear isolation frequencies must be high

compared to the angular control frequencies. An active control system can be used so

that the isolation frequencies can be adaptive in flight. The isolation system will not

remain orthogonal when a large angle pointing is required. Therefore, this system is

limited to angles below 15 degrees of rotation.

Based on an early presentation to MSFC by MMC [33], the system could accommodate a

payload mass up to 600 Kg and achieve a pointing stability of .01 sec. The system would

provide linear motion of tl inch along two-axes and ±10 degrees of angular rotation

about two axes. The gimbalflex system would weigh about 5u lbs.

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3.9 SUSPENDED EXPERIMENT MOUNT (SEM)

The Suspended Experiment Mount (SEM) is a payload isolation and stabilization system

concept that has been under study at MSFC for several years. As shown in Figure 23. the

CMG stabilized structure would mount in the Orbiter payload bay and could be

constructed from Skylab spare parts and other existing hardware. T!:1 system consists of

a spacelab pallet (or pallet train), which is isolated with respect to the shuttle by means

of passive spring dampers or other non- rigid suspension devices. These devices would

Isolate the payload from high frequency disturbances and the CMG system would provide

active control of low frequency disturbances and stabilization of the LOS. Three

(possibly four) Skylab type double gimbal CMGs mounted on the pallet would provide

pointing stabilization to 1 are second for pallet mounted payloads.

Besides isolating the pallet from vibrations, the isolation system allows coarse attitude

control of the Orbiter via the CMG's mounted on the pallet. Thus, the Orbiter Vernier

Control Cystem (VCS) is not required to operate. Although especially suitable for

instruments with the same viewing requirement, instruments with multiple viewing

requirements can be accommodated by the use of simple position and hold pointing

mechanisms. The capability of these pointing mechanisms would be determined by the

payload requirements. In addition to viewing type instruments, this concept also has the

capability to provide a disturbance free environment for other payloads that are sensitive

to relatively small accelerations.

The suspension system performs two functions. It isolates the payload from high

frequency disturbances and allows low frequency control of the orbiter during experiment 

operations. It must also restrain the payload during periods of high dynamic loads.

Therefore, it must consist of a flexible coupling that can be rigidly locked during periods

of high loads. Figure 24 illustrates a suspension/retention method considered by MSFC

during its studies on SEM. It uses standard orbiter active sill and keel trunnion fittings

for rigid attachment. The flexible suspension incorporates a linear actuator to lift the

payload out of the trunnion fittings during operation. This approach has the advantage of

using standard orbiter fittings during the critical flight periods. MSFC has considered

other concepts for active retention devices which have the advantage that the SEM can

be released without requiring a linear displacement of the payload.

56
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57
Helical springs are shown for the flexible isolator In Figure 24, but otiier options have

been considered. These Include: elastrometric isolators, gas filled bellows, wire rope

helical springs, and the solid wire helical spring, Illustrated in Figure 24. Gas filled

bellows and solid wire helical springs are the prime candidates for the flexible isolator,

since they exhibit a more linear "load versus deflection" characteristic.

The SEM can be built using Skylab type Double Gimbal CMGs, although some

modifications to the Skylab CMGs have been considered for a more state-of-the-art

configuration. Reaction wheels can also be used In lieu of CMGs, 6epending on the

torque capability required. The CMG system Is used to control both the payload and,

through the suspension system, the Orbiter. Avoiding Orbiter thruster firings reduces

disturbances to the payload anci reduces contamination, which Is a significant benefit to

many payloads.

Momentum saturation is a concern for any momentum based attitude control system.

After a period of operation, possibly several orbits, non-cyclic disturbance torques acting

on the vehicle, result in net CMG angular momentum build-up. This momentum build-up

eventually causes the CMG to become saturated and not capable of compensating for

disturbances about Its control axis. Desaturation or momentum damping requires that an

external torque be applied to the vehicle to absorb the accumulated angular momentum

(i.e., realign the CMG momentum vector with its vehicle control axis). Use of the

orbiter thrusters is the simplest method of providing this torque. Another possibility is

to use gravity gradient torques either by flying an attitude that produces negligible noncyclic torques or b,^ periodically maneuvering to an attitude where gravity gradient and

aerodynamic torques cause an accumulation of angular momentum in a desired

direction. In the case of gravity gradient dumping of momentum, offset pointing

mechanisms would be required for the payload, since the Orbiter attitude would not be

compatible with the payload viewing direction.

58

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Thp SEM Is intended to provide stabilization and Isolation, not offset orientation, so the

free motion of the suspension system only needs to accommodate the disturbance

motions. These should be relatively small. Because of these relatively small motions of

the SEM with respect to the Orbiter, it should be easy to provide power, communications,

fluids, etc. from the orbiter to the payload. Therefore, It provides t he payload with the

pointing accuracy and disturbance isolation of a free-flying spacecraft and retains

advantages associated with being attached to a base body.

As shown in 'fable 10, the total weight of a complete SEM for the Orbiter/Spacelab

application Is 2090 Kg. A "One Spacelab Pallet SEM" would be capable of

accommodating payloads up to 1675 Kg. A two -pallet train could accommodate up to

3550 Kg of payload.

60
TABLE 10. SEM WEIGHT SUMMARY

ELAB PALLET WITH CMG'S AND SUSPENSION SYSTEM

WEIGHT

Kg (LB)

1 Experiment Support Structure 390 (840)

1 Suspension System 100 (220)

3 CMG's plus CMGEA's 570 (1257)

3 C MGIA's 69 (152)

3 Rate Gyro's 16 (35)

1 Computer do Electronics 32 (70)

1 Experiment Microprocessor 32 (70)

1 Star Tracker Opto-Mech 18 (40)

1 Star Tracker Electronics 15 (32)

1 Signal Conditioner 14 (30)

1 Power Controller 15 (32)

1 Control be Monitor Panel 7 (15)

1 Recorder & high Rate Multiplexer 71 (156)

1 CRT Keyboard System 14 (30)

Subtotal 1363 (2999)

Pallet Equipment Attachment Hardware 70 (154)

Total 1433 (3153)

1 Spacelab Pallet 657 (1444)

:Missicn Total 2090 (4597)

I

y 61
3.10 TALON COLD GIMBAL (TOG)

The Talon Gold Gimbal (TGG) is a three-degree-of-freedom, micro-processor contro!Ied

pointing system, which was developed by LMSC for DOD as part of the 'Talon Gold

Experiment. The experiment, designed to be operated from Shuttle, was to demonstrate

target acquisition, pointing and tracking capabilities. The system, illustrated in Figure

25, included the instrument sensor, a vibration isolation pointing system (VIPS) and the

TOG. The system was required to provide a pointing accuracy of 0.2 u rad (.04 are sec).

The experiment was mounted on the VIPS, which was developed by the Sperry Flight

Systems Division. The VIPS is a magnetically suspended vernier pointing stage, which

provides non-contacting isolation and fine positioni.. a of the payload. The system Is

similar to the magnetic suspension stage used in the ASPS system previously described.

Using a quality sensor, provided as part of the experiment, the vernier stage could

achieve a pointing accuracy of C,1 arc sec and a stability of 0.01 are sec. The Talon Goid

Gimbal provides the coarse pointing capability for the experiment.

The pointing, tracking and stabilization system utilized five Motorola MC68000

micro-computers which were distributed across the system's architecture. One MC68000

was included at each gimbal joint, one was '. l ed on the Instrument mounting plate, and

one on the mounting interface with the Shuttle. The use of these micro-computers which

have an internal data bus structu — rovided a self-contained system which required no

additional electronics or compute, upport from Shuttle. Attitude sensors include

precision rate gyros and star trackers mounted on the payload side of the gimbals. 

Outputs from an accelerometer package, also mounted on the instrument-mounted

surface, were used ire feed forward control loop to compensate for disturbances before

they could propagate to the instrument. The precise pointing and stability was

ultimately achieved by using an error signal generated by a fast steering mirror system

located internal to the payload.

The TOG itself is a precise payload pointing system, which incorporates many state-ofthe-art features to achieve low jitter, high accuracy, quick response, and near-zero

friction.

The gimbal joints use gearless direct drives and are supported by flexure bearings nested

inside rolling element bearings. This eliminates problems associated with backlash in

62
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relative to other systems. The gimbal joints contain rotary transformers, rotary fluid

couplers, capacitive couplers and encoders mounted on the axis of rotation. Electrical

power, instrument signals, fluid-, angle and rate data are transferred across the joints

without cables.

The rotary transformer can transfer up to 5 KW of 120 V at a frequency of 10 KHz to 30

KHz across the gimbals. Use of this device provides several advantages: 1) no wear or

wear products, 2) no arcing, 3) no friction, 4) high electrical performance.

The rotary fluid coupler is a two part assembly having its outer housing fixed and inner

spool rotating. Fluid enters the outer housing through a Swaglok fitting and is ported

Into an annular chamber that interfaces with the inner spool outer surface. The outer

housing annular chamber edges are sealed from the rotating inner spool by O-rings to

prevent axial flow on the spool surface. An opening in the spool surface permits radial

flow into one of several cylindrical chambers adjacent to the spool axis. The fluid exits

this spool chamber through a Swagelok fitting, flowing in the axial direction. The return

flow is a path parallel to the input flow, but in the opposite direction. A rotary coupler

has the advantages of eliminating line flexure fatigue failures, reduced fluid line bulk,

weight, cost and unlimited rotation.

The fluid coupler contains four ports, two for pressure and two for return. Weight is 2.7

kg (6 lbs); construction is steel. The torque required to rotate the spool inside five seals

is approximately 1.02 N-m (9 lb-in). The pressure drop at 11.4 liters/minute (3

gallons/minute) flow is less than 6.89 kPa (1 psia), with 60% ethylene glycol and 40%

water.

Rotary capacitive couplers are used to transfer signals across the gimbal joints. They

provide advantages of minimum wear, low disturbance torque, and high reliability. A

digital line driver transmits a differential charge to a pair of driver capacitor plates. An

opposing differential charge appears on a pair of receiver plates located close to the

driver plates. The two plate pairs make up a pair of parallel capacitors with properties

that are independent of the plates relative rotary motion. A digital line receiver detects

the differential signal transmitted via the capacitor charges. The capacitor plates are

64
matching conductive rings on a pair of circular printed circuit boards. There are four

matching rings which make four capacitors per channel. Two , ring pairs transfer the

differential signal. While the other two ring pairs are grounded guard rings on each side

of the signal capacitors. The guard rings isolate these signal capacitors from the signal

capacitor adjacent channels to minimize cross- talk. The signal channels are further

Isolated by int, leaving primary and redundant channels, the primary and redundant

channels never being used simultaneously. A twelve inch diameter circular printed

circuit board pair provides up to ten signal channels. Electronics are mounted on the

back sides of the capacitor boards. The transmitted signals are digital, with thresholds

set to make the coupler insensitive to plate spacing. Analog signals are converted to

digital for transmission, and then reconverted to analog at the output.

The gimbal encoders generate angles with 21 hits (1/2 arc sec) resolution, and analog rate

with dynamic range equivalent to 14 bits resolution. The rate signal, Is derived directly

m from the encoder optical detector analog waveform, using an analog circuit and discrete

digital logic. The rate signal lag is much lower and the resolution is much finer than a

rate that is computed digitally by differencing the 21 bit digital angle signal. This fine

rate analog signal is instrumental in gimbal low jitter control.

The encoder angle signals are generated by light emitting diode illumination passing

through rotary disk windows onto light detectors. Stationary apertures overlay the

detectors to outline their fields of view precisely. There are 9 inner gray code rings of

windows on the disk and one outer analog signal ring. The analog ring produces 2

triangular signal waveforms that are 90 degrees out of phase.

Absolute angles are assembled from the straight center segments of the analog

waveforms and the gray code. An automatic gain control (AGC) analog loop controls the

illumination to effect coarse regulation of the optical transmission gain. The angle

assembly is done by a microcomputer , which also regulates the fine AGC.

Rate signals are derived from the triangular waveform analog displacement signals, using

an analog rate circuit. For constant rate, the waveforms are square. A demodulating

sampling circuit converts the signals to "dc" rate. The uC controlled fine AGC is

implemented by the digital to analog (DA) and the multiplying digital to analog (MDA)

converters.

65
The TGG can accommodate payloads up to 2400 Kg, and moment of inertia of 11,400 KgM 2 about the gimbal axes. The approximate dimensions of the Talon Gold experiment 

was 10' x 7' x 8' (3m x 2.1m x 2.4m). The pointing system control bandwidth is 0.5 Hz.

The TGG can achieve a pointing accuracy of 3 arc sec and a stability of 0.6 are sec.

