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Produced by the NASA Center for Aerospace Information (CASI)
SSO 83-0094-2

DEVELOPMENT OF DEPLOYABLE STRUCTURES

FOR

LARGE SPACE PLATFORM SYSTEMS

Design DeVblopment Volume 11

NASA/ MSFC CONTRACT NAS8-346# 7

OCTOBER 1983

684- 1011b (rASA-cu-170914) 'UEVILUNt+'' 11 OF DkFLOYNEL1

57bUCTUhZ5 FUb L & bGL SAAC ELAZFUUMS.

VULUAL 1: DUSIGN DLYE1 :. el thl t lnal Report

(ilcx:kw .&1 Later oationsd COL F.) I dj p

UL A

Uac:la E

Shuttle Integration d ®^ Rockwell

SAWIlle Systems D.Vlslon International
6i

SSO 83-0094.2

Development of Deployable

Structures

for

Large space platforms

Design Development

Volume U

Contract NAS 834677

October 1983

t Subm/rted by Approved

H.S. GREENBERG E.E. ENGLER ^r

i .Snrrlr 1lunugt•r .SYUr/1 Col, VA.S^I ^.l/S/^C

Prepared for:

National Aeronautics and Space Administration

George C. Marshall Space Flight Center

Marshall Space Flight Center

Alabama 35812

A^® Rockwell International

Shuttle Integration 8 

Satellite Systems Division

Rockweii International Corpe.rauon

2214 I. jxew..0l Boulevarn 

Downey C^'dorno?0241 
rRFCEDiNG PAGE BLANK NOT FILMED

iii

I

I

S

FOREWORD

Thi.e volume describes the major achievements of Parts 1 and 2 of the study

titled "Development of Deployable Structures for Large Space Platform

Systems." An executive summary of the Part 1 and 2 study is presented in

Volume I, ESD 83-0094-1. An appendix containing the developed design drawings

and supporting stress analysis is contained in Volume III, SSD 83-0094-3.

This study was managed by Marshall Space Flight Center (MSFC) and was

performed by the Shuttle Integration and Satellite SystemF Division personriel

of Rockwell International Corporation located at Downey and Seal Beach,

California. The study COB was Mr. Erich E. Engler. The study manager was Mr.

H. Stanley Greenberg.

The Part 1 stuay was initiated on October 16, 1981 and was completed nine

months later on July 16, 1982. The Part 2 study was started August 6, 1982

and was completed fourteen months later on October 7, 1983.

The major contributors to this study are listed below:

• Design - R. Bart (Lead, Part 1)

R. Barbour (Lead, Part 2)

B. Mahr

A. Perry

J. Keech

W. Wiley

C. Lang

G. Buhler

P. Buc k

T. Clegg

• Stress Analysis - G. Lesieutre

W. Batemnn

• Thermal Analysis - T. Tysor

• Materials Analysis - R. Long

C. Brownfield 

• Mass Properties - C. Griesinger

W. Morgan

• Electrical Power/Date Management - A. Gordon

• Electrical Utilities Integration - A. LeFever 

• Guidance and Control - R. Olgevie

• Technology Development - A. M. Pope
I

,,

t , CONTENTS

Section Page

INTRODUCTION 1

1 DEPLOYABLE PLATFORM SYSTEMS . 1-1

1.1 DESIGN EVOLUT70N 1-1

1.1.1 Part 1 Selected Design . 1-1

1.1.2 NASA/MSFC Requirements . 1-1

1.1.3 Major Goals and Design Philosophy 1-4

1.2 SUMMAR^ OF TEST ARTICLE DESIGN 1-5

1.3 TEST ARTICLE DESIGN 1-15

1.3.1 Major Subassemblies Design Description . 1-15

1.3.2 Ground Test Article Deviations from Prototype 1-35

1.3.3 Assembly Description 1-37

1.3.4 Deployment/Retraction Operations 1-44

1.4 TEST ARTICLE MASS PROPERTIES . . 1-53

1.5 STRUCTURAL ANALYSIS . 1-53

1.5.1 Requirements 1-53

1.5.2 Concept Reviews 1-55

2 DEPLOYABLE VOLUMES . 2-1

2.1 OTV HANGAR DEVELOPMENT . 2-2

2.1.1 Concept 1 - Double-Folded Hard-Shell

(Drawing 42712-1) . 2-5

2.1.2 Concept 2 - Double-Folded Curtain Shell

(Drawing 42712-2) 2-5

2.1.3 Concept 3 - Single-Fold Hard-Shell Design

(Drawing 42712-3) 2-9

2.1.4 Concept 4 - Erectable Hard-Shell

(Drawing 42712-4) 2-9

2.1.5 Concept 5 - Inflatable Deoign 2-12

2.2 OTV HANGAR CONCEPT ANALYSIS 2-15

2.2.1 Micrometeoroid/Space Debris Shielding 2-16

2.2.2 Stess Analysis Review 2-18

2.3 OTV CONCEPT SELECTION . 2-18

2.4 MANNED MODULE DEVELOPMFNT . 2-23

2.4.1 Baseline Space Station Conventional Modules

(Drawing 42712-6) 2-23

2.4.2 Deployable Hard Shell Manned Module

(Drawing 42712-7) 2-23

2.4.3 Inflatable Manned Module Design

(Drawing 42712-8) 2-30

2.5 DEPLOYABLE MODULE MISCELLANEOUS COMMENTS 2-36

2.6 MANNED MODULE SELECTION 2-37

3 REFERENCES . 3-1

PRECEDING PAGE BLANK NOT FILNCD

v
ILLUSTRATIONS

c_.

Z

Figure Page

1 Focus Mission Platform Systems Configurations 2

2 Potential Construction Application of Building Block 2

1.1-1 Selected Design 1-2

1.1-2 Deployment Problem with Guide Rails 1-2

1.1-3 SASP Deployable Ground Test Structure 1-3

1.2-1 Test Article Design Configuration 1-6

1.2-2 Deployable Truss and Utilities Trays Concept 1-6

1.2-3 Deployment /Retraction Mechanism Concept 1-8

1.2-4 Deployment /Retraction Mechanism Details 1-8

1.2-5 Positioning System Motion Profile 1-10

1.2-6 Positioning System Components 1-10

1.2-7 Deployment /Retraction Major Phases 1-12

1.2-8 Precompression System to Eliminate Joint Slop 1-12

1.2-9 Main Housing Structure 1-13

1.3-1 Telescoping Diagonal Joint 1-16

1.3-2 Telescoping Diagonal Center Joint 1-16

1.3-3 Folding Longeron Assembly 7-18

1.3-4 Longeron Center Joint Details 1-18

1.3-5 Longeron Center Joint Additional Details 1-19

1.3-6 Problem with Over-Center Latch Design 1-19

1.3-7 Typical Batten Assembly 1-22

1.3-8 Payload Carrier Structure 1-22 

1.3-9 Batten Deployment/Retraction System Mechanism 7-24

1.3-10 Carriage Locking Detail and Half-Nut to Rtii Matching

Features 1-24

1.3-11 Longeron/Diagonal 'Unlocking Mechanisms 1-27

1.3-12 Jackscrew Support Frame 1-27

1.3-13 27 Nm DC Servomotor Dimensions 1-30

1.3-14 3.7 Nm DC Servometer Dimensions 1-30

1.3-15 Pretension System to Eliminate Joint Slop 1-32

1.3-16 Tension Cable Routing Through Longerons 1-32

1.3-17 Adapter Assembly 1-33

1.3-18 Model Survey Potential Test Set-up 1-33

1.3-19 Main Housing Structure 1-34

1.3-20 Test Article Bay and Batten Identification 1-38

1.3-21 Test Article Assembly Configuration 1-38

1.3-22 Test Article Assembly Configuratioa with Headroom

Restructions 1-43

1.3-23 Precompression System Installation Configuration 1143

1.3-24 Ceplo,vment/Retraction Operations Configuration 1-45

1.3-25 Test i.rticle Packaged Configuration 1-45

1.3-26 Bolt/^iut Analogy at Batten Deployment/Retraction System 1146

1.3-27 Configuration at Deployment of Bay 1 1-46

1.3-28 Configuration at Deployment of Bays 1 and 2 1-49

1.3-29 Configuration at Start of Bay 10 Retraction 1149

1.3-30 Configuration at Start of Retraction of Bay 1 1-50

1.3-31 Positioning System Motion Profile 1-50

1.3-32 Batten 2 and 3 Positions at End of Bay 2 Deployment 1-51

vii PRECEDING PAGE BLANK NOT FILMED

.ry. •

^.. .-

^.^^.. r .e..,uauefYWrtnRx.... . . i: _'..a _. _

_ _ 15..

_^ -

.. .._,.-.. t
i

n' I

Figure Page

1.3-33 Lo. ation of Battens 1 through 7 1-51

1.3-34 Description of Unlocking Systems Required Advancement 1-52

1.4-1 Retraction System Equipment Deletion for Design Without

Automatic Retraction 1-54

1.5-1 Major Structural Requirements 1-56

1.5-2 Candidate Truss Arrangements 1-56

1.5-3 Batten Diagonal Load Paths 1-59

1.5-4 Potential Launch Configuration Features 1-59

1.5-5 Load Paths During Deployment /Retraction 1-61

1.5-6 Adjacent Half-Nut to Jackscrew Clearances 1-61

1.5-7 Test Article NASTRAN Model 1-66

1.5-8 Telescoping Diagonal Spring Activated Plunger 1-68

1.5-9 Telescoping Diagonal Column Loading Considerations 1-68

1.5-10 Telescoping Diagonal Adhesive Joint Test Results 1-69

1.5-11 Longeron Center Joint Column Loading Considerations 1-71

1.5-12 Folded Longeron Column Analysis Considerations 1-71

1.5-13 Diagonal Clevic Deformation Implications 1-72

2.0-1 Potential Space Station Configuration 2-3

2.0-2 Part 1 OTV Hangar Concept 2-3

2.0-3 Part 1 Manned Module Connepts 2-4

2.1-1 Baseline OTV Characteristics 2-4

2.1-2 OTVH Concept 1-Hard Shell Double Fold ( Fully Deployed

Configuration) I 1 2-6

2.1-3 OTVH Concept 1-Hard Shell Double Fold (Stowed and First

Phase of Deployment) 2-6

2.1-4 OTVH Concept 1 -Details 2-7

2.1-5 OTVH Concept 2 -Curtain Shell Double Fold Configuration 2-7

2.1-6 OTVH Concept 2--Curtain Shell Double Fold ( Stowed

Configuration) 2-8

2.1-7 OTVH Concept 2-Curtain Shell Details 2-8

2.1-8 OTVH Concept 3--Hard Shell Single Fold (Fully Deployed

Configuration) 1 2-10

2.1-9 OTVH Concept 3 -Hard Shell Single Fold ( Stowed

Configuration) 2-10

2.1-10 OTVH Concept 4 -Erectable Hard Shell Configuration

Construction 2-11

2.1-11 OTVH Concept 4 -Erectable Hard Shell Configuration

Construction 2-11

2.1-12 OTVH Concept 4 -Erectable Hard Shell (Stowed Configuration). 2-13

2.1-13 OTVH Concept 5 -Inflatable Concept Stages at Deployment 2-13

2.1-14 OTVH Concept 5-Inflatable Concept Stages at Deployment 2-14

2.1-15 OTVH Concept 5 -Inflatable Concept Stowed 2-14

2.2-1 OTV Hangar Man-Made Debris and Micrometeoroid Design

Implications 2-17

2.2-2 Micrometeoroid /Debris Model • Uncertainties for OTVH 2-17

2.4-1 Evolution of Initial to Growth Space Station Configuration 2-24

2.4-2 Potential Space Station Evolution Scenario with Deployable

Manned Modules 0

2-25

2.4-3 Manned Modules Baseline Design Summary 2-25

2.4-4 Manned Module Conventional Design Equipment Arrangement 2-26

2.4-5 Deployable Hard Shell Manned Module Concept 2-27

J

w.

viii

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ix

^S.^J._'.a.

I

C^

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Figure

2.4-6

2.4-7

2.4-8

2.4-9

2.4-10

2.4-11

Deployable Hard Shell Manned Module Concept Details

Manned Module —Deployable Hard Shell Equipment Arrangement

Deployable Module—Hard Shell Alternate Seal Concepts

Manned Module Inflatable Design Concepts

Manned Module Inflatable Design Equipment and Living Space

Arrangement

Manned Module Man-Made. Debris and Micrometeoroid Design

Implications .

Page

2-27

2-29 

2-29

2-31

2-31

2-35

ti,.
1.4-1

1.5-1

2.3-1

2.3-2

2.3-3

2.3-4

2.3-5

2.3-6

2.3-7

2.4-1

2.4-2

I

TABLES

Table Page

1-54

1-66

2-19

2-19

2-20

2-20

2-21

2-21

2-22

2-24

2-26

Test Article Predicted Mass (kg)

Predicted Loads During Modal Survey

OTV Hangar Selection Process Criteria

Relative Cocts ($ Millions)

Design Versatility .

Orbiter Integration

Reliability of Structural Rigidization .

Design, EVA, Technology Development, and Life Risks

Normalized S-,mmary of the Major Criteria

Manned Module Major Requirements

Baseline Conventional Manned Module Mass (kg)

w. 

VG PAGE BLANK NOT FILMED
INTRODUCTION

This volume describes the major achievements of the Parts 1 and 2 study

activities related to deployable structures for large space platforms. These

activities included development of a building-block design for the automatic

construction of deployable platforms such as those shown in Figure 1, and the

development of deployable volumes for manned modules and OTV hangars identified for potential Space Station configurations.

Since the recent accomplishments of Part 2 are of greater interest to the

large space structures community the Part 2 study accomplishments are presented first.

The preponderance of the study effort was devoted toward the deployable

platform systems study which has culminated (Part 2) in a "fabrication ready"

detailed design of a ground test article for future development testing. This

design is representative of a prototype square-truss, single-fold buildingblock design that can construct deployable platform structures in the manner

suggested in Figure 2.

This prototype design was selected (Part 1) through a comprehensive and

traceable selection process applied to eight competitive designs. The

selection process compared the competitive designs according to seven major

selection criteria, i.e., design versatility, cost, thermal stability,

meteoroid impact susceptibility, reliability, performance predictability, and

orbiter integration suitability.

In support of the foregoing, a materials data base, and the development

needs for platform systems technology were established (Part 1).

In the deployable volumes study an e-ectable design of an OTV hangar was

selected (Part 2) and recommended for further design development. This design

was selected from five study-developed competitive single and double-fold

designs including hard-shell and inflatable designs. Also, two deployable

manned module configurations, i.e., a hard-shell and an inflatable design were

each developed (Part 2) to the same requirements as the composite of two Space

Station baseline manned modules. For each of these deployable module designs,

atmospheric sealing suitability was of sufficient concern to offset the

potential launch coat savings. Hence, no further activity was recommended

pertinent to deployable manned modules.

i
WS TEST ART[: .1

ORIGINAL FAGC !:

OF DOOR QUALrTY

i

Figure 1.

NS

Focus Mission Platform Systems Configurations

f

r-_ MAIN MOUSING

UIL01016 (LOCK

ENO

ADAPTER

` OSTOIwFO

CONFIGURATION

PACKAGE

PACKAGE

RENOIIEO FOR

KF? OYIIEI(T

SECOND STAGE

OF DEPLOYMENT BUILDING BLOCK

REORIENTATION FOR

DEPLOYMENT

FIRST STAGE Of

DEPLOYMENT

Figure 2 Potential Construction Application of Building Block

2
1. DEPLOYABLE PLATFORM SYSTEMS

i.l DESIGN EVOLUTION

1.1.1 Part 1 Selected Design 

The Part 2 study activities were airected toward the establishment of a

detailed design of a ground test article that is representative of the Part 1

selected design (Figure 1.1-1).

At this stage of development the deployment mechsnism was not defined but

visualized to be a reciprocating design requiring guide rails to maintain root

strength during deployment. The potential problems of such a design were

recognized during the Part 1 study efforts toward conetrueting and deploying

the generic platforms with each of the competitive building block concepts.

The potential problem associated with the use of guide rails arises during

exteusion of a truss which has a payload or module sufficiently wide that the

guide rails cannot straddle the module (Figure 1.1-2) and, therefore, cannot

be unfolded until the truss has extended and moved the large payload out of

the way. Obviously, if the guide rails are not in position when the truss/

payload is moving, there is no root strength developed. While several

solutions to minimize this problem were established during the Part 1 study,

the mechanism design estab ".stied in Part 2 was developed to completely avoid

this problem.

1.1.2 NASA/MSFC Requirements

The design development was also directed to satisfy the NASA/MSFC

requirements derived from the McDonnell Douglas SAS? study (Reference 1).

The test article configuration was to proximate the SASP deployable structure

configuration shown in Figure 1.1-3. Further, the following additional

requirements were stipulated by NASA/MSFC.

a Provide automatic retraction in addition to deployment

o Use 1.4 m x 1.4 m truss cross section

• Provide mounting for two 3636 kg simulated payload carriers at each of

2 stations each having an attachment bolt pattern of 1.09 m x 1.31 m

(43 x 51.5 in.)

• Provide a design with a minimum natural frequency of 0.10 Hz

o Provide a design that can sustain a 0.04 g load factor applied at the

c.g. of one payload at each station.

o Provide trays in two bays to accommodate four No. 1/0 cables, two No. S

AWG, four No. 12 AWG, twenty No. 22 twisted pair shielded, and four

each coax lines.
URIGINAL PIKE Iw

OF POOR QUAL."

LONOII

UTILITY

ployment Problem with Guide Rails

1-2

POSITION

.S

NED

ING

• ALL MIMUfS AR !,OJND TUNIS

• LONOIRONS WITH CENTIR HIND! JOINTS

*TELESCOPING DUGONALS

•NO OFFSIT LOADS AT JOINT

•ILATTINS CROSS SRACID RY TENSION STM

TRUSS ^

ADAMR

e

°^ 9°

0

e

MAIN HOUSING

^DEFLOYMWT RAILS

UTILRRS INSTALLATION

END VOW FOLDED DITAK OF JOINT

Figure 1.1-1. Selected Design
TEST ARTICLE FUTURES

cawAmn

ORIGINAL !=A(; ,

 rO

OF POOR QUAL17y

PAYLOAD CARRIER

.^ INTERFACES

R

K

DEPLOYMENT CO TIM

PLATFON MECIIAMISN

SUPPORT

Pam

SECONDARY PAYLOAD

CARRIER INTERFACES

Figure 1.1-J. SASP Deployable Ground Test Structure

P. )
t-4

k

N Q 

F 

^ The requirement to mount the 3636 kg (8000 lb) simulated payload carriers

at two sts'.ions was satisfi. nd by using the housing as one of the bases with

the deployable truss providing the seco-d bass. this solution is cheaper than

plac ' ng both bases on the deployable truss and represents the reality of

potential platform systems, wherein payloads or major subsystems will be

mounted onto the housing. With the payload carriers placed am described, the

test article was than designed to be supported, during ground testing, by an

and adopter at thr, end opposite the main housing.

1.1.3 Major Coals and Design Philosophy

The major goals and design philosophy to be used in the development of the

test article were agreed upon between. Nj.SA /MSFC and Rockwell. The major goal

and design philosophy are as follows:

Major Coal: To design the structure and mechanism to suppo I t demonstration of the suitability of the prototype building-block design to

automatically deploy and retract, to have predictable dynamic behavior, and

lc,^tly ti sustain specified limit loads.

Design Philosophy: Minimum test article cost with minimal to no

.v

compromise of the major goals delineated above.

} Major Design Implicat ions:

o Members of the t.usa and joints are aluminum and steel ( composite

designs to be developed later).

r a Test article will be a representative design, but not an optimum design.

o Hear minimum weight approach to structural design only when sensitive

to structural performance otnerwise cost governs.

o Design for NASA /MSFC test facility environment i.e., no provision for

in-space environmental requirements on lubricants, materials, or

positioning equipment.

o Positioning system components are not designed for orbiter launch

survival

o Trays to be provided for future utilities integration and to

accommodate electrical lines with standard insulation and fluid bellows

(uninsulated).

The foregoing directed the test article configuration and detailed design

to that described herein tend by the top assembly and detailed design drawings

contained in the appendix of Volume III.
4

Y ^

y

{

1.2 SUMMARY OF TEST ARTICLE DESIGN

A aummary of the test article design is provided herein. The test article

design is described in further detail in Section 1.3 and by the set of drawings and supportive structural analysis in the app,ndix of Volume III. Figure

1.2-1 illustrates the test article configuration. The major components of

this test article are as follows:

1. The square truss (Figure 1.2-2) containing folding utilities trays-in

Bays 4 and 5 with provisions for future installation of power, data,

and fluid lines.

p w.

2. A mechanization system (Figure 1.2-3) consisting of (a) a batten

deployment/retraction jackscrew system which translates the battens

one at a time, (b) a diagonal latch unlocking system, and (c) a

longeron latch unlocking system.

3. A jackscrew support frame assembly that supports the cantilevered ends

of the batten deployment/retraction jackscrews (Figure 1.2-3)•

4. A positioning system to precisely control the bay-by-bay deployment

and retraction operations (Figure 1.2-6).

5. A precompression system to eliminate structure joint backlash (Figure

1.2-8).

6. An end adapter at the and of the truss with provisions for attachment

to a NASA/MSFC test fixture (Figure 1.2-1). 

7. An aluminum akin and frame main housing (Figure 1.2-9). The housing

and payload carrier frames shown contain inserts for attachment of the

NASA/MSFC simulated payload carriers.

A brief summary of the overall design is presented to enhance an understanding of the individual components detailed designs.

Figure 1.2-2 illustrates the deployable Cruse major design features. The

deployable trues contains square battens stabilized by compression diagonal 

traces. Each batten contains a half-nut at each of the four corners. Through

engagement with each of the four batten deployment/retractior. jackscrews,

counterclockwise rotation of the jackscrews imparts outward linear motion to

the batten (deployment), while the opposite rotation imparts inward motion to

the batten (retraction). With the exception of the first bay, deployment or 

retraction is respectively accomplished by holding the aft batten with detente

while deploying or retracting the forward batten. During deployment, each of 

the four longerons is unfolded and each of the four telescoping diagonals .e 

extended. The longerons and diagonals each have spring-activated locking pine

in latches at their center joints that, upon unlocking, provide axial and

moment structural continuity. Both designs have end rod fittings with

spherical bearings and turnbuckles for precise member length adjustment. The 

afo--9entioned spring-activated pine in the center joints must be unlocked to

permit retraction. This is accomplished with each of the diagonal and

longeron unlockin g eystems (Figure 1.2-4) that contain tripping devices that

rotate exterior caromed surfaces on the latch mechanisms to depress the locking

pine.

I-5

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1 ^

>'

ORIGMAi. PACE 1;

OF POuR QUALf1Y

DEPLOYABLE TRUSS

CONTAINING

UTILITIES TRAYS END

BATTEN NO 9 z^EDAR

IN BAYS 4 A 5 ADAPTER

( OF POSITIONING

SYSTEM MOTORS

PAYLOAD CARRIER

SUPPORT FRAMES

(JACKSCREW

s CTIOM

(TYF)

\

1.09 • 1.4^ PPORT

IAYLOAD CAMIfR

SUPPOAD FRAMES JACKSCREW

1.09 sl

I

SUPPORT OIAOOMAL

BAY 2

2

1.T4 m

r BAY 5 6111_.._

DEPLOYED 

BATTEN 14 •

OEPLOYMENT/RETRACTION

JACKSCREW

^I

min HOUSING STONED

(CONTAINS LONGERON. A

DIAGONAL JOINT UNLOCKING

MECHANISMS. A PRECOMPRESSION SYSTEM)

Figure 1.2-1. Test Article Design Configuration

TELESCOPING LATCH

MECHANISM

INNER TIDE

OUTER TUBE 44.5 ^m

?E

OUTER X 2.1 s

50.8 = O ' OL X 2.1 s

URIbUCKLE DIAGONAL ASSEM"LY

ROD END (2)

TELESCOPING

DIAGONAL — >

FOLDING LATCH I I.

50.8 m. O.D. X 2.1 me MECHANISM L^

INAL

/B/I

Y

I RIACE/

uTr

R00 ENO (2)TURNBUCKLE (2) 

LONGERON ASSEMBLY 44, (STOWED

BATTEN ASSEMBLY

Figure 1.2-2. Deployable Truss and Utilities Trays Concept

1-6

y

-

SAY 1
Y.

!. The truss design also contains trays for Bays 4 and 5 onto which a

generous complement of electrical power, data, and fluid lines can be

mounted. Specifically, space in available for six No. 1/0 cables, three No. 8

AWG, six No. 12 AWG, six coax, twenty-eight No. 22 twisted pairs shielded, and

four 12.5 mm flexible coolant tubes. The trays are hinged from the batten

frames and fold as shown at the lower right (Figure 1.2-2). During the

Shuttle launch of a prototype design the trays would provide lateral support

to the folded longerons. Longitudinal support of the lorgerons is provided by

tight packaging and appropriate and transverse beams in the adapter and main

hoiising.

Figure 1.2-3 illustrates (in the deployed configuration) the major

features and orientation of the test article mechanization system. For

clarity, the ten-bay truss structure is not shown. This system provides fully

controlled bay-by-bay deployment/retraction capability with maintenance of

root strength throughout all phases of deployment. The mechanisms include the

batten deployment/retraction jackscrew system, the longeron unlocking system,

and the diagonal unlocking system. The batten deployment/retraction system

(Figure 1.2-4) consists of four asemblies of guide rail, eplined shaft, and

jackscrew mounted in a slide carriage. These assemblies are located at each

of the four corners of the main housing. In the first stage of deployment,

i.e., Bay 1, clockwise rotation of each of the spline shafts advances the

slide carriage and jackscrews out of the housing into the configuration

shown. Concurrently, the jackscrew support frame assemSly is advanced to the

configuration shown with the automatic locking of the telescoping diagonals.

A controller-driven single motor, slaved to a chain and sprocket system,

drives all four assemblies. The longeron unlocking system (Figure 1.274)

consists of four assemblies of guide rail, jackscrew, carriage, and tripping

device. These systems are located adjacent to the individual batten

deployment/retraction assemblies. Each of the four diagonal unlocking

assemblies are the same as that of each longeron unlocking system, except for

the tripping devices, and are each located at the center of the housing

sidewalls. The longeron and diagonal unlocking assemblies are each

controller-driven by a single motor slaved to a chain and sprocket system to

drive all four assemolies.

Figure 1.2-4 further describes the deployment/retraction mechanism. The

batten deployment/retraction jackscrew shown illustrates one of the four

jackscrew assemblies. The jackscrew, carriage, and spline assemblies are

cradled within a rigid rail. A splined bushing at the aft end of the

50.8--mm-diameter jackscrew- encircles a eplined shaft that extends nearly the

entire length of the jackscrews. The jackscrew splines extend beyond the aft

end of the rails where a chain and sprocket are attached.

Encircling the rotating jackscrew is a carriage fitting that has external

ea-s that engage matching grooves running the length of the rails. The

carriage is pulled forward with the jackscrew, during deployment of Bay 1

(Figure 1.2-3), until a hole in the side of the carriage is engaged by a

spring-operated pin mounted near the forward and of each rail, thereby locking

the carriage. During retraction of this final bay the pin is manually

retracted from the carriage, thus allowing the jackscrew to be retracted into

the housing.

One of the longeron and one of the diagonal

shown (Figure 1.2-4) in the partially deployed

configuration the carriages are entirely withi n

unlocking assemblies is each

configuration. In the stowed

 the main housing.

1-7
'I

r,;I ORIGINAL PAGE 18

OF POOR QUALITY

I.TS DIAGONAL UNLOCKING

SYSTEM (4)

OSNG

HOUSING FWIIN

BATTEN DEPLOYMENT/

aETMCTIOM SYSTEM (4)

1.I6 n I I

^

TELESCOPING DIAGONAL

^. (AUTOMATIC LOCKING -

MAIN NOISING MANUAL UNLOCKING)

IJ•

ACKSCREW SUPPORT

LOMGERON FRAME

UNLOCKING

SYSTEM (4)

Figure 1.2-3. Deployment/Retraction Mechanism Concept

MOTOR. TACHOMETER, ENCODER, GEAR

0011 ASSY

^' SPROCKET

L^

SPLINE SHAFT '" ^16p \^ < N ' NALF HUT AT

J CARRIAGE ALL 4 BATTEN

i I U/AUT014AT IC CORNERS

LOCK

1^ L 50.8 m. 0A

JACKSCREW

i 1 RAIL

SUPPORTED -'

- BY HOUSING

t)

r.,

IKITOR, ^%^" [`

JACKSCREW \ ^TpCHIMIETER,

ENCODER ASSY _BATTEN DEPLOYMENT/ SUPPORT FRAIIE

RETRACTION SYSTEM

CHAIN &

ENA

^

SPROCKET \ r^

(e 1 \ ^,^

Y\

25.4 0 ADEPLOYABLE/RECTRACTABLE 

}= v

m

DEPLOYABLE/ - JACKSCREW CARRIAGE

25.4 wpIA / RETRACTABLE

JACKSCREW

CARRIAGE

,`

LEVER

`_m

LOMGERON UNLOCKING SYSTEM DIAGONAL UNLOCKING SYSTEM

TRIPPING PROBE

Figure 1.2-4. Deployment/Retraction Mechanism Details

1-8
The separate longeron and diagonal unlocking systems are activated only

during retraction and are respectively used to unlock the longeron and

t diagonal center joint latches prior to the start of the batten retraction.

The diagonal and longeron center joint latches are unlocked by forward motion

of the trip lever pine and tripping probes mounted on the

deployable/retractable carriages installed within rails and driven by the

25.4-mm-diameter jackscrew.

The positioning system requirements for this program are a version of .

standard motion control (robotics) used in inrustrial machine control

applications.

Motion profiles are built up as sequenced indexes. Each index consists of

a direction, acceleration time, deceleration time, feed rate, and travel

distance. The controller calculates acceleration rates, deceleration rates,

1 and the position to begin deceleration. The mechanization uses encoder and

tachometer sensing with overrated motors and mechanization to ensure precise

position control without overshoot in the presence of varying output loads.

The motion profiles for the test article are shown in Figure 1.2-5.

The three -axis system selected will allow totally separate positioning of

(1) the batten deployment/retraction system, (2) the longeron unlocking

system, and (3) the diagonal unlocking system.

The batten deployment/ retraction axis controller will use a direct-drive

do servo motor rated at 27 Nm (140 lb-in.) continous operation up to 225 rpm.

The motor will be driven with a standard pulse-width modulated drive.

Positioning resolution will be within 0.001 revolution which is equivalent to

a longitudinal accuracy of 0.0064 mm (0.00025 in.) on the 6.35 mm pitch

jackscrew. The deployment/retraction profile will be achieved as a series of

ten indexes entered into a specific program.

The diagonal and longeron unlocking controllers will be configured with

identical hardware and software. Again, direct-drive do servo motors will be

used rated at 3.7 Nm (33 lb-in.) continuous up to 2400 rpm. The motor

contains an integrally mounted encoder and tachometer. Each of the motors

will have its own pulse-width modulated drive and do drive power supply.

These controllers wSll be to within 0.0025 revolution which is equivalent to a

longitudinal accuracy of 0.127 mm (0.0005 in.) on the 5.08 mm pitch jackscrew

to which they will be mounted. Figure 1.2-6 showi an illustration of the

typical components

Figure 1.2-7 illustrates the key discrete stages of deployment and

C retraction. Starting from the stowed package (View 1) the end adapter, which

` is the foward batten of Bay 1, is forward of the jackscrew support frame. The Y

first stage of deployment positions and locks the jackscrews and the jackscrew

support frame diagonal struts, and develops (View 2) Bay 1. At this point,

the Batten 1 (Figure 1.2-1) half-nuts are engaged with the aft end of the

jackscrew thread. The batten deployment/retraction system jackscrews are

13. reversed to start the deployment of Bay 2 (View 3).

1-9
1

D.

190 REV

PIOTOR DIRECTION 1% REV I I O DENOTES BAY

(TTR EXCEPT CARRIER NUMBER

CLOCKWISE I AS NOTED) BAY 156 REV r

COUNTERCLOCKWISE O r,1 4 o ^, O O - 1Q --

l:i, 

I

LJ ^J

SEQUENCE 0 1 2 3 4 S 6 ' 8 9 10

DNL►YMENT PHASE -- BATTEN SYSTEM

25 REV r-1

ZO REV

TYP. 10 REV +1 r 5 REV

^--^

CLOCKWISE

21 0 REV

COUNTERCLOCKWISE I- 60 REV I I - - -I- ---- -

i 1 1 I I I I

OENOTES

RETRACTION PHASE -- DIAGONAL/L KING SYSTEM UNLOCKING

6 I I tFDIAGONAALS

CLOCKWISE 10 I O O O O

^

CL f-164 REV o

COUNTERCLOCKWISE

I 1

1-1 156 REY 1

F+M P 196 REV (TYP EXCEPT At NOTED) I 1I REVk

9"9.

I

SEQUENCE 0 1 2 7 41 S 6 1 8 R",

10

RETRACTION PHASE -- BATTEN SYSTEM

Figure 1.2-5. Positioning system Motion Profile

ORIGINAL PAGE IS

OF POOR QUALITY

Figure 1.1-6. Positioning system Components

1-10
Batten 2 (Figure 1.2-7) is held in place by spring-loaded detente until

Bay 2 is fully extended and locked, and is later overwhelmed by the jackscrew

starting the deployment of Bay 3. In this manner, each of the bays is

deployed one at a time until the fully deployed truss configuration is

achieved (View 4).

At this point, precompression of the longerons can be appl?ed and removed

by manual activation of the precompression system.

Iu the retraction phase the longeron and diagonal unlocking carriages are

initially positioned such that each of the tripping probes :s 25.4 mm away

from the longeron and diagonal latch trip levers. Tka four iiagonsl and four

longeron latches in Bay 10 are tripped after tan clockwise revolutions of the

unlocking system jackscrews. After a number of milliseconds (to bt determined

in the future ground tests) the batten deploymentirstraction system motors are

rotated clockwise until Batten 9 (Figure 1.2-1) in plsced on the rail. As

each bay is retracted, the carriages on the unlocking systems ere adjaaced to

the next unlocking position (Figure 1.2-5). This proceeds f: g um Lcy 10 chcough

unlocking of Bay 1 (view 6). Upon unlocking the longerons and diagoav 's of

Bay 1, the batten deployment/retraction jackscrews are rotated Counter,

clockwise 32 revolutions. The extended diagonal and longeron unlocking

systems are then retracted into the housing to permit the final retracti % n of

Bay 1.

Figure 1.2-8 describes the major features of the precompression system

provided to eliminate joint backlash in both the longerons and diagonals. A

cable/bungee system, with a cable pretension of 1780 N (400 lb), will apply up

to 1425 N of compression in each of the four truss longerons. This

compression load will, through compatible strain, provide up to 260 N of 

precompression in the diagonals.

The precompression system consists of two spring bungee assemblies mounted

on the aft and of the main housing. From either end of each bungee, threaded

rode are extended that mate with a turnbuckle. From the opposite end of each

turnbuckle is another threaded rod swagged to a long cable. The two from each

turnbuckle traverse laterally until they engage a pulley near the axes of the

longerons. The cables wrap around the pulleys 90 degrees and extend forward

where they enter the longerons located at the four corners of the truss. The

cables continue forward through the longerone of all ten bays. The cables

exit the longerons of Bay 1 and engage another pair of tairleads mounted

within the adapter. These fairleads are canted in such a way that the cables

continue toward the geometric center of the adapter within its diagonal

braces. Swagged balls on the cables attach to fittings whose mounting

locations are adjustable within the adapter.

The bungees are supported on the rear of the housing by two pairs of

brackets that partially encircle the cylindrical body and still allow the body

to move along its axis as the turnbuckles are utilized to pretension the

cables to their final 1780 N load.

Figure 1.2-9 illustrates the major parts of the main housing which is a

combination welded, riveted, and bolted assembly into which all other major

assemblies are installed. A welded frame consisting of 50.8 mm square

aluminum 6061-T6 tubing has numerous skin and stringer subassemblies riveted

to it.

.., cv—Lw+atf..,:^:,.:: at 7 _ .• ^a:.,, '. !;. ,._6'Y. - ti>a t"k — . 
Ct

® ALL N

RAY 1

I

ORIOIPIAL PAUL 1%

OF POOR QUALrN

4^^ 4 

0,^ —17--^

/ m DE►LOrMEMT OF

m fTOYED• ADAPTER

PACKAGE • JACKSCREW SUPPORT

• FIRST RAY (BAY 1)

FRAME ASSY

(3)PARTIAL. DEPLOYMENT,

1 0 A E

S DEPLOYMENT OF

SAYS X THROUGH 10

w

TRIPPING A IAOORAL

TRIPPING DE V ICES

RETRACTED

BAY 10

Figure 1.2-7. Deployment/Retraction Major Phases

3.1 = CIA (7 X 19) CABLE

CABLES EXTEND FORWARD

THROUGH LONGERONS d

ATTACH TO ADAPTER

MANUAL ADJUSTMENT OF

TENSION CABLE PERMITS

TEST DATA FROM

0 TO 1780 M

PROVIDES

Figure 1.2-9. Precompression System to Eliminate Joint Slop
F PL-,0 Q U,4

Lrry

DRIVE MOTOR

SUPW

rxvw "AM AWAWIT TO

ATTAM PITTING

2 3 No ALLIMI

' ^^1}'. _1T7 I _ ^ 1 • ALL S I OCS ii'"w PANELS

ACCESS

PAKU STIFFENERS TO

PROVIDE MINIMUM

FREQUENCY TO 10 "1

HOUSING

lArr[M MAJOR

DETENT FRAMING

ADJUSTPOT

"OUS (TYP)

Figure 1.2-9. Main Housing Structure

1-13
N Panels on the aft side of the housing are removable to provide access to

the precompression system and the three chain-and-sprocket drive systems

located near the center of the housing. Access holes along the four sides of

the housing align with the batten retaining detents to provide adjustment

capability.

A rectangular pattern of threaded inserts is provided on the four sides of

the housing for the future attachment of the NASA/MSFC simulated payload

terrier structures.

The foregoing describud test article design is representative of a

square-truss, single-fold prototype building-block design from which potential

space platforms (Figure 1) or Space Station structures can be constructed.

The protoype building block has the following significant design

characteristics:

• Automatic bay-by-bay deployment and/or retraction to facilitate

identification of problem (in the event this occurs)

• Maintenance of root strength during deployment/retraction operations

which can permit orbiter VRCS firing (if necessary)

• Longitudinal deployment/retraction within the housing cross-section

envelope

• All inter-building-block electrical connections in olace prior to 

orbiter installation

• In-space Inter-building-block structural connections made automatically

without a fixture.

• Square-shaped truss which is most suitable for inter-building-block

attachments; mounting of payloads, docking ports, and propulsion

modules; and redundancy for meteoroid impact

• Circular tubes for all truss members which is the minimum cost

construction for graphite composite construction

• Widt trays for mounting of a large complement of utilities with ease of

initial installation, repair, replacement during the total ground

fabrication period, aid with minimum truss structural design constraints

• Narrow trays for minimal complement of utilities with a longeron fold

angle of 40 to 45°

• Payloads and propulsion modules can be attached using the RMS

h precompression system to eliminate joint backlash

Unlocking systems (for retraction) that are easily removable for

applications in which retraction is not required

1-14
1.3 TEST A'1TiCLE DESIGN

This section contains a detailed description of each of the major subaseemblies that comprise the test articles major components (1.3.1), a listing

of the test articles major deviations from the prototype design (1.3.2), a

description of the test article assembly procedure (1.3.3), and a detailed

description of the method of deployment and retraction (1.3.4).

1.3.1 Major Subassemblies Design Description

The major subassemblies that comprise the major cumponents summarized in

Section 1.2 are described in further detail in this section.

1.3.1.1 Deployable Truss

Each of the major subassemblies that comprise a bay of the deployable

truss are described as follows:

Diagonal Assembly (Drawing 42112-110) - Figure 1.3-1 illustrates tr.e major

design features of the diagonal assembly. Tha outer part of this telescoping

uesign consists of, from left to right, a rod end threaded to a turnbuckle

which is threaded to a transition fitting that is blind civeted to a

50.8-mm-diameter 6061-T6 aluminum tube that is adhesive bonded to an outer

center fitting. The unlocking latch mechanism is attached to this fitting.

The inner sliding part, from right to left, consists of an identical rod end

threaded to a transition fitting that is blind riveted to a 44.5-mm-diameter

tube also adhesive bonded to an inner center f tting. The rod ends are

stainless steel, the transition fittings CRIS 304 steel, and center fittings

(Figure 1.3-2) 6061-T6 aluminum.

Within this inner center fitting is a pair of opposing spring-loaded

locking pins oriented 90° to the axis of the tube. These pins have blunt

points machined on them that are aligned with holes on opposite sides of the

inner and uuter center fittings. A apr • ing between the two pins exerts a

separating force on them. In all but the extended position, the pins are

prohibited from protruding outside the inner tube because of the encircling

wall of the outer tube. During truss deployment, the two tubes are extending

with the spring loaded pins of t'e inner tube pressing on the inner wall of

the outer tube. At within 6 mm of compl__s diagonal extension the spring-load

pins come ..nto axial alignment with the two opposing holes in the outer tube.

The spring forces the pins into the matching holes of the outer tube, thus

locking the two tubes in that position. The longitudinal force provided by

the pair of springs is 82.8 N. This operation is automatic and occurs in each

of the four diagonals of each bay.

a

The two shear pine provide the axial load path between the inner and outer

center fitting. The pins also provide capability to sustain bending moments

acting in a plane containing the axis of the pins. A bending moment acting in

a plane perpendicular to the axis of the pine is sustained by bearing between

i the inner and outer fittings at the stations reinforced by rings (Figure

1.3-2). The tolerances between the mating diameters of the two fittings have

!.' been limited for this reason (Section 1.5). The diagonals will be fabricated

to the dimensional length tolerances shown on the drawings. Upon final

assembly, as described in Section 1.3-3, the diagonals will be adjusted to the

length compatible with trouble-free deployment and retraction by micro--

adjustment of the turnbuckle provided.

1-'i5
1 ^.

URIGINAL PAGE IS

OF POOR QUALITY

1.71 n (TYPICAL IAY)-

5/6

STEELNT-I

AIS a D.C.

46E3=1 v

TIRNIMIEKL[

..TUBE 

60.1 m D.C. TIME 0/B IM.-IA DIA m CIA HOLE STEEL 400 END (2) RICAL

ING (1)

Figure 1.3-1. Telescoping Diagonal Joint

INNER TUBE

4 1-5 . CIA

N 2.1 w• WALL

152 n --^—^1

DOTER TUBE

$0.8 w CIA K

2,1 s WALL

CAM ADJUSTING

SCREW

LATCH UNLOCKING

LEVER (NEAR SIDE ONLY),

.jA

 Z

UNLOCKING CAN

RETURN SPRING

UNLOCKING

/ CAN (2)

11.7 s CIA CITES

LOCKING PIN (2)

17.5 mw D.C.

K IJ s COIL

OJA 17-) PH WIRE

EPDXY ADHESIVE JOINT

SECTION A-A (ROOM TEMPERATURE)

slescoping Diagonal Center Joint

1-16
For the retraction operation, an external cam lever is provided ( Figure

1.3-2) and is actuated by the diagonal unlocking trip lever pin as follows:

The trip lever pin on the carriage is extended by the jackscrew until it

contacts and causes the diagonal ' s external cam lever to rotate. The rotating

lever causes the cam to partially depress the protruding spring - loaded pins.

Subsequently, the batten jackscrew begins moving the batten toward the

housing. This motion induces bearing on the pin holes of the outer tube which

completely compresses the tapered pins inside the center fitting. Continued

batten retraction permits total telescoping of the diagonal. In the stowed

configuration the diagonal nests between the batten assembly members ( Figure

1.2-1).

Longeron Assembly (Crawing 42712-120) - Figure 1.3-3 illustrates the major

design features of the l ongeron assembly. Each longeron is comprised of two

identical halves ( except for provisions for attachment of the latching

mechanism). Each half consists of a rod and (identical to that provided in

the diagonal assembly) threaded to a turnbuckle which is threaded to a CRES

304 transition section which is blind riveted to the 50 . 6 mm (2 in.) diameter

6061-T6 tuhas which are blind riveted to a 2024 aluminum center fitting

(Figure 1.3 -4). The center fitting on each half, in conjunction with the side

plates, comprise the center hinge mechanism required to fold the longerons to

the configuration shown in Figures 1.2-2 and 1 . 3-5. This design was developed

to prevent the increase and subsequent decrease in longeron length that occur

in an over -center latch pivot design at the stage of deployment shown in

Figure 1.3-6. The added complications to the aforementioned diagonal

telescoping tube locking and the positioning system motion profile would be

prohibitive.

Longeron axial loads are transferred from the left center fitting to the

pivot pin into the pair of side plates and then into the right center fitting

through the pivot pin. Bending moments acting in a plane perpendicular to the

planes of the the side plates are similarly transferred across the joints.

Bending moments acting in the plane of the side plates are transferred by

bearing in the pivot pins and the 9 . 5 mm (3 1 8 in.) diameter locking pins.

Shear stiffiness in the plane perpendicular to the side plates is provided by

the web shown in Figure 1.3-4.

Two sector gears are installed on either side of the hinge to eliminate

the instability of the four pivot points of the longeron during depl-)yment.

These sector gears anchor the two pivots of the longeron halves together

causing them to rotate in a singular controlled manner. The result is to make

the two pivots act as one hinge on the center line of the longeron.

Locking of the longeron in the fully deployed position is accomplished by

a pair of spring - loaded pine ( similar to those in the diagonal) in the fold

mechanism. The aforementioned side plates extend beyond their pivot pins a

short distance and provi4v a pair of hole° that are aligned to receive the

spring- loaded pins when the two rotating arms arrive at their fully deployed

position. The engaging pins and pivot pins lock each half of the longeron to

the side plates and, since the plates are continuous across the hinges, the

hinge becomes rigid in the fully extended position.

1-17

.i
ORIGINAL PAGE 15

OF POOR QUALITY

W"

B

I.... ... .

14 mm CIA

17.7 PH WIRE

CENTER FITTINGS

6.4 mm CIA PIVOT BOLTS (4)

/

SIDE PLATE 9.5 mm CIA ORES

_ ZLOCKING PINS

PLATE SIOE 17^mm

CIA PIN SIDE PLATE

SECTION B-B (2024)

Figure 1.3-4. Longeron Center Joint Details

1-18

.uwuax6^..:x1vexvt. ^n..,..cfmaYE/1La^.i/, .0 u.ri,f.. X:.^+S^

.. ...::..3J...

19 mm

LIGHTENING

HOLE ^

SECTION C-C

HOLES FOR 6.4 mm CIA

PIVOT Bola

SECTION A-.A

IB

WEB FOR

VERTICAL

SHEAR

cR[s 1.50 • (TYPICAL BAY)

STEEL 5/B IN.-IB

TRANSITION STEEL TURNBUCKLE (2)

FITTING(2) \`

3/6 M,24 STEEL R00 ENO (2)

7.94 mm CIA HOLE 50.8 w 0.0. TUBE

SPHERICAL

BEARING (2)

Figure 1.3-3. Folding Longeron Assembly
FOR

FOR

BAY

ORIGNAL PAGE 13

OF POOR QUALITY

---"-- ',ARS

TRIP 127 w^--+{

LEVER

TRIP MECHANISM

SPRING

LOADED

SIDE PLUNGERS

PLATES

i

y .-y l - g

LONGERONS STOWED CAN (708 ORES) SIDE PLATE

(TRIP LEVERS7.2 m - 2028

AL

♦ ^

OMITTED)

A . A-A CAI RETURN SPRING

$0.8 = O.D. X 2.1 w WALL AL TUBE

SC

SIX - 4.8 m DIA AL rc^ BLIND RIVETS

TRIP LEVER

LONGERON DEPLOYED

Figure 1.3-5. Longeron Center Joint Additional Details
Like the diagonal struts, there is r taper on the and of the spring-loaded

pins that assists in the final degrees of longeron deployment. This is

accomplished by the points of the tapered pins engaging the edge of the holes

in the side plates with the resulting ramming action forcing the two arms of

the longerons into their final deployed position.

After the spring-loaded pins have fully engaged their respective holes in

the side plates, their aforementioned conical ends extend outside the plates.

To unlock the longeron and allow it to fold into its retracted position, a

pivot hole is added to the side plates. A rotating shaft is inserted through

the pivot hole and a cam plate is installed on each end of the shaft. The

cams are located so as to depress the extended cones of the spring-loaded

locking pins in a manner similar to the aforementioned diagonal. A lever arm

is added to one of the cams. A probe extending from the carriage of the

longeron retraction jackscrew assembly engages this lever and rotates it

enough to unlock the latch by causing the cams to depress the apring-loaded

pins (Figure 1.3-5) and also exerts a side force on the hinge that causes the

hinge to fold off "dead center" as the outboard batten is retracted by the

b.,tten jackscrew assembly. As the batten continues to move toward the

housing, the longeron folds at an angle of 35° inboard from the plane of the

truss. This out-of-plane fold angle provides clearance for the telescoping

diagonal struts to be stowed along side the batten members.

Upon final assembly (Section 1.3.3) each of the longeron halves will be

micro-adjusted to the lengths compatible with trouble-free operation by the

turnbuckles located at richer end. Dual turnbuckles are necessary so the

folding hinge is always located at the center of the longeron to prohibit

binding during the fold sequence. The turnbuckle shown, compared to

conventional turnbuckle designs, provides a more direct load path and, because 

of the reduced length, results in a stiffer member for column applications.

Typical Batten Assembly (Drawing 43712-190) - Figure 1.3-7 illustrates the

major design features of the typical batten assembly that, in conjunction with

four longerons and four diagonals, comprise a typical bay of the deployable

;'. truss (Figure i.2-2). This batten assembly is not to be confused with the

payload carrier batten assembly (Drawing 42712-220).

The batten is comprised of four 35 mm (1-3/8 in.) diameter 6061-T6 tubes

that define the trues square frame, stabilized by a compression-carrying

diagonal brace of the same size and material. The brace contains tapered end 

fittings to preclude interference with the stowed longerons. The four frame

members are joined together by blind riveting to the four corner fittings

shown. The corner fittings contain integral clevises for attachment of the

diagonal and longron rod ends, and the threaded half-nuts that engage the

• batten deployment/retraction system jackscrews. The corner fittings also 

provide gussets for mechanical fastening of the diagonal brace. Each of the

corner fittings contains a pair of pockets into which ball detente, mounted in

the deployment/retraction system rails, enter to retain the stowed batten.

The corner fittings also contain openings, through which the precompression 

1 system tension cables are routed.

The batten corner fitting clevises (for the rod ends of the diagonal and

longeron members) contain shoulders on the inner surfaces to minimize the

axial rotation of the members to + 1.5°. This restriction is necessary to

preclude the possibility of the latch release levers rotating out of the plane

of their respective tripping probes mounted on the diagonal and longeron

unlocking system carriages.

1-20 
The half-nut corner fittings are coated with teflon ME) to reduce

friction while in contact with the rotatii,Z jackscrews during deployment and

retraction.

Payload CarrierBatten Assembly (Drawing 42712-220) - The payload carrier

batten assembly overall configuration and design features are shown in Figure

1.3-8. There are two of these battens onto which the NASA/MSFC simulated

payload carriers will be mounted. The 1/2-20 threaded inserts provided 1.31 m

(51.5 in.) apart along the batten and 1.09 m (43.0 in.) apart between the Bair

of'bettens are the attachment interface for the simulated payload carriers.

This attachment interface is provided on each of the four batten faces.

This batten is designed to locally support installation of a 3636 kg

(8000 lb) simulated payload carrier, for modal survey tests, without

detrimental deformation. provided the frame itself is supported by a pair of

vertical tension cables attached at the bottom of the frame or at points "A"

and "8" (Figure 1.3-8). Therefore, the batten is comprised of 50.8 mm 6061-T6

square tube members. Further, the square tubing can permit future

installation of electrical feedthrough plates.

The corner fittings are configured to accommodate blind riveting for

i w

attachment of the square tubes and to provide threaded inserts for

installation of the payload carriers. It also provides clevises with

rotation-limiting features for the longerons and diagonals, pockets for the

ball detente, threaded half-nuts, and holes for the routing of the

is precompression system tension cables.

g

Like the corner fittings on the regular battens, the half-nuts are also

coated with teflon.

Utilities Support Trays (Drawing 42712-230) - The deployable truss design

includes utilities support trays. For this ground test article the utilities

trays are provided in only Bays 4 and 5. The trays provide support for the

required complement of four No. 1/0 cables, two No. d AWC, four No. 12 AWG,

four coax, twenty No. 22 twisted pairs shielded, and four 12.5 mm flexible

coolant tubes. There is additional space available for two No. 1/0 cables,

one No. 8 AWG, two No. 12 AWG, two coax, and eight No. 22 twisted pairs

shielded.

The tray assembly in each bay consists of two half-trays hinged to each

other at the tray edges in the center of the bay. Each tray half attaches to

hinge fittings mounted on the batten tube frame. 

Each half-tray assembly has a machined aluminum base containing the four

hinge points, tray sides and stiffeners. Three epoxy glass block clamp sets

and a machined aluminum cover are bolted to the base. The block clamp s,ts

p rovide support of the utilities and ample separation of power and data

lines. The base and cover contain cutouts to reduce weight.

One half-tray assembly base in each bay has provisions for four bolt-on

longeron supports. These supports provide lateral stability for the longerons

in their stowed position and have capture guides that assist the longeron 

during retraction.

1-21
I

112 -20 THREADED

INSERT FOR

NASA/NSFC PAYLOAD

1,5 1

}

I

[ DF THREAD

ORIGINAL PAGE 19

OF POOR QUALrN

35 w DIA N 1.2 m

WALL 6061-T6

BLIND AL RIVETS

6061-T651

LORGERON CLEVIS

117 M

A A

n

0.S

t LONGERON CLEVIS

Figure 1.3-7. Typical Batten Assembly

101.6 w

o ^_

A-A 2.0 - 6 ACME - AG

♦uDrin

A

CKET FOR BATTEN

TENTION BALL DETEHTS

A

M OF

1 ACME - AG THREAD

Figure 1.3-8. Payload Carrier Structure

1-22

_. ... .^^y4

1. e .•. •a Y - _^_-l.
When stowed, each utility tray will nest between the diagonal braces of

adjacent battens. Each tray base has two pads that can transmit longitudinal

support from the folded tray to the adjacent tray. For a truss containing

trays throughout, the longitudinal load can be transferred to the support

beams provided in the main housing and adapter.

The length of the utility tray assembly is slightly longer than the

distance between adjacent batten hinges so that the trays center hinges will

be approximately 50 mm (2 in.) above the batten hinge centerline when

deployed. This is to prevent a locking condition on retraction.

1.3.1.2 Deployment/Retraction Mechanism

( Each of the major subassemblies that comprise the deployment/retraction

mechanism is described as follows:

` Batten Deployment/Retraction System (Drawing 41712-130) - Figure 1.3-9

illustrates the major design features of one of the four batten deployment/

retraction system assemblies. Each assembly consists of a jackscrew,

carriage, and spline assembly cradied within a rigid rail that is attached to

a pair of main housing frames. The jackscrew threads and rails match similar

features on the half-nuts of the batten and carrier assemblies (Figure

" 1.3-10). A splined bushing at the aft end of the jackscrew encircles a

splined shaft that runs nearly the entire length of the jackscrew. The

jackscrew spline extends beyond the aft end of the rail where a chain and

;.: sprocket is attached. One of the four spline shafts is coupled to a drive

motor mounted on the aft end of the housing. A chain encircling the four

sprockets drives all of the jackscrews simultaneously. (Figure 1.2-4)

Encircling the rotating jackscrew is a carriage fitting with external ears

that engage matching grooves running the length of the rails. The carriage is

pulled forward with the jackscrew, during deployment of the first bay, until a

hole in the side of the carriage engages a spring-operated pin (Figure 1.3-10)

mounted near the forward end of each rail. During jackscrew retraction, the

spring-loaded pin is manually retracted from the carriage, thus allowing the

carriage and jackscrew to be subsequently retracted into the housing.

kt the cantilevered forward end of the spline is attached a sleeve whose

outside diameter is slightly smalier than the inside diameter of the

jackscrew. The purpose of this sleeve is to maintain concentricity of the

splined shaft with the jackscrew.

j:

` At the forward end of the jackscrew (Figure 1.3-9) a probe is provided

that supports the jackscrew support frame and the end adapter in the retracted

position. The jackscrew support frame is attached by a plate on the frame

engaging an annular groove in the probe. The end adapter is retained on the

" probe by an adjustable spring-loaded ball detent in each probe engaging a

matching groove within each of the four drogues located at the corners of the

w.

adapter.

At eight places along each rail are adjustable spring-loaded detents that

are utilized to offer resistance to deployment of battens stowed along with

the rails while a bay is being formed along the extended jackscrews. Once the

*r... , bay is formed b full extending the lon eron and diagonal struts Y Y Y 8 B B , sufficient

1-23
PIN RETRACTION CAN

IN CARRIAGE

CARRIAGE

, RAIL

LOCKED

CARRIAGE LOCK

,u.

6001

MIL

ORIGINAL PAGE 18

OF POOR QUALITY

BATTEN 106 = P.D. CRES LATCH

HALF SPROCKET DEPRESSION

NUT RAMP

I7 = DIA HOLE

411 - 4 CARRIAGE FOR LATCH PIN

6G NO. 35

100 STEEL) CRES CHAIN

MA

TUN)

A 20.6 s

SPLINE SHAFT

x•

1.67 n 203 s—^

'r II 36 OU 67.now

-T66

FOR MOUNTING OF

JACKSCREW

THREADED

THREADED

INSERT FOR AL TRU55

INSERTFOR

ADAPTER SUPPORT ATTACK FAME

PROBE ATTACH / 

ATTACH FITTING ATTACH

/-IF

ANNULAR FOR IATTEN ' RETENT ION F_F

GROOVE

JACKSCREW ASSEMBLY

Figure 1.3-9. Batten Deployment/Retraction System Mechanism

I

v•:

Figure 1.3-10. Carriage Locking Detail and

Half-Nut to Rail Matching Features

1-24
force can be exerted on the restrained batten by the jackscrew to overwhelm

the resistance of the detents and, thereby, deploy the batten. The same

procedure is repeated for deployment of each bay until the entire truss is

deployed.

Attached approximately midway to the rails are truss attachment fittings

that iaterface with the aft ends of the longeron and diagonal struts that

comprise the last bay. These anchor fittings replace a batten that is

normally used elsewhere in the truss. When the truss is completely deployed

the fittings transfer longeron and diagonal loads to the rails, and to the

housing through the rail attachment to the housing. This frame is adjacent to

the fittings and is stiffened by the shear web shown in Figure 1.2-9.

Longeron. and Diagonal Unlocking Assembly (Drawing 42112-140) - Figure

1.3-11 illustrates the major design features of the longeron and diagonal

unlocking assemblies. The figure at the upper left represents the partially

deployed configuration. The diagonal unlocking assembly is identical except

for the difference in the tripping device.

There are four longeron unlocking assemblies (one for each of the four

longerons) and four diagonal unlocking assemblies (one for each of the four

diagonalsi. These assemblies are located as shown on Figure 1.2-3

Each of these eight assemblies consists of a 6061-T6 guide rail, 25.4 mm

(1 in.) diameter 304 CRES steel jackscrew, 6061-T6 carriage, and appropriate

latch tripping device. The jackscrew has an ACME thread with a 5.08 mm pitch

(0.20 in.). The guide rail is fastened through adjustable shims to the main

housing frames. Sprockets mounted on the aft-extended end of the threaded

carriage rods are interconnected by a chain encircling the sprockets. A drive

motor is attached to one of the extended threaded carriage rods, and the

interconnecting chain drives the three remaining assemblies within etch

system. (Figure 1.2-4).

' The separate longeron and diagonal unlocking systems are activated only

during retraction and are respectively used to unlock the longeron and

diagonal center joint latches just prior to the start of the batten

retraction. The diagonal and longeron center joint latches are unlocked by

fo--ward motion of the trip lever pins and tripping probes mounted on the

deployaLle/retractable carriages installed within rails and driven by the

jackscrews. This is accomplished as follows:

4

i At the center of each telescoping diagonal strut and folding longeron are

internal spring-loaded pins that automatically lock each strut in its fully

extended position. In this locked position, tapered ends of the spring-loaded

pins are extended through the walls of the struts beyond their exterior

surfaces. Lever operated cams, mounted on the exterior of the members depress

x the spring loaded pins, thus unlocking the struts and thereby permitting them

to be folded or telecoped to the etowed position.

A motor controller is programmed to rotate each carriage jackscrew,

thereby extending probes, (mounted on the forward ends of each carriage)

enough to engage the cam levers on each strut and to cause them to rotate

until the cams depress the spring-loaded pins within the members. The

1-25
1.

}

i

,

forward batten of that bay can then be moved toward the main housing, causing

the diagonals to compress and the longerons to fold into their fully-stowed

position. Because the stowed thickness of a typical bay is 101.6 mm (four

inches), the carriage must be advanced this distance to center the carriage

trip probe with the lever cam latches in the subsequent bay. After the

latches of the final bay are tripped, the motor controller is programmed to

retract the extended carriages back into the housing to permit subsequent

retraction of the jack s crew support frame and the truss end adapter.

At first glance one would think that the mechanization of a tripping

system for the diagonals would be identical to one of the longerons because

they both simply rotate levers on the respective latches. A closer analysis

of the two systems kinematics reveals the initial rotation of the latch lever

depresses the locking pins of the latches. In the case of the diagonals, once

the pine are retracted, the extended bay batten can be retracted into the

housing. This action causes the telescoping diagonals to shorten. In the

case of the longerons, however, just retracting the locking pins will not

necessarily allow the longeron to fold. A friction lock could still inhibit

the longeron from folding. An additional force perpendicular to the axis of

the longeron pivots is required to displace the pivot off-center simultaneouzly with the retraction of the extended bay batten. Thus it can be seen

that a single unlocking system to perform all these functions was not

appropriate.

1.3.1.3 Jackscrew Support Frame Assembly (Drawing 42712-170)

Figure 1.3-12 illustrates the major featuras of the jackscrew support

frame assembly. This assembly, consisting of four telescoping diagonals and a

frame, provides a "deep truss" stiffness to the cantilevered ends of the four

batten deployment/retraction system jackscrews during all operations excepting

the deployment/retraction of Bay 1 (Figure 1.2-3). During deployment/

retraction of this bay the four jackscrews and frame behave as a Vierendeel

truss. This is expected to be adequate during the initial stages of deployment of the trues since the length is small; hence, on-orbit moments and

stiffness requirements should be minimal. The frame itself also minimizes any

relative lateral separation between the jackscrews compatible with maintaining

thread engagement.

The assembly consists of the support frame (View A-A) and four telescoping

diagonals. The diagonals are identical to the typical diagonals in the

deployable square truss except for length. The ends of the diagonals are

attached to clevis fittings on the main housing forward frame and to the frame

as shown in detail "A".

The frame is comprised of four 6063-T52 rectangular tubes riveted together

through the fittings shown on Details "A" and "B". The dog-legs in the

fittings are required to permit clearance during the traverse of the

deployable truss longeron latch mechanisms and the batten corner fittings

half-nuts. The dog-leg and requirements of compact stowage directed the

design to have offsets between the jackscrew, diagonal, and frame member

centerlines. This has been accounted for in the design (Volume III).

1-26
U'11 "119,4L PAGE t9

OF POOR QUALITY

91.9E

i-N 4G THREAD 1 S AG

1 S 

AC

HE

6061

CARR IAGE

6061 AL

L

M

n

MIL 6061

Al

NO. 75 (9.S3 = /ITCH

L GERON ON CRIS CHAIN

UNLOCKING

MECHANISM

106.17 v P.D. CRIES

SPROCKET

(AS SIN") 25.1 = DIA JACXSCREWI

CRES STEEL

CAMIAGE

(1.24 n LONG) ^-i ,

RELEASE PROBE

4 ATYACNIfM

rTO MOUSING

fA1RLEAQjOR JACKSCREX SUPPORT

I TO HOUSING

I mo--" 1.57 n .l

UNLOCKING MECHANISM STONED

NOTE • DIAGOML UNLOCKING

NECINNISH 15 ESSENTIALLY

THE SAME EXCEPT FOR

LATCH RELEASE 6 ATTACHMENT

Figure 1.3-11. Longeron/Diagonal Unlocking Mechanisms

a•

F' 51 w CIA TELESCOPING JACKSCRE

DIAGONALS (AUTOMATIC .,6 RETAINER

LOCKING/MANUAL UNLOCKING)

y.a

I d 'I^

I A JACKSCpEN

FERRU LE / 6061-T

DETAIL B

4-6.4 w CIA

BLIND AL

RIVETS (TYP)

4. CL Vls

f DIA

GONAL

I A-A `% DETAIL A

Figure 1.3-12. Jackscrew Support Frame

1-27

v • R

X. 
1.3.1.4 Positioning System Design

During the early part of this study Rockwell obtained the assistance of

PMI Motors (Division of Kollmorgen Corporation) to determine the positioning

system design characteristics. The resulting design is documented in

Reference 10. At the conclusion of this study, PMI Motors deferred the

positioning system responsibility to the Kiowa Corporation. The following

positioning system description has been provided by the Kiowa Corporation.

The positioning system requirements for this program are a version of a

standard motion control (robotics) used in industrial machine control

applications.

Motion profiles are built up as sequenced indexes. Each index consists of

a direction, acceleration time, deceleration time, feed rate, and travel

distance. The controller calculates acceleration rates, deceleration rates,

and the position to begin deceleration. The mechanization uses encoder and

tachometer sensing with overrated motors and mechanization to ensure precise

position control without overshoot in the presence of varying output loads.

The motion profiles for the test article are shown in Figure 1.2-5.

The three-axis system selected will allow totally aeydrate positioning of

(1) the batten deployment/retraction system, (2) the longeron unlocking

system, and (3) the diagonal unlocking system.

In addition, the system includes a programmable output option that allows

the batten deployment axis to sequence the longeron and diagonal unlocking

axes. The system consists of a standard main frame chassis with three'

standard motor control modules, position feedback modules, and digital input/

output modules. In addition, three machine logic simulators are included for

all motion fuctions on any axis; for example, jog, run, hold, and high or low

speed.

The batten deployment/retraction axis controller will use a direct-drive

do servo motor rated at 27 Nm (240 lb-in.) continuous operation up to 225

rpm. The motor will be driven with a standard pulse-width modulated drive.

Positioning resolution will be to within 0.001 revolution which is equivalent

to a longitudinal accuracy of 0.0064 mm (0.00025 in.) on the 6.35 mm pitch

w jackscrew. The deployment/retraction profile will be achieved as a varies of

ten indexes entered into a specific program.