Torque capability is 70 ft-lbs (95 n-m) per axis when primary and redundant motors are

used simultaneously In emergency stewing. The TGG can slew payloads up to 3

degrees/sec and provide acceleration up to 400 are sec/see

3.11 HUBBLE SPACE TELESCOPE

The Edwin P. Hubble Space Telescope (HST) is included here since it includes the highest

precision pointing system developed to date. It demonstrates that the technology

required for highly precise pointing systems is available. In addition it represents one of

the free flyers that may utilize the S.S. for servicing. The ST pointing and control

system (PCS) uses a combination of rate gyros, reaction wheels, star trackers and

interferometrie fine guidance sensors to slew the ST from target to target and

accomplish precise pointing and stabilizing accuracy. The PCS points the ST at the

target object to 0.01 arc seconds and is required to achieve pointing stability of 0.007 are

seconds over a 24 hour period.

The anticipated mission operational lifetime is fifteen years. The scientific instruments

and some of the electronics are modular design; i.e., orbital replaceable units (ORUs), so

that they can be replaced with improved or different models. Periodic orbital

maintenance and altitude reboost via the Shuttle is planned. For these reasons, the

facility could operate in space for decades. If necessary, it will be possible to retrieve

the ST, return it to earth, perform maintenance and repair, and relaunch the facility.

Figure 26 shows the general operational configuration of the ST. It weighs 11,600 Kg, is

14.3 m long, and 4.3 m in diameter. An exploded view of the ST is shown in Figure 27.

Major elements of the ST are: the Support System Module (SSM), the Optical Telescope

Assembly (OTA), Science Instruments, Science Instrument Control and Data Handling,

and the q-lar Array (SA).

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The SSM shown in Figure 28 encloses the OTA and the Sls, supports the Solar Array (SA)

and Haigh Gain Antenna (HGA) and provides all the interfaces with the Shuttle Orbiter.

The module contains a very precise pointing and stabilization control system, thermal

control system, data management system. Instrumentation and Communication (I&C)

system, and electrical power system. The electrical power to operate the telescope is

provided by batteries which are charged by the two SA panels during the sun side of its

earth orbit.

The PCS provides the ST with the capability of pointing at designated targets with

sufficient accuracy and for the necessary period of time to achieve mission objectives.

A PCS functional diagram is shown in Figure 29. The PCS operates in a digital mode and

uses software modules resident in the flight computer and hardware items physically

located throughout the ST.

The PCS equipment (Figure 30) can be divided into three categories: sensors, computer,

and actuators.

There are five different types of sensors that record ST attitude and rate

measurements. These sensors are: Fine Guidance Sensor (FGS), Fixed Head Star Tracker

(FHST), Coarse Sun Sensor (CSS), Magnetic Sensing System (MSS), and the Rate Gyro

Assembly (RGA).

The flight computer calculates the control law, attitude updates, momentum

management law and the command generators. To limit structural mode acceleration,

the command generators shape the acceleration and incremental angle commands to the

control systems in an accelerate, coast and decelerate pattern.

Finally, actuators turn the control signal into torques that move the vehicle. There are

two kinds of actuators: the Reaction Wheel Assembly (RWA), and the Magnetic Torquer

(MTs).

Redundant hardware elements are provided to allow a timely resumption of science

operations follow.ng a detected anomaly in system behavior.

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74
The components used in the design of the ST PCS represent the leading edge of todays

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technology. Particularly noteworthy are the Fine Guidance Sun Sensors (FGSs) and the

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t Rate Gyro Assembly (RGA). The FGS measures the angular position of the stars and is

the most accurate and the most complicated of the sensors. It uses the principle of

interferometey to obtain an error signal that is accurate to fractions of an arc. sec. Two

of the three FGSs are used by the PCS to perform stellar pointing or low rate scanning.

These sensors are used for attitude correction of the telescope Line of Sight (LOS) with

rate gyros used for rate and short term attitude Information. Control about the LOS Is

provided by rate integrating gyros with information from the two FGSs being used to

provide attitude corrections.

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The FGSs share the telescope Field-of-View (FOV) with the scientific instruments. Each

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FGS has a FOV of about 60 arc min squared. Four photomultiplier tubes are used for

both acquisition and fine pointing. For fine pointing, the output of the tubes is

differenced for use in an interferometeric mode. Because of the limited dynamic range

of the interferometer (approximately t.02 are sec), the FGS uses the it,±m output of the

photomultiplier tubes to achieve initial star acquisition after the PCS maneuvers the ST

from one target star to another. The FGS contains star selectors, which are rate control

servos with rotating prisms, to allow the 5 arc sec squared instantaneous FOV to be

maneuvered over the total 60 are rein. squared FOV of the FGS.

The RGA is a strapdown, reference gyro package. It senses vehicle motion using two

modified, ultra low noise Rate Integrating Gyros (RIG) to provide two-channel digital

attitude and analog rate information. During normal vehicle operations, these rate and

short term attitude updates are provided by the ST RGAs for fine pointing and spacecraft

maneuvers. The RGA system is also used to produce inputs for Pointing Control System

(PCS) operation during vehicle safemode.

A single RCA is composed of two separate subassemblies: the Rate Sensor Unit (RSU)

which contains the RIG gyros, and the Electronics Control Unit (ECU). The RSU weighs

24 pounds, and the weight of the ECU is 17.5 pounds. Both units are designated as Orbit

Replaceable Units (ORU).

73
The input axis of each single-degree-of-freedom gyro Is skewed relative to the ST axes.

The RDAs sense vehicle motion and output Incremental attitude data to the Data

Management System. The data is received by the Data Management Unit (DMU) at 830

micro sec intervals and accumulated into 25-msec samples to be used by the PCS for

vehicle attitude control.

The rate gyros have a dual rate range. The high rate scale range is greater than 1800

aresec/sec, the low rate scale is greater than 20 aresec /sec. One of the gyro sensor units

is redundant and kept in a dormant state unless needed.

74

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ON DESCRIPTION

The Space Station configuration used as a baseline for this study is the "Power Tower"

configuration as described in the "Space Station Reference Configuration Description"

document (1). A summary of the 1007. SS configuration described In [1] is presented in the

following paragraphs. Some aspcets of the SS not directly involved with instrument

pointing are included. This Is P g Intended provide an undevstandin P k of the overall g

environment in which instrument pointing must be performed and to facilitate a

subsequent discussion of instrument servicing (Section 6.0).

4.1 SPACE STATION CONFIGURATION SUMMARY

Figure 31 illustrates the Space Station Reference Configuration as described in 111. It is

composed of a set of deployed linear trusses to which pressurized modules, subsystems,

and user equipment are attached. The principal structural components are a keel and

three booms at right angles to the keel. These arc refered to as the upper, transverse

and lower booms. The coordinate system is Z parallel to the keel, positive toward nadir;

X perpendicular to the keel and booms, positive in the direction of flight and Y parallel

to the booms, positive toward starboard. The coordinate system origin (i.e., navigation

base) is at the intersection of the transverse boom and the keel. The Attitude Control

ASSEMBLY (ACA) houses the navigation base components and is a rigid nine-foot cube at

the Intersection of the keel and transve rse boom. It contains the Control Moment Gyros,

Star Tracker and other Guidance, Navigation and Control components, needed for control

of the station from the initial launch. The Space Station will be incrementally assembled

in orbit using the STS Orbiter. It is anticipated that the permanently manned IOC Space

Station will become operational in the early 19901x.

The IOC Space Station will have five pressurized common modules.. Two Habitability

Modules (HM), two Laboratory Modules, and a Logistics Module. The kee` is divided at

the bottom to allow installation of the pressurized modules on the centerline of the

keel. This maintains the principal axis in the orbit plane while the Orbiter is being

berthed. Four of the pressurized modules are arranged in a quadrangle to allow crew

movement into or out of all modules, even if one becomes unusable. The quadrangle is

also referred to as the "racetrack configuration." The module arrangement is shown in

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evacuation of any mocule. The laboratory modules are a Materials Processing

Laioratory (LM 2) and a Life Sciences Lab (LM 1). Lab 2 Is located at the top of the

quadrangle to keep it as close to the center of mass of the station as possible. The HMs

are adjacent to each other, to eliminate unnecessary movement by the crew through the

j Laboratory Modules. A control station for berthing is at the front of HM2 and allows a

direct view of operations. The primary Orbiter structural interface is at the end of

HMI. An alternate position at the end of Lab 1 is provided for emergency access in case

the primary port is inoperable. The logistics module is berthed to HM 2 to permit

unloading with minimum disturbances to laboratory operations.

Hatches are provided on all operational ports and crew airlocks. They are hinged so that

crew and equipment can traverse through the open port as required. Each may be closed

to serve as a pressure bulkhead between modules. Hatch weight has not been defined.

The minimum hatch diameter is 50 inches. The logistics module has two end hatches.

The other modules have four radial and two end hatches. External airlocks are attached

to the port side hatches of HM1 and HM2. The electrical system, including power

generation, storage and conditioning, is installed on the transverse boom outboard of the

alpha gimbal joints. Each array wing is gimballed individually for beta adjustor ant to

minimize principal axis shifts and simplify assembly and deployment. Primar, heat

rejection is provided by radiators mounted on the lower keel. The radiators are rotated

to maintain, an edge toward the sun. They are rewound during the dark potion of the

orbit, to avoid a continuously rotating fluid joint. TDRSS and GPS antennas are mounted

at the ends of the upper boom to allow a maximum upward view. Tracking and

i rendezvous radar antennas are mounted on the lower keel under the transverse boom for

a clear view along the flight path.

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Identical sets of propulsion thrusters are mounted at four locations at the ends of the

lower boom and on the lower keel. These are capable of firing forward, aft, and

outboard. The lower thrusters (on the lower boom) normally provide all backup attitude

control. The upper thrusters normally fire only aft, for orbit maintenance. A Mobile

Remote Manipulator System (MRMS) can move on the nodes of the truss to any location

on the forward faces of the keel and booms. By moving outboard of the alpha joint and

rotating the transverse boom, the MRMS can also move along the aft face of the

structure.

4.2 PAYLOAD ACCOMMODATION

Two primary locations are provided for observational users. Solar and stellar instruments

are mounted along the upper boom to maximize their view of that part of the sky not

occulted by the Earth and Station. Solar viewing is possible on every orbit because the

Sun is always within 52 deg. of the orbit plane.

Earth sensors are mountec, on the lower boom and the local vertical orientation allows

continuous viewing limb-to-limb with no ititage rotation. Some solar sensors may also be

mounted on the transverse boom outboard of the alpha gimbal to simplify pointing if

their mass characteristics do not adversely affect the Station 's dynamic properties.

Satellite servicing provisions are located along ti-.e keel. Two storage /servicing bays are

situated parallel to the upper keel. This location minimizes contamination of optical

surfaces and sensitive components. Related ORU and tool storage is provided on the

transverse bo^m. A refueling bay is located on the lower keel. Satellite propellants and

cryogens are stored at the top of the keel extension near the refueling bay. The OMV and

OMV kits are berthed alongside the port and starboard keel extensions.

79

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4.3 SPACE STATION ORIENTATION

In normal operation, the Station is oriented with the keel approximately aligned with

local vertical and the Y principal axis (approximately parallel to the transverse boom)

held perpendicular to the orbit plane (POP) as shown in Figure 33. A constant pitch

attitude is maintained with respect to the local vertical to balance gravity gradient and

average aerodynamic torques. CMG's are used for momentum management to

compensate for aerodynamic torque variations around the orbit. Long-term (more than a

few orbits) variations in mass properties or atmospheric density are accommodated by

small adjustments in the flight attitude of the Station.

The solar array is maintained normal to the Sun vector by a continuous rotation about the

alpha gimbal at orbital rate and a slow adjustment about the beta gimbal to track the

Sun's motion relative to the orbit plane. The radiators are also rotated at orbital rate,

but are "rewound" during the dark part of the orbit to avoid continuously rctating fluid

joints.

The Space Station orbit characteristics were defined in the Space Station Definition and

Preliminary Design Request for Proposal [2]. These parameters are identified in Table 11

for the Space Station and its associated co-orbiting and polar platforms.

TABLE 11. SPACE STATION ORBIT PARAMETERS

NOMINAL DESIGN ALTITUDE RANGE

ELEMENT ELEMENT INCL. HIGH LOW

SPACE STATION 50OKM 28.50 555KM 463KM

(270NM) (300NM) (250NM)

CO-ORBITING PLATFORMS 50OKM 28.50 555KM 463KM

(270NM) (300NM) (250NM)

POLAR PLATFORMS 705KM 98.250 900KM 40OKM

(381NM) (486NM) (216NM)

POLAR PLATFORM SERVICING ALTITUDE IS 276KM (149NM)

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Other orientations for the station are also under evaluation. Reference i42) describe an

"airplane mode" or "power arrow mode" in which the longitudinal axis is aligned with the

velocity vector as shown on figure 34. The preliminary analysis contained in (42) shows

that the "airplane mode" offers advantages over the baseline orientation in terms of the

momentum storage requirement, reboost control, and the level of "micro g" environment

available for payloads. With respezi to astrophysics instruments located on the upper

boom It will have the effect of decreasing observation time and Increasing instrument

contamination.