The diagonal and longeron unlocking controllers will be configured with

identical hardware and software. Again, direct-drive do servo motors will be

used rated at 3.7 Nm (33 lb-in.) continuous up to 2400 rpm. The motor

contains an integrally mounted encoder and tachometer. Each of the motors

will have its own pulse-width modulated drive and do drive power supply.

These controllers will be to within 0.0025 revolution which is equivalent to a

longitudinal accuracy of 0.0127 mm (0.0005 in.) on the 5.08 mm pitch jackscrew

to which they will be mounted.

1-28

--
The controller is an Intel-8085 based eight-bit microcomputer with 2K of

random access memory ( RAM), 16K of e-raseable programmable read-only memory

(EPROM), and 1K of electrically alterable read-only memory ( EAROM). For

extended programming capability, the EAROM may be expanded to 5K. EAROM is

nonvolatile, having at least ten years of storage life for the user input

parameters. Operating system software integral with the controller simplifies

the task of user programming. The configuration is arranged through a

prompting routine displayed from an alphanumeric display requiring simple

command input from a keyboard.

Interface between the controller, motors, and sensor feedback is done

through 24 input / output lines defined for specific functions.

All of the hardware, except the motors, the remote programing panel and

cabling will be mounted in a standard 914 m by 914 m by 305 m enclosure.

Figure 1.2-6 shows an illustration of typical components, and Figures

1.3-13 and -14 show they key dimensions for the selected motors.

1.3.1.5 Precompression System Assembly ( Drawing 42712-150)

Figure 1.3- 15 illustrates the major design features of the precompression

system provided to eliminate the slop in the 280 pivots of the ten-bay

deployable cruse.

The system consists of a pair of spring bungee assemblies mounted on the

aft end of the main housing. From either end of each bungee threaded rcds are

extended that matt with a turnbuckle. From the opposite end of each

turnbuckle is another threaded rod swagged to a long cable. Th y two 3.2 mm

(1/8 in.) diameter wire rope ( 7 x 19 strands) from each turnbuckle traverse

laterally until they engage a pulley aligned with the axes of the longerone.

The cables wrap around the pulleys 90 degrees and extend forward where they

enter the longerone located at the four corners of the trues. The cables

continue forward through the longerone of all ten bays. The cables exit the

longerone if Bay 1 and engage fairleads mounted within the adapter. These

fairleads are canted in such a way that the cables continue toward the

geometric center of the adapter within its diagonal braces. About midway to

the center of the adapter, swagged balls on the cables attach to fittings

whose mounting locations are adjustable within the adapter.

r The bungees are supported on the rear of the housing by two pairs of

} bracke\ (Figure 1.3 - 15) that partially encircle the cylindrical bodies

(6061-T ) and still allow the body to move along its axis as the turnbuckles

E ! are ati zed to pretension the cables to their final 1780 N load.

The cable pretension is accomplished by utilizing a GSE tool (described on

nril the drawings) as follows. Slots located in each leg of the tool ' s yoke that

straddle the turnbuckles are engaged with the flats on the cable terminal and

- the rods extending from either side of the bungee into the turnbuckles. A

4! 25.4 m crescent wrench is simultaneously engaged with the center of the

turnbuckle to turn it to shorten the effective length of the cable and i

wi thereby, compress the spring within the bungee. The same procedure is applied

to the other turnbuckle to totally foreshorten each cable 101.6 m ( 4 in.),

for a total cable pretension of 1780 N ( 400 lb).

1-29

.. u... .. _ ^...

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ORIGINAL PAGE IX

OF POOR QUALITY

177 w---I + 79.1 . 211.2 w— u.s w

7,N s 1.19 w I 7.06 w 1.1 w

UYW1T

111.7 w 201 w

111.27 w CIA

' I

') ', - -^, 171.7. 11

71.92 w

' 70.6 w

10.2 w

31.90 r C^7

17,1 w 7e 6 w

fUl1

112 w $0 17.5 w VIA 1 NOLES 211 w

ON A 200 w CIA I.G.

176 w

4.7 w16 UK 21 16.5 w O7.

1 4xts CO. V.

ON A 119 w CIA B.C.

Figure 1.3-13. 27 Nn, DC Servomotor Dimensions

1N

1.67 w UNC-28

1 IDL[s EO. T ON

A 119 w MOLT CIRCLE CIRC 1.1E 1 78.1 w LOW KIMY

15.9 w CIA

SM TT 111 in CIA

AAeeET 

1!7 11FI , .4

.7 = WT /1 ►E TN

DISPLACES

I5 DE

r

-121

AS

17.5 w 1^J7e w-----«^

Figure 1.3-24. 3.7 Mm DC Servometer Dimensions

1-30
The bungee utilizes a 79.4 mm (3-1/8 in.) G.D. compression spring

fabricated from 8.7 mm (.343 in.) diameter 17-' Ph wire. The compression

spring is stabilized by the inner wall of the bungee cylinder body. The

spring is designed to provide a spring constant of 87.5 N/cm (50 Win.) to

minimize pretension cable load fluctuations that could occur in a space

article. Figure 1.3-16 illustrates the routing of the tension cables between

adjacent longerons through the batten corner fitting guide openings, and the

routing through the longeron rod end transition fittings past t:.e center joint

mechanism. The openings provide adequate clearance for bending of the cable,

to permissible bend radii, during folding of the structure. The openings at

the longeron center joints also permit bending to adequate radii and confine

cable excursions to an acceptO)le envelope.

1.3.1.6 End Adapter Assembly (Drawing 42712-180)

Figure 1.3-17 illustrates the major design features of the end adapter

which is the forward batten for the first bay and has provisions for

attachment to a NASA/MSFC test fixture for stiffness, modal survey, and limit

load strength testing (Figure 1.3-18). The end adapter also is the mounting

for the swagged balls on the four tension cables of the precompression

system. Representative of future flight article needs, during orbiter launch,

the diagonals stabilize the main housings forward frame and vertical beams

provide rigid longitudinal support for the folded longerons and trays.

The end adapters square configuration is constructed from four 50.8 x 76.2

mm (2 x 3 in.) 6061-T6 rectangular tubes mechanically fastened together by

corner jackscrew drogue fittings (View A-A). The drogue is shaped to capture

the ball detente mounted on the end of the batten deployment/ retraction

jackscrew probes. The corner fittings provide the clevis fittings to receive

the rod ends of the diagonals and longerons. The corner fittings also contain

two inserts shown (Figure 1.3-17 - right view) for attachment of eight

brackets that are mounted to the outside face of the main housing forward

frame. The forward face of this corner fitting (View B-B) each contains

threaded inserts for attachment to the NASA/MSFC teat fixture. The inserts

are centered on the projected longeron axis.

A pair of back-to-back 76.2 mm (3 in.) deep extruded channel sections are

provided for each of the adapter diagonals. The spacing between the channels

is used for routing the precompression system tension cables, and provisions

for the adjustable fittings to which the swagged end of the cables are

attached. Each end of the diagonals are attached to the corner fittings by a

pair of shear clips and a pair of gussets as shown in Figure 1.3-17. The same

type of structural attachment is provided at the diagonal crossings at the

vertical beams, and at the inersection of the diagonals.

The square tube 6061-T4 beams are attached to the adapter square frame

with shear clips and gussets.

1.3.1.7 Main Housing Assembly (Drawing 42712-160)

The main housing structure, Figure 1.3-19, is a combination welded,

riveted and bolted assembly into which all other major assemblies are

installed.

1-3i
ORIGINAL PAGE IS

OF POOR QUALITY

M. s OIA (7 % IB) CABLE \

LonDERBN 

CABLES EXTEND FORWARD

LONGERONS

ATTACWW TO AWTER L

1

Figure 1.3-15. Pretension System to Eliminate Joint Slop

LONGERON

CLEAR CABLE

TNRU CENTER

i rBATTEN

C= —

a a

MIEN A-A A

Figure 1.3-16. Tension Cable Routing Through Longerons

I-J2
ORIGINAL PAGE 19

OF POOR QUALITY

7

r ^.

7

CABLE\

^^ 1

WLE 76 r o A• K 4) RETAINER 5,7 (6061-N) ^.

FCC f!'.XAR BEAMS TO ^.

RETAIN TRAYS B

I

i LOIGER9NS

E

JACK

1.5 n I __ •_:r - I %, SCREW

1,5 n

1/2 IN.-20 THREAD INSERT (A)

76.2 so DEEP

BACK TO BACK 

TO SECURE SNIPPING STRAP

MANNELS .

A

ADJUSTMENT HOLES

I

1.5 n RETAMER TO REMGVE f

f CABLE SLACK

76 m DEEP

$0.8 m K 2 m A vn •

CHANNELS 6061-T6

WALL AECTk-4, 13 *IBE. °

CABLE GUIDE (6066-T52) "^ B r

° ° • / 7.2 m DIA PRETENSIONED

1/2 IN.-20 THREADED ^ ^./ VIEW A-A CABLE

INSERT (8) FOR JACKSCREW

INTERFACE SUPPORT _

V IEW B-B B OROGUE FITTING

Figure 1.3-17. Adapter Assembly

i
01tGINAL PAGE 1a

OF PO()R QUALITY

DRIVE

MOTOR

SUPPORT

FRAME

_

ca.... ^^'r

ATTACH F ITTING

TO

J ^A

' Z.O s ALUMINUM

SKIN PANELS

ALL SIDES

ACCESS

K^

' PANELS " /STIFFENERS TO

^_.,. PROVIDE MINIMUM

FREQUENCY TO 10 H1

`/

/ I HOUSING

MAJOR

DETENT

C

/I FRAMING

AOJUSTMEHT

F' HOLES (TYP) _ )`

Figure 1.3-19. Main Housing Structure

1-34

^LF

-. .: ...

.. . S'SrvET^.a..i.... ,.. _id.....r: s.^r._,._^... .3'm;...«..^..,....:
^ E

A weld assembly of three frames and connecting stringers form the housing

framing structure. The frames are made of 50.8 mm (2 in.) square, 6061-T6

aluminum tubing. The corners are braced with a 9.5 mm (.375 in.) thick

6061-T6 aluminum gussets welded in place. The forward and center frames

provide the mounting for the rails of the four batten jackscrew assemblies,

and the four diagonal unlocking and four longeron unlocking assemblies. Two

insert mounts welded in each external frame side provide the 1/2-20 threaded

inaert pattern for payload carrier mounting. The center frame has two 76.2 mm

(3 in.) square tube 6061-T4 aluminum members to accept longeron and utility

tray longitudinal loads when stowed. The center frame also has a mounting

otructure for the precompression system. The aft frame provides mounting

structure for the three drive motors and three chain drive idlers.

The center frame is closed out with two 2.3 mm (.090 in.) thick 6061-T6

aluminum sheets riveted in place. The sides of the main housing and aft drive

motor frame are covered with 2.3 mm thick 6061-T6 aluminum skin panels that

are bolted in place. All large skin panels have 44.5 x 44.5 x 3.2 mm 6061-T6

aluminum angle stiffners spot-welded in place to provide a minimum modal

frequency no less than 10 Hz (Volume III). Access holes are provided in the

skin for batten detest adjustment and jackscrLe carriage release.

The 1/2-20 threaded insert pattern on each side of the main frames are

also used for accepting eye bolts to attach support structures during ground

handling and deployment/ retraction operations. The forward inserts are used

to secure the adapter to the housing during shipping.

f"

1.3.1.8 Corrosion Protection

Corrosion protection on all components is in accordance with the standard

industry practice, i.e., anodize for aluminum, cad plating for low carbon

steel parts, TFE and dri Tube for sliding and rotating surfaces.

1.3.1.9 Dimensional Accuracy

Dimensional accuracies with regard to detail fabrication are in accordance

with specification ANS1 Y14.5- 1973. In additicn, adjustments have been

designed into minor and major assemblies via turnbuckles and shims.

1.3.2 Ground Test Article Deviations from Prototype

The ground test article design is " fabrication ready" and, hence, clearly

defined. The prototype building -block design that is represented by the

ground test article has not been designed in detail, and is not as clearly

understood. However, to aid reviewers interested in future space applications

the known ground teat article deviations from the protype design are

delineated as follows:

o Adapter-to-Housing Retention Mechanism - The test article is preaeutly

secured in the stowed position by four S-fittings connecting the end

adapter to the main housing forward frame. Unlocking is done manually

by removLi of the four fittings. A prototype design will require

substitution of a remotely controlled latch system.

I-35 
o Batten Jackscrew Carriage Locking Pins Assembly - Prior to retraction

of Bay 1, the current design requires manual release of each of the

four batten deployment / retraction jackscrew carriages locking

assemblies. In the prototype design a remote motorized system is

required.

o Jackscrew Support Frame Diagonal Unlocking - Prior to retraction of Bay

1 the test article requires manual release of each of the four telescoping diagonal latches that position and support the jackscrew

support frame assembly. For the prototype design a remotely operated

unlocking system will be required.

o Positioning System Motors, Tachometers, Encoders - The test article

servo motors, encoders, and tachometers are heavy -duty commercial

designs. The prototype design will require flight -qualified designs,

i.e, thermal, vacuum, radiation, Shuttle launch.

o Precom esio + stem Cable Tensioning - The test article tension is

applied 

re

removed manually with GSE wrenches. For the prototype design

the tension will be applied / removed by a remotely controlled motor.

o Motor Redundancy - The test article contains no redundancy. For the

prototype, redundancy is required for all the motorized components.

• Deployable Curtain - Reliable deployment / retraction operations on-orbit

may require a curtain around the four open fEces of Bay 1 (Figure

1.2-3) The curtain can minimize fluctuation in the temperatures of the

batten deployment / retraction jackscrews and in the longerons and 

diagonals during deployment, and with time permitting, during

retraction. The curtain can be foldable in the same manner as that

shown in Figure 2.1-7 for the Concept 4 OTV hangar. The need for a

curtain depends on the range of temperatures that would occur in the

affected members during deployment without it. The desirability of a

coarser ACME thread in the jackscrew which would provide an increased

tolerance for thermal induced length changes is also dependent on

thermal analysis results (Section 1 . 5.2.5).

Depending on the particular future system-impoaed requirements, the

following design modifications may be necessary for a prototype design.

• Main Rousing and End Adapters - The ground teat article design may be

modified fs^ • the unique end attachment applications, and stowage in the

orbiter.

• Mate Critical Designs - A mass critical design would require modifications such as reduced housing akin gauges, more extensive machining

of the batten deployment / retraction rails, and reduced gauges on the

longerons, diagonals, battens, jackscrew support frames, and end

adapter.

• Stringent Figure Control Design - Stringent dimensional stability

requirements may require the use of low coefficient of thermal

expansion ( CTE) longerons, diagonals, and battens. Such a design in

conjunction with an invar deployment / retraction jackscrew and a low CTE

main housing would preclude the need for a curtain or the concerns of

Section 1 . 5.2.5.

1-36
Finally, thermal coatings may be required on all non-rubbing surfaces of

the truss structure.

1.3.3 Assembly Description

The assembly of the ground test article (Figure 1.3-20) is to be

accomplished in the following manner:

1. Support the completed main housing assembly and orient as shown on

Figure 1.3-21 and remove the aft skins. The aft face of the housing is

to be approximately 1.5 m above the floor. (There are 16 handling lugs

that will screw into the housing payload/carrier attachment inserts.)

2. Install the four batten deployment/retraction jackscrew assemblies and

position (with shims) to satisfy the dimensional requirements shown on

the top assembly drawing.

3. Install the four loogeron and four diagonal unlocking jackscrew

assemblies and position (with shims) to satisfy dimensional requirements shown on the top assembly drawing.

4. Install the jackscrew support frame assembly such that the underside of

w the frame rests on the housing and attach the four telescoped diagonals

to the clevices on the housing and frame. This represents the stowed

condition of the jackscrew support frame assembly.

5. Manually extend together the jackscrew support frame assembly and four

batten deployment/ retraction jackscrews and carriage to the deployed

configuration shown on the top assembly drawing. Secure the four

carriages to the housing by installing the four carriage locking pins

through each rail into each carriage. Adjust the turnbuckle on each of

the four diagonal tubes to the required dimensions shown on the top

assembly to achieve squareness on all planes.

6. Support (as required) the jackscrew support frame assembly to permit

removal of the four carriage locking pins and remove the assembly

together with the jackscrews, splined shafts and carriages. During

removal, protect flexible parts with wood bracing or equivalent.

Install the seven typical battens and two payload carrier support

battens in the proper sequence into the housing, one on top of the

other. (Batten 9 is supported by the four clevis fittings attached to

:he batten deployment/retraction jackscrew rails.) Partially lower the

issembly removed in Item 6 into place until the aft ends of the four

)attec, deployment/retraction jackscrew threads contact the forward

.aces of the four Batten 1 half-nuts. Orient the lead thread on each

if the four batten jackscrews with its corresponding lead thread on the

:our batten half-nuts. Lower the entire assembly by simultaneously and

ianualiy rotating the four jackscrews clockwise until the holes in the

:arriage line up with the corresponding holes in the rails. Reinstall

:he four carriage locking pins.

1-37

7
E

m

ORKMNAL PAGE 18

OF POOR QUALITY

MY a SAY

NNSINS) T 7

^ur e

MrNi

6^"^ " BATTEN

L BATTEN I

^-BATTEN Z

BATTEN

BATTEN 1

BATTEN B

6

B TYPICAL DEPLOYABLE TRnS SAYS 6

BATTEN B

I ..JA MY (MY 6)

w.

MgIN HWING CON AMINO

^,. MECHANISMS

r Figure 1.3-10. Test Article Bay and Batten Identification

SAY 16

(INSIDE

'• a

^ 3

^

J^

^^.eAr 1 Ir

`

6
r

8. Install the drive sprockets, and chain assemblies and shaft collar;, as

shown on the top assembly drawing. Pretension the chains as specified.

While installing and pretensioning the chains be sure to maintain the

thread indexing of the batten jackscrews with the Batten 1 half-nuts.

Upon completion of chain pretensioning, tighten the setscrews that

secure the sprockets to the splined shafts.

9. Install the end adapter on the four jackscrew probes.

10. Install all four Bay 1 longerons and diagonals to complete Bay 1. Any

length adjustment of the longerons is to be shared equally between each

of the two turnbuckles per longeron.

11. Provide external vertical support that will permit retraction of Bay

1. Remove the four carriage locking pins and manually unlock the four

jackscrew support frame diagonals, and the four longeron and four

diagonal latches. (The sprockets have spanner wrench holes. A spanner

wrench detail is provide on the top assembly drawings.) Use the

spanner wrench to manually rotate the batten jackscrews until the

jackscrew support frame rests on the main housing. Deploy and retract

Bay 1 until trouble-free operation is achieved.

12. Using the clanner wrench, manually redeploy Bay 1. Continue

counterclockwise revolutions until the forward face of the Batten 1

half-nuts are flush with the aft face of the jackscrew support frame.

Position Batten 2 so that the forward faces of its half-nuts are

contacting and radially aligned with the aft face of the batten

jackscrew threads. Manually rotate the batten jackscrews countercockwise eight revolutions to obtain a 50.E' mm (2 in.) engagement of

the Batten 2 half-nuts.

13. Install all Bay 2 longerons and diagonals. As before, any length

adjustment of the longerons is to be equally shared by each of the twc

turnbuckles on each longeron.

14. Manually unlock all four diagonals and longerons of Bay 2. Retract and

deploy Bay 2 until trouble-free operation is achieved.

15. Install Bays 3 and 4 longerons and diagonals in the same manner as Bay

2 until trouble-free operation is achieved.,

16. Redeploy Bay 4. Continue manual counterclockwise rotation of the

batten jackscrews until the forward faces of the Batten 4 half-nuts are

25.4 mm (1 in.) below the aft face of the jackscrew support frame.

Position Batten 5 (payload carrier support frame) so that the forward

faces of its half-nuts are contacting and radially aligned with the

jackscrew threads. Manually rotate the jackscrew 12 revolutions

counterclockwise.

17. Install the Bay 5 longerons and diagonals and adjust as described in

Item 13.

18. Repeat Item 14 for Bay S.

1-39

:dT'iR^JT . •

Tip
19. Redeploy Bay 5. Continue manual counterclockwise rotation until the

forward face of the Batten 5 half-nuts are 254 mm (10 in) below the aft

face of the jackscrew support frame. Position Batten 6 (payload

carrier support frame) so that the forward faces of its half-nuts are

contacting and radially aligned with the jackscrew threads. Manually

rotate the jackscrews 32 revolutions counterclockwise which is a

203.2 mm (8 in.) travel.

20. Install the Bay 6 longerons and diagonals and adjust as in Item 13.,

21. Repeat Item 14 for Bay 6.

22. Redeploy Bay 6. Continue manual counterclockwise rotation until the

forward faces of the Batten 6 half-nuts are 25.4 mm above the aft face

of the jackscrew support frame. Position Batten 7 so that the forward

faces of its half-nuts are contacting and radially aligned with the

jackscrew threads. Manually rotate the jackscrews 8 revolutions

counterclockwise.

23. Install the Bay 7 longerons and diagonals and edjust as in Item 13.

24. Repeat Item 14 for Bay 7.

25. Install Bays 8 and 9 in the same manner as Bay 2.

26. Redeploy Bay 9. Continue deployment until the forward faces of the

Batten 9 half-nuts are 1.35 m (53 in.) from the forward face o£ the

rail attach fittings.

27. Perform Item 13 for Bay 10.

28. Perform Item 14 for Bay 10.

29. Manually retract Bay 10 through Bay 1 and record the relative longitudinal and lateral contact points for each longeron and diagonal

retraction probe in each bay. Adjust each of the eight probes for best

compromise. Redeploy and retract to verify suitability of probe

adjustment.

30. Redeploy the entire structure Bays 1 through 10.

31. Manually and concurrently, using spanner wrench, rotate clockwise ten

revolutions the diagonal and longeron retraction jackscrews, thereby

unlocking the latches in each longeron and diagonal for Bay 10.

Manually rotate the batten deployment/retraction system jackscrews

clockwise to retract Bay 10.

32. Manually and concurrently rotate the unlocking system clockwise 20

revolutions, thereby unlocking Bay 9. Retract Bay 9.

33. Repeat Item 32 for Bays 8 and 7, except prior to retraction of Bay 7; 

manually and concurrently rotate the unlocking system counterclockwise

five revolutions.

1-40
1

r ^'

34. Manually and concurrently rotate the unlocking system five revolutions

clockwise, thereby tripping the longerons and diagonals of Bay 6.

Retract Bay 6.

35. Rotate the unlocking system 60 revolutions clockwise, thereby unlocking

the Bay 5 diagonals and longerons. Retract Bay S.

36. Repeat same procedure for Bays 4, 3, and 2 as for Bays 9 and 8, ending

with retraction of Bay 2.

31. Rotate unlocking ayste: a 20 revolutions clockwise, thereby unlocking the

Bay 1 longerons and diagonals. Manually unlock the four jackscrew

support frame diagonals and pull the four carriage locking pins.

Retract Bay 1 through 32 counterclockwise revolutions of the jackscrew

systems. Retract the unlocking system by rotating the jackscrew system

through 210 revolutions counterclockwise. The retraction probes should

be in the initial stowed position. Complete the retraction of Bay 1 by

rotating the unlocking assembly jackscrews counterclockwise through an

additional 164 revolutions.

38. Deploy Bays 1 through 10 and install the precompression system as shown

on the top assembly drawing.

39. Manually deploy and retract all bays to assure tension system cables do

not interfere with deployment and retraction.

40. Deploy Bay 1. At Batten 2, using GSE wrench No. VW62 inserted through

detent access holes in the housing skin, advance all eight detents (two

at each batten corner half -nut) until contact with the half-nut is

made. Note the depth of insertion of the wrench. Advance all eight

detente 1 . 52 to 1.65 mm (0.060 - 0.065 in.). This position of the

detent will impart 6 total of approximately 445N ( 100 lb) force

resistance to the batten before the detents are compressed, thereby

allowing the batten to move. If this amount of force appcars to be too

little or too great, the detente can be readjusted until the desired

batten resistance is achieved.

41. Install the three drive motor assemblies.

42. perform positioning system motor and computer adjustments.

43. Upon completion of Item 42, deploy Bays 1 through 5 and install the

tray assemblies in Bays 4 and 5, respectively.

44. Retract and deploy Bays 4 and 5 to assure trouble - free operation.

Retract the remaining structure and install the adapter/housing four

"S" fittings.

The following further defines the assembly operations requirments:

(a) During deployment ani retraction the trues structure weight

exerted upon the jrckscrews shall be counterbalanced to within +

the weight of one bay. Appropriate lateral support safety guides

shall be provided.

1-41

p

Q^
U

(b) The assembly of the test article is to be performed in an environment in which temperatures are controlled to 27 + 10°C (75` +

18°F).

(c) The manual deployment and retraction operations will be performed

by applying torque to a spanner wrench that attaches to any one of

the four batten deployment/retraction jackscrew sprockets. A

torque wrench attached to the end of the spanner wrench will

permit measurement of the system's required torques prior to

Installation of the positioning system.

In the event the vertical deployment is restricted such as shown in Figure

1.3-22, the following assembly variations relative to the foregoing are

imposed.

(a) Pertinent to operations 15 through 37, the upper bays can be

manually folded, provided the aforementioned counterbalancing

requirement is satisfied during deployment and retraction,

adequate support is provided for the bay being manually folded,

and upon retraction the folded bays are manually redeployed prior

to retraction with adequate support provided.

(b) Folding will be accomplished by manual tripping of the diagonal

and longeron latches, with latch retention maintained by a

suitable GSE device that will free the technicians for other tasks.

38. Deploy Bay3 1 through 5 and install the tray asazmblles in Bays 4 and

5, respectively.

39. Retract and deploy Bays 4 and 5 to assure trouble-free operation.

40. Perform item 40 as previously described pertaining to adjustment of the

detente.

41. Reorient the stowed configuration and manually deploy Bays 1 through 10

to the configuration shown in Figure 1.3-23.

42. Install the precompression system and manually deploy and retract all

bays to assure tension system cables do not interfere with deployment

and retraction.

43. Install the three drive motor assemblies. Reorient the configuration

to that shown in Figure 1.3-21.

44. Perform the positioning system motor and computer adjustments. Upon

completion retract the remaining structure and install the

adapter/housing four "S" fittings.

1-42

i y
I^, ^ t ti

OF POOR 

PAGE

POOR Q

UAD v

Figure 1.3-22. Test Article Assembly 

with Headroom Configuration

Restructions

VIEW A-A

Figure 1.3-23. 

Precompression System 

Installation .on configuration

t-43 

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TNNEAOEo

Noo

I , Z x 1 INCH

Moon
A

,

a

1.3.4 De p loyment/Retraction Operations

This section provides a detailed description of the means by which the

previously described design accomplishes controlled, automatic, bay-by-bay

deployment and retraction. While much of the forthcoming figures illustrate

the test article in a horizonal attitude for drawing convenience, the i

deployment and retraction are most easily accomplished in the vertical mode as

suggested in Figure 1.3-24.

1.3.4.1 Stowed Configuration

Figure 1.3-25 illustrates a section through the test article in the stowed

position. The adapter is nested to the jackscrew support frame which is

captured between the adapter and housing. The adapter is fastened to the

housing with eight eye-bolts into the four "S" fittings. Battens 1 through 9

are captured (Figure 1.3-10) between the four rails as shown at the right,

with each of the four half-nut threads engaged with each of the four stowed

jackscrews. Battens 2 through 9 (all except Batten 1), have the eight detents

per batten engaged. Finally, Battens 1 through 9 are positioned fore and aft

such that they bear against each other, with Batten 9 bearing against the

truss attach fitting.

1.3.4.2 Deployment of Bay 1

The deployment operation starts with removal of the eight eye-bolts and

the start of clockwise rotation of the drive motor that rotates the four

batten deployment/retraction system spline shafts which rotate the threaded

jackscrews. The kinematics is best explained by the analogy of a bolt and nut

(Figure 1.3-26). Clockwise rotation of the bolt with the nut constrained from

moving aft will result in forward motion of the bolt. Since the batten

half-nuts are constrained from moving aft the jackscrew and attached carriage

are pushed outward, i.e., to the left (Figure 1.3-25). Continued rotation

will continue to drive the jackscrew and carriage to the left. At 196 + .001

revolutions, the carriage locking pin, for each of the four carriages will be

automatically engaged locking the carriages fore and aft. Concurrently, the

jackscrew support frame and adapter have been driven forward. The forward

motion of the adapter unfolds the four longerons and extends the four

telescoping diagonals attached to it. The forward motion of the jackscrew

support frame assembly extends the four telescoping tubes attached to it. At

carriage locking, through the adjustments described in the assembly

operations, the 12 members previously described will all be locked. The

configuration is as shown in Figure 1.3-27 with Bay 1 developed, and the

jackscrew support frame assembly rigidized. It is pertinent to note that the

jackscrew threaded length is such that the Batten 1 half-nuts are on the aft

101.6 mm (4 in.) of the jackscrews. This phased development is noted as (1)

on Figure 1.3-31.

t

1-44
y

ORIGINAL PAGE 1.9

OF POOR QUALITY

Figure 1.3-24. Deployment/Retraction

Operations Configuration

Figure 1.3-25. Test Article Packaged Configuration

1-45

^^ . .^s}4xfsf^..1., ♦. ,... ..fiLL.. .^.ffcti
0RUNAL PAGE 19

OF POOR QUAUTY,

FORWARD

Figure 1.3-26. Bolt/Nut Analogy at Batten

Deployment/Retraction System

JACKSCREW SUPPORT FRAME

CARRIAGE LOCK

FIRST

SECOND

JACKSCREW ^ COAFIGURATION AT COMPLETION OF FIRST BAY

SUPPORT FRAME

Figure 1.3-2:. ..a-.figur^tion at

Deployment of Bay 1

1-46

..J. s

.. .. • .f ,,...wa^.n n Au3^.,I.! 
L.3.4.6 Precompression System Operations

Upon completion of deployment the precompression system can be manually

activated to apply and deactivated to remove the precompression loading.

1-47

1.3.4.3 Deployment of Bay 2

The analogy of the bolt and nut can also illustrate the deployment of Bay

' 2. For this stage of deployment the analogy consists of the boit restrained

longitudinally. A counterclockwise rotation of the bolt will advance the nut

forward. The carriage locks restrain the jackscr,twe; however, there is no

constraint against forward moLion of Batten 1 which is engaged with the

jackscrew. Batten 2 is held in place by the eight detente. Deployment of Bay

2 commences, therefore, by the commuter commanded drive motor rotating the

four jackscrews counterclockwise. This Totation and the engagement with

Batten 1 drive Bay 1 forward overwhelming the ball detente (on the drogue)

that previously restrained the adapter. Batten 2, however, is held in place

by the detente. The rotation is continued until 196 + .001 revolutions are

counted by the encoder. It is pertinent to note thee the encoder counting

accuracy of + .001 revolution represents a longitudinal positioning accuracy

1 to .0064 mm T.00025 in). At completion of ;.he 196 + .001 revolutions •'s

longeron.s and diagonals are locker' completing the deployment of Bay - to the

configuratioa shown in Figure 1.3-18. This phase of deployment is represented

as (2) in Figure 1.3-31. In this configuration Batten 1 is fully engaged with

101.6 mm (4 in) of the jackscrew forward end. Specifically, the forward face

of the half-nuts is at the forwara face of the jackscrew thread. The forward

: w face of the four half-nuts on Batten 2 is at the aft edge of the jackscrew.

1.3.4.4 Deployment of Bays 3 and 4

i The deployment of Bay 3 starts with the counterclockwise rotation of the

four Batten deployment/retraction jackscrews .

Thro.: 7 ' engagement of Batten 1

with the jackscrew, as in Bay 2, the rotation drives the completed Bay 2

forward by overwhelming the eight detente restraining Batten 2. The total

restraining force (longitudinal component) of the detente can be set to vary

r. from 180 to 1780 N. Since the half-nuts are 101.6 mm (4 in) wide during the

first 16 revolutions both Batten 1 and 2 half-nuts are engaged with the

jackscrew. After 16 revolutions only Batten 2 is engaged. Upon completion of

the 196 + .001 revolutions Bay 3 is complete and is designated by (3) on

Figure 1.3-31. At completion the Batten 2 and Batten 3 (held by detents) are

located as shown in Figure 1.3-32

y, 1.3.4.5 Deployment of Bays 4 through 10

The deployment of Bays 4 through 10 is accomplished in the same manner

except that the completed bay is located further and further aft, i.e., closer

to the aft attach fitting. This is because the aft Batten at any completed

bay is located as shown in Figure 1.3-33. The only other exception is that

the carrier bay being 1.09 m (43 in), i.e., center to center of the Battens

and containing 151.4 mm (6 in) half-nuts needs to advance only .94 m (37 in)

to deploy. Hence, the number of revolutions shown on Figure 1.3-31 for phase

(6) is 146 revolutions.
i

1.3.4.7 Retraction of Bay 10

The retraction operation starts with activation of each of the diagonal

and longeron unlocking systems (Figure 1.3-29). Since the tripping devices

are set 25.4 mm aft of the bay longeron and diagonal latch faces, rotation of

the jackscrews with a pitch of 5.08 mm (.2U in.) advances the probes to

contact the latches. Continued rotation advances the tripping devices which

rotate the latch mechanisms that depress the locking pins. At the proper

phasing (to be determined during the ground test deployment/ retraction

operations) Lhe batten deployment/retraction jackscrews will be rotated

clockwise (Figure 1.3-31). The rotation through engagement of the four

half-nuts of Batten 9 will drive Batten 9 aft towards the housing. In the

initial stages the batten closure further depresses the locking pins. The

rotation continues until Batten 9 is driven into the housing and over the

detents and repositioned. The spring activated detente again retain the

batten. At 225 rpm the retraction of Bay 10 will occur in approximately 52

seconds. During the first 16 seconds (approximately) the unlocking systems

continue to advance forward until 20 revolutions are completed and the

tripping devices are 25.4 mm from the face of the Bay 9 longeron and diagonal

r.. latch mechanisms.