4.4 ATTITUDE CONTROL SYSTEM DESCRIPTION

The Space Station ACS is an inherent part of the Guidance, Navigation and Control

Subsystem (GN&C). The GN&C subsystem controls the attitude and orbital position of

the Space Station, Its payloads and appendices. In addition. it supports proximity

operations and traffic control during all mission phases. The GN&C design is responsible

for accomplishing:

A) Attitude estimation and control of the Space Station.

B) Momentum Management with aerodynamic variations.

C) Navigation base/payload relative alignment.

D) User interface with the Space Station crew.

E) Payload disturbance isolation.

F) Solar Array, antenna and radiator pointing.

G) Navigation and control during rendezvous and docking.

H) Stability and control during build up and growth.

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Specific performance requirements, such as; 0.01 deg. attitude determination accuracy

at the navigation base, and measurement of the payload relative alignments to .002 deg.

accuracy with respect to the navigation base, are major design drivers.

The baseline control system approach per reference [11 is a centralized system of colocated attitude sensors and CMG Control Actuators, as shown in figure 35, which

provide inherent stability of the flexible body. The navigation system utilizes Inertial

Sensor Assemblies (ISA) composed of rate gyros and accelerometers to provide attitude,

attitude rate, position and velocity. Star Trackers will be used to update attitudes, and

GPS satellites will be used to update position and velocity. The navigation processing

will be performed in the Navigation and Traffic Control Computer.

The primary attitude control actuators will be CMGs, which may be supplemented by

magnetic torques to perform continuous momentum management. The use of magnetic

torquers will probably be eliminated [43,44] In favor of a torque equilibrium attitude

(TEA) which uses the gravity gradient torque to minimize CMG momentum buildup. The

overall momentum management strategy will be to reduce the roll/yaw disturbances to

near zero by placing the vehicle principle axis (Z) in the orbit plane. A Z-axis offset

(pitch) within the orbit plane will nominally balance aerodynamic bias torques with the

gravity gradient torque.

The ICS will be used to perform all translational maneuvers of the Space Station and to

backup the CMG system when the momentum or torque exceeds the capability of the

CMG system. The RCS will also be used for primary attitude control (with reduced

pointing) in the event that the CMG system becomes inoperative.

The RCS and CMC systems will have the capability to operate simultaneously or

individually. An attitude transfer system will be used to determine the relative

alignment between the navigation base and the mounting surface of the pointing

instruments.

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The GN&C equipment will be located as follows:

The CMGs and the Inertial Sensor Assemblies will be located in the Attitude Control

Assembly, which in turn is located at the intersection of the keel and the transverse

boom. The Star Tracker triad location has riot been finalized, but it will be located

either at the upper boom or in the ACA.

If used, the magnetic torquers will likely be on the main keel and/or on one of the

booms. The solar array and thermal radiator drive unit electronics will most likely be

placed at the location of the drivers themselves.

The c. N&C processors and the interface devices will be located in the ACA.

The navigation/traffic control processors and the other miscellaneous GN&C electronics

will be located in HM2. Each interface device (ID) that supports remote sensors will be

located with the sensor Itself.

4.5 SPACE STATION DYNAMICS

The Space Station as described in reference [11 is a quite flexible structure having

frequencies in the range of .0965 Hz to 0.948 Hz. The frequency of the fundamental

vibration mode of the IOC SS is approximately 0.1 Hz. A review of the deflection

characteristics at the intersection of the keel and the upper boom reveals that a pointing

system with three degrees of freedom ( two translational and one rotational) will be

required to decouple the payload from disturbances, propagating through the flexible

body. Response of the upper boom in terms of peak accelerations and settling times due

to Shuttle bArthing, RCS Reboost firings, MRMS operations and crew motions are also

given for variou., operational con figu-r tions. Peak acceleration at the end of the upper

boom when the MRMS is located there, is found to be 0.0248 g (without payloads

attached). This implies a significant control problem for the st. 'on during MRMS

operations on flexible portions of the SS. With payloads attached to the upper boom, the

peak acceleration is reduced to 0.0003 g's.

86 
Peak acceleration levels at the upper boom root with payloads attached, due to orbiter

berthing and RCS reboost firing, are 0.0028 g's and 0.0062 g's respectively. It responds to

a crew motion (translational kick-off) with a peak acceleration of 1.42 x 10-4 g's. A

pointing mount required to provide stability In the arc sec or sub arc sec range will need

to be isolated from the effects of the various disturbances on Space Station.

4.6 DISTURBANCE SOURCES

The Space Station environment from an instrument pointing system point of view will

encompass both high frequency and low frequency disturbances. Disturbances arise both

from internal and external sources as shown in Table 12. The effect of these disturbances

on the pointing system will, In large part, be determined by the structural dynamics

(flexibility) of the base body. The latest technical information with respect to Space

Station disturbances is contained in Reference (1,45,461.

4.6.1 ENVIRONMENTAL DISTURBANCE SOURCES

The major environmental torques that will effect the Space Station attitude, as listed in

Table 12 are:

A) Aerodynamics torque caused by the motion of the Space Station through the upper

atmosphere.

B) Gravity gradient torques caused by the small difference in gravitational attraction

from one end of the Space Station to the other. 

C) Geomagnetic_ torques, due to the interaction between the Space Station magnetic field

and the earth's magnetic field.

D) Solar radiation torques, due to both the electromagnetic radiation and particles,

radiating outward from the sun.

The relative strength of the various torques will depend on bath the Space Station

orientation and the form and structure of the space station itself.

1

87

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88

- 

TABLE 12. DISTURBANCE SOURCES

INTERNAL

• RANDOM DISTURBANCES

MAN MOTION

OUTGASSING/VENTING

FLUID SLOSH

• ARTICULATED ELEMENTS

- MOBILE REMOTE

MANIPULATOR SYSTEM

- SOLAR PANEI:S, RADIATORS

- ANTENNAS

- PAYLOADS

• ORBIT MAINTENANCE

• ATTITUDE CONTROL

EXTERNAL

• ENVIRONMENTAL

- GRAVITY GRADIENT

- AERODYNAMIC FORCES

AND TORQUES

- GEOMAGNETIC FORCES

AND TORQUES

- SOLAR RADIATION

PRESSURE

- SOLAR HEATING

• DOCKING/BERTHING

a POST DOCKING OPERATIONS
The largest environmental disturbances with the Space Station in Its baseline orientation

are caused by gravity gradient and aerodynamic torques. Figure 36 shows an or`^Ital

torque profile, due to the combined effects of aerodynamic and gravity gradient torques,

which will act on the Space Station reference configuration. The torque profile assumes

a principle axis of Inertia offset from the local vertical, causing the yaw and roll cyclic

gravity gradient and pitch bias gravity gradient torque. The large pitch double bulge is

caused by aerodynamic torques on the solar arrays. These torques will be overcome by

careful vehicle re-orientation, using t he attitude control system CMG's to generate the

momentum profile, which Is also shown In the figure. Depending on the control

philosophy implemented, the momentum shown can be absorbed entirely by the CMG's

resulting in no disturbance at the upper boom (experiment platform).

Figure 37 Illustrates an example of a several degree sway about the Y-axis, which can

occur under specific aerodynamic density conditions. The amount of sway varies and L

dependent on the number of CMG's (momentum capacity) resident on the Space Station.

The figure shows that a complement of CMG 's capable of 19,000 nms of momentum

storage will control the station with zero sway under all anticipated conditions, while a

lesser momentum capacity will result In an associated vehicle sway, as shown. For

example, a 12,000 nms CMG momentum capacity will result in a 1.5 degrees, once per

orbit sway, if the diurnal bulge (variation of the aerodensity over the sunlit and dark 

portions of the orbit) Is 506 considering a .67 x 10 kg/m3 aerodynamic density. This is a

low frequency disturbance to the astrophysical experiments. In addition, the Space

Station Is already rotating at 4 degrees/minute to maintain the local vertical orientation.

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4.6.2 DISTUVBANCES DUE TO OPERATION AND INTERNAL SOURuHS

In addition to the environmental disturbance sources, disturbance from sources internal

to Space Station or arising from Space Station Operations, will also effect the pointing

system performance. These include:

A) Shuttle Docking/ Berthing Operations

B) Post Docking Operations

C) Motion of articulated elements, such as the Mobile Remote Manipulator System

(MRMS), motion of the solar arrays, radiators, antennas, and other payloads.

D) Random events, such as astronaut motion, outgassing/venting, fluid slosh, vibrations,

from pumps, CMGs, etc.

E) Space Station Orbit Maintenance and Attitude Control.

4.6.2.1 EFFECT OF ASTRONAUT DISTURBANCE AND ORBITER ATTACHMENT

The response of the Space Station to disturbances will differ as a function of its mass,

stiffness and inertia. Figure 38 illustrates the effect of a typical astronaut disturbances,

of 70 lb. of force applied at the Space Station Habitability Module 1. Two forces are

shown. F I I is a 70 lb force applied in the flight direction, causing the Space Station to

pitch. F 1 2 ;s a force in the - y direction and causes the Space Station to roll. The effect

of these forces at the upper boom (experiment platform) is shown for two cases, with and

without the orbiter attached. The effect of orbiter attachment reduces the motion of

the upper boom from ±.01 degree to ±.0025 degree as shown. This analysis was

performed by the Lockheed Missile and Space Company, during the Space Station Phase E

Proposal, using their Terflex 20-body dynamic simulation.

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Early studies for Space Station control indicate that, with a centralized control system,

large upper boom base motion results from astronaut disturbance. This is true, as shown

In Figure 39 for both low bandwidth and sigh bandwidth systems, even when active

control is implemented. Figure 39 shows that, when distributed actuators are employed,

motion at the payload is reduced. In this case, however, the control torque is large and

requires single gimbal CMGs. A study is required to trade the complexity of the Space

R t

Station Control System versus the complexity of the astrophysical platform configuration

as a function of control.

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5.0 POINTING MOUNTS FOR SPACE STATION

The Space Station is required to simultaneously accommodate instruments that have

Earth, anti-Earth, Solar and Stellar viewing requirements. Satisfaction of this

requirement from the Space Station flying in the baseline attitude orientation (LVLH)

will require a gimbal system. In the baseline orientation, the Z-axis is aligned to the

local vertical and the station rotates about the Y-axis, which is perpendicular to the

orbit plane, once per orbit. The Space Station phase B RFP (2] recognized the

requirement. It provides for up to 9 attachment points for externally mounted

payloads. Some of these nine (9) attachment points will be articulated attachments

which would provide structural support, positional control, electrical, thermal and data

Interfaces for the attached payloads. These coarse pointing/tracking gimbals would be

designed such that attitude control would be between ±1 to ±5 degrees and attitude

determination to ±.01 degrees. Maximum angular rates would not exceed ±.02 degrees

per second. In addition, the Space Station would provide knowledge of the instrument

mounting surface relative to the navigation base within ±0.002 degrees. Users or

customers with instruments that require higher pointing accuracy will be required to

provide their own pointing mounts.

Conceptually, as shown in Figure 40, the payload at',ach points could incorporate Alpha

and Beta gimba:s at the upper boom similar to those used to point the solar arrays and

radiators to provide the general pointing guaranteed by SS. Pointing systems, such as the

IPS, would then fine point the instruments, provided that the disturbance environment is

within its operational capabilities.

An alternate concept, shown in Figure 41, is to provide an attach point with a Beta

gimbal on the transverse boom outboard of the Alpha gimbal. A pointing mount could

then be attached at the Beta gimbal for fine pointing.

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The RFP [21 identified several levels of pointing accuracy required by the various

Instruments identified in the SSP Mission Data Base [3] and the percentage of customers

satisfied with each level of pointing accuracy. These data are listed in Table 13.

TABLE 13. POINTING CAPABILITY VS. CUSTOMER SATISPACTIO14

LEVELS OF ATTITUDE ERROR SATISFIED

POINTING ( PLACEMENT) STABILITY JITTER CUSTOMERS

ACCURACY (ARC S"CONDS) (ARC SEC) (ARC SEC) ( PERCENT)

1 100 20 2 40

2 30 6 0.6 55

3 10 2 0.2 60

4 3 0.6 0.06 85

5 1.0 0.02 0.02 95

6 0.15 0.002 0.002 100

The coarse pointing/tracking gimbal to be provided by SS would satisfy less than 40 per

cent of the customers. The survey of existing pointing mount technology indicates that

100% of the pointing error requirement and up to 35% of thy;• stability requirement could

be met by existing technology if the environment was comparable to the Shuttle

environment for which these systems were designed. The ASPS could satisfy up to 100%

of the attitude error (accuracy) requirement and 85% of the stability requirement for

payloads less than 600 Kg. The AGS could satisfy 100% of the pointing accuracy

requirement and up to 60 % of the stability requirement for payloads up to 7200 Kg. The

IPS could satisfy up to 90% of the pointing accuracy requirement and up to 80% of the

stability requirement. The ST capabilities illustrate that over 95 % of the attitude error

and stability requirements can be achieved by a free flyer. As shown previously, a

generic pointing mount capable of satisfying the ASO requirements would have to

accommodate a 7000 Kg payload, provide a pointing accuracy of 1.0 are sec and a

stability of 0.1 arc sec. Such a system would satisfy up to 90% of the customers. The

remaining customers would then have to include additional pointing capaoility within

their instruments for higher pointing accuracy or stability.

li
The coarse pointing/tracking provided by the SS per the RFP (2) would require that over

6096 of the customers provide their own pointing mount. This may not be cost effective,

since existing technology Indicates that it is possible to develop a generic pointing mount

that will achieve a much higher per cent of customer satisfaction.

Current or nea,. .erm pointing systems like the IPS were developed for shuttle and their

performance is predicated on operation in the Shuttle environment. The Space Station

environment is expected to be significantly more noisy and complex than the shuttle

environment. It is not likely that current systems will maintain their performance if

operated without design modifica' ions on Space Station. Development of a pointing

mount insensitive to or isolated from the effects of disturbances is the key to providing a

generic pointing mount for Space Station.

Figure 42 identifies approaches which have been taken previously to disturbance isolation

or compensation. Disturbance isolation or compensation can be incorporated either on

the pointing mount outside the payload or by the payload using image motion 

compensation. Image motion compensation, whether by image data processing, or by

movable optical elements is usually tailored to a specific Instrument and what works with

one !,istrument may not work with another. This limitation may be overcome. MSFC has

developed an IMC system to provide more stringent pointing stability for two of the

Astro-1 telescopes than could be provided by the IPS [47]. The ;ASFC IMC system

t consists of an Astro Star Tracker, IMC Electronics, and the DRIRU If. There components

replace the IPS computers and sensors.

The IMCS, in conjunction with the telescope secondary mirror actuators, provides the

ability to improve pointing stability by sensing the IPS motion due to external

disturbances and issuing commands to the secondary mirror actuators. The DRIRU

senses the high frequency (up to 15 hz) errors and the AST is used by the IMCS software

i to correct for drift of the DRIRU II. The IMCE provides for a dedicated HRM channel

interface to down link IMCS house keeping and pointing performance data. The hardware

complement can be considered portable, since it can be easily adapted to other

applications.

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A generic pointing mount, which provid-s a high de ree of pointing stability, will :educe

the requirement for payloads to employ these or similar techniques.

Conside r pro, iding disturbance compensation outside the payload on or near the pointing

mount. Reference 1 481 developed fundamental stability and disturbance rejection

characteristics for simple and two stage gimbal systems such as the IPS, ASPS and

Gimbalflex, as well as a soft mount inertially rsacting concept.

The si mple gimbal system considered in 14 8 1 is shown in figure 43. The two stage system

considered was a model of the ASPS which has a magnetically isolated vernier stage on

top of a gimbal system. The soft mount inertially reacting concept is described as a

payload mounting body, equipped with both actuators and sensors connected by a soft

Interface to the base body. The ictuators could be reaction wheels or CMGs and the

sensors would be star tracke ► s and gyros. An inertially referenced proportional plus

derivative control law was assumed. Conclusions derived in 1481 are directly applicable

to selection of a pointing concept for SS and are reiterated here.

Simple gimbal systems will always have the problem of dynamic interaction with the base

body. There is a frequency range determined by the base body natural frequencies in

which the pointing system control frequency must not lie in order to maintain system

stability. This is an unacceptable restriction for SS since there are many flexible modes

below 1 Hz.

Simple gimbal systems with end mounted payloads also will be affected by base body

translational motions at frequencies above the control bandwidth. If the pointing system

Is offset from the base body center of mass it will also be affected by high frequency

rotational motions.

Systems which employ an isolation or soft suspension stage on top of the gimbal will

perform well from a rigid base body such as Shuttle. However, they may still have the

same dynamic interaction problems as simple gimbal systems when operating from a

flexible base body like SS.

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A soft mount, inertially reacting system will minimize dynamic interactions and

approach the performance of a free flying spacecraft. I t could still obtain required

resources from the base body.

The soft mount, inertially reacting concept is equivalent to t^.e SEM which has been

under study at MSFC for several years. Figure 44 identifies its potential advantages over

other concep ts. Hardware is available to implement the pointing and control p,-- t }^^ -°

the SEM but implementation of the suspension system needs to bid studied. The coneept

also has applications to other payloads which require low acceleration environments. The

interaction of the SEMs control system with the Space Station ACS must also be

considered.

6.0 INSTRUMENT FOCAL PLANE SHIRTSLEEVE ACCESS

Maintenance, servicing and repair of attached paylnads, free flyers or instruments on coorbiting platforms, will be accomplished by EVA and MRMS operations. The primary

maintenance technique will be by changeout of OF Us supplemented by a capability to

perform some troubleshooting and repair at an elemental level (1).

A major find of the "Task Force Study on Utilization of the Space Sta t ion" (TFSUSS) was:

"The most important new capability to be provided by the Space Station Program for

astronomy and astrophysics is that for routine maintenance and repair of co-orbiting

facilities."

The "TFSUSS" also commented that:

"in addition, a large hangar-like module providing a'shirtsleeve' environment would

greatly augment the maintenance capability and be very helpful in the assembly, test,

and checkout of instruments delivered to the Station."

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In the sense used, the term "shirtsleeve environment" would require a pressurized volume

or module in which an astronaut in normal work attire could perform hands-on focal

plane Instrument change-out, repair or maintenance. The payloads to be serviced would

include the attached payloads, Free- flyers and instruments located on co-orbiting

platforms.

6.1 IOC SPACE STATION "SHIRTSLEEVE ENVIRONMENT"

As described previously, the IOC Space Station will have five pressurized modules.

Provisions for maintenance and repair of equipment have been provided in HM2. The

module contains a maintenance work station which will provide the capability for some

level of troubleshooting and repair. The work station will be outfitted with controlled

lightning and magnification capabilities. A dedicated microprocessor with a printer and

display capability will serve as a diagnostic tool supplemented by hand-held

instruments. Electrical test equipment, a maintenance tool kit an y a soldering kit will be

provided. The maintenance sta lJon will be equiped with various small item restraints: a

particle control system, a measurement kit (tape, micromet,-r, gauges, etc.), an adhesive

kit (tripe, cement, etc.), and a fastener kit (nuts, bolts, washers, etc.). If it is desired to

perform troubleshooting and repair at this work station, the unit would have to be

transported to It. In the case of the astrophysics payloads, an EVA would be required to

obtain the unit requiring servicing and transport it back to the HM2 work station. The

MRIMS may be used to perform or assist in this operation if possible. The instrument or

component size will be limited by the module hatch diameter and available space within

HM2.

If moving a unit to the work station is not practical, an alternate scheme is provided. A

portion of the maintenance work station in HM2 is removable and can convert into an

EVA compatible work station. It can be transported, along with necessary equipment to

perform troubleshooting and repair, to any location on the Space Station requiring

servicing. It will be designed so that it can be positioned FLnd rigidly restrained at the

work site.

106
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6.2 [CC SPACE STATION SERVICING FACILITY

An external Servicing Facility will be provided on board the manned IOC Space Station.

It will be used to service free- flying satellites, co-orbiting platforms and attached

payloads. The Servicing Facility will also provide for the storage of ORUs, instruments

and tools needed for use during servicing. It will also provide for the maintenance of the

OMV, OMV kits and storage of payloads, OMV and OMV kits.

These operations are performed in two separate bays. One bay is used for servicing

operations and anotoer for refueling operations. These two separate bays were

determined to be necessary to prevent degradation of instrument optics by contaminants,

which may be present due to refueling operations. This concern with potential

contamination also dictated that the servicing bay be located away from the pressurized

modules and the Orbiter berthing area.

6.3 PRESSURIZED HANGAR DESIGN CONSIDERATIONS

The servicing bay for the IOC S.S. is estimated to be a cylindrical volurne, which is 30 ft.

In diameter and 70 ft. long. This would allow for the berthing of a 15 ft. diameter by 60

ft. long satellite, with clearances all around for movement of an EVA crew and the

placement of work stations.

A pressurized hangar, in which to perform services or provide access to an instrument

focal plane, would be of similar dimensions. Consider the HST, its dimensions are

approximately 15M x 5M x 5M (49 ft. x 1.6 ft. x 16 ft.). Approximate dimensions of the

SOT are 7.5M x 4.4M dia. (25 ft. x 15 ft. dia.). The SOT attached to the IPS, along with a

frame to secure them, wouic: have approximate dimensions of 13.4M x MM dia. (44 ft. x

15 ft. dia.). It is not unreasonable to assume that a hangar of the dimensions cited will

be required to contain these large payloads and still provide working room, especially if

more than one payload requires servicing at the same time.

Consider the location of the Hangar. Concerns with respect to instrument contamination

would dictate that it be located on the upper keel above the transverse boom. This
location would either require crew EVA for entrance/exit and the associated

requirements that go with EVA, or alternate means of entrance/exit which preclude

EVA. A pressurized tunne! running from one of the pressurized modules on the lower

keel to the hangar could provide entrance/exit; so could a pressurized "elevator," which

would traverse from a module on the lower keel to the Hangar. If crew access to the

Hangar is the driving concern, then the Hangar should be located near the pressurized

modules where a short crew transfer tunnel or module to Hangar interconnection would

provide access. Regardless of the location, the Hangar, if Incorporated into S.S., would

have an effect on the ACS. The mass, size, shape and location of the Hangar would

certainly force adjustments in the ACS so that it could orient and control the S.S. as

required. The effects on the S.S. capability to provide required micro-g environment and

point payloads must also be considered. The Hangar would have to have a large access

port , enable the payloads to be brought In and out. Allowing the gas, required for a

shirtsleeve environment, to escape each time the payload access port is opened, is

unacceptable both from a contamination and resource point of view. A system will be

required that will pump down the Hangar before the door is opened and repressurize it

after the door is closed. The hangar environment itself should be controlled to some

acceptable level to prevent contamination of the instruments . Obviously, the Hangar

will have a major impact on the Environmental Control and Life Support System

(ECLSS). It will require a separate ECLSS or be interfaced with the Space Station

ECLSS, potentially requiring re-sizing of this system. Incorporation of the Hangar would

have a major impact on most SS subsystems, including: Power, Thermal Control, Attitude

Control, Environment and Life Support, information and data management, propulsion,

structures and mechanisms, and crew systems. The Hangar would have to be assembled

in orbit because of she orbiter constraints on cargo size and weight. Figure 45 illustrates

the envelope for the orbiter cargo bay [491. The envelope is of cylindrical shape with a

diameter of 15 feet (4.572 M). Length available for the cargo would be 47' (14.266 M).

The orbiter has the capacity to place up to 29,484 Kg (65,000 lbs.) into orbit.

Maximum cargo landing weight is 14,515 Kg (32,000 lb.). In essence, the cargo weight is

constrained to 14,515 Kg, although the orbiter can return and land with cargo weight up

to 29,484 Kg (65,000 lb.) under abort conditions.

These considerations identify that incorporating a module on Space Station which would

provide a shirtsleeve environment would have many impacts. Incorporating such a

facility would require a major program commitment.

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6.4 PAYLOAD DESIGN CONSIDERATION FOR SHIRTSLEEVE SERVICING

The performance of on-orbit servicing tasks will impact the design of the hardware

requiring service. On-orbit servicing tasks would include:

- observation and inspection

- test, fault detection and isolation

- instrument mod!fication or change-out

- component repair or rep ► acement

- alignment and calibration

- surface restoration

The requirement to perform these tasks should be included as part of the system or

hardware design and performance specifications. These specifications would incorporate

the requirement for accessibility to the portions of the hardware which will require

servicing. The hardware design should facilitate access to -emoval and installation of

components, which are planned to be replaced on orbit. it should provide test points for

connection of Electrical Support Equipment (ESE) required to perform fault isolation and

detection.

From a mechanical or packaging point of view, the design should consider:

crew visual and manual access

- tool insertion, engagement and movement envelope

- electrical cable identification, and positioning

- electrical connector location, indexing/ keying and ease of mate/demate

- access to guides, rails, alignment aids

- access to fasteners, hold-down and release devices, clamps, etc.

- placement of grounding straps

- hardware corners and edges

- protection or safing of any sensitive items

- placement of hand-holds or tethers

- effect of crew-induced loads on the hardware

- crew safety

iacement, attachment, and thickness of any protective blankets, wrappings

or covers

110
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Electrical designs should facilitate on-board testing to validate payload /system

M interfaces and performance of fault detection and isolation.

( Since there is a potentially large demand for on -orbit servicing, the designs of the

( payloads should be coordinated to the degree possible to maximize the benefits of the

servicing facility. Designs should be modular to allow easy removal and replacement of

components. Designs should Incorporate the use of common or standard mechanical and

electrical interface devices. These Include the use of standard latches, fasteners,

electrical connectors, electrical circuits, data bus, electrical power and utility

connectors to simplify the performance of on-orbit servicing. Achievement of modular

payload designs using standardized connections, mounting provisions and configuration

restrictions, will influence payload designs in terms of integration compatibility with

t Space Station, as well as enhance the on -orbit servicing capability. These designs should

treduce the requirement for provision of payload unique special support equipment.