1.3.4.8 Retraction of Bays 9 and 8

The retraction of Bays 9 and 8 is accomplished in the same manner as

Bay 10.

1.3.4.9 Retraction of Bays 7 and 6

The retraction of Bay 7 is the same as Bays 8, 9, and 10, except as shown

in Figure 1.3-31, the unlocking systems are returned to the same position as

at the start of retraction. This is due to the difference in length of the

next bay, i.e., the carrier bay, and the Batten 6 half-nut length of 157.4 mm

(6 in). This is illustrated by Figure 1.3-34, and the following calculations

for As which represents the required advancement of the tripping devices.

The advancement from the start of Bay 7 to the start of Bay 6 retraction is

determined from the equations:

H6 H7 R7 R6

As- 2 +2 +2 -2

p a a 50.8 + 76.8 + 495.3 - 622.9 - 0

The advancement from the start of the Bay 6 retraction to the start of the

Bay 5 retraction is similarly determined as:

Arr - 76.8 + 76.8 + 622.9 - 495.3 - 281.2 mm (11 in)

The 55 revolutions shown in Figure 1.3-31 accomplishes the required

281.2 mm (11 in) of travel.

1.3.4.10 Retraction of Bay 5, 4, 3 and 2

The retraction of these bays is accomplished in the same manner as that

described for bay 10.

1-48 
j

w

ORIGINAL PAGE 19

OF POOR QUALITY

Figure 1.3-28. Configuration at

Deployment of Bays 1 and 2
URIW14AL PAGE IS

OF POOR QUALITY

Figure 1.3-30. Configuration at

Start of Retr^!Z tion of Bay 1

196 REV

MOTOR DIRECTION 196 REVI I I O DENOTES BAY (TYP EXCEPT CARRIER NUMBER CLOCKWISE 1 "AS MWED) r BAY 156 REV

COUNTERCLOCKWISE O 03 O ( O7 O Oq 10

SEQUENCE 0 1 2 3 0 5 6 7 8 9 10

DEPLOYMENT PHASE -- BATTEN SYSTEM

25 REV

20 REV

'^ h"""I r'TYV. 10 REV 5 REV

CLOCKWISE

21

r`--^

COUNTERCLOCKWISE ^ 160 REV (,

I I I I I I Iu„ DENOTES I I I I I I I I L^J 

RETRACTION PHASE -- DIAGONAULONGERON UNLOCKING SYSTEM UNLOCKING

CE LONGERON

6 8 DIAGONALS

' CLOCKWISE 10 O9 O (5 O Oj O+160 REV

yy' i COUNTERCLOCKWISE pI 1► 156 REV

J2 REV-I k^7h6 1196 REV (TYP EXCEPT AS NOTED) I I I 11

SEQUENCE 0 1 2 3 A 5 6 7 8 9 196 REV

E ” RETRACTION PHASE -- BATTEN SYSTEM

to

b Figure 1.3-31. Positioning System Motion Profile

1-50
w

ORIGINAL PAGE IS

OF POOR QUALITY

Figure 1.3-32. Batten 2 and 3 Positions

at End of Bay 1 Deployment

Figure 1.3-33. Location of Battens 1 thzough i 

1-51
^ +

UMUINAL PAGE 19

OF POOR QUAUTY
,r

1.3.4.11 Retraction of Bay 1

The retraction of Bay 1 is unlike that of the other bays. The retraction

starts with manual unlocking of the four carriage locking pins and four

jackscrew support frame diagonals and advancement of the unlocking systems to

unlock the Bay 1 four longerons and four diagonals. At the proper phasing (to

be determined in the ground tests) the deployment/retraction batten system

motors are rotated counterclockwise 32 revolutions thereby retracting Bay 1

103.2 mm (8.0 in). Any further retraction is delayed until the extended

unlocking system carriages are retracted back into the housing (210

revolutions). At the appropriate phase the remaining 164 batten system

jackscrew revolutions are accomplished thereby completing the retraction of

the structure. At this point insertion of the eight eye-bolts can reattach

the adapter frame to the housing.

1.4 TEST ARTICLE MASS PROPERTIES

To support the planning of a future ground test program and future space

applications studies a mass properties analysis was performed. The results of

the analysis are summarized in Table 1.4-1. The masses shown were determined

by detailed mass properties analyses of all the drawings, except for the

adapter, utilities tray, and jackscrew support frame assembly which were

estimated from the original layouts. The values shown represent the

calculated masses with no allowance for tolerances.

It is also pertinent to note that minimum cost of the test article was

regarded as more significant than minimum mass, provided there was no

compromise in the repreL atation of concept performance. For example, the

four rails in the batten deployment/retraction system are heavier than

necessary to minimize ;machining cost without any deleterious impact on

performance. The same is true for the main housing structure which is

fabricated with 2.3 mm skius to minimize the addition of stringers for panel

stiffness.

Table 1.4-1 also illustrates the potential mass reduction associated with

removal of the systems required for retraction. Figure 1.4-1 illustrates the

scope of those reductions.

1.5 STRUCTURAL ANALYSIS

The structural analyses performed during the Part 2 phase of this contract

directed the subject ground test article's conceptual and detailed design.

This section describes the structural requirements concept reviews, and an

overview of the major analyses performed to assure design compliance with the

requirements. A copy of the detailed structural analyses, performed to

support the drawing production, is contained in the Volume III Appendix.

1.5.1 Requirements

Figure 1.5-1 illustrates the major limit-loads location and orientation

that established the design internal loads. These are the loads that will be

applied to the ground test article at the conclusion of deployment/retraction,

stiffness, and modal survey tests.

1-53

.. _...,u •fin '._.^_. ...,A ." N '...J.. ,.: ♦. a ^. a.
iby

w.GERON

UNLOCKING

SYSTEM

REMOVABLE SYSTEM

ORIGINAL PAGE IS

OF POOR QUALRY

Table 1.4-1. Test Article Predicted Ales (kg)

COWON14T NO. OF PMTS TOTAL MASS

MASS WITHOUT

RETRACTION SYSTEM

KPLOYANLE TRUSS 1 791 719

• LONGEAONS 10 178 111

• OIAGONALS N I37 122

•BATTENS 9 63 87

-CARRIER 2 70 70

ADAPTER • 1 17 17

DEPLOYMENT/RETRACTION 1 I85 185

SYSTEM

JACKSCREW SUPPORT FRAME 1 26 26

ASSEEM9LY

WAIN HOUSING 6 EQUIPMENT 1 715 277

• STRUCTURE 1 200 199

-UNLOCKING SYSTEMS 8 SS 0

PRETENSION SYSTEM 1 14 11

• 7IOTORS 6 EQUIPMENT 7 16 20

UTILITIES TRAYS* 21 21

• CARRIER 1 10 10

-TYPICAL 1 11 11

TOTAL 979 861

DIAGONAL UNLOCKING SYSTEM

G i

1, r.

SHADED AREA DENOTES DELETEU PARTS

ASSOCIATED MOTORS, ENCODERS, TACHOMETERS, W^p • „^^,,, '^ ° •"1p i"^ .^..c w.

MICROMOTION MOUNTS, CHAIN 6 SPROCKETS '„^.^^ nwr ..r.

NOT REQUIRED

Figure 1.4-1. Retraction System Equipment Deletion

for Design Without Automatic Retraction

1-54
The significant structural requirements are listed as follows:

o The structure shall sustain without detrimental deformation:

(1) The loads occurring during deployment in the vertical or

horizontal mode (with appropriate supports during deployment).

(2) The worst combination of the internal Loads derived from the

loading shown in Figure 1.5-1 in conjunction with limit

precompression system loads.

(3) The placement of the 3636 kg simulated payload carriers supported

as shown in Figure 1.3-18.

(4) A 6g acceleration applied to the stowed configuration along any of

the three major axes (one axis at a time). This is representative

of the Shuttle quasi-static launch load environment.

o The structure shall sustain the above conditions upon application of

ultimate load. The ultimate loading is the above described limit

loading mutiplied by a factor of 1.5.

o Provide a minimum natural frequency, in the stowed configuration, of 10

Hz. This requirement is to provide frequency separation with the

orbiter to minimize quasi-static loads to no more than 6 g.

o Provide a design minimum natural frequency, in the deployed

configuration shown in Figure 1.3-18, of .10 Hz. This requirement is

applicable to the configuration with the precompresaion system cables

tensioned to 1780 N.

o Perform trouble free deployment/retraction operations in a controlled

environment, i.e., room temperature + (10°C)

1.5.2 Concept Reviews

The significant concept structural analysis reviews that directed the

design to its present form are described herein.

1.5.2.1 Truss Configuration

The selected truss member arrangement was determined from consideration of

the two major options shown in Figure 1.5-2. Configuration (a) represents a

design in which the member arrangement permits all the joints to be

identical. This configuration also has all the diagonals oriented in the same

sense with regard to the approach of the diagonal unlocking tripping devices.

These two advantages, however, were leas significant than the advantages of

the selected configuration (b). The most significant advantage, and alone

sufficient to favor selection of configuration (b), is that precompresaion of

the longerona, by the prer_ompression system, will also induce compression in

the diagonals. In configuration (a) there can be no precompresaion of the

diagonals since it would create an unbalanced torque. In this case the

shortening of the longerona is accommodated by torsional rotation of the

batten frames. The selected configuration also experiences reduced longeron

loads due to torsional moment. For the applied limit loads shown in Figure

1.5-1, the total longeron load is reduced by 17%.

1-55

V

• ^k (^

1` / t
Figure 1.5-1. Major Structural Requirements

10

:

UnINurAWn 1b)

r Co or seIg 'lerrdt 'eld0"

ORIGINAL PAGE 18

OF POOR QUALf7y,

s.

Figure 1.5-2. Candidate Truss Arrangements

1-56

Ll ^--w
f

With the selected configuration no loads are imposed on the batten members

by the diagonals, which, in conjunction with the reduced longeron loads,

provide increased torsional stiffness and hence a higher torsional frequency.

The arrangement and configuration of the diagonals within the battens were

also studied, particularly in relation to the development of precompression of

the major truss diagonals. A pair of X-braced tension cables, as configured

in the Part 1 selected design, does not provide the required AE to develop

sufficient precompression in the truss major diagonals. The phenomenon is.

illustrated in Figure 1.5-3. The development of precompression in the

diagonals is dependent on the extecsional stiffness of the batten since any

induced compression in the diagonals would have the resultant loads shown in

Section A-A. The most direct approach is to place the diagonal along the ling

of the resultant load, thereby negating the flexibility of members 1-2, 2-3,

3-4, and 4-1. Hence, placement of the batten diagonal ;Tong the lines of the

resultant force is the selected design for each batten frame. The sensitivity

of the developed diagonal compression force to variations in the design are

shown in the table of Figure 1.5-3.

1.5.2.2 Precompression System Concept

A precompression system utilizing a pair of bungeea and tension cables was

chosen as the method to remove joint slop because it is current technology,

and, hence compatible with the goal of FY 1986 technology readiness. The

removal of joint slop is required not only to preclude a loss of stiffneas but

also, and most important, because a structure with slop is dynamically

unpredictabi.e. Another significant advantage of the precompression system is

that it permits joint slop during deployment. This is very advantageous to

the in-space deployment of the prototype design, which is indeterminate, since

it can accommodate the member length changes (graphite composite members)

during deployment. An indeterminate structure with no joint slop during

deployment, experiencing thermal-induced length changes, will not deploy

unless adequate force is used to accommodate the dimensional changes by

internal load induced deformation. A determinate structure can deploy with

tight joints, provided there is no binding, without any forces required.

The precompression system is designed to induce and maintain sufficient

precompression in the longerons and diagonals, throughout the application of

control system loadings, in systems for which precise control is required.

Rockwell's large space structures studies (References 2 and 3) have indicated

these forces generally result in longeron and diagonal loadings below 1600 and

200 N respectively. A simple example that illustrates this is that of a

cantilevered square truss 50 m long and 1.25 m deep. RCS thruster forces can

be 4.45 N (1.0 lb) or at most 44.5 N (10 lb). A 44.5 N force shear exerted at

the end of this beam would result in a longeron load of 890 N (neglecting

shear relief due to acceleration of the payload mass). The worst diagonal

forces would be due to torsion. To exceed the 200 N precompression, two equal

and opposing thrusters of 44.5 N would have to be 7.90 m apart. The exception

to the foregoing are loads incurred during hard docking. The control

requirements for hard docking, feasibility of berthing, or attenuation is

beyond the scope of this study. Assuming that the control requirements during

hard docking do not require precompression throughout the structure, a tension

in the precompression system cable of 1780 N was established. This does not

mean that higher values are not possible. However, higher precompression

loads result in higher total longeron and diagonal design loads.

1-57

e t
- w

.,

The system contain: a pair of compression springs that each have a spring

constant of approximately 88 N /cm. The low spring constant will minimize

variations in the precompression of a future flight article despite differential strains between the longeron and cable due to either load, thermal

variations or creep. For example, a JO'C uniform temperature change in an

aluminum longeron (worst case) would cause a total length change of 9.5 mm,

and a load change of 83 N which is a percentage change of 4.6, which is

certainly acceptable.

1.5.2.3 Concept Suitability-Launch Loading

Figure 1.5-4 illustrates the pertinent featu, : es of the prototype design in

the stowed configuration. Any of the three orientations relative to the

orbiter are possible. Referring to the references axes shown in the figure,

the following comments are appropriate to the support of the components for

loads directed along the X-axis.

• For a prototype design the adapter will be latched to the main housing

with latches designed for the peak X -axis loading.

• The jackscrew support frame is ceptured between the adapter and main

housing.

• The screwjacke and carriage are constrained by the pawl (Fig 1.5-4).

• The battens are constrained (forward) by the half-nut engagement with

the screwjack and (aft) by bearing on the truss attachment fittings.

The half-nuts of all the adjacent battens are in contact during launch.

• The longerons are constrained by bearing on each other with transfer of

load to the transverse beams on the adapter (forward) or on the

transverse beams (aft) on the housing.

• With a full complement of utilities trays, the trays are constrained by

bearing on each other with transfer to the same end beams used to

support the longerons.

• The diagonals and battens are constrained by their end clevis

attachments to the half-nuts and each acts as aiuple beams between the

ends.

following comments are appropriate to the support of the components

da directed along the Y and Z axes.

For a prototype design Y and Z directed loads will be transferred from

the adapter to the jackscrew support frame and into the main housing

through bearing on suitable shoulders between the components.

The jackscrews are constrained at their ends by bearing on the adapter

and attachment to the carriage which is laterally supported by the

guide rail. The jackscrew acts an a beam between these two ends.

The battens are constrained by the half-nuts bearing on the faces of

the rails (Figure 1.3-10).

1-58
Ole

Or 1

Ix

V

Y

w

ORIGINAL PAD! 19

OF POOR QUALny

oiAF,^"AL

A-A

MINIMUM DIAGONAL C0WRESSION (N)

OUE TO 1780 N PRETENSION

DIAAGON^L

1.WALL 3 .05LL

AT 1-2 187 N

AT 2-4 258 N 220 N

Figure 1.5-3. Batten Diagonal Load Paths

EXTENSION

^1 .

ADAPTER ROUSING

FRAME

SCREN JACK

SUPPORT FRAME

Figure 1.5-4. Potential Launch

Configuration Features

1-59
i

o The longerons are constrained by the fittings provided on the utilities

trays (Figure 1.2-2).

o The trays are constrained by the support provided by the barten mew.-a a

(Figure 1.2-2).

o The diagonals and batten individual members are constrained by their

end clevis attachments to the half-nuts and each act as simple beams

between their ends. The longeron halves act as simple beams between

their end clevis attachment and the attachment to the utilities trays.

1.5.2.4 Boot Strength During Deployment and/or Retraction

Figure 1.5-5 illustrates the Load path during deployment operations

subsequent to deployment of the jackscrew support from assembly. (The load

paths during deployment of the jackscrew support frame assembly were discussed

in Section 1.3.1.3.) Bending moments about the pitch or yaw axes are

transferred from the truss longerons of the last deployed bay to the half-nuts

and onto the jackscrews through bearing on the engaged threads. This results

in axially applied loads t.) the jackncrew which are transferred to the

carriage and into the rails through the carriage locking pins. The rails

transfer the load into the housing through attachment to the frames. Lateral

shear, or torsion induced shears, are transferred from the diagonals of the

last deployed bay to the jackscrews through bearing from the half-nuts. The

lateral load is transferred through bending of the jackscrew to the carriage

and to the jackscrew support frame. The lateral load at the carriage is

transferred into the rails through bearing on the carriage ears. The lateral

load at the jackscrew support frame is transferred to the housing frames

through the diagonal braces. During retraction operations, prior to 

retraction of the jackscrew support frame assembly, the same load paths are

available.

1.5.2.5 Dimensf_:s1 Compatibility - Spacing of Adjacent Batten Half

Nuts/P_, h o: Jackscrew Thread

A critical requirement of the mechanization system is the maintenance of

the longitudinal distance between the corresponding points on the half-nut

threads of adjacent battens (for etch of the four pairs of half-nuts) with the

mati.ig threads on the jackacrewa. Failure to satisfy this requirement would

preclude engagement of the aft batten half-nuts during deployment, and

engagement of the forward batten half-nuts during retraction. The requirement

and the design solution are illustrated in Figure 1.5-6. The threes of the

half-nuts has a radius 1.57 mm (.063 in) larger than the jackscrew. The

larger radius with a 1.57 mm radial separation will provide a 0.76 mm (.030

in) longitudinal clearance when assembled as described in Section 1.3.3. This

is accomplished through adjustment of the longeron and diagonal turnbuckles.

This assembly will be accomplished in a room temperature environment of 24 ±

10°C (75 ± 180F).

1-60
ORIGINAL PAGE 15

OF POOR OUALI7y

Y

APPLIED

LOAD

JACKSC7

SOPPORI

FPAIIE

JM.N1rM1[II nv... n.

Figure 1.5-5. Load Paths During Deployment/Retraction

1.35 n

TYP BA'

—1.35 —

JACKSCREN LENGTH

1.57 m

RADIAL

CLEARANCE

1.57 m t RADIAL

SEPARATION

PD.76^CLEARANCE PO-76 m CLEARANCE

INSTALLATION OF HALF

HUT TO JACKSCREW

Figure 1.5-6. Adjacent Half-Nut to Jackscrew Clearances

' 141

^' "rs_^ ..Ya« ....:.^._..._,._... saw .... :ii..._.^^•^....iN.',rtlasw.Y.^ ._:L ., i.[._ ^. ... - J_^.. .. .:. <"r^..,.,.,..^..^.. .... ..;:a....... ...._._. .^ ... 
Deployment and retraction operations should be performed in the same

environment. The three potential future deviations from the assembled

condition can be due to length changes from differential thermal expansion,

jackscrew pitch variation as the battens move along the jackscrew, and

longeron joint slop affects on longeron length.

o The worst case for thermal expansion would , aszemble at 34*C and

deploy / retract at 14 0 C or the opposite extreme ndition. A 20°C

temperature change will result in a 0.30 mm (0.012 in) differential

length change over the 1.35 mm length.

a

o For the typical bay of 1.35 m the half-nuts adjacent battens are both

engaged for a length of 50.8 mm ( 2 in) while for the carrier bay this

length is 203.2 mm ( S in). In both cases this ..s half the full travel

length but representative of the deviation from the assemb l ed condition

(Section 1 . 3.3). Reference 4 indicates the maximum jackscrew pitch

variation across 203.2 mm is 0.018 mm ( 0.007 in).

o The potential change in length, of the longeron, due to joint slop at

the two end clevises and two center pivots pine (4 pins, 8 holes) is

+- 0.16 mm. This was determined by calculating the maximum gap between

each smallest pin and the twu largest holes and taking the square root

of the sum of all the squared values of the gaps.

^• Numerical addition of the 0.30 mm, 0.018 mm, and 0.16 mm deviations T

results in a total of 0.48 mm, which is less than the initial value of 0.76 mm

and acceptable.

From the foregoing discussion, however, it is evident that it is critical

to maintain the radial separation of 1.57 mm between the threads of the half-

_ nut and jackscrew. A closer view of this problem indicated a small increase

(0.34 mm) in the radial separation is permissible since it does not decrease

the longitudinal clearance of 0.76 mm, and engagement is maintained at the

sides and most of the half -nut. A decrease in the radial separation of the

threads does decrease the longitudinal clearance.

i During deployment this concern is applicable at the aft face of the

jackscrew thread, where the aft batten comes onto the jackscrew. During

ii retraction this concv -n is applicable at the forward face of the jackscrew

thread, where the forward batten must mate with the jackscrew.

For the aft region, during assembly, the rails can be shimmed laterally U

to 0.05 mm of true position. Arithmetic additions of the tolerances between

the rail to the carriage and the carriage to the jackscrew add up to ±

0.175 mm ( laterally). For t he forward end the jackscrew support frame,

ferrule locations ( Figure ].3-12) and .ferrule to jackscrew thread tolerances

arithimetically add up to 9 .165 mm la:., i lly. The batten half-nut location

thread tolerances add up to • 0.30 -gym ( lateral). A statistical calculation of

the resultant radial (not 14Cerc'.' caviation is 0.34 mm (0.0136 in).
Since all the structure affecting the lateral stability of the ground test

article is aluminum, no significant differential thermal distortions are

anticipated. For a prototype flight article using low CIE materials the same

conclusion applies.

However for a flight article with an aluminum housing and truss, a detailed

thermal analysis is required to determine the magnitude of differential

thermal distortions. This will determine whether it is necessari to decrease

the pitch of the jackscrew thread to 9.5 or 12.7 mm (0.5 in.) to achieve a

thread depth up to 6.35 mm. This would permit an increased radial separation

and longitudinal thread gap up to 1.5 mm. This does not preclude the need for

a thermal curtain (Section 1.3.2).

A radial closure of 0.34 mm will result in a reduction of the 0.76 mm

longitudinal gap by 0.14 mm to 0.62 mm. Since the foregoing analysis

indicated a total longitudinal closure of 0.48 mm, the design is adequate.

1.5.2.6 Torque Requirement for Deployment/Retraction System

The output torque required for the batten deployment/drive system of the

test article was determined for vertical deployment without any ballast. The

torque requirement was determined to lift a truss mass of 454 kg and overcome

` ' a batten detent resistance of 1780 N. The equivalent force is 6230 N (1400

lb). The classical jackscrew analysis equations extracted from Reference 5

defines the required torque as

T . Wr tan (4 +0) where tan 6 - pitch/2rt r

't

and W is equal to 6230 N, r in the jackscrew radius of 25.4 mm, and is

determined from the coefficient of friction (tan .14). Based on the

foregoing a peak torque requirement of 26 Nm was determined. The use of a

coefficient of friction of 0.14 is considered to be conservative. The

determination of the required output torque for the unlocking systems is

presented in Volume III.

ss

1.5.2.7 Drive Chain Deformation Implications

i The drive chain used is a Browning No. 35 (3/8 in. pitch) stainless-steel

roller chain. The tensile strength capability of this chain is 9345 N (2100

lb), with a life of 15000 hours. These capabilities are greatly in excess of

1. the requirements. Of concern was the magnitude of chain slop and axial

^`• deformation in regard to its affect on the relative accuracy of the positioning of the jackscrews. The slop will be eliminated by a chain pretension from

• 223 N (50 lb) to 446N (100 lb). The peak maximum load in a chain due to the

^. maximum output torque is 845 N (Volume III). This is less than the chain

allowable working load of 934 N.

A total axial deformation of 0.467 mm was deter '.ned for the loads

developed in the chain due to the applied torque. This deformation, and an

additional 0.24 mm for wear (100 operations), will cause a deviation between

the longitudinal position of the worst jackscrew relative to the motor-driven

jackscrew of 0.014 mm (0.0006 in) which is acceptable.

1-63

...,__. _...uu.u., .. r:..:..,.v ...•. .3a-ft 7;".xa^3.4,^.. ..

... ^_.ti._^1 „.. ^__J. 3 ' _.... <T .. ^ - —:-]
1.5.2.8 Chain Drive Considerations

During operation of the three chain drive configuration shown on the top

assembly drawings, it is essential that adjacent chains do not touch. Hence,

any dynamic amplification must be limited to the clearance envelope provided.

The forcing frequency of the chain is determined by NS/60 where N is the

number of teeth on the sprocket and S is the sprocket revolutions per minute.

The frequency at 225 rpm is 131 Hz and at 75 rpm is 44 Hz. A chain pretension

of 445 N provides a minimum chain frequency of 12.8 Hz, i.e., an order of

magnitude below that of the forcing frequency. Therefore, it is not practical

to provide a chain frequency above the forcing frequency.

It is expected that buildup and decay of the shaft rpm will occur in no

more than 2 seconds. This will result in a forcing frequency of 8 to 18 Hz

for .15 second, i.e., the time for two cycles of chain vibration at 12.8 Hz.

This is not expected to cause any significant amplification, in the presence

of the IS static loading. However, if the implication is significant guide

supports can easily be added during the assembly and checkout stage of the

teat article.

1.5.2.9 Test Article Internal Loads

A NASTRAN model of the ground test article from which the internal loads

were established is presented in Figure 1.5-7. The model represents the

configuration of Figure 1.2-1 and the loading shown in Figure 1.5-1 plus the

addition of four 1780 N end loads representative of the precompression system

loading. The design limit compression loads, obtained directly from the

program are delineated as follows. The ultimate design values used are also

shown.

Longerons 13,800 N limit, 20,700 N (4650 lb) ultimate

Diagonals 6,750 N limit, 10,125 N (2275 lb) ultimate

Typical Batten Members 0, 2,140 N (480 lb) ultimate

Batten Diagonal Brace 0, 1,000 N (225 lb) ultimate

Payload Carrier Batten 0, 13,350 N (3000 lb) ultimate

The 13,350 N column load capability is provided to permit use of a

spreader beam (if necessary) below the truss to support the simulated payload

carrier. It is derived from 1/4 the 3636 kg mass, and a safety factor of 1.5.

The model also indicated a peak lateral deflection of 43 mm (1.7 in), at

point A (Figure 1.5-7) in the direction of the applied loads. The angle of

twist, of the plane at point A, about the longitudinal axis was 0.9 degree.

Further, for future ground test planning, the predicted spring constant is

400 N/cm.

t

d

w

t-64
1-65

1.5.2.10 Modal Analysis

A modal analyses was performed using the NASTHAN model shown in Figure

1.5-7. The 3636 kg simulated payload carrier c.g. was located as shown in

Figure 1.3-18. The first, second, and third modal frequencies were 0.40,

0.405, and 1.4 Hz.

The first mode exhibited a combination of lateral bending and torsion. An

analyses t, determine the sensitivity of the first mode value to payload

carrier moment of inertia variations indicated the following:

First Mode

Ixx I^ I2t Frequency Hz

3320 nm2 3018 =2 1210 om2 0.410

6640 =2 6036 nm2 2420 nm2 0.408

13280 =2 12072 nm2 4840 om2 0.400

1.5.2.11 Internal Loads - First Mode Behavior Modal Survey

For future test planning the internal loads anticipated for the longerons

and diagonals during a modal survey are a-mmarized in Table 1.5-1. The

magnitude of the precompreaeion in the longerons and diagonals depends upon

the initial compression ina.uced by the precompression system and the amplitude

of steady state modal excitation. It is assumed that during the modal test

only the lowest cantilevered mode of the truss will be excited. In order not

_ to introduce any non-linearity in the system, i.e., sloppy joints, it is

u required that the longeron and the diagonals should always be in compression.

o, The NASTBAN math model of Figure 1.5-7 gave the internal load distributions

for cantilevered deflections of the trues. From this various amplitudes of

i

excitation were selected, such that the longeron and the diagonal always

remained in compression during the test. The mass of the truss was considered

to be neglible compared to the mass of the payload.

Xf A review of the table indicates all the members are in compression if the

peak deflection at point A does not exceed 6.25 mm.

1.5.2.12 Diagonal Assembly Structural Analysis Issues

The diagonal assembly structural analysis reviews contained the following

significant structural analyses and design concerns.

(1) During deployment, how much spring force is required during the last

6 mm of stroke to ensure full deployment of the spring activated locking pins?

(2) How much outward spring force can be tolerated by r _ outer tube wall

design?

(3) What are the load implications of the radial tolerance between inner

and outer tube walls and the associated local tube stiffness affects?

(4) Was a room temperature cured adhesive joint suitable for attachment

of the telescoping tube to the center joint fittings?
ORIGINAL PAGE 19

OF POOR QUALf1'Y

NASiRAN MODEL

PAYLOAD

PAYLOAD CARRIER

CARRIER BAY

ATTACH

Z

X

I< Y

HOUSING PLANE OF

ATTACHMENT

POINT A TO NASA/MSFC

FIXTURE 

Figure 1.5-7. Test Article NASTRAN Model

Table 1.5-1. Predicted Loads During Modal Survey

CABLE PRECOMPRESSION (N)

0 450 goo 1800

ELEMENT MINIMUM AND MAXIMUM LOADS CONTROL MODAL DISPLACEMENT

CRITICAL ELEMENTS

+ 1.56 mm

1. LONGERON ±189 -645 -1000 -1713

-67 -423 -1135

2. DIAGONAL ±42 -106 -171 -300

-23 -87 -216

+ 6.25 mm

1. LONGERON ±1157 -1513

+801

-1869

+445

-2580

-267

2. DIAGONAL ±167 -232 .296 -425

•102 +38 -91

1-66

•
S<' , ,

These questions were resolved as follows:

(1) A conservative estimate of the potential force required during the

last 6 mm of stroke was determined a follows:

The major concern g as applicable to a potential flight article using a low

CTE diagonal of graphite composite. ( ( An aluminum design is expected to

require a deployable curtain on the jackscrew support frame assembly ( Section

1.3.2)). The concern is generated from the potential thermal-induced change

in-one diagonal members length during on-orbit deployment. For a 60°C change,

the change in length is estimated to be 0.13 mm ( 0.005 in.). For all other

compocents being as-built the force capability in the diagonal needed to

overcome this length variation was determined for the following NASTRAN Models

(assuming no joint slop).

lot bay deployed 0 (structure is determinate)

1 In 2nd bay with 2 bays deployed 49 N

In 9th bay with nine bays deployed 73 N

In 10th bay with 10 bays deployed 124 N

The larger force in the 10th bay is due to the rigidity of the hcusing.

In the foregoing section, a potential longeron length change, due to joint

slop was determined to be 0.16 mm. Using only 0.04 mm a virtual work analysis

indicates that such a slop can accommodate the entire 0.13 mm variation.

Therefore, the 124 N value was considered too high. A design goal of 73 N

appeared more reasonable and was established ( See Volume III). Friction

t effects are expected to be small by comparison.

A

(2) Analysis of an outer tube wall of 1.75 mm (0.068 in.) exposed to a

force of 90N indicated a stress leval of 1310 N/cm2 ( 1900 psi)., A

gauge of half that value would experience four times that stress

(still acceptable even for a composite constriction design). The

analysis was accomplished with a NASTRAN model. THe 90 N force

represents the outward load of the compressed spring ( Figure 1.5-8).

The plunger head is rounded to reduce friction and contact stresses.

(3) Figure 1.5-9 illustrates the main features of the analysis to

determine the implication of tolerance between the inner and outer

center fittings and local stiffness affects on the strength of the

diagonal tube. The concern is the potential magnification of the

initial deflection as ultimate load is approached. The initial

center deflection due to a radial gap of 0.125 mm (0.005 in.) between

the inner and outer sleeve is 1.27 mm (.05 in). With provision of

rings at the bearing lines, and accounting for ring local deflection,

a peak deflection at ultimate load of 5.3 mm (.21 in) was obtained.

The acceptable stresses compatible with this deflection are

illustrated in Volume III.

(4) Figure 1.5-10 illustrates the test configuration and results of

development testa performed upon three simulated bonded joints. The

lowest developed strength was 5.3 times greater than the ultimate

diagonal load.

1-67

....^ ... a -,.,.. ..

^._,y _., t.,,.. _ ..b.a ..,.......: ..,. ^ ...:-
ORIGINAL PAGE IS

OF POOR QUALM

17-7PN SPRING

Is .I M OU,

1.70 M WIRE DIAMETER

D

Figure 1.5-8. Telescoping Diagonal

Spring Activated Plunger

1.27

1012E N ,

^^

« 10124 M

r..

1793

r

END RING LOADING

STMI

P

Figure 1.5-9. Telescoping Diagonal

Column Loading Considerations

1-68

I
ORIGINAL PAGE 19

's OF POOR QUALITY

PLUG 1.0.

41.91 1 0

PLLG 0•we R

0.254

TUBE I

7.178 s WALL 19.05

e 0.254 ^w

TUBE 0.0.

49.55 s

49.50 eN

M PLUG 1.0.

p IUG 0. D. 19:81 W. 55.50 

49.16 s

1 0.254 w

1

NOTE: APPLY MB0120 .00B BOND PER

MA0105-705 GRADE A TO

TAPERED AREA OF PLUG AND TUBE

RESULTS

SPECIMEN ! - 54775 N

SPECIMEN 2 - 64970 N

SPECIMEN S - 53400 N

(DESIGN BOND ALLOWABLE1'9.2 N/ n2 - ACTUAL STR SS

1 0 N/cot TO 2207.2 N/cW)

VENDOR WILL DO SANE

PESTS ON SAME SPECIMENS

PROVIDED ROCKWELL

Figure 1.5-10. Telescoping Diagonal Adhesive

Joint Test Results

1

E'

d^

rc

a. 1-69

r

& 

2

. ,. .`.:... it r..^.-- ... - t..•....--. .. S` ^ 1 
V '

Y

1.5.2.13 Longeron Assembly (Drawing 42712-120)

The longern assemblies structural analysis reviews contained the following

significant structural analysis concerns.

(1) During deployment are the spring forces provided during the last 1/2

degree of longeron rotation adequate to ensure side plate locking?

(2) Does the center joint side plate design, in conjunction with the

reduced flexural stiffness in the rod ends, provide sufficient rigidity

compatible with the required ultimate load?

(3) Now much initial deflection is associated with the tolerances of the

center joint design, and what are the consequences at ultimate load?

These questions were resolved as follows:

(1) The spring torte in the plungers provide a lateral force of 47.9 N

(10.7 lb). The couple arm is 110 mm (Figure 1.5-11) hence, a torque

capability of 527 n cm is provided. The two major sources of opposing

friction torque are the two rod end clevises and pivot pins. An upperbound

estimate of the friction torque at each source is provided by the torque

breakaway data in rod and bearings (Reference 6). For the rod ends used this

value is 170 N cm thus requiring a total for breakaway of 680 N cm. Since the

longerons are already in motion, the 527 n cm provided is expected to be

adequate.

(2) Figure 1.5-12 summarizes the analysis method performed to verify the

column suitability of the longeron design. A NASTRAN model of the center

splice joint was made and used to join two halves of a constant

moment-of-inertia tube. The column capability determined was used to define

an equivalent flexural stiffness E2I2 as shown in the figure.

E I II represents the equivalent flexural stiffness of the earlier

rod-end turnbuckle design. This hand analyses was performed using the

equation shown where P is the critical column load. A value of 29,700 N was

obtained. This analysis is based on the improved rod end to turnbuckle design

which provides an increased EIII and a shorter effective length $ 1

(Figure 1.5-12).

(3) Figure 1.5-11 illustrates the tolerances upon which an analysis of

the worst possible initial deflection, at the center of the longeron, was

made. The initial deflection calculated was 3.63 mm by assuming the pivot

pins and spring plungers have the minimum values and the sleeves the maximum

values. Using an ultimate column capability of 29,700 N the magnified

deflection at an ultimate load of 20,700 N was determined to be 11.95 mm from

p f A : 1

1 Pult

1-70
ORIGINAL PAGE' 19

OF POOR QUAMY

110 M BOLT DIA 5.299 +0,000

-0.051 n.

— 'i 6675 6675 N^

N 9.7 m

20697 N' ^T0697 p 206

20697 N

6675 N

PIN DIA 17.435 •O.ODO LOADS FOR TOTAL LONGERON

.0.015

NOLE DIA 17.475 +0.025 4. 000

-0.000 O PIN DIA 9.484 -.017 w

77 N HOLE DIA 9.525.027 NN

-0.000

10346 N 10746 N

7778 Q, , HOLE DIA 6.712 ♦.027 M

7 N 7778 N L

-0.0141

LOADS FOR EACH OF 2 SPLICE PLATES

Figure 1.5-11. Longeron Center Joint

Column Loading Considerations

1 E i t I

{L

.

MSTRAH COLUMN STABILITY ANALYSIS

^fI

I 41 

^L2^42^t1 

I

E I I IE212 E I I 

I

(TAN P/EIIII II (TAN P^212) E212/E01

NANO ANALYSIS

Figure 1.5-12. Folded Longeron Column

Analysis Considerations

1-71
LUG NOT ADJACENT

10 LONGERON

LUG ADJACENT

TO LONGEAON

PREDICTED

REDUCTION

IN ST4iNESS

10S

Sul

aNe0w +2

E(

NeC4 .2

E!

Vrdw

GA

DIAGONALBENDING)

(POST (LUG

BENDING)

(LOO

SHEAR)

• .556 ♦ .023 • .015 ♦ .020 • .614 va

where Pult and Pcr are respectively the applied ultimrte load and Euler

column critical load and e i is the initial deflection. The stress

} analysis based upon this condition is presented in the appendix of Volume III.

1.5.2.14 Diagonal to Batten Attachment Stiffness Analysis

While all the center lines of the loagerons, diagonals, batten members,

and batten diagonal braces intersect at a common point, the clevis to which

the main diagonals are attached sustains shear and bending deformation during

transfer of diagonal loads (Figure 1.5-13). The clevis configuration is as

shown to provide clearance for the longerons. To assure that no significant

loss of torsional stiffness would be incurred due tr 1 Ir aforementioned lucal

effects the vertical deflection due to vertical sh as determined as shown

on the figure.

The percentage due to local affects (i.e., the last three terms) was less than

30% which is certainly acceptable.

4460 N I(a 1450 N

r

I

ri*

1

]204 N

PER BAT VERTICAL DEFLECTION DOE TO SHEAR IS

AEGION

p CONTRIBUTING TONTR16

TO

MTI

oEPMaAnaN

Figure 1.5-13. Diagonal Clevic Deformation Implications

ORIGINAL PAGE IS

OF POOR QUALITY

t

1-72
2. DEPLOYABLE VOLUMES

Space Station configurations developed in NASA and industry studies have

depicted OTV hangars and manned modules (Figure 2.0-1). The OTV hangars are

too large to be packaged into the orbiter, hence, the need for either a

deployable cr erectable design. The size ( approximately 8 m in diameter and

12.2 m long), however, does not invite the use of a space - fabricated design.

Each of the manned modules prosently configured is approximately 4.5 m in

diameter and 12.5 m long and, therefore, can be placed into the orbiter

without folding. This is accomplished at a significant increase in launch

cost compared to a deployable manned module. For example, two manned modules,

such as previously described, require two Shuttle launches compared to only

one Shuttle launch of an equivalent deployable module. However, the

deployable modules structural systems are more complex, require additional

specific sealing techniques, and will require on-orbit installation of

partitions and systems.

The goals of this deployable volumes study, for an OTV hangar, therefore,

were:

• To establish the major design driver requirements for the OTV hangar 

design

• To develop candidate deployable and erectable OTV conceptual designs,

that satisfy these requirements, utilizing hard shell and inflatable

constructions

o To select for future development the most suitable of the developed

designs. The selection was to be accomplished through a

systematic / traceable process, similar to that used in Part 1.

The goals of this deployable volumes stt l y for manned modules were:

• To obtain from the Rockwell Space Station study (Reference 7) and the

NASA study documentation ( Reference 8) the major design driver

requirements.

• To define, through collection and collation of Rockwell ' s Space Station

study, definition of a manned module, a description of a baseline

non-deployable (conventional) manned module design.

• To develop a hard-shell deployable and an inflatable design that

provide the same combined capability as the two baseline modules shows,

in Figure 2.0-1.

• To compare the developed deployable designs with the baseline design

and to select the most suitable design through a systematic / traceable

selection process similar to that used in the Part 1 study.

2-1

.—....v..:vu...v.^..iu,..w.^,^ . ftlfi.-^412...i i. r._.. ... .. ^i,C:: danx_.,t ... _..:,3. .._..... .'...:_...^..«..^.. ._, .. ^ ..^..... .. c ... ._.... ..,. 
The OTV hangar concept development utilized the following information

obtained from the Part 1 study:

o The concept, developed in Part 1. of providing a Space Station docking

port structure with ingress/egress between the station and hangar, with

integral provision of an astromasc for OTV ingress/egress, and integral

umbilicals for fueling the OTV, was maintained.

o Concepts 1, 3, and 6 developed in Part 1 (Figure 2.0-2) were developed

in further detail.

o The OTV hang-.r unpressurized, deployed ccnfiguration concept was

maintained in which with the aft and is open and oriented toward earth.

o The pertinent questions raised in the Part 1 study concerning

unpressuriznd inflatable structures were presented to Goodyear

Aerospace Corporation for resolution (Reference 9).

The manned module development utilized the following information obtained

from the Part 1 study:

o The concept of a large rigid strovgback, onto which either a deployable

hard-ohell or an inflatable shell i.e to be mounted, was maint"ined.

within this strongback are mounted the conglomeration of ECLSS air

revitalization, humidity control, water reclamation, waste management,

t: food preparation, electrical power, thermal control, and crew systems

equipment.

v .'

o Concepts 3 and 5, developed in Part 1 (Figure 2.0-3), were developed in

„ further detail.

The study was directed by the following major goals:

• Development of a design for the early 1990's i.e., a design requirini, a

minimum of technology development that can wisfy the requirements

• Strive for minimum usage of orbiter cargo bay length

• Utilize EVA where complexity can be significantly reduced

2.1 OTV HANGAR DEVELOPMENT

The development of OTV hangar design concepts was directed by the

established requirements shown in Figure 2.1-1. These requirements represent

the beat information available at the time, which was during the early months 

of the aeries of Space Station contracts.

The drawings describing the concepts developed are contained in the

appendix cf Volume III.

All the concepts have an integrated central hub which contains the Space

Station decking port with egress/ingress between the Space Station and hangar,

an astroms.it , and OTV refueling probes. Upon OTV approach the astromast is

extended outward to the forward edge of the hangar. Three extendible struts

stored in the docking device, mounted at the forward end of the astromast

2-2

-
ORIGMAL PAGE 19

OF POOR QUALfV

Figure 2.0-1. Potential Space Station

Conf igruration

(D RIC"D DEPLOYED 8.2M HANGAR

m WED 0 %, OIA

21610 DEPLOYED

HANG"

PSTOW 14 . 4A0 ED IIA

HANOM

^7 M7

STOWED

-4

1101D MLOYw 4.5M DIA

HANGAR

SPACE LAUNCHkO

STA^oN r.-- 0 wrw ON

CUM

= IL M

RICHO .22A Uh

(D

WGID PwLs

E.V.A WONX STATION

$TOWED vOLuml

6.ZA v 4.JM DIA

RJOID KANdA&

ROTATIN 1

WO" $1 VION AND

PANEL SIOWAGE s

PACE

it ^^^STATION

6 .. J

Figu^,^ -2. Part 1 OTV Hangar Concept

2-3

--7
®1_

IT^

7.07M D10 u RAW e • . OTy 10.7M

T

NANGAN .^.

^.I

1 —I2.7M

sM ou

L11. DIA ^I -

FDOD IANOTN

T TOWS

II.AMDupf

L

0 1

AM Du

^.^.IY.BM DeRWID

a. PM sTO++to

® mou

^I11.I

M S1

DQFLOYDT

ADD I I.I0MtD

76

Secim"

—"( 1^7.OSM DN

VAA

NI OIA

—^'1 I^7.OSM 0

DIA

LH2

APS LHI TANK L0I TANK

BASE LINE OTV

Figure 2.1-1, Baseline OTV Characteristics

2-4

URIGINAL PAGE 19

OF POOR QUALITY

i; h

01

HUA

Figure 2.0-3. cart 1 Manned Module Concepts

• LIFE OF IO YEARS

• DOCKING PROVISIONS FOR ATTACHMENT TO SPACE STATION

• PERMIT CREW INGRESS/EGRESS FROM SPACE STATION

• PROVIDE FOR OTV BERTHING ON DOCKING, AND INGRESS/EGRESS

• PERMIT CAPABILITY OF ATTACHMENT OF OTV SERVICING, LIGHTING,

ELECTRICAL POWER EQUIPMENT

• PROVIDE WORK PLATFORMS AND CLEARANCE (i TO 1.5 m) FOR WORK SPACE

• PROVIDE CAPABILITY TO STORE SERVICE EQUIPMENT AND/OR SPARE PARTS

• PROVIDE DEBRIS/MICROMETEOROID PROTECTION FOR OTV

• PROVIDE RADIATION SHIELDING FOR CREW AND STORAGE EQUIPMENT

• PACKAGE WITHIN ORBITER DYNAMIC ENVELOPE AND SUSTAIN LAUNCH ENVIRONMENT

11.am

WII TANKS

a

APS LOX TANK

FUEL TANKS
extend radially to the inner face of the hangar platforms until bearin 6 on the

platform is developed. The astromast is laterally supported and therefore has

sufficient flexural stiffness to sustain OTV docking. Upon completion of the

docking maneuver the astromast draws the OTV into the hangar for servicing.

All the OTV platforms require longitudinal bracing and in-plane structural

attachments to develop frame behavior in the platforms. In all the concepts

this structural continuity of the frames is accomplished by EVA. In all the

concepts, excepting only Concept 3, the longitudinal bracing is installed 4y

EVA.

A detailed description of each of the concep t s developed is as follows:

2.1.1 Concept 1 - bouble-Folded Hard-Shell (Drawing 42712 - 1)

Figures 2 . 1-2 through 2.1-4 describe the major design features of Concept

1. Figure 2 . 1-2 illustrates the fully deployed configuration. The central

hub fixture described above in this case also contains six telescoping braces

used to laterally deploy the stowed configuration ( Figure 2.1 -3) to the first

stage of deployment. The braces also provide longitudinal support to the ON

hangar shell. The shell is comprised of twenty - four 50 mm-deep graphite

composite faced honeycomb sandwich panels packaged as shown in Figure 2.1-3.

The panels utilize graphite composite face sheets to minimize thermal

distortions for assurance of post-deployment locking of the latches shown in

Figure 2.1-4.

In the folded configuration, which has a length of 5.6m, platfoi +r panels A and

B are in different planes to permit folding. Platforms are provided at the

two ends and as shown in Figure 2.1-2. To accomplish the folding shown and

permit deployment, the panels in Bays 2 and 3, during deployment, are attached

only along the common circumferential edges. Hence, the latched details are

provided as shown on Figure 2.1 -4. Details 1, 2 and 4 present the concept of

the latches required for structural continuity of the shell. Detail 3

presents a concept for EVA installation of the struts as shown in Figure

2.1-3. The configuration is supported in the orbiter by the fittings shown.

A band clamp is also provided.

2.1.2 ConceRt 2 - Double-Folded Curtain Shell (Drawing 42712-2)

Figures 2 . 1-5 through 2.1-7 describe the major design features of Concept

2. The deployed configuration is shorn on Figure 2.1-5. Eight double-ended

continuous longeron astromasts are used as the longerons to deploy the curtain

shell construction shown in Figure 2.1-7 which uses 6.25 mm deep aluminum

panels. The astromasts are laterally stabilized by sidewall frames at the

ends and at the middle. During deployment the frames unfold radially as shown

in Figure 2.1-5. A system of 16 sets of X-bracing is provided between the

astromasts and frames to provide a trussed configuration. The X-bracing

cables are tensioned by the final stages of astromast deployment. Discussions

with AstroResearch personnel confirmed the feasibility of this approach.

During and subsequent to deployment, the curtains span circumferentially to

tension cables attached to storage r..els mounted on the frames. Three sets of

work platforms are attached to the folding frames as shown in Figure 2.1-6. 

Continuity of the frames is provided by self-latching devices or the frames.

Installation of the longitudinal bracing supports requires EVA.

2-i

h

• • Y

• ems.: .• u >.^. \
ORIGINAL PAGE I'l

.t

OF POOR QUALIV

f

1

Figure 2.1-2. OTVH Concept 1 —Hard Shell Double Fold

(Fully Deployed Conficuration)

HINGE 8 PLAT DAN

FITTINGS FOR LAUNCH

PANEL A i

REFUELING

PROBE

HOUSING

HINGE A `

BI GRAPHITE

STANDARD

SIUIDNICH PANELS DEVICE

STONED BAND CLAMP KEEL FITTING

^3.6 s-+i

WORK PLATFORM-4

HINGE B PLATFORM PANEL A rte A

` WORK PLATFORMS f TELI

HINGE A BALI

/

BTTACF@ffNTS

PLWORN

PANEL B STRUT

/!1 ATTACHMENT

^Y BY EYA

DETAIL 3

\ (FIGURE 72)

- -1

AST

DEPLOYMENT ARM -- CAN

A PLACES (MOTOR

DRIVEN)

A

7.9 FIRST DEPLUYMENT_PHASEr

A-A

Figure 2.1-3. OTVH Concept 1—Hard Shell Double Fold

(Stowed and First phase of Deployment)

2-6
SIDEYALL

FAMES

DEPLOYED

D (EVA)

1/4 OF STRDCTUCE -*J A - -

Figure 2.1-5. OTVX Concept 2—Curtain

Shell Double Fold Configuratior

2 -7

ORIGINAL PAGE IJ

OF POOR QUALnY

PIN FORCE

1

G LINK

SELF ALIGNING

REARING

DETAIL LLOCKING BALLS

QUAIL

AO REO'D

IALL-SCREW LINO

MOTION DEVICE

ACTUATOR

TORSION

SPRING..

TAPERED HOLE \x/ PITTING

V

CONICAL NOSE-PIECE DEPLOYING

PULL 0

12 REO'D PCTAIL O

2/ REO'D

Figure 2.1-4. OTV9 Concept I—Details

ELECTRIC

MOTOR

ROTALY

ACTUATOR

7).

k

s;

t

i
• DIA

^. 1

CURTAIN

SIDENALL

STOWED 10. um

FRAME

STORAGE

REEL

ORIGINAL PAgE IS

OF POOd QUALrry

0.6

STOWED CURTAIN

CENTER

ASTROMAST

CANISTER

SIDEWAII 1

FRWS

STONED

";i A4. OTY

DOCKING

^

/

PORT

ORBITER EKTENDAKE

RETENTION STRUT (8 PLACES)

CURTAIN CRADLE

DEPLOYMENT PLATFORM RADIAL

ASTROMAST STONED DEPLOYMENT

% (8 PLACES) CYLINDERS

(8 PLACES)

Figure 2.1-6. OTVH Concept 2—Curtain Shell

Double-Fold (Stowed Configuration)

jM^

s

IULK

ALL

WRING \ ,l 1 ! ELONGATION

CABLE CURTAIN GUIDE CURTAIN

SIDEKALL

FULLY

ELONGATED 

Figure 2.1-7. OTVH Concept 2—Curtain Shell Details

2-8

4. a •^.^.

ALI

,8,

.....
i

Similar to Concept 1, extendable Struts (telescopic braces) are provided

to first laterally extend the packaged structure as shown in Section B-B of

Figure 2.1-5. For this design eight struts are required;' i.e., one for each

astromast. Upon completion of this phase, and locking of the frame joint

latches, the astromauts are mechanized to longitudinally deploy the structure.

The packaged configuration length as shown in Figure 2.1-6 is 3.6 m. A

cradle is provided as shown in Figure 2.1-6 for support during orbiter

launch. Fittings are also provided on the hub structure.

2.1.3 Concept 3 - Single-Fold hard-Shell Design (Drawing 42712-3)

Figures 2.1-8 and 2.1-9 illustrate the major design :eatures of Concept

3. This concept is primarily the simplification of Concept 1 achievable by

deleting the longitudinal folding requirement at the cost of a significant

increase in orbiter stowage length (14.5 m). While the significant launch

cost penalty is appreciated, if a significant part of the available space

shown can be used to place OTV spare parts, or other Space Station equipment,

that penalty may be acceptable in consideration of the significant reduction

of hangar complexity. In this concept the entire outer shell is monolithic

except for the longitudinal hinges that permit the lateral folding shown. The

panels are 50-mm-deep aluminum faced honeycomb sandwich structure panels.

Thermal distortion In this design can be tolerated by the design of the common

longitudinal hinges.

The platform panels A, are folded up against the flat outer shell panels

to permit folding as shown in Section A-A, Figure 2.1-9. In the deployed

configuration these panels A and B are in the same plane. Longitudinal braces

can be ground-installed; however, EVA is required for on-orbit attachment of

panels A to B, and adjacent panels A to each other.

As in the Concept 1 design six telescopic braces are provided for lateral

deployment. The stowed configuration is shown in Figure 2.1-9 with provision

for orbiter attachment.

2.1.4 Concept 4 - Erectable hard-Shell (Drawing 42712-4)

Figures 2.1-10 through 2.1-12 describe the major features of this design.

The outer shell of this design is built up from eighty-eight 18.8 mm (0.75

in.) deep by 1.2 m x 2.9 m graphite composite faced honeycomb sandwich

panels. here too, graphite composite construction is used to ensure minimal

thermal distortion during the erection process. The deployed shell is

stabilized by work platforms at the ends and center of the configuration built

up from frame panels as shown in Figure 2.1-11 (Stage 6). All the panels are

stowed into the storage container space within the work platform enclosure, as

shown in Figure 2.1-12.

The construction is accomplished in the manner shown in Figures 2.1-10 and

_ 2.1-11, Stages 1 through 7. Extension of the platform for construction is

accomplished by the astromast shown, which also is used to provide

ingress/egress for an OTV.

The construction is accomplished by two astronauts and will proceed as

follows:

Referring to Stage 1, the plattorm container is extended to the position

2-9

a

i
T

1.

ORIGINAL PAGE 19

OF POOR QUALITY

DEPLOYMENT

MECHANISM

GUIDE MILS

WITH BACK

PINION DRIVE

(A PUS AT 90'

A-A

LONOIIUOIML

BRACES T+n

i•

^ I J,^

RIO,

L►

END VIEW

Figure 2.1-8. OTVH Concept 3 —Hard Shell Single Fold

(Fully Deployed Configuration)

ATTACH Z REFUELING

FITTING ATTACH R 1 Z M--Er A

PROBE 1 FITTING 1 1

HOUSING

H

Y

STANDARD

H OOCKING

DCVICE

Y ATTACH M p

--f— 14.5 n

RN SI s ALUMINUM

FACED PANELS

AVAILABLE SPACE

FOR ATTACHMENT

OF EQUIPMENT

-

..^. — .

ALL A SECTORS B

. + -

A•

ALL A'BAYS

INNER

PANELS

STORED

I OUTER PANELS OUTER

SHELL STOWED WORK

PLATFORM TIED

SS! TOGETHER BY

55 SEPARATION JOINT

B

-B SERRATED PLATE A-A (LOCKHEED SUPER ZIP)

Figure 2.1-9. OTVH Concept 3—Hard Shell Single Fold

(Stowed Configuration)

2-10
ORIGINAL PA GC• IJ OF POOR OUALfry

/:rlwwrl

-^IAtwin ISt SIT

r 1101 IM LI

i

Iran

ul^

IM[l MIIIRt

p..

r r

ae

Iurralll

IIMI NItMIMt

rurraw olnoro

Figure 2.1-10. OTVH Concept 4—Erectable

Hard Shell Configuration Construction

k v

Figure 2.1-11. OTVH Concept 4—Erectable

Hard Shell Configuration Construction

2-11

ma.ae „tAmir..,1 Adt.... .r
shown in Stage 2. The end-hinged stowed panels are deployed as shown in Stage

3 and joined to each other. The periphery of the end panels have provision

for mounting of the first set of side panels as shown in Stage 4. The panels

are fastened together by the studs shown in Figure 2.1-11 that are mounted

onto a cable prior to attachment to the panels. Numerous other means for

attachment are possible. The construction continues as shown in Stage 5. The

placement of the panels is assisted by the guides and tabs shown in Figure

2.1-11. In Stage 6 the installation of the platforms is shown.

This construction has been conservatively estimated to require 105 manhours for each of the two astronauts for a total of 210 man-hours. This

includes 18 man-hours for the work platform deployments. The estimate is

composed of a nominal time plus a 30% contingency time to account for delays

and uncertainties. A total of five hours nominal time for suit donning and

doffing is included for each six hours of EVA (or less when breaks seemed

advisable in the sequence). With the eight-psi pressure suit, such don-doff

time may be reduced to two hours or less per six hours of EVA. The

preliminary concept for procedures and assembly hardware could probably be

significantly reduced (perhaps by 25Z to 502) following further study and

application of concepts using simple rods, tools, and latches, which would

reduce the need for some crew relocation activity during assembly. Assuming a

labor rate of $15,000 per man-hour the approximate cost for the worst case

would be $3,150,000. With the use of a rotating platform, it is quite

possible that this time can be reduced to 48 man-hours each, i.e.,

construction in six eight hour days. The cost would be reduced to 81,440,000.

This concept results in the smallest packaged length, i.e., 3.0 m.

Support in the orbiter is provided at attachment fittings on the docking port

and platform container.

2.1.5 Concept 5 - Inflatable Design (A drawing was not developed)

A drawing was not developed since adequate design definition is provided

by Figures 2.1-13 to 2.1-15. This design was developed by the Goodyear

Aevospace Corporation (GAC) under contract to Rockwell (Reference 9). The

construction recommended by GAC is shown in Figure 2.1-14 and further defined

by the following table.

MATERIAL WEIGHT

ITEM GM/CM2 LB/FT2!-

Outer Cover (With Thermal Coating) 0.034 0.068

Interlayer Adhesive 0.005 0.010

Two Plies of Fiberglass 0.036 0.071

Interlayer Adhesive 0.005 0.010

Gelatin - Impregnated Scott Foam (64 mm) 0.375 0.750

Interlayer Adhesive 0.005 0.010

Two Plies of Fiberglass 0.036 0.071

Inner Coating 0.020 0.040

TOTAL 0.516 1.030

2-12
AIR SUPPLY \

k4

10

f... INFLATABLE

HANGER S10YE0

' " 1 PAYLOAD CONFIGURATION RADIAL SUPPORT AIM

HANGER SHELL

OF POUR QUALITY

I FORT CREW RESTRAINTS

YORK PLATFORMS

ASTAONAST

CANISTEA

i

s..

1.1

mill

/ i

S.0 s

KEEL

FITTING FOR OTV

COCKING

Figure 2.1-12. OTVH Concept 4—Erectable

Hard Shell (Stowed Configuration)

ORBITER ^

TRUNNION

(1 PLACES)

Figure 2.1-13. 0TV11 Concept 5—InflatableConcept Stages at Deployment

c1^^y 2-13

r
MANGE STABILISATIC:

DURING RADIAL STRAPS INSTALLED

DEPLOYMENT BY EVA _t

TWO PLIES OF

FIBERGLASS

ORIGINAL PAGE It

OF POOR QUALITY

RADIAL SUPPORT ARMS

ROTATE AS HANGER IS

THICKNESS

1j

POLYURETHANE MESH

CORE IMPREGNATED WITH

GELATIN RESIN FOR

RIGIOIZATIDN

TWO PLIES OF FIRERGLASS

WITH THERMAL COATING

CONSTRUCTION

COMPLETED

OTV HANGER

CI RCUMIERENT IAL

WORK PLATFORMS

Figure 2.1-I4. OTVH Concept 5—Inflatable

Concept Stages at Deployment

PLATFORMS

STOWED HINGE (8 PLACES)

A TOROIDAL TANK

AIR SUPPLY RIB (8 PLACES)

OOCKIHG PORT IF

INFLATABLE

_

COVER Ih I STOWED

KAMM ^ ^, ^

;

,, . HANGER

STOWED f . P

r DOCKING PORT I

i

HANGER

ASTRONAST

CANISTER

END RIBS ^•^

A t 4.6 M CLR DIA

--4.0 M

SECTION A-A

Figure 2.1-15. OTVH Concept 5—Inflatable

Concept Stowed

2-14
2-15

I

The packaging and development of this structure are complex. The

following concept is the result of a very preliminary study and should be

considered as such. Inflation is used for controlled deployment and a gelatin

system is proposed for rigidization. The work platforms are one meter wide

and held in place axially by straps. Open areas are provided in the platform

areas to permit suited astronauts to move axially io the hangar. Velcro,

snaps, hooks, and straps can be installed as required to enhance maintenance

operations. Rigid subsystems can be added once the hangar is deployed.

Deployment of the structure will proceed as follows from the stowed

configuration (Figure 2.1-15). The stowed configuration, which is best

visualized as a "rolled-up sock," is pushed out of the stowage container by

extension of the astromast which drive, the covered docking port against the

folded configuration (Figure 2.1-13 Stages 1 and 2). The air supply system

(toroidal tanks) shown is used to inflate the "Chinese Whistle" longitudinal

tube (Stages 3, 4, and 5). At this stage, continued inflation of the 64 mm

aandwtch develops the configuration to Stage 6. Subsequent to completion of

Stage 6 the chemical process of the gelatin syutem provides final

reigidization of the shell. EVA is required to connect the tension straps

that stabilize the work platforms.

A potential problem is posed by the process used to package the structure

into the container. A vacuum bag is positioned over the sandwich structure of

" the hangar shell. The vacuum bag is used to compress the foam sandwich. This

may be the least understood part of the process. The rigid radial support

arms are stretched along the hangar axis while the hangar skin is pleatfolded into a diameter compatible with the storage container. The inflatable

tube is added to the pleated skin. The collapsed hangar skin is now rolled up

like a window shade in th y direction of the storage container. The rolled

material is placed into the container and the rigid radial support arms are

folded into position to hold the stored hangar in the rigid ^Tn. The

potential p problem is how to get ril of the vacuum bag after aangar deployment.

A further unknown is what is the best way to initiate the rigidization process.

A major concern is the lifetime of such a structure. However, CAC (hence

most likely any other firm) does not have sufficient test data to predict the

lifetime of the proposed hangar materials in the LEO environment. Past

projects have indicated a 5 to 10 year terrestrial storage and service life

for these flexible structures; however, ±ti' ac'ial on-orbit test data are

evaluated it is rather diff,i:ult co pc• I izt n Lij yea: lifetime.

Stowage in the orbit6e is accomplished as shown in Figure 2.1-15.

Supports can be provided on the docking port and the stowage container. The

packaged length is 4 m.

2.2 OTV HANGAR CONCEPT ANALYSIS

The analyses performed to support the concept evaluation process are

briefly described herein.

5:..

tr , t

LL

` L

4 -
r 2.2.1 Micrometeoroid/Space Debris Shielding

Figures 2.2-1 and 2.2-2 illustrate the significant issues pertinent to 07'V

s hielding from the meteoriod and space debris environments. Figure 2.2-1

ilustrates the vulnerable areas computed for the surface of the OTV. For the

wironments and probabilities of no impact the particle sites, for which

-ielding i s required, are shown. Unquestionably, there is great uncertainty

the validity of the debris model. Also, many Space Station system

-.gin-- cs believe advanced detection of such large debris particles and

a.oidsoce is preferable to attempts at shielding. The prediction of required

shielding has the additional uncertainty C lustrated by Figure 2.2-2, Section

A-A. For one m wide platforms, the OTV shell is approximately 1.2 m away from

the hangar wall. All the existing ballistic data for micrometeoroid particles

are bated on constructions in which the shielding is appreciably cl.ser than

1.2 m, i . e., 0.05 to 0.1 m. There is no data at all for debris. In view of

all the foregoing uncertainties, no analyses for debris was attempted.

For the micrometeoroid environment, the following analysis results are

pertinent.

• The aluminum single-sheet thickner <. for Po - 0.99; for no penetration

by a 7.4 mm diameter micrometeorr_,;, is 21.6 mm

• The combined effective sheet thickness of the OTV expected space-based

construction shown is 3.1 mm

• For the total hangar wall cn;,struction the required bumper ballistic

efficiency is 21.6 / 3.2 - 6.74

• The current ballistic efficiency for 50 mm spacings is 5; hence for

1.2 m, a value of 6.74 is readily achievable

The imp;. ications of the above are as follows:

o It is expected that all designs can be suitable for micrometeoroid

protection

o The outer- shell mass required to provide protection may be leer for

designs : iith the greatest construction depth, thus providing Sreater

cargo bay location flexibility. This is demonstrated as follows:

• A linear mass variation of 1078 to 2156kgim satisfies

the lauch envelope

• A T-bar comparable to 1078 to 2 7 56 kg /m stowed in 3 m

varies from 3.6 to 1.2 mm ( 0.26 in.).

Hence lightweight designs such as a T -bar of 3.6 mm can be placed anywhere.

2-16

C _'

4 ^
• LLMERACLE

AREA: 270 02

• VULNERABLE

PROJECTED

70ARf2

M_ S.D.

VELOCITY (Yi/f6^c) 0 10

DENSITY (MVcn') 015 2.8

DIRECTION OMMI III.

• PARTICLE SIZE (PRON. OF NO IMPACT • Po)

25.1

SPACE DEBRIS

0

20.3 . ( PO • 0.99)

d 115.2 MICROMETEOR010

lPO 0.99)

v

V

10.2

lb

5.1

0

10 20

DESIGN TIME (c) YEARS

I

10 .8 -6 -1 -2

LOG MASS (n ) GRA45

.6 F

g -9

n

V-10

^g

57 17'

G!

p; -11

J -15

12

NICROMETEORDIO

(S► 8017)

LOO M•-11.71

- 1.217 LOG n --

SPACE MINIS

LOG 8 • .12.00

. 0.25 LOG •

(KESSLER. 1990)

OF 'aNAL PAGE 19

POOM QUgLry

n PARTICLE ENVIRONMENTS

i

I

Figure 2.2-1. OTV Hangar Man-Made Debris

and Mic':ometeoroid Design Implications

HANGER J L

WALL V

DESIGNS HAVE DEPTHS

Q 

_

FROM 6.75 TO 50.8 m

I

1219 mm

12 7 MI NIMUM

SHELL (SPACE-BASED;

E-BA E0

--r' —^— SHEL (SPAE-BASED;

YAII ES

,072 9/m7 o GR CORE

0.20 mm GRAPXITE

FACE SHEET (2)

ALUM TANK WALL 0.76 mm

SECTION A-A 

Figure 2.2-2. Micrometeoroid/Debris Model

Uncertainties for OTVH

2 -17

Y 

..r. ... .. ...,. ^ _. u. .. . ... f. ... . ^._^ .. ^..f lo- _. ^..' ... ^..w a.^ .. .-.:.4-:`... 1. i.r .i.u.:e..... _.. .. .. w.. 
^g

C

2.1.2 Stress Analysis Review

All of the described designs were reviewed to satisfy structural design

suitability. It is apparent the OTV hangar, during operation has minimal

strength and stiffness requirements that are achievable with all the designs.