Some payloads, the HST and AXAF, :or example, have or are incorporating these design

considerations, since they are required to be serviced on-orbit via EVA. Other payloads

and systems have not been specifieally designed for on-orbit maintenance. The 1PS

maintenance concept, for instance, is to remove and replace to the functional line

replaceable unit (LRU) level, either on a scheduled or unscheduled basis [25]. A LRU is

defined as a vombination of parts, a component, a subassembly or an assembly

manufactured and/or assembled into one unique unit, or so arranged that, together the

combination is common to a single mounting, and in addition provides a complete

function to the higher level equipment within which it operates. Examples of possible

LRU's are: fuses, harness, circuit boards, electrical units, thermal blankets, mounting

struts, integrated assemblies, control panels, etc.

Previous demonstrations of on -orbit EVA servicing and repair (e.g., Solar Max Repair

Mission) indicate that on -orbit maintenance of systems like the 1PS in a shirtsleeve

facility sould be possible to some extent, depending on the spares, tooling and test

equipment provided.

111
7.0 CONCLUSIONS AND RECOMMENDATIONS

The Space Station "local--vertical local-horizontal" orientation is not conducive to

performance of solar or stellar pointing experiments without additional pointing

mechanisms. The Space Station RFP (21 recognized this and included the requirement to

provide articulated attachments for pointing instruments. The station would provide

two-axis pointing with an accuracy of ±1 degree and stability of ±0.02 deg/sec.

Customers, who require more precise pointing, would have to provide their own pointing

mounts. The study indicates that the technology required to provide precise pointing for

payloads exists or is rapidly becoming available. It is recommended that a trade study be

performed to determine the cost effectiveness of requiring the individual instr iment to

provide precise pointing versus the Space Station providing a generic precise pointing

capability.

The dynamic characteristics and the disturbance environment of Space Station are the

challenges with respect to pointing mount design. Existing systems and concepts were

designed for operation from the Space Shuttle and it is unlikely that they will sustain

performance when exposed to the Space Station disturbance environment. Development

of a disturbance Insensitive pointing mount is the key to providing a precise generic

pointing mount for Space Station. An inertially reacting pointing system Is less sensitive

to dynamic interactions with the base body than gimbal systems. It is recommended that

the Suspended Experiment Mount concept be Investigated for use as a potential pointing

mount for Space Station.

The Spacelab IPS Is an existing flight system. It represents a large investment of time

and money both from ESA and NASA. It should be considered for use as a generic

pointing mount for Space Station. It is recommended that a study be made to determine

if the Spacelab IPS can be modified for ise as a generic pointing mount on Space Station

and to determine the nature and extent of any required modifications. The Talon Gold

Gimbal is a mature system, which provides for a high degree of disturbance rejection in

the gimbal drives. Its ability to provide power, signal, and fluids across the gimbals

without the use of cables or hoses also enhances its performance. It is recommended

that the Talon Gold Gimbal be studied for potential use as a Space Station pointing

mount. The technology incorporated into Talon Gold should also be considered for use in

other Space Station mechanisms.

112

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Availability of a module which will permit instrument change-out, repair and

maintenance, is particularly desirable from a customer's view point. Inclusion of a

shirtsleeve environment module would impact most Space Station subsystems. Addition

of a facility of this type would require a major program commitment. It is recommended

that a study be performed to determine the cost impact of providing a "shirtsleeve

environment" facility versus the benefits gained by the customer with respect to longer

orbital lifetime, increased observing time and less complicated EVA tasks.

I'.

113
:r ..

APPENDIX 1

ASTROPHYSICS PAYLOADS

OBJECTIVES AND DESCRIPTIONS
SAAX0001 SPECTRA OF COSMIC RAY NUCLEI

OBJECTIVE:

To help explain the characteristics and distributions of galactic cosmic ray propagation

through interstellar space.

DESCRIPTION:

CRNE exists. It was developed through the spacelab program: a one to two year

exposure on the Space Station will enhance the scientific return by two orders of

magnitude. Power, telemetry, contamination, and pointing requirements will be easy to

satisfy on the gravity gradient stabilized station. The flight duration could be made

variable, dependent on data gathered, but a two year exposure is baselined.

SAAX0004 SHUTTLE INFRARED TELESCOPE FACILITY

OBJECTIVE:

To conduct definitive high-sensitivity infrared photometric and spectroscopic studies of a

wide range of astrophysicu! phenomena.

DESCRIPTION:

The infrared platform will cam the Space Infrared Telescope Fdcility (SIRTF), which is

being planned as a 15-year mission with refurbishment at 2-year intervals. Consideration is being given to two options for the infrared platform: one as a platform

derived from the Space Station's subsystems and the other as a platform derived from

other astro platform subsystems. In either case, the infrared platform and Its SIRTF

payload will be serviced at the Space Station. This entry in the data base contains

information delineating requirements on the platform, as well as requirements for

servicing.

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SAAX0005 TRANSISTION RADIATION do ION CALORIMETER

OBJECTIVE:

To observe electrons, protons, and helium nuclei in order to determine if there is a major

change in the composition of cosmic rays at very high energy.

DESCRIPTION:

The transition radiation and ionization calorimeter experiment is an out-growth of a

spacelab experiment. Long observation times improve its capability by a factor of ten.

It would make a good candidate for a station resident experiment, since pointing

accuracy and contamination are not primary factors; some preference for high

Inclination, but not critical.

SAAX0006 STARLAB

OBJECTIVE:

To investigate the visual, ultra-violet, and near infrared properties of extra galactic

space, the milky '.ray galaxy, and the solar system with high spatial resolution and wide

angle imaging.

DESCRIPTION:

'rhe first Starlab mission is to be by shuttle.

A 1-2
SAAX0007 HIGH THROUGHPUT MISSION

'TIVE:

To investigate high energy processes in astrophysical systems, including the very

compact sources associated with black holes and neutron stars.

DESCRIPTION:

HTM is an X -ray telescope which obtains good resolution by trading off mirror quality for

quantity. Duration can continue beyond year 2000.

SAAX0008 HIGH ENERGY ISOTOPE EXPERIMENT

OBJECTIVE:

To search for rare nuclei and exotic particles like magnetic monopoles and to measure

the composition of - ltra-heavy nuclei.

DESCRIPTION:

a the HEIE is selected as a polar platform payload, it will be placed in a 57 deg. (or

higher) orbit by the shuttle.

A 1-3
SAAX0009 PINHOLE OCCULTER FACILITY

OBJECTIVE:

To study non-thermal phenomena of plasma dynamics in the solar corona and to observe

the acceleration of nonthermal particles in solar flares and in coronal disturbances with

both x-ray and coronagraphic instruments.

DESCRIPTION:

The POF consists of a 32 M boom that separates an osculating mask from an array of

detectors and telescopes. The name 'pinhole' derives from the use of a remote occulter

(for coronagraphic studies), containing an array of small apertures, to obtain high angular

resolution of hard x-radiation.

SAAX0010 SOLAR CORONA DIAGNOSTIC MISSION

OBJECTIVE:

To determine the cause of solar corona heating.

DESCRIPTION:

The SCDM has been envisioned as a possible free flyer. However, the SCDM should be

studied as a station resident experiment, operated and maintained by the station

personnel. Also seen as the third user of a common explorer bus, replacing Fuse.

Replacement effected at the Space Station.

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SAAX0011 ADVANCED SOLAR OBSERVATORY

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OBJECTIVE:

To carry individual instruments capable of examining solar phenomena that can be

pointed to regions of interest on the solar disc or throughout its atmosphere.

DESCRIPTION:

The ASO consists of the extreme ultraviolet telescope, the solar soft x-ray telescope,

pinhole/occulter, and the Solar optical telescope operated attached to the Space Station.

SAAX0012 HUBBLE SPACE TELESCOPE

OBJECTIVE:

To learn of the evolution of stars, of ours and other galaxies, and to explore quasars,

pulsars, gas clouds and other planets.

DESCRIPTION:

The Hubble Space Telescope (HST) will be in service fr!: about 5 years before a station

based OMV will be available for servicing. Although not in the planned baseline for HST,

it could be returned to earth for updating and refurbishment and replaced in orbit for

service by the OMV. From then on, the OMV could be used for periodic servicing and

reboost. HST Space Station support operations are assumed to begin in 1992.

ti

A 1-5
SAAX0013 GAMMA RAY OBSERVATORY

OBJECTIVE:

To make the next major step in gamma ray astronomy by providing the first opportunity

for comprehensive observations covering over five decades of energy from 0.05 to 30,000

MEV.

DESCRIPTION:

The observatory will be a three axis stabilized free flyer, capable of pointing to any

celestial target for long periods of time. The attitude control accuracy is 0.5 degrees

and the attitude determination accuracy is 2 arc minutes. All telemetry and command

will be done by TDRS. GRO will consist of 4 experiments that collectively span the

photon spectrum from 0.05 to 30,000 MEV.

SAAX0014 X-RAY TIMING EXPLORER

OBJECTIVE:

Measure temporal variability of x-ray objects for time scales ranging from microseconds

to years.

DESC RIPTION:

Free-flyer explorer-class spacecraft, which houses a largq area proportional counter and

a wide field optical camera. Could be first of a series of replzceable telescopes on a

single free-flyer bus. Replaced at Space Station by Fuse in 1993 in thi' scenario.

A 1-6
SAAX0016 SMALL SOLAR PHYSICS

OBJECTIVE:

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Investigate causes and effects of solar flares.

DESCRIPTION:

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MMS with telescopes covering UV to gamma rays.

SAAX0017 ADVANCED X-RAY ASTROPHYSICS FACILITY

( OBJECTIVE:

I

X-ray astronomy research in the areas of source location and structure, spectroscopy,

polarlmetry, and temporal behavior.

11 DESCRIPTION:

X-ray platform will carry the Advanced X-ray Astrophysics Facility (AXAF). The X-ray

platform will have spacecraft and instrument modules that will be replaced at regularly

scheduled intervals at the space station. AXAF will be a national x-ray telescope

facility with advanced capabilities in energy range, sensitivity, angular resolution and

mission lifetime. AXAF instruments have been selected, and the platform is in Phase B

definition. AXAF is a candidate for a payload on a space station platform.

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SAAX0018 VERY LONG BASELINE IN"ERFEROMETER

OBJECTIVE:

To provide maps of compact celestial radio sources with finer resolution, less ambiguity,

and more efficiency than earth-bound VLSI techniques.

DESCRIPTION:

t The very long baseline interferometer mission requires correlation of measurements with

the ground. The orbiting portion is placed by the shuttle. Co-habitation with the station k

is a possibility, if pointing and contamination requirements can be met.

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SAAX0019 FAR UV SPECTROSCOPY EXPLORER

OBJECTIVE:

To carry out high and low resolution spectroscopy of distant stars, galaxies and

intersteller matter in the 90-120 NM spectral range.

DESCRIPTION:

The FUSE is being planned for a shuttle launch and injection into geo-orbit. Another plan

is to make fuse the second in a series of explorer missions, which use a single spacecraft

bus. The replacement would take place at the Space Station. It follows XRTE and

preceeds SCDM in this scenario.

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SAAX0020 LARGE DEPLOYABLE REFLECTOR

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OBJECTIVE:

To conduct infrared astronomical investigations for the survey of a wide variety of

astrophysical phenome:ia throughout the infrared spectral region.

DESCRIPTION:

LDR will be brought to the Space Station on two shuttle loads. It will be assembled and

tested at the station. It gets placed in o,)erational orbit by the OMV and is periodically

serviced by the OMV.

SAAX0021 SUPER CONDUCTING MAGNET 

OBJECTIVE:

Antimatter and energetic isotopes contain information on the propagation of cosmic rays

through the interstellar medium.

DESCRIPTION:

A superconducting magnet and charged particle track detectors form a facility in which

the charge to mass ratio of particles can be measured. Associated detectors depend on

the specific objective and are changed for different experiments. A bucking coil cancels

the large scale field. A liquid helium cryostat is located between the two coils, attached

to the top of a gravity-gradient stabilized space station. Power, commands, and data

will be handled automatically through the station.

Al-9
SAAX0022 THROUGH SAAX0025, SPACE STATION SPARTAN PLATFORM

OBJECTIVE:

Low cost astrophysics and space science payloads to perform scientific investigations and

test new instruments and concepts. Low cost, minimum turn around time is a prime

objective.

DESCRIPTION:

The Space Station based Spartan will be an evolution from the shuttle Spartan program.

Scientific payloads will be integrated with the spartan carrier at the space station. The

Spartan will be then transported into an appropriate orbit by the OMV and operate

essentially independent of the Space Station for up to 3 months before being retrieved by

the OMV, de-integrated on the station, serviced and re-deployed with a new payload for

another 3 months.