The main output of the review was the inclusion, in the design, of all

necessary structural components to satisfy astronaut foot loads, hand loads,

placement of lighting and black box equipment, and OTV docking loads. The

most stringent structural requirements occur during launch. A loads and

stress analysis of the configurations was obviously beyond the study scope.

However, during the development of each concept, the design was reviewed and

driven to provide assurance of launch suitability. For example, Concept 2

required a retention cradle.

2.3 OTV CONCEPT SELECTION

Table 2.3-1 illustrates the selection process criteria used to evaluate

the relative suitability of the five concepts developed. Tables 2 . 3-2 through

2.3-6 present the actual evaluations performed for each of the major and

subselection criteria listed in Table 2 . 3-1. Table 2 . 3-7 provides a

normalized summary of the major criteria 1 through 5.

r .y

Concept 4, the erectable design was selected on the basis of the data

contained in Table 2 . 3-7 and the following judgmental observations:

i

t. o While Concept 1 is close ( 343) it.. total point value, its cost is $8.8

• million higher

o In comparison with Concept 1, the erectable design has the highest

potential technology transfer to other large deployable volumes

o While the inflatable design has great growth potential, the suitability 

of the proposed rigidization system is uncertain with no apparent

potential for on-orbit remedial action in the event of a problem.

Further the on-orbit life of the inflatable construction is not known.

It is also pertinent to mention the following additional observations.

• The current erectable design demonstrated the advantage of an erectable

OTV hangar, but may not necessarily be the beat erectable design. 

• The on-orbit estimated construction time of 210 hours, with an EVA cost

of $3.15 million, can be reduced through use of automatic latches and a

rotating platform. 

• A future study to resolve the above, directed toward development of a 

point design, represents the next logical step

2-18

._. .... ,l ..^-•r-.-^^j,. .:: .... 

'. ..

._. a.— ^ AY i 3:.......nom ..» a...... ... .2 ............ ..h a 
OPP QRjQfjV

PAQF

Table 2.3-1. OTV Hangar Selection Process Criitte iaaUAL,ry

(1) RELATIVE COSTS(IMILLIONII

• LAUNCH • TECHNOLOGY DEVELOPMENT

• FABRICATION w EXTRA VEHICULAR ACTIVITY

(2) DESIGN VERSATILITY

• CAPABILITYFOR

- GROUND INSTALLATION OF EQUIPMENT

- ON•UR/IT INSTALLATION OF EQUIPMENT ON SHELL

- ON-ORBIT INSTALLATION OF EGUIPMENT ON PLATFORMS

- ADOEOPLATFORMS

• POTENTIAL FOR SIZE GROWTH

(3) ORBITER INTEGRATION

• CONSTRUCTION SUITABILITY

• CRADLE COMPLEXITY

• LAUNCH LOCATION FLEXIBILITY

'(4) RELIABILITY OF STRUCTURAL RIGIOIZATION

• SIMPLICITY OF JOINING

• CAPABILITY OF REMEDIAL EVA

• CONTROL DURING CONSTRUCTION

• NUMBER OF MOVING PARTS/FUNCTIONS

• CONFIDENCE FROM GROUND TESTS

IS) OESION/EVA/TECHNOLOGY DEVELOPMENT/LIFE RISKS

• LEAST DESIGN RISK

• MINIMUM EVA RISK

• MINIMUM TECHNOLOGY DEVELOPMENT RISK

• MINIMUM STRUCTURE LIFE RISK 1

Table 2.3-2. Relative costs ($ MillionsJ

TOTAL

LAUNCH POINTS

PACKAGK MAXIMUM LKNOTII Teal. A

CUIICEPT (•) CO"* IAO. nEV. KVA 70TAL 100

(D 5.8 53.5 2.0 - 0.5 28.0 81

O

3.0 15.0 2.0 - O.B 17.8 99

03 14.5 80.8 1.5 - 0.3 02.0 0

O 3.0 12.5 1.0 0.5 3.2 17.3 100

4.0 16.8 0.5 2.5 0.5 2) 3 93

'Launch cost for FY )984 Is $77 pillion.

2-19
ORIGINAL PAGE IS

OF POOR QIIALfTY

Table 2.3-3. Design Versatility

w..

iNMILITTCNMIIITV— tNMIIItT— OM-ONOIT INSTALL. CNMII ITT fON WIO INSTALLATION ON-00411 INSTALL. 0710;I/NIMf 04010 POTIIITIAI IN Of liquip11tut Of QU IP ON SN1LL ON HAT101110 ►LATTOANS 1110 Callum TOTAL NAMN POINTS MOIL POINTS NANK f01N1S AIMA ro1N1f -NA-ft- -POINTS POINTS

MN. POINTS Ml. POINIl MX. I+p 1N1s MAX. POINis MT. roofs WINIA

COINSPT 10 90 90 0 BO

01 7 5 1 16 1 10 4 4 0 0 47

O S 0 0 4 S 16 0 9 S 0 79

O0 1 10 l 70 1 70 1 10 0 ' 0 00

O 0 0 0 16 0 16 N 0 2 l4 31

O ] 0 4 14 7 l0 9 0 l X0 So

Table 2.3-4. Orbiter Integration

gti, 1

CONSTRUCTION CNADLK LAUNCH LOCATION

SUITAOILITY COWL9XITT FLEXIBILITY- TOTAL

SANK POINTS SANK POIN78 HANK POINTS pill NTS

MAY. POINTS MAX. POINTS MAX. POINTS MAXIMUM

CONCXPT 10 10 00 SO

U1 1 10 1 to 1 00 SO

O S 9 S 0 4 10 IS

O 1 10 1 30 S 0 ZO

O 1 10 1 10 X 70 90

O 1 10 1 10 1 30 SO

*Due to might Implication from motsoroid shloldinM requirements.

2-20 
2-21

i

r,

ORIGINAL PAGE 19

OF POOR QUALf1'y

Table 2.3-5. Reliability of Structural Rigidization

I

CONTROL cow 11 INCI

SIMPLICITV OT CAPASILITV Of DINING w. OF MOVING TWO GROUND

M MMES REMEDIAL OA an TION AIRS 1 a IS

MNK FOISTSNTt RAMX POINTS AANK ►OINT4 RANK POINTS RAMIt ►DINTS

TOTAL

POINTS

EMI, POINTS MAX. POINTS MK. POINTS INN. POINTS MAN. POINTS INAINYI

COwIPI 10 20 10 20 20 100

O S 2 2 24 a 2 4 10 2 26 64

O 2 a 4 1 n 1 to O a 2 20 as

O 2 a 2 24 1 10 2 10 1 20 Be

O 4 0 1 30 1 10 2 to 2 26 00

O I l0 a 0 4 7 1 20 a 10 41?

Table 2.3-6. Design, EVA, Technology Development,

and Life Risks

LEAST MINIMUM MIN. TECIINOLOOV MIN. STRUCTURX

0ESIGN RISK KVA RISX DEVE1.01NBNT RISK LIVE RISK TOTAL RAIIK POINTS HANK POINTS RANK POINTS RANK POINTS POINTS

MAX. POINTS MAX. POINTS MAX. POINTS MAX. POINTS MAXIMUM

Co"CEPT 10 20 20 70 so

1O 4 a 1 20 1 20 2 20 71

O 4 A 1 20 1 20 1 20 75

O l 10 1 20 1 20 1 20 SO

O l 10 S 0 1 20 2 20 OS

O 2 0 1 20 S a S 10 42
A

ORIGINAL PA9E 19

OF POOR QUALRY

Table 2.3-7. Normalized Summary of

the Major Criteria

RESIGN ORBITER TOTAL

PORT VERSATILITY INTEGRATION RELIABILITY RIBS IRIS

MAX. POINTS MAX. POINTS MAX. POINTS MAX. POINTS MAX. 1401178 MAXIMUM

COMCBPT

100 JOB 50 100 100 50

O EI 50 50 J 04 SU 943 J

O

00 J 2E l0 66 94 J 305

O

0 75 J 20 58 J 100 J 283

IOU J 05 40 00 J 70 365 J

O03 70 J 50 J 47 54 314

J BIRDIES CONCEPT HAS FIRST OR SECOND NIGIIEST POINT VALUE

r.,

y

S'

μ 1

2-22
I

f

a

2.4 MANNED MODULE DEVELOPMENT

The manned module development consisted of definition of the two baseline

modules portrayed in the Rockwell Space Station Configuration Studies

(Reference 7) and the development of hard shell and inflatable designs that

each accomplish the total functions of the two baseline modules. This

approach was taken since it is compatible with the study goals, and ensurez

that the developed concepts could indeed be used in a Space Station

configuration such as is shown in Figure 2.4-1. This approach also assureq a

p-bper comparison of the designs. For example, the two baseline modules

equipment ( as best defined) was placed into the two deployable concepts. The

major requirements were, of course, the same as tabulated in Table 2.4-1.

Attention is directed to the two goals listed in the table which are extracted

from Reference B. Both of these goals are satisfied by the two deployable

concepts developed. Figure 2.4 -2 illustrates a potential Space Station

scenario utilizing the deployable modules. This scenario and the designs

developed were coordinated with Space Station project and design personnel to

assure satisfaction of the major requirements,

2.4.1 Baseline Space Station Conventional Modules (Drawing 42712-6)

Figure 2 . 4-3 illustrates the major design features of the baseline manned

t +^ module design. Figure 2.4-4 illustrates the distribution of ECLSS, water

purificatiou, power, and hygiene systems equipment.

k.: The baseline module is a classical welded aluminum cylindrical shell

containing at each end a toroidal transitior section to a conical shell. Each 

r'

end of the conical shell contains a docking port. A horizontal floor is

provided that is constructed of integrally machined floor plates and supported

by longitudinal beams that span to three transverse beams. The three

transverse beams are attached to three external frames that are welded or

mechanically fastened ( bosses) to the shell. The floor plates will be

fastened to the shell to provide a shearflow path for the longerons that

attach to the Orbiter. The pair of longerons are provided at the floor level

and attach to the extension of the floor plates.

The outer shell construction is as shown in Figure 2.4-3. The array of

equipment mounted above and below the floor is shown in Figure 2.4-4. One of

the major design goals is essentially a smooth inner surface of the module

wall to permit inspection and repair of microweteoroid punctures. Therefore,

much of the equipment, 1 to 2 m above the floor, may have to be supported by

temporary bracing down to the floor. This remains a future problem area since

the scope of the Space Station design has not evolved to that level of detail

as yet. A summary of the preliminary mass properties prepared for the

baseline module is presented in Table 2.4-2.

2.4.2 Deolovable Hard Shell Manned Module ( Drawing 42712-7)

Figures 2.4-5 through 2.4-7 illustrate the major features of the hard

shell manned module concepcual design that is equivalent to two conventional

baseline modules. For design commonality with an energy module, and for

docking requirements, a pair of conical shells using toroidal transition

sections are provided at each end of a central deployable section. If

necessary, side docking ports can be installed on these end structures to

2-23

. 1._

F

f 
INITIAL

• 3 TO 4 MAN CREW

am moon

rr40mru191 ureawte

s /&ante ruanr

a. 0

satrau

nraraa

""A

vur

mmss r

LAS

„a mama in

ORICJNAL PAGE IS

OF POOR QUALffy,

ALL UP

• LEO • 20 years

• LAUNCH/PACKA GING • Compatible with tho orbiter

• CREW • 4 to 8 efseemen for 90 days

+21 days emergency

is EN VIRONMENT • 10.5 n/sm 2 and 18 to 240C

is REOUNOANCY • Two separate pressurized compartments

• Two route for in rass/soffass

• MAXIMUM LEAKAGE RATE • 0.22 Kelday

• EQUIPMENT • Life support equipment

m Command Control Center

• Two docking ports

• METEOR010 PROTECTION • 1990 Keefer spree debris and

NASA SPSO 3 micromajoafrold m odel

• RAOIATION PROTECTION a 0.50 grsm/cm2

• Manned module design commonality (dodgn god)

• Compatibility with rvalutlon from initial to growth configuration

Figure 2.4-1. Evolution of Initial to Growth

Space Station Configuration

Table 2.4-1. Manned Module Major Requirements
2-25

E

OR1OIN

OF POpR 

QUFl I-rl'

i

4

DEPLOYABLE NODULE

(DEPLOYED)

ft

LOGISTICS MODULE

BE ME

OERTHED TO DEPLOYABLE MODULE D

Figure 2.4-2. Potential Space Station Evolution

Scenario with Deployable Manned Modules

r

MANNED MODULE

CONSTRUCTION

I ENVIRONMENTAL SHIELD

BIRDER I71 ra AL DUMPER/RADIATOR

SUPPORT I I SWPORT

25 w MILTI-LAYER (1 NEAT PIPE INSULATION (MLI) IPRESSURE SKIN

M (0.6 m ALUMINUM)

DESIGN COMMONALITY

6:i

HABITAT NODULE - 12.2 a LONG

COMMAND MODULE - 12.2 s LONG

4i

ENERGY MODULE - 6.1 n LONG

Figure 2.4-3. Manned Modules Baseline Design Summary

Ar

I

fJ!)

¢1
INTERDECK ACCESS

STORAGE E UTILITIES

WASTE COLLECTOI

HYGIENE FACILITY

PANG WASHER

POTABLE N

( O TAINTS)

ORIGINAL ''PACE 19

OF POOR QUALITY

TABLE

WARDROOM

CLOTHES RASHER A DRYER /1 I

iISHWASHER OVEN .MLi EY/FODC PAEPARA Y ION AREA

RCFRIGEPATOR E FREEZER

\E—' cTATCR 'M

S TATENOAN

— M _ STORAGE E UTILITILS

INTERUECK ACFQS

STATEROOM

S 61—L43 TR T E RDOl1

40RRSIgP/LABORATORY

EQUIPMENT ABOVE FLCOR

EMERGENCY WATER TRASH COMPACTOR

—^ NOT ATEP- -COLD WATER EMERGENCY MYGiN

SUPPLY CXYhEN

.! PRESSU RIZATICI

r^ SYSTEM

PURIFICATION EVAPORATOI' Q -T

V SUPPLY NITROGEN

w WASTE WATER

R EMERGENCY NITROGEN (] TANKS) r^ ^ —

COLD WTCR -- hex AIR COOLING

NOT WATER —il CO2 REDUCTION

DOOR CONTROL CATALYTIC OXIDIZER

oEMAM 1DIF rzR

EQUIPME NT BELOW FLOOR

Figure 2.4-4. Manned Module Conventional

Design Equipment Arrangement

Table 2,4-1. Baseline Conventional

4 Manned Module Mass (kg)

SYSTEM

MASS

(4)

BODY STRUCTURE 1375

INDUCED ENV. PROTECT. 657

LAUNCH, RECOYERY 6 DOCK. 222

ORIENT, CNTLS/SEP. A ULLAGE AA

PRIME POWER 0

PWR. CONY. A DIST, 159

GUIDANCE E NAVIGATION 0

INSTRUMENTATION 207

COMMUNICATION AI

E NV IRONMENTAL CONTROL Z048

PE(,,ORNEL PROVISIONS 1267

CREW STATION CNTLS E PNLS 98

RANGE SAFETY A ABORT TBD

WEIGHT SROWTH ALLOW, 1682

PERSONNEL TOO

CARGO 0

BALLAST TOO

RESIDUAL PROP. A SERV. ITEMS TOO

RESERVE PROP. 6 SERV. ITEMS TBO

INFLIGHT LOSSES TBD

TOTAI GROSS MASS (X9) 10.186

2-26

i

dr
STONED `—+^ (SIDE PANELS FOLDED)

COWIGURATICN C

IDIATORS MAIN FRAMING BEANS

D

F'_IP-WT " FOLD OM PANEL

PANEL

-

_4..

'I I^ \HINGE LINE B r

C

RADIUS 2.08 N

RD CYLINDER

DEPLO'. cD

C ONE IGUMTION

--HINGE LINE FLANGES WITH

SEAL 6 LATCHES 7l_

A VIEW B—B E --

SIDE HINGES SECT

Figure 2.4-5. Deployable Hard Shell

Manned Module Concept

-CEILING LINE

FLOOR LINE

I

FULLY DEPLOYED

BASIC

0.127 o AAPTON CONFIGURATION

POLYPROPYLENE SILICONE BONGED

POLYPROPYLENE SEAL HEAT

TO STRUCTURE SIDE PANEL SEALS

IN A CONTINUOUS STRIP

PONDEO TO SHELL

HINGE

AMIS

^.

"ZIPLUCA' BEAD SEAL

*SIOEHING^ESEAL

INSTALLED MANUALLY

AFTER FOLD DOW PANEL

IN PLACE

POLYPROPYLENE LINER

HINGE AXIS FABRICATED BY HEAT

SECTIUN B—B BONDING

Figure 2.4-6. Deployable Hard Shell

Manned Module concept Details

2-27

SEAL

INSTALLED

ORJ

B

Op POOR Q

Aac i

IB N 'C B

A, A

A STRONGBACA I.B . A 4.2

BOA SE:TION-ALL :IDES

B CAPABLE OF SUSTAIMING

SCTION A-A DIFFENENTIAL PAC aURF

METEOROID BUMPER

25.E .• MILTI-LAYER

INSULATION DELI) NEAT PIPE

50.8 R^ s 

PRESSURE SHIN

(3.6 Rr. ALUMINUM)
^V

accommodate the Space Station configuration shown in Figure 1.4-1 (bottom).

The deployable central section is comprised of a 1.8 m high x 4.2 m wise x 9.2

m long strongback structure onto which the pair of deployable cylindrical

shells is mounted at shown. The radius of the shells are equal. to 2.08 m with

hinge lines A, B, C, and D oriented along the entire cylinder langth. The

folding of the cylinder to suit the orbiter cargo bay diameter envelope is

shown in Section A-t. The four end bulkhei,l • that provide closure at each end

of the pair of cyLinders are shown in Vied B-B. Folding of the flat bulkheads

is accomplished such that each pair of flip-out panels is folded inward about

the vertical hinge line and then together with the :old-down panels are folded

inward and tucked under the ceiling. Upon completion of all four bulkheads

the cylindrical shells can be folded downward through rotation about the hinge

lines A, B, C, and D. Deployment would be conducted in the reverse manner,

with deployment of the cylindrical .:hells accomp),Lshed by a pair of jackscrews

for each shell mounted in a non-obstructing area of the floor. Fuld-duwn and

flip-out of the panels and locking are accomplished by astronauts standing on

the strongback floors, securing latches such as shown in Section C-C and D-D

of Figure 2.4-8.

The structural design of such a structure to withstand prissure induced

loads appears entirely feasible. The structural behavior of the cylindrical

.. section is essentially that of a double-bubble pressure vessel. As such, the

cylindrical shell experiences essentially PR loads, except at the ends where

local discontinuity moments will exist. The flat bulkhead design presents no

foreseeable unsurmountable problems. There will be, of course, a need to

extend the strongback structure into the transition conical shell. The

^.- strorgbac:c structure must be designed to sustain any Combination of to:s of

press-!- pence, the flat surfaces of the strongback are designed to sustain a

oeak atmospheric pressure differential. It is expected that numerous

cra.r• ,t.se bulkheads will be used to permit satisfaction of this requirement.

The aunch attachment fittings will also be mounted onto the strongback

^s structure.

The construction ueed for the outer shell is identical to that defined for

the baseline design, and is shown in Figure 2.4-6. The end bulkhead panels

and strongback will be integrally machined panels stiffened by mechanically

fastened beams similar to the flat bulkhead of the STS orbiter crew module.

Figure 2.4-7 illustrates the packaging of all the identified baseline

configuratiun equipment. Much of the equipment is packaged in the volumes

between the ceiling and cylindrical shell and between ;he two tloors, i.e., in

the strongback. No problems pertinent to packaging of this equipment in these

regions was uncovered. However, it is apparent from Section A-A that equipment can only be packaged between the ceiling and strongback as shown on

Figure 2.4-7. This equipment must be rearranged and secured after deployment

by astronauts.

Any of the foregoing design concerns all app--ared to have acceptable

potential solutions. However, the requirement of maintaining the total module

leakage to no more than 0.44 kg/day (Reference 8) appears most difficult in

view of the following comments.

2-28
Q OR!CINAI- PACs 13

OF POUR QUAL.rry

ARRANGEMENT

EQUIPMENT A:TEP EVA

DEPLOYED ARRANGEMENT 1CTIVITY

Figure 2.4-7. Manned Module—Deployable

Hard Shel-- Equipment Arrangement

STAINLESS STEEL

BRAZED TO ALUMINUM

SHELL

i; A S 5

t

1( ALTERNATE TO SIDE HINGE

FLIP OUT FLa>

LEFT SECTION B-D -1

END PANEL --

SIDE PANEL

f L___

^^

SECTION A-A

FLIP OUT —

/ PANEL

r /END PANFL

FOLDDOWN

°STOWED 

t

FOLD DOWN PANEL _J LATCH

LA

F7Il

SECTION D-D SECTION C-C

ll^DEPLOYED ALTERNATE TO SEALS INSTALLED BY EVA

Figure 2.4-8. Deployable Module—Hard

Shell Alternate Seal CnnCeptS

tj

f

2-29

^J^w4^vcui
o The STS orbiter crew module structure me,nsured leakage rate is 4.9

kg/day, i.e., more than 10 times the requirement. While it is

appreciated that the crew module has a large number of electrical

penetrations and many large area windows the minimization of leakage to

0.44 k&/day represents a significant challenge.

o The magnitude of the challenge can be visualized by noting that the

total equivalent leakage area in a module at =. t mospheric pressure

cannot exceed 0.12 mm (0.0047 in) in diameter to maintain a leakage,

rate of no more than 0.44 kg/day. Some of this may be used up at the

docking port seals.

During the concept development several conceits to seal the cylindrical

hinge linos, end bulkhead hinge lines, and fold-out panel to cylinder mating

surfaces were investigated. These concepts are summarized in Figure 2.4-8.

These designs, however, cannot satisfy the above discussed requirements at the

intersection of two hinge lines. Therefore, the aeal recommended is the

design shown in Figure 2.4-6. This seal consists of a continuous polypropylene membrane across the flat bulkhead and ends of the cylindrical shell as

shown at the left and in detail as shown in Section B-B. The cylindrical

hinge lines A, 8, C, and D ar sealed by the hinge sideseal shown. This seal

concept nevertheless involves EVA installation of the membrane after

deployment of the structure.

An alternative seal configuration that would be in place during launch

would have to accommodate the inward motion of the fold out panels, i.e., be

sufficiently oversized. This may be the best solution yet, but is beyond the

scope of this study.

2.4.3 Inflatable Manned Module Design (Drawing 42712-8)

The major fnatures of this design concept are summarized by Figure 2.4-9

and 2.4-10. The same concept as that of the hard shell structure previously

described is used in regard to the end conical shells and center section

strongback. However, in this design the center structure consists of a pair

of inflatable shells attached to each side of the strongback by an end

attachment such as shown in Section B -B. The configuration of the center

structure is a pair of cylindrical shells capped by an ellipsoidal dome. The

cylinder radius is 2.08 m, and the dome deptn is 63: of the shell diameter.

The construction, recommended by GAG is a fabric consisting of Nomex

uniuirectional cloth coated with a Viton B -50 elastomer. Goodyear has

qualified this material to the requirements of the NASA Space Shuttle orbiter

crew cabin and Spacelab module.

An even number of multiple layers of the Nomex/Viton B-50 qualified

material will be laminated together until adequate strength is obtained. the

aluminum foil layer was defined as a flame barrier in earlier GAG studies.

2-30
nuleu sul uLM

/

SEAL

COMP RE$SEN 9 PLIE, OF MOMEx

CLOTH COATED WITH

VITON B-50 CLAnIOMER

(0 w THI(K)

( INFLATABLE CONSTRUCTION

;OIL

IAYCN

^r

V

46i

,^TC.O • I'

STONED

CONEIGUAATtUN

ORIGINAL PACE 19

OF POOR QUALI7Y

STONED INrLA;RBLE

DEPLOYABLE RroR AOU

STRONG BACK

-A

OUTER rjvrR/

TNERYAL COATING

POLYURCTNANE FOAM,

SECTfUN B-B

Figure 2.4-9. Manned Module Inflatable

fesign Concepts

AY ItN LVA ACTIV;TT

STONED MAIN DECK

EQUIPMENT B PARTITIONS

Figure 2.4-•10. !-fanned Module Inf1^ a Design

Equipment and Living Space Arrar• it

2-31
The mass distribution of this construction is shown below.

MATERIAL MASS

ITEM GM/CM2 LB/FT2

PLUMINUM FOIL LAYER U.UO2 0.0041

ADHESIVE 0.005 0.0102

STRUCTURAL LAYER 1.310 2.810

TASLAN INTERLOCKIN

LAYER 6 ADHESIVE 0.024 0.049

ONE I14CH POLYURETHANE FOAM 0.080 0.164

ADHESIVE 0.005 0.0102

OUTER COVER AND TBLRMAL

COATING 0 . 031 O.C54

TOTAL 1.52 3.11

Figure 2.4 - 9 also illustrates the deployable radiator configuration which

s.rves as a meteoroid bumper. This is necessary since radiators cannot be

mounted on the flexible structure.

Figure 2.4- 10 illustrates the mounting of the miscellaneous equipment into

the module. While less space is available in this module than the foregoing

hard-shell design because of the domes, all the baseline equipment identified

was packaged.

Pertinent to the inflatable design the issues of crew protection from

radiation material flammability and offgassing, construction life, crack

prevention propagation, and sealability are discussed as follows.

2.4.3.1 Radiation Protection

Since the inflatable shell material is projected to have a mass of 1.52

gm/cm2 , its mass is considerably above the 0.50 gm / cm2 considered adequate

for radiation protection of the crew.

2.4.3.2 Flammability and Offgassing

Since fluorocarbons have been tested successfully in the past for

flammability and offgassing, it was believed that Viton was a good choice for

an elastomer. The above items, in conjunction with the fact that the early

development units made by McDonnell Douglas had used Viton end had been

approved by NASA, led to the selection of Viton as the elastomer for the

Goodyear flex element. lue to processability the specific Viton selected was

Viton B - 50. Several comtinations of elastomer and cloth were submitted to

NASA via MDTSCO for evaluation. The results showed that the Viton B-50

2-32

2 

a
elastomer and Nomex cloth would pass the flammability and offgassing test as

called out in NHB8060.1A (Reference 9) by being subjected to an au^oclave

cure, post cure and vacuum bakeout. Therefore, it is anticipated that all

flammability and offgassing requirements of the program can be met.

2.4.3.3 Lifetime

None of the environmental conditions which the Spacelab tunnel flex

element must withstand were very severe compared to that of a Space Station

module. Most elastomers could survive the operational environment imposed on

the flex element. There were, however, two significant requirements that led

to the selection of Viton as the elastomer: 10-year life and offgassing/

flammability requirements. Viton is known for its long life at elevated

temperatures; and, since there was no elevated temperature requirements on the

flex element, the life would be well in excess of ten years. (Elevated

temperature is used for accelerated aging testing of rubbers.) Characteristic

dry heat resistance of volcanizates of Viton in continuous service is

considered to be greater than three years at 230°C.

For comparison, conventional elastomers would be brittle after one day at

230°C.

w

Other characteristics of Viton which make it an excellent selection are

its weatherability and fungus resistance.

GAC quotes their elastomeric products as being good for ten years under

reasonable, uncontrolled terrestrial storage conditions. This includes such

items as diaphragms, seals, pillow type tanks, fuel cells, flotation bags,

uprighting bags, impact attenuators and similar structures. After this 

storage, these products can provide five years of service, generally.

Diaphragms and seals usually provide 18 to 20 years of service.

Obviously, further tests are required for on-orbit life but it is believed

that the Viton B-50 elastomer has the best chance to meet the desired 20-year

lifetime.

2.4.3.4 Crack Propagation Prevention

GAC has conducted pressure tests on the flex section for the Spacelab

transfer tunnel to verify the constructions stability in the presence of

detectable damage. That same construction is proposed for this manned module

but with more plies since the limit circumferential loading is 2100 N/cm (120U

lt;in). With the increased number of plies in this manned module the stress

levels should be comparable. For a damage totally through the construction

the crack propagation resistance should be similar to that of the tested

design.

The purpose of the leak-before-burst test was to demonstrate that the

design of the flex element is stable under limit conditions with readily

detectable damage.

The leak-before-burst test involved inducing a 12.5 mm cut completely

through the flex element while at an internal pressure of 11 N/cm 2 . The cut

was oriented perpendicular to the yarns of the ply that was loaded.

2-33
t

The blade used was double edged, 12.5 mm wide, and came to a point. It

was mounted on a pneumatic actuator such that once the blade penetrated the

unit, it could be withdrawn and the pressure maintained for four minutes. The

leak-before - buret test was filmed for documentation. Visual examinatiou of

the cut after the test revealed that the cut did not grow beyond the initial

cut length.

The design was stable under limit pressure and displacement conditions

when a readily detectable damage consisting of a 12.5 mm slit through both•

plies was intentionally inflicted on the unit. This represented a successful

completion of the leak - before-buret test.

From the foregoing the suitability of the construction to sustain

micrometeoroid punctures is quite likely (particularly over the surface

shielded by the radiators. The suitability to resist space debris particles,

as for the metals, is not known.

2.4.3.5 Sealability

The two potential sources .f leakage are through the basic construction

and/or at the flexible seal. While GAC has demonstrated by test the leakage

j' of the flex section of the Spacelab transfer tunnel to be no more than 0.050

' kg/day that structure has a surface area approximately 1/100 that of the

manned module. In its favor this construction has an increased number of

plies. On the other hand this capability after 5, 10, 15, or 20 years in iM1

space is not known. While the perimeter ratio of the flexible seal is only

1/6 that of the manned module the radial deformation tending to cause leakage

E - is (for the same stress) approximately 20 times greater. The uncertainty is

apparent. What is certain is that the requirement of 0.44 kg/day is

representative of a total integrated hole size no larger than 0.12 mm (.0047

in.) if no leakage occurs elsewhere (at the access hatches). Until

F development teats are accomplished that are directly related to this

M

application the leakage rate is an unknown.

2.4.3.6 Meteoroid/Space Debris Issue

It was stated earlier that a micrometeoroid puncture is not expected to

precipitate a catastrophic failure. Figure 2.4-11 illustrates the pertinent

data concerning the potential for puncture and repair of the shell. Since the

radiator configuration serves as a bumper over most of the projected area

(Figure 2.4-9) the analysis is performed on that basis.

For the surface area quoted, and a 20-year exposure, there is a

probability of 0.99 that no impact will occur for a particle size greater than

9.9 mm in diameter. This particle, for no puncture, requires an aluminum

single sheet thickness of 29.5 mm. For a bumper ballistic efficiency of 5,

the equivalent thickness required of this construction is 5.9 mm (0.23 in.).

The polyurethane foam, having a density of 0.031 gm/cm 3 (2 lb/ft3),

behaves like an aluminum sheet 4.3 mm thick. This is based on the following

data provided by GAC.

"Based on hypervelocity particle impact tests conducted by Goodyear

Aerospace and on tests conducted at the micrometeorite testing facility at

Wright-Patterson AFB, flexible foam with a .015 gm/cm 3 (1.0 lb/ft3)

density has been selected as a suitable barrier material. The tests at

Wright-Patterson were conducted with a particle that had an average mass of

0.005 g, traveling at 8232 m/sec (27,000 ft/sec).

2-34

..^.- .._ .. -.. - .^.-... .. .. n.^...^ .^._.. ... [u^^... eve iv..... ^.4..
'O

DIRECT'('

g

-9

N

-11

s MICROMETEOROID

- : I2

(SP 8013)

LOG/ •-19.37

13

-1.213 LO G n

LL

-lA

-IS

-12 -10 -8 -6 -A -2 0 2

LOG MASS (A) C UMS

OINI UN I.

SPACE DEBRIS

LOG B • -12.00

- 0.15 LOG •

(KESSLER. 1"0

ORIGINAL PAGE 9

OF POOR QUALITY

• PA

N S. D.

VELOCITY (km/t ) 20 10 [DENSITY (q^/c•^1 0 5 2 B

y

• YULIIERABLE

SURFACE AREA;

600 n2

• VU.NERIOLE

PROJECTED

AREA:

110 .2

• PARTICLE SIZE (PROS. OF NO IMPACT • Po

25.4

r

v 20.3 SPACE O%S

(PO • 0.99(

= 15.2 MICRONETEOROID

p 0 .99)

10.2

—(P

R <

-

,^.. 5,1 ^DEBRIS

(PO • 0.9L')

0 

ID 20

DESIGN TIME (t) YEARS

Figure 2.4-11. Manned Module Man-Made Debris

and Micrometeoroid Design Implications

2-35
Both series of testa indicate that a foam barrier has the same effectivene99 as a single sheet of aluminum with 15 times the mass per unit area."

It is expected that provision of a foam with increased density can provide

the required stopping power. Based on this it can be said that for a life of

20 years the probability of a puncture occurring that requires repair is

0.01. Such a puncture can be repaired.

The problem of space debris for this structure as the hard shell, is

ill-defined. To assure that no catastrophic crack prupagar:un w^u'd

initiated by impact of a space debris particle, the size, density, and speed

properties of the particle must be explicitly defined and b :11..dt,c eats nest

be performed upon candidate constructions in pressur; .ed conf1yu,1tir,-,j

A brief summary of the foregoing is as follows:

o There is no on-orbit data to a confirm 20-year-.life suitability of t r,

wall construction utilizing a Viton B-50 elastomer.

o CAC prior test data do net resolve the question of the 10 to 20 year

seal capability of the basic wall construction (maximum leakage

,.. requirement of 44. kg/day).

• Micrometeoroid impact is not expected to be a problem with the use of

radiators as a bumper.

• The construction suitability fur space debris is not known. 

• Crew radiation shielding, material outgassing, and nonflammability

requirements are satisfied by construction.

2.5 DEPLOYABLE MODULE MISCELLANEOUS COMMENTS

The following general comments are applicable to both deployable module

designs.

o Both deployable configurations can satisfy the manned module crew

safety, volume requirements, design commonality, and growth evolution

requirements.

o Each configuration requires 16 m of cargo bay, compared to 25.3 m for

the two modules. The potential launch cost savings is $39.0 million

(1984 dollars).

o An extrapolation of the Space Station baseline manned module Wass

(Table 2.4-2) results in a potential mass of the deployable modules of

23 kg to 25 kg.

o For the c.g, at orbiter Station 982, the maximum cargo launchable mass

is 19 kg, but an aftward shift of the c.g. of 0.89 m can be

accomplished with 4 to 5 kg ballast or additional equipment. This

permits the launch of a 29.5 kg mass.

o Both designs require additional on-orbit cost of installation of

partitions and systems at a rate of $15,000/hour.

2-36
€mil

2.6 MANNED MODULE SELECTION

o The capability of either hard-shell or inflatable deployable designs to

satisfy the minimum leakage requirement is very questionable. A 2.2

kg/day additional leakage amounts to 16,600 kg over 20 years which

alone negates the initial launch savings.

o While the inflatable design has the beat potential, with technology

development, to satisfy the leakage requirements, no data exist to

confirm a 20-year life of the construction.

o Any launch cost savings will be reduced by increased structural

complexity, seal technology development, and on-orbit cost of EVA for

equipment installation.

Hence, it is recommended that the conventional baseline design be

maintained until the technology development tf seals can provide mssurance

that the requirement of .44 kg/day can be satisfied. Also, no further

conceptual effort is recommended until Space Stotion configurations,

requirements, and criteria are more accurately defined.

2-37
3. REFERENCES

1. SASP Deployable structure Ground Test Project, Test Article Design

Requirements, papared by MSFC Preliminary Design Office, Program

Development, January 1980.

1. Satellite Power Systems (SPS), Final Report (Exhibit E Report), LSST

Systems Analysis and Integration Task for SPS Flight Test Article,

Rockwell Report, SSD 80-0102, dated August, 1980.

3. Space Construction Systems Analysis, Part 2, Final Report, Platform

Defir,ition, NAS9-15718, Rockwell International, April 1980.

4. Federal Standard FED-STD-H28/12, Screw Thread Standards for Federal

Services, August 31, 1978.

5. Engineering Mechanics, Ferdinand L. Singer, Harper and Brothers, 2nd

Edition.

6. SD&P Mechanical Systems Design and Parts, Rockwell Volume III Publication

543-G-25, April 1983.

7. Space Station Needs, Attributes, and Architectural Options Study,

SSD83-0032, April 22, 1983, Contract NASW 3683.

8. Space Station Program Description Document, System Requirements and 

Characteristics, Book 3, JSC, November 1982.

9. Preliminary Study to Adapt Inflatable Structures to a Space Station 

Habitat and OTV Hangar, GAC 19-1585, February 1983.

10. Midterm Re^ l iew - Development of Deployable Structures for Large Space

Platform Systems, SSD 83-0004, January 28, 1983.

3-I

ar a
CONTENTS

i.

I

Section page

1 CONCEPT DEVELOPMENT 1-1

1.1 DEPLOYABLE PLATFORM SYSTEMS REQUIREMENTS 1-1

1.1.1 Strength and Stiffness Requirements 1-3

1.1.2 Power and Data Utilities 1-10

1.1.3 Fluid Utilities 1-15

1.1.4 Control System 1-18

1.1.5 Structural Temperatures 1-20

1.1.6 Servicing 1-21

1.1.7 Orbiter Integration 1-21

1.1.8 Environment 1-22

1.1.5 Payloads, Propulsion Modules, ACS Modules 1-24

1.1.10 System Fainting Accuracy 1-25

1.1.11 Requirements Perspective 1-25

1.2 GENERIC DEPLOYABLE SPACECRAFT CONFIGURATIONS 1-27

1.3 DEPLOYABLE PLATFORM SYSTEMS CONCEPTS 1-31

1.3.1 Deployable Trusses . I 1 1-31

1.3.2 Utilities Installations 1-40

1.3.3 Deployment Mechanization Concepts 1-52

1.3.4 Main Housings and Adapters 1-58

1.4 DEPLOYABLE PLATFORM SYSTEMS CONCEPT INTEGRATION 1-60

1.4,1 Structural Sizing Method and Approach 1-62

1.4.2 Eight Candidate Building-Block Concepts 1-65

1.4.3 Orbiter Packaging 1-77

1.4.4 Thermal Analysis 1-84

1.4.5 Mass Properties Analysis 1-8'

1.4.6 Cost Analysis 1-87

2 MATERIAL DATA BASE 2-1

3 IDENTIFICATION OF TECHNOLOGY DEVELOPMENT NEEDS 3-1

4 CONCEPT SELECTION 4-1

4.1 M..JOR CRITERIA 4-3

4.2 METHODOLOGY 4-4

4.3 DESIGN VERSATILITY OF STRUCTURAL CONCEPT 4-5

4.4 COSTS FOR GENERIC PLATFORM ('EO and GEO) 4-12

4.5 THERMAL STABILITY OF STRUCTURAL CONCEPT 4-13

4.6 METEOROID IMPACT SUITABILITY OF STRUCTURAL CONCEPT 4-13

4.7 RELIABILITY OF DEPLOYMENT OF BUILDING BLOCK 4-15

4.8 PREDICTABILITY OF PERFORMANCE OF STRUCTURAL CONCEPT" 4-17

4.9 ORBITER INTEGRATION SUITABILITY 4-18

4.10 SUMMARY OF POINTS AND GRADING 4-20

4.11 CONCEPT SELECTION 4-23

5 DEPLOYABLE VOLUME ENCLOSURES 5-1

5.1 HABITAT MODULES 5-1

5.1.1 Habitat Requirements 5-1

5.1.2 Radiation/Meteoroid Shielding Review 5-2

5.1.3 Habitat Design Approach 5-3

5.1.4 Candidate Concepts for Habitat Modules 5-4

5.1.5 Preferred Concepts 5-5

v

PRECEDING PAGE BLANK NOT FILMED
Section Page

5.2 TUNNELS 5-9

5.3 uTV HANGAR 5-9

5.3.1 OTV Hangar Requirements 5-9

5.3.2 OTV Hangar Design Approach 5-10

5.3.3 Meteoroid Impact Analysis 5-11

5.3.4 Candidate Concepts for OTV Hangar )-13

5.4 MAJOR DESIGN ISSUES 5-1/

5.5 POTENTIAL TECHNOLOGY DEVv LOPMENT NEEDS 5-18

5.6 CONCLUSIONS 5-19

6 REFERENCES I•-I

I

vi

L'4
4

ILLUSTRA'T'IONS

Figure Page

I Deployable Platform Systems Concept 3

2 Focus and Miscellaneous Configurations 4

3 Study Logic-Part I 5

4 Summary of Deployable Platform System Accomplishments C

5 Square Truss with Modified Longerons (Concept 6A) 8

6 Possible Configuration (Concept 6A) 9

7 Summary of Deployable Volume Enclosures Accomplishments 10

1.1-1 Four Configurations Extracted from Three Focus Mission

Studies 1-1

1.1-2 1.0dules 1 and ', f GSP Alternative 4 Configuration 1-2

1.1-3 Flexural Stiff- s Vs. Frequency (Hz) for ASASP 1-4

1.1-4 ASASP Configuration-Srationkeeping Thrust 1-5

1.1-5 Module 1 (Alternative 4) Attitude Control Thrust 1-6

1.1-6 Module 2 (Alternative 4) Attitude Control Thrust 1-7

1.1-1 GEO Trinsfer T/W and Number of Perigee Burn Implications

on Propellant Mass . 1 1-9

1.1-8 Platform Power Distribution Subsystem 1-12

1.1-9 Pump Characteristics . I 1 1-16

1.1-10 Maximum and Minimum Temperatures _ I-20

1. 1 -11 Radiation Dose Rates at CEO-Functi;ns of Shield

Thickness I 1 1-22

1 1-12 Natural Van Allen Belt Dose Rates, 0° Inclination Orbit 1-23

1.1-13 Natural Time Averaged Meteoroid Flux at i. AU form the Sun. 1-24

1.1-14 kockwell Communications Configurations and SASP 1-26

1.2-1 Generic Linear Platform 1-28

1.2-2 Generic Srea Platform 1-30

1.3-1 Deployable Platform S y stem Components -

Basic Building Block 1-32

1.3-2 Review of Applicability of Industry Existing and Rockwell

IR&D Designs 1-33

1.3-3 Candidate Deployable Platform Structure: Cuv-epts 1-34

1.3-4 Concept 1-Pentahedra! Truss 1-36

1.3-5 Concept 2-Warren Truss 1-36

1.3-6 Concept 3-Tetrahedral Truss 1-37

1.3-7 Concept 4-Truss 1-37

1.3-8 Concepts 5 and 7-Truss 1-38

1.3-9 Concept ^,-Truss 1-39

1.3-10 Concept 8-Truss 1-39

1.3-11 Utilities in One Plane 1-43

1.3-12 Utilities with Overlap and Increased Bend Radius 1-43

1.3-13 Maximum 3ending Strain-No. 2 Conductor 1-46

1.3-14 Maximum Bending Strain-No. 0 Conductor 1-47

1.3-15 Minimum Bending Curvatur^ Ratio-1 x 7 Structure 1-•48

1.3-16 Minimum Bending Curvat4ie '.atio- 1 x 19 Structure 1-49

1.3-17 Copper Filament Fatigue D.:a 1 1-49

1.3-18 Latching Mechanisms between No. 2 and No. 3 Battens 1-54

1.3-19 Longeron Latching Mechanisms . I 1 1-54

vii
M

c /sI

Figure Page

1.3-20 Deployment Rails and Mechanism 1-55

1.3-21 Detail of Deployment Rail 1-55

1.3-22 Deployment Rail Issue 1-56

1.1-23 Fixed Rail Method 1-57

1.3-24 Penetration Method 1-57

1.3-25 Spar Boom Method 1 -5 7

1.3-26 Candidate Housing and Adapter Concepts 1-59

1.4-1 Concepts for Packaging of Generic Platform 1-60

1.4-2 Structural Concept Dimensiuns . 1 1-61

1.4-3 Final Concept for Packaging (Concepts 2, 6, and 3) 1-62

4-4 Example of Column Sizing Data 1-64

1.4-5 Concept 1 1-65

1.4-6 Installation of Utility Trays (Concept 1 ) 1-66

1.4-7 Concept 2 1 67

1.4-8 Installation of Utilities (Concept 2) 1-67

1.4-9 Concept 3 1-68

1.4-10 Installation of utilities (Concept 3) 1-69

1.4-11 Concept 4 1-70

L.4-12 End View of Folded Configuration 1-70

1.4-13 Utilities Installation (Concept 4) 1-71

1.4-14 Concepts 5 and 7 1-71

1.4-15 Concept 7 1-72

1.4-16 Utilities Installation (Concepts 5 and 7) 1 -74

1.4-17 Concept 6 . 1 1-74

1.4-18 End View Stowed 1-75

1.4-19 Utilities Installation (Concept 6) 1-75

1.4-20 Concept 8 . I I 1 1-76

1.4-21 Utilities Installation (Concept 6) 1-76

1.4-22 Orbiter Packaging-Concepts 1 and 4 1-77

1.4-23 Platform Deployment-Concepts ' and 4 1-78

1.4-24 Orbiter Packaging-Concepts 2, 6, and 8 1-79

1.4-25 Platform Deployment- :oncepts 2, 6, and 8 1-80

1.4-26 Orbiter Packaging-Concept 3 1-80

1.4-27 Platform Deployment-Concept 3 1-81

1.4-28 Orbiter Packaging-Concepts 5 and 7 1-82

1.4-29 Platform Deployment-Concepts 5 and 7 1-83

1.4-30 Structure Artangemeet for Thermal Gradient Analysis

(Dimensions in Meters) 1-85

1.4-31 Thermal Analysis Configuration 1-85

1.4-32 Approach to Reduce Thermal-Induced Truss Rotation 1-87

1.4-33 Orbiter Payload Mr.w3-Launch/Landin 6 Constraints 1-89

1.4-34 ASASP Modal Frequency Variation with Diagonal End Joint

Design 1-93

1.4-35 Model for "Implication of Joint Slop" Analysis 1-94

1.4-36 Equivalent Structure Frequency due to "Joint Slop" 1-95

3.0-1 Logic Diagram-Special Technology Needs 3-2

4.2-1 Point Evaluation Methodology 4-4

4.11-1 Concepts 6 and 8 14-24

4.11-2 Concept 6A 4-25

5.0-1 Deployable Volume Enclosures 5-1

i

viii
r'

Figure cage

5.1-1 HaA tar Radiation/Meteoroid Implications 5-3

5.1-2 Candidate Deployable Habitat Module Concepts 5-4

5.1-3 Preferred Habitat Module Concepts 5-5

5.1-4 Habitat Module (Concept 3) 5-6

5.1-5 Habitat Module (Concept 4) 5-7

5.'-6 Habitat Module (Cc,icept 5) 5-8

5.3-1 OTV Construction 5.11

5.3-2 OTV Meteoroid P rotection/Hangar Cost 5-12

5.3-3 OTV Hangar Concepts 5-13

5.3-4 OTV Hax,gar (Concept 1) 5-15

5.3-5 OTV Hangar (Concept 6) 5-16

5.3-6 OTV Hangar (Concept 7) 5-17

ix
TABLES

Table Page

1.1-1 Focus Mission Limit Strength and Stiffness Requirements 1-3

1.1-2 Adopted Loads (Limit) and Stiffness Requirements 1-8

1.1-3 Generic Power and Data Utilities Requirements 1-10

1.1-4 Reference Payload Group 1-11

1.1-5 Effects of Differing Power Load Assumptions Upon Wire Count 1-13

1.1-6 TCS Requirements for 2-cm Fluid Lines 1-17

1.1-7 Control System Equipment Requirements 1-19

1.3-1 Matvix of Truss Variations 1-35

1.3-2 Comparison of Utilities Installation 1-42

1.3-3 Electrical and Fluid Utilities Characteristics 1-44

1.3-4 Permissible Electrical Conductor D/d Bend Ratios 1-45

1.4-1 Structural Sizes to Adopted Strength and Stiffness

Requirements 1-63

1.4-2 Thermal Gradients between Longe:ons 1-84

1.4-3 Candidate Concepts Generic Platform Mass Summary 1-88

1.4-4 Comparative Mass (kg x 10-3 )-Generic Platform 1-90

1.4-5 Fabrication Cost Data ($ Million) 1-91

2.0-1 Glossary of Terms for Data Base . 1 2-2

2.0-2 Metallic Materials for Structural Members 2-3

2.0-3 Structural Member Candidate Composite Materials 2-4

2.0-4 Metallic Materials for Fittings and Springs 2-5

2,0-5 Non-Metallic Materials for Fittings and Springs 2-6

2.0-6 Materials for Tension Cables 2-7

2.0-7 Materials for Electrical Conductors 2-8

2.0-8 Diffusion Barriers for Fluid Lines 2-9

2.0-9 Diffusion Barriers for Bellows 2-10

2.0-10 Vibration Damping Materials 2-11

2.0-11 Thermal Control Coatings 2-12

3.0-1 Special Technology Needs Summary 3-1

3.0-2 Summary of Potential Technology Requirements 3-3

3.0-3 Validation Questions 3-3

3.0-4 Priority Rating Criteria 3-5

3.0-5 Priority Rating Results 3-6

3.0-6 Sensitivity Trade Results 3-7

4.1-1 Major Criteria Used in the Selection Process 4-3

4.3-1 Electrical Accommodations (GEO) 4-6

4.3-2 Electrical Accommodations (LEO) 4-7

4.3-3 Fluid Utilities Accommodation (LEO and GEO) 4-8

4.3-4 Structural Materials Variation (LEO and GEO) 4-9

4.3-5 Strength and Stiffness Accommodation (GEO) 4-10

4.3-6 Strength and Stiffness Accommodation (LEO) 4-10

4.3-7 Platform Construction (LEO and GEO) 4-11

4.4-1 Costs for Generic Platform (LEO and GEO) 4-12

4.5-1 Thermal Stability of Structural Concept 4-13

4.6-1. Probability of Meteoroid Damage 4-14

4.6-2 Meteoroid Impact Suitability 4-14

4.7-1 Reliability of Building-Block -)eployment 4-15

xi

PP.ECEDING PAGE BLANK NOT FILM173
w

i

Table Page

4.7-2

4.8-1

4.4-1

4.10-1

4.10-2

4.10-3

4.10-4

4.10-5

Number of Joints in Basic Structure for Generic Platform

Predictability of Performance

Orbiter Integration Suitability .

Total of Normalized Points (LEO)

Total of Normalized Points (GEO)

Total of Normalized Points (LEO) —Sensitivity Trade .

Total of Normalized Points (GEO) —Sensitivity Trade .

Summary of Points and Grading, LEO and GEO .

4-16

4-17

4-18

4-20

4-21

4-22

4-22

4-23

xii

E

Y
ki

INTRODUCTION

During the next decade, a revolut 4 . on in spacecraft design will occur,

resulting in large apace platforms the; will accommodate multiple payloads.

Cost savings to users will occur through sharing of spacecraft utilities, ease

of servicing, and the ability to change payloads. In addition, for

geosynchronous communication payloads, platforms will reduce the crowding of

thd•a important orbital location.

The development of deployable platform systems is the most significant

technology step in the direction of realizing these platform capabilities.

Althougb the Shuttle allows payloads with much larger dimensions than other

launch systems, the large dimensions of the platforms will, nevertheless,

require extensive structural deployment to package it within the orbiter.

Much of the previous industry effo't in large structuraa concentrated on

orbital construction or erection. A recent Rockwell study, Space Construction

System Analysis Study (NAS9-15718), developed the detail necessary to

understand the difficulty of joining machine-made beams and integrating

spacecraft utilities in orbit from the Shuttle. These studies pointed up the

difficulty of erecting or constructing large platforms from the Shuttle.

Consequently, ground integration of utilities into a deployable structure was

selected by NASA as the first logical approach to platforms. NASA /MSFf has

preparad a five-year plan to achieve technology readiness of deployable

platform systems by FY 1986. Phase I of that plan is to identify the one or

two most suitable deployable platform systems (Pest 1) and establish all the

information necessary to plan and execute a follow- on-hardware development

teat program (Part II). On October 16, 1981 Rockwell initiated the study

activities in support of Part I, wi'.h completion 9 months later - on

July 16, 1982. Sections 1 through 4 of this interim report describe the

pertinent study accomplishments for Part I.

Future missions such as the manned apace platform require both pressurized and

unpressurized volumes, respectively, for crew quarters and manned

laboratories, and maintenance hangars. Deployable volume enclosures can

minimize launch costa and enable use of volumes greater than those which can

be transported by the Space Shuttle orbiter. On April 16, 1982, Rockwell

initiated a 3-month add-on study of Deployable Volume Enclosures with the

objective of identifying generic concepts for manned habitats, tunnels and OT"

hangars. The accomplishments of that study are described in Section 5 of this

interim report.

During the course of this conceptual study, 31 drawings were completed. The

drawings are provided in a separate document, i.e., Volume II, SSD 82 -0121-2,

August, 1982.

1

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DESIGN APPROACH

The design approach employed in this study of deployable platform systems is

developed to satisfy the objectives and guidelines as follows:

Objectives

- o Development and evaluation of generic deployable platform system

concepts applicable to the focus mission LEO/GEO applications and

foreseeable applications for the 1990 to 2000 time period

• Establishment of a materials data base for structures and utilities

systems compatible with LEO/GEO applications

• Identification of the new technology development needs, schedules,

and costs

• Systematic/traceable selection of the one or two most suitable

generic concepts

Guidelines

o Automatic deployment - minimum EVA

o Platform system - not just a structure

o FY 1.986 technology readiness - test-proven hardware

o Generic - not a basepoint design

• Versatility - can be used to build spacecraft of different

configurations; "building block" approach (self-contained modules)

• Distinction between LEO and GEO designs

o Adaptable for a wide range of payloads

Of the abova, the guideline of automatic deployment was the major design

driver.

The entire concept development is directed toward automatic deployment of the

entire platform system without use of a construction fixture or EVA. This

approach is illustrated in Figure 1. The deployable platform system is

comprised of several building blocks which are preassembled before flight to

provide the defined spacecraft confl iration. Each building block contains

the deployable truss; tntegrated power, data, and fluid lines; a main housing;

adapter; and deployment mechanization system. Attachments for payloads,

reaction control system (RCS) modules anR docking ports are provided on the

main housings or adapters. The packag ,,d platform system can be integrated

into tie orbiter with a pallet that serves as a control system module and

contains batteries and telemetry, tracking, and command (TT&C) equipment. The

system can be removed from the orbiter by the remote manipulator system (RMS)

and placed on a handling and positioning aid (HAPA) or, depending on

configuration, size, and shape, utilize both RMS and HAPA during the initial

2

l .... .. fir... _... .
ORIGINAL PAGE 18

OF POOR QUALITY

TYPICAL

^jLDING

BLOCKS

deployment stages and during control system checkout. Subsequent to control

system checkout, the platforms can be translated away from the orbiter (100 to

200 meters) for completion of the automatic deployment phase.

T:.MI

sl.,.

SnM

BUILDING

BLOC ``^I??^

,'

,YY5^4I1eeJJ

,r ADAPTERS

STRUCTURAL ^Q

CONCEPTS

UTILITIES

INTEGRATION HOUSINGS MECHANIZATION

_ Figure 1. Deployable P1 'form Systems Concept

T

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ice. i

Focus Mission Cont

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SPS TEST ARTICLE

eous Configurations

ROCKWELL

comm, 

ORIGINAL PAGE IS

OF POOR QUAL"

A major factor in the study approach is the requirements. Unquestionably, the

regime of strength, stiffness, and utilities requirements is a major

consideration. The requi rements are extracted directly from the three focus

mieeions (Figure 2, References 1, 2 and 3) and supplemented by Rockrell

analysis and recognition of other potential applications.
w

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ORIGINAL PAGE 19

OF POOR QUALITY

The study approach was directed according to the study plan shown in Figure 3.

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Figure 3. Study Logic—Part I

SUMMARY

This section briefly summarizes the ma j or study accomplishments.

Figure 4 summarizes the deployable platform system acccomplishments. In

Subtask 1 . 1, the strength, stiffness, electrical, and fluid utilities

requirements and additional requirements encompassing structural temperature

limits, guidance and control, pointing accuracy, and propulsion were

established. In 3ubtaek 1.2, the generic configuration ( to serve as a study

tool) consisting of linear members was configured depicting the platform

size,general arrangement, utilities distribution and docking ports. Also,

investigation of an area platform constructed in such a manner that plate

behavior is developed, resulted in termination of that concept for the reasons

5
,.:.;1.......

• SUBTASK 1 . 1

REQUIREMENTS -.Ems""`-"""^ mw _sy,

ORIGINAL PAGE 19

OF POOR QUAMY

• SUBTASK 1.2 GENERIC PLATFORM

• SUBTASKS 1.3 AND 1.4

8 INTEGRATED BUILDINGBLOCK CONCEPTS

• SUBTASK 1.4

ANALYSIS FOR COMPARISON

ti. ar

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I+ ^^ 1►^ *TASK 2.0 MATERIALS

1

^: 1 y> SELECTION/DATA BASE

m %Ls-%aq=l m sasr F

^•.^ ^.^ss_zxa na¢:rs.. TECMNnLO

SELECTED DESIGN r-riwr+.rcr^

•TASK 4.0

CONCEPT SELECTION

PROCESS '--

Figure 4, Summary of Deployable Platform System ACCOmplishments

6
delineated in Section 1.2. In Subtasks 1.3 and 1.4, eight candidate structure

concepts with integration of utilities, and concepts for deployment

mechanisms, houainge to contain the mechanisms and structure, and adapters for

payloads were developed.

All the concepts were integrated into eight building-block concepts. The

eight building blocks were compared on the basis of packaging the generic

platform into the orbiter, with supporting parametric structural, thermal,

Maas properties, meteoroid impact, and cost analyses. The results of these

analyses are tempered by the consideration that the generic platform and

requirements represent one design condition in the spectrum of platform

applications. 7n fact, the generic platform and adopter. strength, stiffness,

and utilities requirement are at the u-)per end of the spectrum of foreseeable

requirements for most platforms.

In Task 2, candidate materials most suitable for the deployable platform

system components were identified, with establishment of a data base for these

candidates. The data base is comprised of 10 tables of mechanical and

physical properties.

In Task 3, new technology development needs were identified, prioritized, and

scheduled, including the development coat estimates. Of the 16 new technology

items identified, no show-stopper is apparent.

The foregoing data were used in the concept selection process of Task 4. The

concept analyses compared the designs on the bdsis of design versatility,

coat, thermal stability, meteoroid impact suitability, reliability,

performance predictability, and orbiter integration suitability. This

selection process methodology, in conjunction with judgmental evaluations,

resulted in the selection of Concept 6A, (Figure 5). The major features of

Concept 6A are summarized as follows.

• Building-block approach for automatic deployment of platform systems

• Square shaped trues - most suitable for inter-building-block

attachments; mounting of payloads, docking ports, and propulsion

modules; and provides redundancy for meteoroid impact.

• Circular tubes for all truss members - minimum cost construction for

graphite composite construction

• Minimal complement of utilities mounted on longerons

o Trays for mounting of large complement of utilities - ease of

initial installation, repair, replacement during total ground

fabrication period - minimum trues structural design constraints

imposed by utilities integration.

o Square-shaped housing with reciprocating deployment mechanism

I

o Bart -by-bay deployment (to facilitate identification of deployment

problem, if it occurs)

d. f a

;

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UTILITIES INSTALLATION

9

ORIGIN .,AL FAUN

OF pppR QUA_rri

• ALL MEMRER AR! ROUND TUNS MAIN HOVSNG

• LONGERONS WITH CENTER HINGE JOINTS

*NO OFFSET LOADS AT JOINT ICJ

• MTTENS CROSS PACED N TENSION STRA ►f

ADAPTER

^" ^_h^ R 

TRl1SS \ `

NT RAILS

END VIEW FOLDED DETAIL OF JOINT

Figure 5. Square Truss with Modified Longerons (Concept 6A)

o Rail system for root strength during bay-by-bey deployment permits

orbiter berthing and orbiter vernier reaction control system (VRCS)

firing (if necessary).

o Adapters for mounting of payloads with automatic electrical

connector interface

o Payloads and propulsion modules attached using RMS

An example of a possible configuration achievable with this concept that is

constructed of 1.75 and 2.75 m deep trusses is shown in Figure 6.

The major conclusions drawn in this study of deployable platfoTme systems are

as follows:

• Deployable platform systems technology readiness for FY 1986 period

is quite feasible ( with appropriate funding)

• Tha building- block concept utilising Concept 6A can effectively be

used to construct LEO/GEO platforms

• Deployment is accomplished with orbiter RMS and /or NAPA without use

of construction fixture

o NASA /MSFC goal of automatic deployment of platform system (not

including payloads, RCS modules) is achievable

8
7

I ORIGINAL PAT, 131

OF POUR QUAL"

PAYLOAD RAY

1.75M SQUARE TRUSS ^TUWE^

PAYLOAD BAY 7.I5M SOIIARE TRUSS IOM

^4.57M DIA

S0M

i

J ~^ 90M DEPLOYED

{ -NO REOR IE NTAt ION OF

4 ` BUILDING BLOCK REQUIRED

Figure 6. Possible Configuration (Concept 6A)

o All candidate building-block concepts, applied to generic platform,

are packageable into one orbiter and are well within 20,000 kg

launch capability (28.5 - inclination, 210 ami)

o The selected concept accommodates the upper regime of platforms size,

'Itilities, and adopted strength requirements and is applicable to

arge range of reduced requirements

o commodation of adopted stiffness requirements is dependent on the

ryoolution of "Joint Slop" issue ( moat platform applications will

have stiffness requirements well below adopted values)

o No foreseeable unresolvable technology development requirements

o Major extent of technology development for selected concept is

applicable to alternate candidate concepts

Figure 7 summarizes the major accomplishments of the Add - on Deployable Volume

Enclosures Study. The Space Station studies currently being performed at

Rockwell are used for the establishment of requirements and applications for

manned habitats and OTV hangars. Eight habitat and seven OTV hangar candidate

configurations were developed. Three preferred configurations were identified

for the manned habitat9 and developed in further detail ( Section 5 )• Three

9
OTV CANDIDATES

O

C) ..

Figure 7. summary of Deployable Volume Enclosures Accomplishments

10

' 4

F

1

UNIG)NAL PAGE 19

OF POOR QUALITY

DEPLOYABLE VOLUME

curt nC11GFS

HABITAT REQUIREMENTS

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CANDIDATES

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OTV hangars were also developed in further detail. Further, the major design

issues and new technology development requirements were identified.

The major conclusions drawn from this study are an follows:

o The application to habitats appears attractive - large useful

volumes are achievable

• Metallic structures (can be sealed) can provide ample meteoroid,

radiation protection, and equipment mounting surfaces, but are

constrained by pressure loade/packaging requirements

• Inflatablee alone are not sufficient - hard structure is required

for mounting of consoles, orbiter and apace station integration, and

heat rejection

o Inflatatles in conjunction with rigid core module provide a variety

of feasible large volume designs provided

Materials are suitable to crew safety/apace environment; foam

micrometeoroid stopping power is comparable to existing data;

repair of punctures or use of meteoroid bumper is feasible; and

adegiate protection of the crew from radiation can be provided.

• Hangar requirements are ill-defined - OTV meteoroid protection alone

is not sufficient justification

• Most attractive OTV hangar concepts appear to be metallic

deployable/erectable or inflatable with foam core (provided

stiffness is adequate)

a

11
SPS TEST ARTICLE

GSP ALTERNATIVE 4

Figure 1.1-1. Four Configurations Extracted from

Three Focus Afission studies

1-1

7

I)

ORIC=INAL PAGE 0

OF POOR QUALM

1. CONCE?T DEVELOPMENT

T;iis section describes the study concept development accomplishments which

include establishment of platform system requirements (1.1), development of a

generic spacecraft configuration (1.2), development of the building-block

component concepts (1.3), and integration of thee* components into eight

candidate buildirg blocks (1.4). Section 1.4 also describes the appl'_cation

of the building blocks to the generic platform; packaging of the platform into

the orbiter; and the comparative structural, thermal, mass properties, and

cost analysis performed for the Concept Selection Trade Study ^4).

1.1 DEPLOYABLE PLATFORM SYSTEMS REQUIREMENTS

This section describes the generic platform system requirements extracted

aither directly from the t..rae focus mission studies (References 1, 2, 3

and 4) or determined by supplemental analysis from the data provided therein.

In the course of performing the supplementary analysis pertaining to the ASASP

and GSP requirements, the study managers of these studio , sere contacted for

additional information and/or clarification.

The three focus mission studies provide four platform configurations r. q sho,,n

in Figure 1.1-1. Since Geostationary Space Platform (GSP) Alternative 1

715 m
MODULE ;

ORIGMIAL PAGE IS

OF POOR QUALITY

represents a configuration entirely nonr . table with either astromaets or

super-masts, emphas ! , was placed upon the uSP Alternative 4 configuration which

represents the high . . of the spectrum is terms of size, strength and

stiffness requirements. This configuration is comprised of three modules

individually transferred to and joined together in GEO. Two of the largest of

these three modules are shown in Figure 1.1-2. The information shown in

Figure 1.1-2 ,d mass distribution data were extracted from the several books

of Vplume II, deference 5•

^o

Figure 1.1-2, Modules 1 and 2 of 

GSP A1t#vxnative 4 Configuration

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OF POOR QUALrN

Additional to the supplementary analysis required to completely define the

requirements for the missions, it was necessary to understand the derivation

of the most severe requirements in orler to assess the applicability of these

requirements. The aaceesment was tempered by the consideration that this

study is a tec:,"ology development contrast, with extension of the technology,

as practicable, being a major goal. Further, many of the requirements stem

from conditions where two equally possible alterna'ivea to the design can

exist and are dependent on the particular spacecraft configuration and systems

trades. For example, thermal control of payloads can be accomplished with

dedicated rndiatora at each payload or with the use of a central radiator and

coolant lines integrated into the deployable struetu:e. In such cases, in the

the absence of any other information, the alternative requirement that

advences the technology (in this case, integration of coolant lines into the

structure) is included.

1.1.1 Strength and Stiffness Requirements

Table 1.1-1 summarizes the strength and stiffness requirements for the focus

mission spacecraft configurations. The data shodn were obtained by a

combination of direct extraction from the focus mission documents and

supplemental hand calculations. The data directly extracted are identified

with a subscript "d".

Table 1.1-1. Focus Mission Limit Strength and Stiffness Requirements

PLATFORM I

PARAMETER SPS ASASP OSP ALT. 1 GSP ALT. 4

DEPLOYABLE STRUCTURE. MODULE.