SAAX0026 LEASED PLATFORM 1 (EXPLORER)

OBJECTIVE:

Exploxer missions are low-cost missions for special astrophysics, space plasma physics,

and atmospheric investigations from space. The first payload would be XRTE. 

DESCRIPTION:

The leased platforms will provide spacecraft services to the explorer payloads. Specific

payloads are to be selected by OSSA as part of the approved explorer program. XRTE

uses a large-area proportional counter, a scanning shadow camera, and a hard x-ray

timing experiment to learn about the physics of neutron starts, black holes, and other

compact celestial objects through variations in their x-ray output. Payload has been

selected and the instruments are being defined. ?ayload will be replaced by another

' explorer after 2.5 years.

Al-10
SAAX0027 LEASED PLATFORM 2 (EXPLORER)

OBJECTIVE:

Explorer missions are low-cost missions for special astrophysics, space plasri ► a physics,

and atmospheric investigations from space.

DESCRIPTION:

The leased platforms will provide spacecraft services to the explorer payloads. Specific

payloads Rre to be selected by OSSA as part of the approved explorer program. Payload

could be the FAR Ultraviolet Spectroscopy Explorer (FUSE).

SAAX0028 LEASED PLATFORM 3 (EXPLORER)

OBJECTIVE:

Explorer missions are low-cost missions for special astrophysics, space plasma physics,

and atmospheric investigations from space.

DESCRIPTION:

The leased platforms will provide spacecraft services to the explorer payloads. Specific

payloads are to be selected by OSSA as part of the approved explorer program. First

payload could be the Solar Corona Diagnostics Mission.

Al-11
SAAX0029 LEASED PLATFORM 4 (EXPLORER)

OBJECTIVE:

Explorer missions are low-cost missions for special astrophysics, space plasma physics,

and atmospheric investigations from space.

DESCRIPTION:

The leased platforms will provide spacecraft services to the explorer payloads. Specific

payloads are to be selected by OSSA as part of the approved explorer program.

SAAX0030 SPACE STATION HITCHHIKER 1

OBJECTIVE:

Very specialized science and applications investigations will be conducted with small

payloads attached to the outside of the Space Station. Specific missions are to be

determined by OSSA.

DESCRIPTION:

Space Station Hitchhikers are small, low-cost investigations attached to the outside of

the Space Station. Payloads would fit into the shuttle Hitchhiker envelope -- gas can, sill

and bridge carriers. Power, command, and data will go through the station to the

ground. This Hitchhiker describez a get-away-special cannister with Hitchhiker power

and telemetry.

Al-12
SAAX0031 SPACE STATION HITCHHIKER 2

OBJECTIVE:

Very specialized science and applications investigations will be conducted with small

payloads attached to the outside of the Space Station. Specific missions are to be

determined by OSSA.

DESCRIPTION:

Space Station Hitchhikers are small, low-cost investigations attached to the outside of

the Space Station. Payloads would fit into the Shuttle Hitchhiker envelope - - gas cans,

sill and bridge carriers. Power, command, and data will go through the station to the

ground. This hitchhiker describes the MSFC hitchhiker concept for the shuttle.

SAAX0032 SPACE STATION HITCHHIKER 3

OBJECTIVE: t

Very specialized science and applications investigations will be conducted with small

payloads attached to the outside of the space station. Specific missions are to be 

determined by OSSA.

DESCRIPTION:

Space Statio► , Hitchhikers are small, low-cost investigations attached to the outside of

the Space Station. Payloads would fit into the Shuttle Hitchhiker envelope - - gas cans,

sill and bridge carriers. Power, command, and data will go through the station to the

ground. This hitchhiker describes a concept developed at GSFC for use on the shuttle.

Al-13

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APPENDIX 2

POINTING MOUNT DESIGN CONSIDERATIONS

16
I

TABLE OF CONTENTS

SECTION NO. TITLE

1.0 PAYLOAD POINTING MOUNT DESIGN CONSIDERATIONS ... A2-1

1.1 KEY PAYLOAD REQUIREMENTS ...................... A2-5

1.2 PAYLOAD POINTING SYSTEM TERMINOLOGY ............ A2-5

1.3 A2-7

1.4

DISTURBANCE SOURCES

POINTING MOUNT ARCHITECTURE ............................ A2-7

1.4.1 PAYLOAD MOUNTING .............................. A2-8

1.4.2 LOS ARTICULATION OPTIONS ........................ A2-9

1.4.3 SENSING OPTI ONS ............................ . .... A2-11

1.4.3.1 ATTITUDE SENSING HARDWARE ...................... A2-11

1.4.3.2 SENSING HARDWARE LOCATION ..................... A2-13

1.4.4 CONTROL METH01iOLOGY OPTIONS .................. A2-14

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LIST OF ILLUSTRATIONS

FIGURE NO. TITLE PAGE

A-1 PLATFORM POINTING SYSTEM CONCEPTUAL DIAGRAM ... A2-2

A-2 MULTI—LEVEL ARTICULATION ARCHITECTURE .......... A2-10 

L 

LIST OF TABLES

TABLE NO. TITLE PAGE

I

A-1 POINTING SYSTEM DESIGN CONSIDERATIONS ........... A2-1

A-2 DISTURBANCE SOURCES ........................... A2-14

A2—i
1.0 PAYLOAD POINTING MOUNT DESIGN CONSIDERATIONS

In simple form, an instrument can be accurately pointed from a base body If a pointing

system is provided that can compensate for base body motion, slew from object to

object, and track the object, as necessary, to compensate for orbital rotation rate when

r ` the base body is operated in a local vertical orientation, as in Space Station. These

x.

simple tasks rapidly become complicated as the designer asks questions regarding size,

J

shape, accuracy, time, amplitude, frequency, cost and schedule, in order to produce the

j hardware with the required capability. A list of the various studies and task, that need

to be performed to arrive at answers to the pointing mount designers questions, are

Identified in Table A-1.

TABLE A-1. POINTING SYSTEM DESIGN CONSIDERATIONS/TnSKS

• PAYLOAD REQUIREMENTS • ERROR ANALYIS

• BASEBODY DISTURBANCES • INTERFACE DEFINITION

(ELEC., MECH, ETC.)

• BASEBODY DYNAMICS 0 COMMAND FUNCTION

• ENGIRONMENTS • TECHNOLOGY DEVELOPMENT

• ACTUATOR STUDY • IN-FLIGHT ADJUSTMENTS

• SENSOR STUDY • GROUND OPERATIONS

• CONTROLLER STUDY • COST

• ISOLAT'^" MOUNT STUDY • SCHEDULE

• ESTI,.1ATOR STUDY • OTHER

A simple block diagram of a Space Station mounted platform pointing system is shown in

Figure A-1.

A2-1

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The block diagram is a convenient and well known method for representing a physical

system or the set of mathematical equations that characterize its components. The

blocks are the transfer elements which represent the functional relationships between

the various inputs and outputs. The operations of additions and subtractions are

represented by the small circle, called a summing point. The output of the summing

point is the algrebraic sum of the inputs, each with an appropriate sign. The

platform/payload is the part of the system that needs to be controlled. Inputs to the

system come from three sources:

A) the command generato,

B) dynamics of the base body (base motion)

C) environmental disturbances.

t Ideally, we wish the pointing system to respond only to the command generator. The

command generator provides a command profile for moving the instrument line of sight

(LOS), either between observations or during observations, to compensate for the

station's orbital rate motion due to its local vertical orientation. Inputs resulting from

the base body dynamics and the environment are disturbances which produce unwanted

effects. Table A-2 lists the major sources of disturbances which arise both from the

i dynamics of the base body, events occurring on the base body and from the environment

a

and events external to the base body. The effect of these disturbances on the pointing

system will, in large part, be determined by the structural dynamics (flexibility) of the

3

base body. Refer to the simple block diagram of Figure A-1. The command generator

provides a command to move or maintain the instrument LOS at some desired

d location/direction. This is compared to the actual location of the LOS, as best can be i

determined by the inertial sensor assemblies, which may include gyros, accelerometers,

gimbal position encoders, and star trackers. The resultant difference or error signal is

utilized by the controller, which computes commands according to a control law and

issues these commands to a torquing device or actuator. A control law is a principal on

which the controller is designed to achieve the overall system performance. The torque

actuator which can be a gimbal torque motor, reaction wheel or CMG provides the torque

required to move the platform/payload. Platform/ Payload rate, position and attitude

data are then fed back and compared to the original input.

A2-3
TABLE A-2. DISTURBANCE SOURCES

INTERNAL EXTERNAL

• RANDOM DISTURBANCES

- MAN MOTION

OUTGASSING /VENTING

- FLUID SLOSH

• ARTICULATED ELEMERTS

• ENVIRONMENTAL

- GRAVITY GRADIENT

- AERODYNAMIC FORCES

AND TORQUES

- GEOMAGNETIC FORCES

AND TORQUES

- SOLAR RADIATION

PRESSURE

- SOLAR HEATING

• DOCKING/BERTHING

- MOBILE REMOTE

MANIPULATOR SYSTEM

- SOLAR PANELS, RADIATORS

- ANTENNAS

- PAYLOADS

• POST DOCKING OPERATIONS

a ORBIT MAINTENANCE

• ATTITUDE CONTROL

The attitude update control law removes known systematic errors, which are introduced

by known or measurable sensor inaccuracies by means of state estimation techniques.

These are normally implemented by means of Least Squares Estimators or Kalman

Filters.

A2-4 
rr

1.1 BEY PAYLOAD REQUIREMENTS

The payload or instrument requirements, which most directly influence pointing system

design, are payload mass, shape, pointing accuracy, stability, acceleration, tracking rate

and settling time.

Payload weight and shape affect the pointing system structural design for withstanding

both launch and operational loads and, because of its direct influence on moment of

Inertia, it becomes a major factor in designing mechanical isolators, gimbal bearings and

actuators. Pointing accuracy requirements affect allowable mechanical alignment

accuracies and thermal distortions. Stability Requirements primarily affect control

system design, sensor selections and allowable control system bandwidth. Payload

requirements for acceleration, tracking rate and settling time directly influence the

selection of mechanical isolation devices, torquer/actuator design and overall control

system design.

1.2 PAYLOAD/ POINTING SYSTEM TERMINOLOGY

The definition of several terms used to characterize payload requirements or pointing

system capabilities are provided. "Pointing Error" - Defines the total angular error that

will be allowed by the Pointing System. Pointing error is normally the sum of pointing

accuracy and pointing stability. Statistical combinations are permitted, but the method

must be defined.

"Pointing Accuracy" - Defines h^w close the Pointing System must initially point to the

desired target. The numerical value of pointing accuracy is usually set by an instrument

field-of-view (FOV). Pointing accuracy is static or at Ir.,ast quasi-static relative to

experiment operation time. Examples of errors that contribute to pointing accuracy are

sensor readout, etc. Pointing accuracy is often referred to as bias accuracy or bias

error. Pointing accuracy numerical values are considered to be 3 sigma values because

they set the probability of acquisition.

"Pointing Stability" - Defines how close the Pointing System must stay, during an

experiment operation, pointed to the initial point. The initial point is set by the pointing

accuracy requirement. Errors that contribute to pointing stability are dynamic in nature,

such as sensor noise, structural bending vibrations, limit cycles, and perhaps thermal

A2-5 
distortions. Pointing stability is often thought of as two conditions; either quiescent or

disturbance. Quiescent Stability is the case where there are no external disturbances

acting on the Pointing System. The quiescent stability values are considered as 1

sigma. The disturbance stability errors are due to external disturbances acting on the

Pointing System. Crew motion and RCS thruster firings from the Space Station are

examples of the external disturbances. The disturbance stability errors are considered as

peak values.

"Stability Rate" - Alternate definitions for the term are used. U is sometimes defined as

the first derivative of the stability error time history. Alternately, it is defined as the

maximum total attitude error that can be tolerated during any one (1) second of time

When defined as the latter, stability rate is not the time derivative of attitude error.

Stability rate is often referred to as jitter or sometimes jitter rate.

"Slew Rate" - Defines the maximum angular rate required during experiment operation or

between operations as in slewing to another target. Since s11 5!wing is dynamic, a settling

time (the time to reach steady state) must be defined. This also implies an acceleration

level.

"Settling Time" - Defines the allowable time for the pointing control system to reacy

steady state following any slew. When steady state is reached, it is understood that the

de r '---d pointing accuracy is met.

"Line of Sight (LOS)" - The LOS is a v.ictor wti-h considers with the payload optical

axis. For many payloads, the LOS intersectF the pointing system center of rotation

(coordinate system origin) and is perpendicular to the payload interface plane.

"Servo Bandwidth" - Bandwidth is defined as the natural frequency of the controller

position closed loop transfer function (-3db crossover).

A2-6

Now v 
1.3 DISTURBANCE SOURCES

Once the payload requirements have been defined, the disturbance environment in which

a pointing system will perform needs to be defined. Disturbances arise both from

int, .rnel and external sources as shown in Table A-2. The effect of these disturbances on

the poi nting system will, in large part, be determined by the structural dynamics

(flexibility) of the base body.

1.4 POINTING MOUNT ARCHITECTURE

An understanding of current pointing systen. technology is facilitated by a review of the

multilevel architecture currently utilized. Reference (511 provided such a review and it

is summarized here. Pointing mount architectural considerations can be placed into

three categories:

a) LOS Articulation Options (LOS)

b) Sensing Options

e) Control Methodology Options

LOS Articulation Options refer to the number of stages of isolation between the base

body and the payload. Sensing Options refer to the source of the attitude reference,

which, in turn, will influence the choice of the system attitude sensor (s). Control

Methodology refers to the manner in which the control system loop is closed.

The method used to interface or physically attach the payload to the pointing mount will

have an impact on the choice of the basic architectural options.

A2-7
1.4.1 PAYLOAD MOUNTING

Previous systems and concepts have concentrated on three pointing system

configurations with respect to payload mounting or interfacing to the pointing system.

These are:

a) Conventional Girth Ring or Center of Mass Mount Pointing System

b) Base Plate Mount, End Mount or Cantilever Mount Pointing Systems

c) A Suspended or Floated Mount Pointing System

Center of Mass mounting systems require the payloads' center of mass to coincide with

the center of rotation (COR) of the pointing system gimbal axis. This method of payload

mounting reduces the effects of high frequency disturbances, since it decouples the base

body translational motions from the payload rotation. Mounting every payload at its

center of mass becomes impractical, since the payload mounting system at least, and

perhaps the entire pointing system, must then be custom made for each payload. The

Skylab Apollo Telescope Mount experiment pointing system represents a typical center of

mass mounted payload pointing system.

Base Plate, End Mount or Cantilevered systems provide a more universal mounting

interface that is advantageous when a wide variety of payload shapes and sizes are to be

accommodated by the same pointing system. Other advantages over the center-of-mass

method include a generally larger pointing envelope, a more compact and rigid mounting

structure, and simplicity of deployment and stowage. Disadvantages include: an increase

in the gimbal torque and torquer size required to provide the same angular acceleration;

and difficulty in ground testing [521. The major disadvantage of end mounted pointing

systems (also referred to as cantilever mounts) is that translational motions of the base

body at frequencies above the control bandwidth disturb the payload pointing angle. High

frequency base body rotational motion (jitter) will also affect payload pointing when the

pointing system is offset from the base body center of mass. This is almost certainly to

be the case on Space Station. An end mount system then can easily accommodate a large

selection of different instruments, but this is subject to increased disturbances from base

body motion [48). The Dornier Instrument Pointing System (IPS) represents an end

mounted pointing system.

A2-8
t

A2-9

Another concept, which has been under study at MSFC for several years, is to provide an

instrument mounting plate, which is attached to the base body by a flexible suspension

system. Base plate stabilization is accomplished through CMGs or Reaction Wheels

mounted on the base plate. The resulting system isolates the payload from high

frequency disturbance originating from the base body and provides a high degree of

stabilization for the payload. The payload may still acquire power, communication, and

other services from the base body. The concept has applications to payloads that require

a low-acceleration environment (e.g., materials processing experiments) and for payloads

that require a high degree of stability. Slewing of payloads could then be accomplished

by relatively coarse pointing mounts.

1.4.2 LOS ARTICULATION OPTIONS

LOS Articulation Options refer to the number of stages of disturbance isolation between

the payload and the base body implemented to achieve the desired pointing and stability

characteristics. In the case of a free-flying spacecraft which is dedicated to the pointing

of its payload, there may be no articulating elements external to the payload. The

Infrared Astronomical Satellite (IRAS) is an example. High precision is achieved by

making the telescope an Integral part of the spacecraft, constraining disturbances to a

very low level and pointing the whole assembly, using a set of reaction wheels.

LOS Articulation becomes necessary to satisfy multi-payload pointing requirements and

to compensate for base body disturbances if the spacecraft base body is noisy. More than

one level of articulation allows for a veinier approach where each level of articulation

becomes more precise. A three-level system is shown in Figure A-2 . Note that the

various levels of articulation can serve distinctly different functions. The coarse gimbal

rejects low frequency disturbances, the veinier stage acts as an isolator, attenuating the

amplitude of the disturbance, and the Image Motion Compensation (IMC) stage rejects

high frequency disturbances. A multi-level articulation architecture provideb the

designer with the options as to the number of levels, the arrangement of levels, and the

component complement at each level.

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A2-10
1.4.3 SENSING OPTIONS

Sensing Options pertain to the choice of attitude sensing hardware and to the location of

the sensing hardware or source of attitude reference.

1.4.3.1 ATTITUDE SENSING HARDWARE

Hardware used for attitude determination and sensing includes Sun Sensors, Horizon; Sensors, Magnetometers, Star Sensors and Gyroscopes (Gyros)(50). Sun Sensors are the most

widely used type; one or more have flown on most spacecraft. This is due to several factors. For low earth orbit operations the sun's angular radius is sufficently small that a

point source approximation can be made and the Sun is bright enough to permit the use of

relatively simple equipment without discriminating between sources. This simplifies the

design of the sensor and the attitude determination algorithim. Most space missions are

concerned with the orientation and time variation of the Sun vector in spacecraft body

coordinates. The wide variety of applications have led to the development of numerous

sensor types with fields of view (FOV) from several square arc minutes to steradians and

resolutions of several degrees to less than 1 are sec.The basic classes of Sun C,nsors are

Sun Presence Sensors, Analog Sun Sensors and Digital Sun Sensors. Analog Sun Sensors

have a continuous output signal that is a function of the sun angle. Sun Presence sensors

provide a continuous output whenever the sun is within the FOV. Digital Sun Sensors

provide an encoded digital output which is a function of the sun angle.

Horizon Sensors provide a means for directly determining the orientation of a spacecraft

+ with respect to the earth. Earth looking payloads typically require pointing accuracies of

0.05 deg. to 1 arc min. which exceeds the capability of present horizon sensors. Earth

= oriented spacecraft control systems utilize error signals from horizon sensors with accuracy ranges from 0.5 to 1 deg. Thus, although payload requirements may not be met by

horizon sensors, control system requirements are easily satisfied. For the most part,

Horizon Sensors are designed to operate in the infrared spectral band.This is done to

avoid problems which would be introduced by visible light sources. It avoids triggering of

the sensor by visibl( light reflected off clouds, and sunlight reflected from spacecraft

surfaces. They are normally classified with respect to the scanning method used to

search the celestrial sphere. These include rotating mirrors, body mounted sensors, and

wheel moi!.nted sensors. Body mounted sensors are fixed at a known angle to the structure

of a rotating spacecraft. Wheel mounted units are fixed to a momentum wheel and the

wheel rather than the spacecraft provides the scanning motion.

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A2-12

Magnetometers sense the direction and magnitude of the earth's magnetic field. They are

lightweight, have low power requirements, operate over a wide temperature range, and

have no moving parts. They are not extremely accurate since the earth's magnetic, field

Is not completely understood and models to predict the magnetic field magnitude and

direction at the spacecraft position are subject to substantial errors. Residual spacecraft

magnetic tlases restrict the use of magnetometers to altitudes below 1000 KM due to the

earth's magnetic field decreasing with distance from earth as 1/11

Star Sensors measure known star coordinates in the spacecraft frame and provide

attitude information when the observed coordinates are compared with the known star

locations obtained from a star catalog. Star sensors are the most accurate of attitude

sensors achieving accuracies in the arc sec range.On the other hand they are heavy,

expensive, and require more power than other sensor types. More extensive computer

software is required since measurements must be processed and Identified before

attitude information Is obtained. There performance Is affected by interference from the

earth , sun and other bright sources and by occultations of the sun and earth. There are

three main types: star scanners, gimballed star trackers, and fixed head star trackers.

Star scanners use the spacecraft motion to perform the searching and sensing function.

Gimballed star trackers use mechanical action for the search and track functions. Fixed

head star trackers utilize electronic searching and scanning over a small fielc' of view.

Gyroscopes are instruments that use a rapidly spinning mass to sense and respond to

changes in the inertial orientation of its spin axis. Basic types used as attitude sensors

are Rate Gyros, Rate Integrating Gyros, and Accelerometers. Rate gyros measure

spacecraft anngular rates and are frequently part of the control loop for attitude

stabilization. The angular rates may also be integrated via an on board computer to

provide an estimate of spacecraft attitude with respect to some inertial reference. Rate

integrating gyros measure angular displacements directly.
6

A2-13

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1.4.2.2 SENSING HARDWARE LOCATION

Payload attitude determination schemes fall into thre- categories:

A) base body referenced

B) inertially referenced

• C) target referenced

In the base body referenced scheme, the attitude sensing hardware is located on the

vehicle or spacecraft base body. The payload is responsible for determining its position

fwith respect to the base body. This is normally accomplished by use of polentiometers,

resolvers or encoders. The attitude of the payload is measured with respect to the base

( body. The attitude of the base body is measured by using the attitude °ensors located on

l the base body. The method has certain limitations in that the payload attitude can never

be as good as the base body, since sensing the relative position of the payload with

( respect to the base body introduces errors. It is difficult to sense relative position with

high precision. In addition, structural deformations between the base body attitude

sensors and the Payload LOS are not sensed and can take on significant values.

An inertially referenced method of attitude sensing utilizes attitude sensors mounted on

the payload side of the gimbals. The payload attitude is measured with respect to known

star locations. The accuracy of this method is limited by the errors inherent in the

'a sensors themselves and the accuracy of the star catalog oeing used. Current star

catalogs are accurate to about 1 arc/sec and are predicted to be accurate to 0.1 are sec

during the next decade. An inertially referenced attitude determination scheme can 

satisfy the requirements of most pointing payloads, provided the attitude sensors are

rigidly attached to the instrument 's focal plane and their alignment to the instrument

LOS is known and maintained to the required accuracy.
If an inertlally referenced scheme is not sufficiently accurate, target referenced attitude

sensing can be employed. In target referenced attitude sensing, the pointing control

system uses an error signal provided by the payload's focal plane sensors to close the

pointing loop. Errors introduced by structural bending modes and misalignments are

bypassed. This method of attitude sensing does complicate interfacing the payload to the

pointing system and utilization of the pointing system to accommodate multiple

payloads.

1.4.4 CONTROL METHODOLOGY OPTIONS

Control methodology options refer to the manner in which the control loop is closed or

the control system design approach. The basic choices in this regard are classical control

or modern control techniques. Classical control techniques primarily treat single input

and single output systems and numerous frequency domain procedures have been

developed to handle such systems. When more than one input or output are present, they

are assumed to be decoupled or independent and separate single input and single output

systems are provided for each. Modern control techniques are particularly well suited

for use in complex systems where coupling effects are important, such as cases where

several degrees of freedom are affected by a single actuator or measured by a single

sensor. Modern control techniques, primarily formulated in the time domain, utilized the

computational capability of digital devices to handle non- linear control problems.

Whether a classical or modern approach is taken, the concepts of feed-ahead commands,

adaptive control, and distributed control can be used. Feed-ahead or feed-forward

commands utilized knowledge of a previously calibrated anticipated disturbance to

cancel the disturbance torque before it propagates through the system, resulting in

attitude motion, which then must be sensed and counteracted via a low bandwidth control

system. Adaptive control systems include a parameter adjustment loop which changes

control law parameters, such as: control gains, bending mode filter parameters, and

command generator commands to adapt to changes to the configuration of the base

body. Adaptive control is based on knowledge of previous input/output history, and a

prior knowledge of the system configuration (i.e., an accurate system model or system

identification) and the performance requirements. The adaptive control system

continually adjusts itself, based on the system model and what it perceives through its

sensors.

A2-14

I
Distributed control Includes systems where two or more control systems having their own

sensors and actuators operate in a coordinated manner (i.e., master/slave or hierarchical

control) or a system which utilizes distributed sensors and/or actuators which interface

with a centralized controller.

Whether a classical or modern control approach Is used, an accurate system model

(system Identification) must be available. Due to the dynamically Interactive

environment of Space Station, a precision pointing system may have to have a model not

only of itself, but of the Space Station as well. Changes to the Space Station's mass and

system properties during its evolutionary growth, and due to shuttle docking, will

necessitate changes In the control law gains. The new mass properties can be estimated

on-board, based on the properties of the indivi. 'al elements added to the station. The

controller gains may be updated via ground command, astronaut inputs or autonomously

on board.

A2-15

I^
GIST OF ABBREVIATIONS

AAS American Astronautical Society

AC Alternating Current

t ACA Attitude Control Assembly

41 ACS Attitude Control System

AEPI Atmospheric Emissions Photometric Experiment

AGC Automatic Gain Control

AGS Advanced Gimbal System

AIAA American Institute of Aeronautics and Astronautics

APCS Attitude and Pointing Control System

APM Articulated Primary Mirror

ASM Assembly

ASO Advanced Solar Observatory

fASPS Annular Suspension and Pointing System

ATM Apollo Telescope Mount

ATMDC Apollo Telescope Mount Digital Computer

AXAF Advanced X-Ray Astrophysics Facility

BEA Backup Electronics Assembly

BGSD Bendix Guidance Systems Division

CDMS Command and Data Management System

CEA Control Electronics Assembly

C.G. Center of Gravity

CGA Coarse Gimbal Assembly

CIP Coordinated Instrument Package

CMG Control Moment Gyroscope

CMGEA Control Moment Gyroscope Electronics Assembly

CMGIA Control Moment Gyroscope Inverter Assembly
e

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LIST OF ABBREVIATIONS (Cont'd)

Conf. Configuration

COR Center of Rotation

CPS Cycles per second

CRT Cathode Ray Tube

CSS Coarse Sun Sensor

C&T Communications and Tracking

DA Digital to Analog

DARPA Department of Advanced Research Projects Agency

dB Decibel

DC Direct Current

DCU Digital Control Unit

DDU Digital Display Unit

DEA Digital Electronics Assembly

Deg Degrees

DEP Dedicated Experiment Processor

DMU Data Management Unit

Doc Document

DRIRU Dry Rotor Inertial Reference Unit

ECLSS Environment Control and Life Support System

ECU Electronics Control Unit

Elec Electrical

EPC Experiment Pointing Control

EPCS Experiment Pointing Control System

EPDS Electrical Power Distribution System

EPDU electrical Power Distribution Unit

EPEA Experiment Pointing Electronics Assembly

EMC Electromagnetic Compatibility

ESA European Space Agency

EVA Extravehicular Activity

Exp Experiment
LIST OF ABBREVIATIONS (Cont'd)

FGS FinA Guidance Sensor

FHST Fix, d Head Star Tracker

FOV Field of View

FSS Fine Sun Sensor

FUSE Far Ultraviolet Spectroscopy Explorer

G Gravity

GEA Gimbal Electronics Assembly

GN&C Guidance, Navigation and Control

GPC General Purpose Computer

GPS Global Position Satellite

G RO Gamma Ray Observatory

GSFC Goddard Space Flight Center

Gyro Gyroscope

I

,

HEIE High Energy Isotope Experiment 

HGA High Gain Antenna

HM Habitability Module

F HRM High Rate Multiplexer

HST H•.,bble Space Telescope

HTM High Throughput Mission

Hz Hertz

I&C Instrumentation and Communication

ID Interface Device 

I/F Interface

F IMC Image Motion Compensation

IMCS Image Motion Compensation System

I/O Input Output

l,r
LIST OF ABBREVIATIONS (Cont'd)

IOC Initial Operational Capability

IPS Instrument Pointing System

IRAs Infrared Astronomical Satellite

ISA Inertial Sensor Assembly

IVA Intravehicular Activity

JEA Jettison Electronics Assembly

i

Kg Kilogram 

Kg-- M 2 Kilogram Meter2

R

Kg-M3 Kilogram Meted

kHz Kilohertz

Km Kilometer

Kw Kilowatt

Lb Pound

LDR Large Deployable Reflector

` LLLTV Low Level Light Television

LM Laboratory Module

LMSC Lockheed Missiles and Space Company

LOS Line of Sight

LRU Line Replaceable Unit

LVLH Local Vertical Local Horizontal

e

m Meter

MAST Modified ATM Star Tracker

4 Max Maximum

MBA Magnetic Bearing Assembly

MDA Multiplying Digital to Analog
4

LIST OF ABBREVIATIONS (Cont'd)

MD1Y, Multiplexer Demultiplexer

Mech Mechanical

MEV Million Electron Volts

MJA Mounting and Jettison Assembly

MMC Mount manual control

MMC Martin Marietta Company

MPC Manual Pointing Controller

,MPM Miniature Pointing Mount

MRMS Mobile Remote Manipulator System

MSFC Marshall Space Flight Center

MSS Magnetic Sensing System

MT Magnetic Torquer

NASA Nationei Aeronautics and Space Administration

N-M Newton Meter

N-M-S Newton Meter Second

NSSC NASA Standard Spacecraft Computer

OA Orbit Adjust

OMV Orbital Maneuvering Vehicle

OPTO-MECH Optical Mechanical

ORU Orbital Replaceable Unit

OSP Optical Sensor Package

OSSA Office of Space Science and Applications

OTA Optical Telescope Assembly

OTV Orbital Transfer Vehicle

OWS Orbital Workshop

,. ;7 ^XXY 
LIST OF ABBREVIATIONS (Cont'd)

PCA Payload Clamp Assembly

PCE Power Conditioning Electronics

PCs Pointing Control System

' PDU Power Distribution Unit

PEA Payload Electronics Assembly

P/L Payload

PMS Payload Mounting Structure

POF Pinhole Occulter Facility

POP Perpendicular to the Orbit Plane

RAU Remote Acquisition Unit

REF Reference

RG Rate Gyro

RGA Rate Gyro Assembly

RIU Remote Interface Unit

RMS Root Mean Square

RPM Roll Positioning Mechanism

RSU Rate Sensor Unit

RWA Reaction Wheel Assembly

SA Solar Array

SAAX Science and Applications Mission Data Base

SCDM Solar Corona Diagnostic Mission

SCRN Spectra of Cosmic Ray Nuclei

SEC Second

SEM Suspended. Experiment Mount
LIST OF ABBREVIATIONS (Cont'd)

SEUVTF Solar Extreme Ultraviolet Telescope Facility

SI Science Instrument(s)

SIPS Small Instrument Pointing System

SIRTF Shuttle Infrared Telescope Faciiity

SOT Solar Optical Telescope

SS Space Station

SSM Support System Module

SSP Space Station Program

SSXRTF Solar Soft :-Ray Telescope Facility

ST Space Telescope

STRUC Structure

STS Space Transportation System

SUPERMAG Super Conducting Magnet

TACS Thruster Attitude Control System

TDRS Tracking and Data Relay Satellite

TDRSS Tracking and Data Relay Satellite System

TEA Torque Equilibrium Attitude

TFSUSS Task Force Study on Utilization of Space Station

TGG Talon Gold Gimbal

TRIC Transition Radiation and Ion Calorimeter

TV Television

Typ. Typical

UV Ultraviolet

V Volts

VAC Volts, Alternating Current

VCS Vernier Control System

VDC Volts, Direct Current

VIPs Vibration Isolation Pointing System

VLBI Very Long Baseli ne Interferometer

VPA Vernier Pointing Assembly
Y

LIST OF ABBREVIATIONS (Cont'd)

WJU Wiring Junction Unit

XRTE X-Ray Timing Explorer

ay..`^
u

REFERENCES

1. Space Station Reference Configuration Description, Space Station Program Office

Document - JSC-19989, August, 1984

t a 2. Space Station Definition and Preliminary Design, Request for Proposal, National

Aeronautics and Space Administration, September 15, 1984

3. NASA Langley Space Station Data Base, September 1984

4. NASA Langley Space Station Data Base, May 1985

5. Science and Applications Space Platform Payload Accommodations Study, Space

s Platform Payload Data, Teledyne Brown Engineering, SP82-MSFC-2583, March

1982

6. Pinhole Occulter Phase A Study Report, NASA, Marshall Space Flight Center,

January 1984

7. Report on the Solar High Resolution Telescope Cluster, Notes - W. Roberts, MSFC

8. The Solar Optical Telescope, Brochure, NASA, GSFC

9. Space Station Attached Payload Accommodation Requirement Definition for the 

Solar Optical Telescope, Final Report 

10. Skylab Program Payload Integration Technical Report, Attitude and Pointing

Control System (APCS) Definition Document, Volume 1, Revision F, Martin

Marietta Corp., ED-2002-984, March 1973
11. Skylab Program Payload Integration Technical Report, Attitude and Pointing

Control System (APCS) Definition Document, Volume II, APCS Hardware

Complement, Rev. A, Martin Marietta Corp., ED-2002 - 984 Rev. A, March 23, 1973

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Technical Memorandum NASA TMX-64817, July 1974

13. Space Shuttle Experiment Pointing Mount (EPM) Systems • An Evaluation of

concepts and Technologies, Volumes I and 11, JPL 701-1, April, 1977

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15. Keckler, Claude R. "High Accuracy Pointing for Earth Observation Experiments"

AIAA 84-1880

16. Anderson, Willard W. and Woolley, Charles T., "The Annular Suspension and

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Systems, May 11-12, 1977

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Addendum), NASA Contractor Report 3343, 1980

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Flight Systems, MSFC-SPEC-1505

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Contract NAS8-33413, DR 8-3, Item 10

21. Final Report AGS/IPS Pointing Systems Development and Performance

Comparison Study, Volumes I and II, Teledyne Brown Engineering, SP82-MSFC2607, August 1982

x

^^ 
22. Earth Observation Mission One, Instrument Interface Agreement for Atmospheric

Emissions Photometric Imaging, George C. Marshall Space Flight Center, JA393C, September 1984

23. Atmospheric Emission Photometric Imaging (AEPI-1NS003), Test and Checkout

Requirement and Specifications, George C. Marshall Space Flight Center, MSFCRQMT-613, October 1979

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Issue 4A, June 1, 1984

25. IPS Interface Specification/Verification Document, Dornier Systems IF-IS-0001,

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26. ESA/NASA Spacelab IPS Sequences Handbook, PL-ER-0052, Vol. 2, AAP-A-IPS,

Issue 2, Oct. 1984

27. J. Me Million, Instrument Pointing System, MSFC Presentation July 1982 

8. Graf, Jochen G., "Manned Operations of a Large Pointing System"

29. Spacelab IPS Systems Manual, NASA, Lyndon B. Johnson Space Center, Systems

Training Branch, SIPS-SYS-2102, December 1984

30. Osborne, N. A., "The Gimbalflex System" AAS 78-116

31. Osborne, N. A., "Gimbalflex 5-Degree-of-Freedom Inertially Stabilized Platform"

AAS 80-015

32. Osborne, N. A., "Gimbaled Fine Pointing in Noisy Environments" AAS 80-024

33. Presentation to MSFC, "The Gimbalflex Concept," Martin Marietta Corp., circa 1978

34. The Floated Pallet Definition Study, Evaluation of Alternate Telescope Pointing

Schemes, Volume 1, The Bendix Corporation, Guidance Systems Division, Final

Report NAS8-30889, May 1975
35. The Floated Pallet Definition Study, Fin r i Report, Volume II, Suspended Pallet

Pointing Performance Study, The Bendix Corporation, Guidance Systems Division,

NASA Contract NAS8-30889, April 1975

36. Nicalse, P. D., "The Suspended Experiment Mount (SEM)," MSFC Presentation

Material, October 1983

37. "The Suspended Experiment Mount," MSFC Presentation Material, May 10, 1984

38. Bailey, Wayne L; DeSanctis, Carmine E.; Nicaise, Placide D.; Schultz, David N.,

"Payload Isolation and Stabilization by a Suspended Experiment Mount (SEM)"

39. Jacquemin, G. G.; Rusk, S.J., "Drag Compensated Precision-Powered Hinge

System" 19th Aerospace Mechanisms Symposium, NASA Conference Publication

2371

40. Space Telescope Systems Description Handbook, ST/SE-02, Lockheed Missiles and

Space Company, Inc., LMSC/D 974197A November 30, 1984

41. Dougherty, H.; Rodoni, C.; Rodden, J.; Tompetrini, K., "Space Telescope Pointing

Control" AAS 83-365

42. Carpenter, A.; Hopkins, M., "Space Station Attitude Options," MSFC Memorandum

ED-15. 85-47, ,august 1, 1985

43. Space Station GN&C Technical Review Meeting, Presentation Material, Volume 1

& 11, McDonnell Douglas Corp. July 16, 17, 1985

44. Woo, H.H., "Attitude Control and Momentum Management (Project 26)," Space

Station Advanced Development (A/D) Meeting, Presentation Material, May 14,

1985

45. Sirlin, S. W., "Space Station Dynamic Disturbance Environment," JPL IOM 343-84-

1082, October 29, 1984
46. SOT Accommodation on the Space Station, Pointing anC Vibration Control Task,

Final Report GSFC, March 1985

47. Shelton, Harvey L., "The MSFC IMCS as a Portable Pointing System," Working

Papers, March 11, 1985

48. Sirlin, S. W.= Laskin, R. A., "Payload Isolation and Precision for the 1990's,"

Annual Rocky Mountain Guidance and Control Conference, AAS 85-011, Feb. 2-6,

1985

49. Space;ab Payload Accommodation Handbook NASA/ESA, SLP/2104 Issue 1.,

Revision 7, May 6, 1983

50. Wertz, James R. (Editor), "Spacecraft Attitude Determination and Control," D.

Reidel Publishing Co., 1978

51. Mettler, E. and Sirlin, S.W., "Space Station/ Space Platform Science Pointing

Requirements and State of the Art Pointing Technology," JPL Memorandum 343-

84-1235, December 3, 1984

52. Cunningham, David C., and Wilson, George W., "Zero Gravity Simulator for

Dynamic Testing of Pointing Mounts," AIAA 83-2300