*FLEXURAL STIFFNESS (NMI ) 17.3 x 106 d V/9,0 x 'GA d2.9 x 10 6 d 1/ 2.6 x 10 4 d

*TORSIONAL STIFFNESS (NM I ) 4.4 x 106 d 3 1.1 x 10 1 9.2 x 104 3 1.1 x 107 

•BENDING MOMENT (Nm) R08d 1/9000 6570d 3 1.0 x 10 5 d

• TORSIONAL MOMENT (Nm) 18d 3 4900 3500 1/1.8 x 104

'AXIAL LOAD (N) 200 500 3 4880 3700

a SREAR (N) 400 1 200 1

880 k 3 5400

PAYWAD INTERFACE.

•BENDING MOMENT ( Nm) 1800 1/ 4900 NEGLIGIBLE 90

•TORSIONAL MOMENT ( Nm) NEGLIGIBLE V4900 NEGLIGIBLE 110

•AXIAL LOAD (N) 200 375 NEGLIGIBLE 90

•SHEAR LOAD (N) 100 500 NEGLIGIBLE 110

PBOPUL910N MODULE INTERFACE

•RENDING MOMENT (Nm) 1200 1080 6190 1/1.3 x 104

•AXIAL LOAD (N) 200 17 

go

1.05 x 10 4 3 1.) x 104

*SHEAR LOAD (N) 400 250 1050 3 7950

I. Attachment of orbiter to platform axc mpllRhed by berthing.

2. Berthing loads are: bending momrnt - 1830d (Nm), axial load - 1900d (N).

3. Subscript "d" denotes data obtainod directly from focus mission documents.

Z4. 3Denot PR maximum valu"s,
ORIGINAL PAGE IS

OF POOR QUALITY

The strength requirements are very sensitive to configuration and payloads

size, shape, and mass distribution; and locations of propulsion modules for

LEO stationkeeping or transfer to GEO.

The stiffness requirements are not only dependent on the payload and

configuration size, shape and mass distribution but also the inte7.play between

pointing system requirements, attitude control system ( ACS) thrua , ter levels,

and control system design. For a basepoint design, trades between these

systems can be made along a wide variation of parameters. Since this study is

generic, such trades are not possible. However, the implications are

recognized in the establishment of the requirements.

1.1.1.1 ASASP Flexural and Torsional Stiffness Requiremee.ts

The flexural stiffness value of 6 x 10 8 Nm 2 for the ASASP ( Advanced

Science and Applications Space Platform) configuration 13 extracted directly

from Reference 4 as shown in Figure 1.1-3 and was determined to provide a

minimum first modal frequency of 0.10 Hz. Since no information was available

reqarding the 5equired torsional stiffness, a hand calculation resulted in a

value 1.1 z 10 Nm 2 for the same first modal requirement of 0.10 Hz. The

total mass of the platform was 80 , 553 kg. The individual payloads, and mass

moments of inertia were extracted from Reference 2.

V1

1010

n

Z

c

_W

109

H

W

ZLL

LL_

H

108

D

X

W-a

LL

107

0.1 0.2

FREQUENCY (Hz)

0.3

Figure 1.1-3. Flexural Stiffness

VS. Frequency (Hz) for ASASP

1-4

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ORIGINAL PAG- 19

OF POOR QUALITY

It is pertinent to note that during the orbiter packaging Investigations, a

NASTRAN modal analysis using an EI - 2 x 10 8 Nm2 , and GJ a 0.5 x 10 7 Nm2,

i.e., stiffnesses respectively 1/3 to 1/2 of the above quoted values,resulted

in a first modal frequency of 0.043 Hz. Since flexural stiffness is the main

driver, a frequency of 0.057 Hz would correspond exactly. However, within

the broad context of this study, and the derived generic requirements, the

model sufficiently confirms the validity of the stated requirements.

1.1.1.2 ASASP Strength Requirements

The ASASP strength requirements of 9000 Nm (bending) and 4900 Nm (torsion)

occur during stationkeeping maneuvers. The propulsion module thrust of

890 N, directed as shown in Figure 1.1-4, induces the above specified loads.

It is pertinent to note the relatively large torsional moment results from

the large offset of the center of mass of the payloads.

STATION

KEEPING

THRUST

a

890 N

J%

Figure 1.1-4. ASASP Configuration

—Stationkeeping Thrust
Vt

ORIGNAL r%,c7 n

OF POOR Q'_^.L Lrf-Y

1.1.1.3 GSP Alternative 4 Flexural and Toraioaal Stiffness Requirements

The stiffness data shown in Table 1.1-1 are extracted directly from Table 3-6

of Reference 3. The requirements shown in Table 1.1-1 are the maximum

requirements for members B and E (Figure 1.1-5). The requirements are based

upon a restriction of the relative lateral displacement between the reflector

and feed of 0.20 m due to RCS thruster induced loads. A hand calculation of

the deflection due to inertial loads directed as shown compatible with an

"approximate acceleration of .0003 g" (from Reference 3) indicated a

deflection of 0.14 m (using amplification factor of 2). This proximity

between 0.14 m and 0.20 m is quite adequate for this study. Of primary

importance is the fact that the torsion&; induced portion of the deflection

was more than 95% of the total deflection.

55 m

ATTITUDE

^j CONTROL

THRUST i

60-m — I 90 m REFLECTOR

MEMBER

MEMBER B

FEED FOR

60-m REFLECTOR

Figure 1.1-5. Module 1 (Alternative 4)

Attitude Control Thrust

Consideration was also devoted to the stiffness requirements to provide

frequency separation with the control system. The following is an extraction

from Reference 3 pertinent to this concern.

"A NASTRAN finite element model was generated for the Alternative No. 4

platform based on the individual module orbit transfer strength requirements.

The model was comprised of 65 grid points, 64 structural elements, and 390

structural degrees of freedom. Natural modes and corresponding natural

frequencies were determined for the system. The fundamental natural frequency

of the system based on strength requirements is 0.019 Hz. A similar analysis

of the Alternative No. 4 platform resized to comply with stiffness

1-6
x_0.035

ORIGINAL PAGC IS

OF POOR QUALITY

requirements would yield significantly higher natural frequencies. Again,

caution must be exercised to ensure that the lower frequency vibration modes

do not interact with the RCS and cause instability. As noted previously,

ccntrol techniques can obviate this possibility."

1.1.1.4 GSP Alternative 4 Strength Requirements

The maximum bending moment extracted from Table 3-5, Reference 3 is 101,427 Nm

for member H of Module 2 (Fig. 1.1-6). Torsional momenta are not provided in

Reference 3. The bending moment results from the GEO orbit transfer thrust of

6000 N directed as shown. The data presented in Reference 5 were used to

construct a mass distribution model of Module 2 from which hand calculations

confirmed the specified moment of 101,427 Nm, or more approximately stated,

provided confidence in the mass distribution model used. A hand calculation

of the torsional moment resulted in a value of 18,000 Nm. Both calculations

include a factor of 2 for dynamic amplification.

Figure 1.1-6. Modula 2 (Alternative 4)

Attitude Control Thrust

1-7
A

1.

Cr'

r

a

•

UaiIG114AL PAG" K5

Of- POOR QUALrTY

Table 1.1-2 illustrates the strength and stiffness requirements adopted for

the concept development and structural analyses to be performed in this

study. One set of requirements, i.e., the adopted requirements, is used to

size the structure for the concept development drawings. However, the

implication of order of magnitude variations of these requirements is studied

in section 4.

Table 1.1-1. Adopted Loads (Limit) and Stiffness Requirements

PARAY6r{R

PI.ATFOM BPB A8A3P GBP ALT. I GOP ALT. d

DEPLOYABLE STRUCTURE MODULE

•FLEXURAL STIFFNESS (NU= ) 17.7 x 106 d 270 x 10 J 2.8 x 106 d 7.0 x 10

•TORSIONAL STIFFNES P

':n7 ) 4.4 x 104 d .50 x 10 8.2 X 10, 0,5 n 707

• BENDING MOMENT (MM) B06d 9000 6570d .25 + 103

• TORSIONAL MOMENT (NO) lad 7060 7500 1 0 x LOA

• AXIAL TOAD (N) 200 500 4660 7700

-SHEAR (N) 400 200 650 00

PAYLOAD INTERFACE

• BENDING MOMENT (NM) 1600 00 NEGLIGIBLE BO

•TORSIONAL MOML'NT (NM) NEGLIGIBLE 4800 NEGLIGIBLE 110

*AXIAL LOAD IN) 200 r3-7-51 NEOLIGIBLE BO

e 8H8AR LOAD (R) 100 co NEGLIGIBLE 110

PROPULSION MODULE INTERFACE

S PENDING MOMENT (NM) 1200 1080 8190 1.7 x l0

*AXIAL LOAD (N) 200 1780 1.05 x LOA 1.1 X 10

•SHEAR LOAD (R) 400 250 1050 7950

0 DENOTES ADOPTED STRENGTH AND STIFFNESS

The adopted stiffness requirements are based upon the following rationale:

o The highest EI value listed for the ASASP of 6x10 8 Nm2 is quite

arbitrary, i.e., to achieve a first modal frequency of .10 Hz for a

platform mass of 80,553 kg. It is unlikely a platform in excess of

40,000 kg will be required. Further, a first modal frequency of .03

Hz is still 100 times the LEO orbit disturbance frequencies.

o Of the val es (from GSP 4 Module 1) of EI - 2.6x10 8 and GJ -

l.lxl07 Nmq ,only the GJ is the major requirement to limit the

required deflection. That requirement is due to placement of the

feeds for the 60 m reflector as shown in Fig. 1.1 -5. Antennas with

their own feed columns would preclude this requirement.

o In view of the foregoing, and the technology advancement goals of

developing designs up to the maximum practical/realistic

requirements, the adopted stiffness values of 2x10 and •5x107

Nm , respectively, for El and GJ were selected.

r

1-8
141 i

ORIGINAL PAW- Q

OF POOR QUALITY

The adopted strength requirements were based on the following:

o The values of 1x105 and 1.8x104 Nm, respectively, for the

bending and torsional moment obtained from the GSP alternative 4,

Module 2, are regarded as unnecessarily high since the module is

20,600 kg (payloads to GEO by the year 2000 are not expected to

exceed 6000 kg). Further, the loads can be reduced by reduction of

the orbit transfer T/W (thrust to weight ratio), and/or minimization

of dynamic amplifications. The latter two items are technology

development needs, for antennae as well as platforms. The potential

reduction of T/W from 0.035 to a possible 0.0137 is discussed as

follows:

Low-thrust liquid propellant systems are indicated for orbit tranfer

applications for current and future missions to GEO as the need for

maneuvering, start-stop operations and especially low thrust levels

predominate as desirable characteristics.

Low thrust and the resultant long burn times can mean larger gravity

losses and increased propellant weight (Figure 1.1-7) for Single

burns. However, multiple perigee burns minimize gravity losses by

reducing the burning arc and theta (9), the angle between the

velocity vector and the local horizontal in the gravity loss, term,

got sinO (simplified).

r . PAYLOAD MASS • 6818 kg

I, • SPECIFIC IMPULSE • 470 sec

• STAGE MASS FRACTION • 0.90

• CONSTANT THRUST

20

PROPELLANT1 PERIGEE BURN

MASS

(1000 kg)

-2 PERIGEE

BURNS

158 PERIGEE

BURNS

0.002 0.004 0.01 0.02 0.04 0.1 0.

MAXIMUM THRUST-TO-WEIGHT RATIO, T/W (END BURN)

Figure 1.1-7. GEO Transfer T/W and Number of

Perigee Burn Implications on Propellant Mass

1-9
ORICANAL ;

OF POOR QUALM i

The use of multiple burns to minimize gravity lose (24 for eight

burns versus 144 for single burs) is attractive in exchange for a

somewhat longer transfer coast time (22 hours for 8 burns versus 6

hours for single burn). '.n many cases, this will be an accepta'ale

compromise. The reductiol in propellant requirements with multiple

perigee burns at low T/W la dramatic, ea shown for the payload

indicated. Generally, propellant requirements increase as T/W

decreases below the value of one, due to the aforementioned gravity

loss effects. The propellant weight is decreased, however, as the

number of perigee burns is increased above a nominal of ous perigee

burn. Eight perigee burns provide a substantial reduction in

propellant requirement at low T/W; the maximum reduction occurring

in the vicinity of IN (final) from 0.01 to 0.02. These savings are

sharply reduced, however, at T/W (final) values below 0.008.

o In view of the foregoing, and the technology advancement goals of

I; developing designs up to the maximum practical and realistic

requirements, the strength values of .25x10 5 and 1x104 Nm,

respectively, for bending and torsional moments were used.

In summary, the rationale for establishment of the adopted strength and

stiffness requirements is presented above. In consideration of the overall

range of requirements throughout the focus missions shown on Figure 1.1-1, the

justifiable departure from the maximum values is minimal. Hence, it may be

stated that the adopted requirements are at the high end of the total

requirements spectrum.

1.1.2 Power and Data Utilities

Table 1.1-3 summarizes the adopted power and data utilities alongside of the

corresponding data extracted from the focus mission studies. This requirement

is uaed throughout the concept development.

Table 1.1-3. Generic Power and Data Utilities Requirements

SYSTEM

FUNCTION ASASP GPS SPS ADOPTED COMMENTS

POWER (KW) 50 10 490 50 SPS DEFINED AS

SPECIAL CASE

DATA 20 MOPS (D) 50 MOPS (D) 20 MOPS (D) (SCIENTIFIC)

50 KBPS (0) NOT DETERMINED 216 KBPS (C) 50 KBPS (D) (HOUSEKEEPING)

25 KBPS (C) 25 KBPS (C)

4.2 MHz (A) 4.2 MHz (A) (TV)

INTERFACES

POWER 6 NO. 0 6 NO. 3 396 NO. 10 6 NO. 0

28 NO. 2 4 NO. 13 16 NO. 2

4 NO. 14 20 NO. 14 4 No. 14

DATA 35 F.O. 34 NO. 19 TSP 4 NO. 22 TSP 90 NO. 22 TSP TO BE ACCOMMO58 NO. 26 TSP 2 COAX DATED BY

144 F.O. 8 F.O. DEPLOYABLE

100 F.O. STRUCTURAL

(OPTION) ELEMENTS

1-10
Q

ORIGMAL PFr ' a

OF POOR QUALITY

The power lines are bared upon a 50 kw total spacecraft power requirement and

distribution system. The data utilities are based upon the scientific,

housekeeping, and TV data needs se shown in Table 1.1-3 and discussed herein.

The power and signal distribution requirements are based upon analysis of

possible payload combinations and altitudes. Table 1.1-4 presents a summary

of the payloads considered. Included in the table are various physical

characteristics and interfaces required to support the selected payloads. In

a general sense, the distribution system is required to accommodate two

classes of signals. The first is power at a relatively high level approaching

20 W. The major power capability in distributed at 124-164 VDC, with lower

loads at 30 VDC and 110 VAC, 400 Hz. The seccnd class of signals are at

fairl3 low levels and consist of command and data signals, digitally coded

data transfer and low level cloaed-circuit TV (CCTV). The digital and video

signal bandwidths are estimated to be 1 Mbpa and 6.0 MHz, respectively.

Table 1.1-4. Reference Payload Group

DEPL. DATA MGMT

MASS SIZE ALTITUDE ORBIT COMMAND DATA

PAYLOAD GROUP (KG) (M) (KM) (DEG) POWER (KBPS) (KBPS) NOTES

I. ATMOSPHERIC GRAVITY 3,000 100 (D) > 250 56-90 33 <25 10 SEE m

WAVE ANTENNA (kWh)

2. PARTICLE BEAM 3,000 100 x 100 400 56-90 3.3 <25 200 + SEE (D

INJECTION (SQUARE) (kWh) 4.2 MHz

(TV)

3. ASTROMETRIC 4,500 2(D) x 18 400 28 1,000 <25 19

TELESCOPE (W)

1,00:

4. LARGE AMBIENT DIS- 16,000 15(D) x 35 400-700 28-56 1,000 <25 7,000 SEE CD

PLAY IR TELESCOPE I I I I

I (W)

NOTES m EVA FOR MAINTENANCE AND DEPLOYMENT

--- m VIBRATION AND GAS PARTICLE SENSITIVE

® VIBRATION AND GAS PARTICLE SENSITIVE ♦ EVA

1.1.2.1 power Utilities

The power utilities requirements are based upon the distribution system shown

in Figure 1.1-8 and line lengths compatible with the generic platform (Section

1.2).

The wire size selection was conservatively based upon the worst case

conditions at an operating temperature of 200'C, and line losses of

approximately 5%. (Refinements in the analysis made at a later date to

account for reduction of the operating temperatures to 20'C are discussed

later.)

The 16 No. 2 requirement shown in Table 1.1-3 was derived from provision of

124 VDC at 290 A. For this condition, the primary path wire bundle was chosen

to be eight No. 2 gauge (stranded) for power input plls eight No. 2 gauge

(stranded) for returns.

1-11

l
6A

ORIC NAL t'POL i

OF POOR QUALrr,

^-

Yr,F YwIiY L -- -.J rr.fr, ,1-,..^Y.^.

Nn.

ftmh . LM/ IMMIN,tir hnl

ININ Yr 11 tt w Vr 11w truer

N Vt 1AM

110'VA RA he

10111040

IN-IN Vw 38 kWMIM N m a^ww^

rYY WWW 

nlLOMW

•M4 IIM

MM/ Y.rl.

1 YNM

L_

II M. 7

twArl ^ N^. 11

r.rN ' 7MY t1wN m.l,rl

Y W.1 •04,.04

-

^ - FNVInr

0 Va

Uw",wt

I MllNr ,

1 ^rIK^ 1 17741 M WE

oubw

blw 1MN

E,4r /n"

Figure 1.1-8. Platform Power Distribution Subsystem 

Six No. 0 for the 30 VDC lines, and four No-14 for the 110 VAC, 400Hz lines

were similarly determined. Later in the study, as part of an (valuation to

determine the significance to the power utilities requirements of a 250 kW

total spacecraft power requirement, refinements in the above analysis Were

made. The essence of this analysis was reduced electrical wire sizes due to

replacement of the 30 VDC by 124-164 VDC, or replacement of both '10 and

124-164 VDC by 110 VAC, with converters and inverters at each port. For this

reason and since the adopted requirement could be integrated into the designs,

the adopted requirement was maintained. The investigation for a total

spacecraft power load of 250 kW indicated essentially the same total circular

mile as that of the adopted values provided the selected power lines operate

at 460 VAC.

The basis for the foregoing conclusions is discussed in the following

paragraphs.

The original evaluation assumed that the 30 VDC and the 110 VAC voltages were

derived at a common location, located some distance from the solar array

module. The effect of this assumption was to require that the deployable

truss be capable of accommodating both the entire 50 kW at 124- 164 VDC as well

as accommodating 5.6 W at 30 VDC and 0 . 4 kW at 110 VAC. In this original

evaluation it was also assumed that the primary 124 VDC wire bundle

temperature could rise to 200'C. Subsequent analysis indicated that a 200C

design limit could be maintained and, therefore, it was decided to redo the

analysis of the basic concepts as well as all the other distribution concept

options at the 20'C design point.

1-12 
V

ti

ORIGINAL ri *I'

OF POOR QUALITY

The result of the revised basic analysis (summarized in Table 1.1-5) is to

reduce the number of No. 2 AVG wires to eight. However, in the re-evaluation

of the original analysis, the number of No. 0 AVG wires is increased to

siylt. Both numbers include thoir respective return wiroa or grounds. The

total crone-sectional area of the wire bundle was reduced from 1,707,000

circular mile (in the original analysis) to 1,384,000 circular mile - a

reduction of 18.44.

An alternate to the basic wiring layout wan also evaluated. In this case, the

conversion to 30 VDC was accomplished within the solar power module with the

resulting reduction in the power handling requirement of the 124-164 VD! wire

bundle to 36 W. At the same time, it was noted that the 30 V"C circuit may

be required to accommodate 9.6 kW rather than 5.6 kW, so this increase in

pow*.r level was also considered. The results of these changes are shown in

the second column of Table 1.1-5. In this approach, the number of No. 2 AVG

wires is reduced to six, reflecting the lowered power level, but t?,

identified increase. in 30 VDC power increases the number of No. 0 AWU wires to

12. The end result is a net increase in wire arms of 21.5%. If the 5.6 kW

requirements at 30 VDC are retained, a wire cross -sectional area of 9.5% is

realized..

Four new options were considered during this Study. Two of these options,

Option A at 45.6 kW and Option C at 250 kW, aeeumed that all of the energy was

provided at 124-164 VDC. The other two options considered the power is to be

delivered at 460 VAC. Option B was rated at 50 kW while Option D provided 250

W. The wire sizes for the ac power analyses were taken from a standard ac

power handbook and adjusted for the voltage and current levels identified.

Table 1.1-5. Effects of Differing Power Load Assumptions

Upon Wire Count

BASIC IIASIC(ALT)6' nm no CID nO

POWrIt 50 kW w 06 kit '• 45.6 kW s 290 kit ,.

UDAD 4-(_64 VDC _I_.44_[64 VPC_ }24 -164 VOC_ —_ 124-164 VUC

R ►Aw ry• 5.6 kRnn 6.6/9.fi k n — _`

WIRE 30VDC._.

SIZE 0.4 kW 6.4 kW w60 kit 6___' ()7 kW o 250 kV n a 0.4 kW n

(AW6) 110 VAC Il0 VnC 110 VAC 460 VAC 110 VAC 460 VAC

14 4 4 4 _ 4

4 _ _ _ AID —. lA

2 R (WAS 16) 6 A — 511

0 A (WAS 6) 8/12 0

M Ll.ctranlr. Inr,c.d In P••••.r wMul..

Apulr.n 124 to In Yn0 ca n.. rter, at ..rh Part, nr r.-panalhllltr Pi a,rr n.

11"Or., In.Irt.r, A r,dlnenr, In - 9n%) In pttir. ..nh:l., and rrrtlrlera In 00;) ;q rnrh P••rt

nr rlthln u,ern' p,d Carat. Is.,,., a! Im at Inverter .111 rnr ro n•r rllrrtl.r .Irr al ,:Bar

.rr.Y. (Inrrm.. to 55 k4)

0 32 1In- va l,nW. Aran P. rwlt Vd' IT - 200C)

1-13
The spacecraft configuration applied to all of the new options considered that

the derivation of any voltage other than the primary distribution voltage,

with the exception of the relatively low-powered ac in Options A and B, is to

be acccmplished at the user port or within the experiment system. Thus, in

Options A anti C the deployable trues moot accommodate only two voltage levels,

124-164 VDC and 110 VAC. In Options B and D, only high voltage (460 V) ac

must be accommodated. The advantage of eliminating tae low-voltage

distribution system is apparent, but it does require the addition of more

equipment at each port.

The question of utilizing an ac distribution concept at a higher voltage was

addressed in Option B. Again, it is possible to further red , ice `.he total

number of wires within a deployable truss at the expense of additional

equipment, specifically rectifiers and transformers, utilized to provide the

lower ac voltages and the various do voltages. A further coat factor that

should be appreciated, when using the ac concept, is the need to compensate

for the additional losses introduced by the added equipment. With an

estimated efficiency of 90% or lose, the option will require a 10-15% increase

in solar array to generate the necessary power.

Options C and D were evaluated to determine the impact of increasing the 

spacecraft experiment support power to 250 kW. Specifically, the no

distribution power level was increased to 250 kW at 124-:64 VDC, while the ac 

system power level is specified at 250 kW at 460 VAC.

The effects of the higher power love's will, as expected, result in a larger

number of wires. Total cross-sectional area of the wires varies from

approximately 3,870,000 circular mile (in the case of Option C) to

approximately 1,619,000 circular mile (Option D). In the case of the ac

distribution system, it will again be necessary to appreciate the additional

losses in the system caused by the part-located conversion equipment.

A final point of discussion is to examine the reasons why, if higher do

voltage will reduce lose effects, a higher do prime,y distribution voltage

(greater than 124-164 VDC) was not selected for the basic concept and for

Options A and C. The major rationale for not selecting a higher voltage

distribution system is the state of switching technology, particularly a*, the

relatively high power levels specified. Switch devices capable of switching

high do voltages at high power levels are not yet available to the confidence

levels needed to assure reliable operation. These switching devices include

these simply used to control the power distribution as well as those used in

do-dc converters supplying the various lower voltages required. Several

attempts at initiating technology programs to develop these equipments have

been made with various degrees of success, but none have been completed.

Accordingly, no hardware is available at the present time nor in the immediate

future.

The second reason for not selecting a higher voltage is the impact upon the

solar array. Solar arrays are comprised of many small solar cells

interconnected in a series-parallel matrix to provide the needed power/voltage

combination. Increased voltage levels increase the complexity of the solar

array with the attendant reduction in reliability, life, and in poorer

operating characteristics resulting from increased internal losses.

1-14
1ki

ti

1.1.2.2 Data Utilities

The data management, control path, and the CCTV patha a.e essentially low

power and are provided using three different forme of signal paths. Discrete

signals or commands are rowed utilizing 22 gauge, twisted, shielded pairs

(TSP). Ninety pairs are included in the design of deployable structure:

elements; sixteen of which are preassigned as emergency control originating in

the 'subsystem control module. The remaining 74 pairs are unassigned and may

be used to provide interconnections between berthing stations.

The placement of these, or any signal wires, immediately -.djacent to power

cables is to be avoided. If the designs do not permit the separation of these

cable groups, it is necessary to provide metallic separation (shields) to

avoid elertromagnatic iz.terference caused by power switching.

The data management concept selected for this study presumes the use of an

integrated, on-board data processing approach that uses a data bus design to

minimize the overall number of discrete paths between aatelli.e or platform

communications interfaces. The proposed data bus link consists of four pairs

of fiber optic cables. Each pair consists of an independent command and data

channel. Four pairs are provided to accommodate reliability concerns (e.g.,

redundant paths in the event of channel failure) as well as pro-iding for

poeeible req uirements calling for independent links to selected payloads. It

is recommended that provisions to add up to 100 additional fiber optic cables

to permit future system expansion be included in the element design.

The final requirement identified in prior studios is the need to provide at

least two channels for routing of CCTV or high bit rate data to the satellite

downlink communications system. The suggested coaxial is type RG-303/U.

In summary, it is pertinent to note that the data requirements delineated

above are considered to be a generous complement of number and sizes. It is

generally, however, consistent with the ASASP and GPS requirements. Further,

spacecraft data needs during any program always press to the limit the ability

of the vtructure to accommodate data needs. Hence, again, the generous

complement of data utilities in the adopted requirements.

1.1.3 Fluid Utilities

The adopted fluid utilities requirement established for this study is two 2.0

cm coolant lines. Propellant lines are not a requirement for the reasons

discussed subsequently.

1.1.3.1 Coolant Lines

Provision of fluid coolant lines is imposed as a requirement since location of

radiators adjacent to a heat source may not always be practical. The

requirement of two 2.0 cm lines was determined from the payload requirements

of Table 1.1-4, which were extracted from Reference 2• For these payloads,

the maximum power level was 25 kW.

1-15
3

x

E

E

1i

ORIGINAL PAGE IS

OF POOR QUALITY

The use of a central radiator system would require that fluid lines be run

along the structure be'ween the heat source and the radiator and return. The

pumping power required to circulate the coolant varies directly with length

and power dissipation level and inversely with line diameter. For the

reference mission payloads, a practical line diameter is 2 cm. Figure 1.1-9

presents the pump power to circulate Freon coolant through a 2 cm line ovir 40

motors and return for a range of power dissipation levels. As shown, a

pumping power of .lose than 0.2 kW is required to circulate coolant to reje^t

25 kW of payload power.

Payload Power Level kW

Figure 1.1-9. Pump Characteristics

1-16

„ — -
ORIGINAL PAGE 19

OF POOR QUALITY

Insulation will ba required for fluid lines used on the deployable platform to

prevent freezing of the coolant during quiet periods. A quantitative estimate

of the thermal control system ( TCS) requirements for fluid lines is summarized

in Table, 1.1-6 for a range of typical TO coating techniques. These

calculations assume the heat transfer area to be that of a 2 - cm-diameter

straight tube. The calculations show that a radiation barrier insulation

(c w0.04) is required to keep the drop in fluid temperature, due to line heat

losses, at reasonable levels. A bare metal, such as aluminum, is

unacceptable. These materials have high solar absorptance compared to its

emittance. This would result in high tube surface temperatures that could

cause localized boiling in the fluid line. A multi-layer insulation (MLI)

blanket of about 5 to 6 layers would provide the proper emittance. The

blanket could consist of concentric wraps of embossed metallized foil. A

non-metallic outer layer may be desirable.

^f

Table 1.1-6. TCS Requirements for 2-cm Fluid Lines

LOSS OF TIME HEAT

TEMPERATURE (°C) TO LOSS AT

EMIT- Fog 40-m LINE FREEZE FREEZ.

TEMPERATURE ('C)

HEAT LOAD SUN SURFACE TANCE FLUID TEMP.

COATING E I W 5 kW 25 kW HR M/m AVERAGE PEAK

PAINT 0.9 52 10.4 2.0 0.5 3.3 24 122

COATING, 0.5 28 5.6 1.2 1.3 2.0 24 122

BARE METAL 0.1 6 1.2 0.3 6.4 0.3 24 122

POLISHED

METAL 0.04 2.4 0.5 0.1 13 0,15 219 381

(a - 0.3)

MLI (a - 0.4) 0.04 2.4 0.5 0.1 13 0.15 -37 42

An alternate to the fluid coolant system is a heat pipe system. The heat pipe

is a sealed heat transport device and does not require a pump to maintain its

operation. A typical installation would use a 2 cm line to transport vapor

and a 0.63 cm line to transport li q uid. The technology of high capacity heat

pipes is developing rapidly. Pipes with the capacity to transport heat loads

of a few kilowatts over distances of a few meters are currently in development.

1.1.3.2 p ropellant Lines

Integration of propellant lines into the deployable structure was not included

as a requirement, although the use of distributed thrusters for either GEO

orbit transfer or active modal control was considered.

The use of distributed thrusters has been suggested to reduce the bending

loads imposed on the structure during orbit transfer. This advantage is not

v regarded as sufficient to offset the numerous disadvantages discussed below:

o Since the design must provide for failure of a thruster, each of the

thrusters lines of force must pass through the platform center of

G mass for the engine out condition. For the sizes of GEO platforms

1-.17

w
s

r

Y

h

r{'

}

this would be applicable to possibly 2 to 4 thrusters. (Concepts

utilizing more than 10 distributed thrusters as shown for solar

power satellite (BPS) structures are not applicable.) Provisions

for this condition would require excessive propellant.

o The additional coat of providing and installing the additional

thrusters, and integrating the cross-feed propellant lines would be

excessive. In particular, the folding and thermal control of

propellant lines is a significant technology problem.

o The reduction of the bending and possibly torsional moments, while

reducing the individual member loads, may not reprosent any

significant weight reduction for stiffness-critical designs. The

weight savings, if any, may be limited to the ,joints. The reduced

loads on the ,joints and individual members could permit increased

packaging efficiency. However, the loads can be reduced by stowage,

which appears to be a simpler task.

It is pertinent to note the Large Spacecraft .°yetems/Propulsion Interaction

Bork Shop, held on October 22 and 23 in 1981, recommended the use of a single

orbital transfer vehicle (OTV) with clustered thrusters for orbit transfer of

GEO platforms. For the range of platform mass up to 6000 kg, a single OTV has

the capability.

The use of distributed control thrusters is not considered app_, opriate for

platform control in view of the following:

o The major spectrum of LEO/GEO mission requirements is achievable

without special distributed sctuators. Sufficient stability is

attainable without, distributed thrusters.

o For the special cases where unique payload precision pointing and

high levels of stability are required special mounts such as initial

pointing system (IPS) or annular suspension pointing system, gimbal

system (AGS) will be provided.

If for some reason, distributed control is required, the preferred location

for rotary or linear actuators is at the attachments between building blocks.

Since the payloads are mounted only at the main hcising or adapters, shaping

of the basic truss is not required.

1.1.4 Control System

Control system requirements can be satisfied without the mounting of control

system equipment directly onto the deployable truss. A control system module

is provided to contain control system equipment other than that located at the

payloads.

This review examined the three focus mission requirements with the basic

platform functions and control philosophy as follows:

1-18

ell
ORIGINAL PAGE i3

OF POOR QUALITY

The platform provides a general -purpose base to support

special-purpose (and multidiel; iplinary) payloads whose pointing and

control requirements can be quite diverse.

o The platform provides gross stabilization and pointing control.

Nominal attitude will be a local level and/or inertial orientation

selected to minimize the platform disturbance torques and the

resulting control system requirements.

o Payload precision pointing and high levels of stability will be

provided by special pointing mounts ( such as IPS anc AGS).

o Specialized passive or active control for stmctsral dynamic or

fig-.,re control augmentation can be located at the housings or

adapters ( rather than on the deployable trues).

o Phe control system is designed to meet the most common recurring

control problems - not rare or ill-defined special situations.

As evident from Table 1.1-7, it is feasible to locate all attitude control

system equipment either in the control system module, the payload package, or

at the building - block-to- building- block interfaces. Much of the equipment

listed as mountable on the " control system module or payload package" can

probably be mounted in the control system module for most NASA missions in

which structural deformations are well within pointing requirements. 

Table 1.1-7. Control System Equipment Requirements

S/C USED ON POSSIBLE LOCATIONS

CONTROL

SYSTEM BUILDINGGENERIC DEPLOY. MODULE OR BLOCK

COMPONENT ASASP GSP SPS ONLY STRUCTURE P/L PKG, INTERFACE

• SOLAR TRACKERS (ST) x x x

• SOLAR ASPECT SENSORS (SAS)

• FINE SUN SENSOAS (FSS)

K

x

R x x

x x

• COARSE SUN SENSORS (CSS) x x

• HORIZON SENSOR (MS) x x x

• INERTIAL REF. UNIT (GYROS)-08U) x x x x

• MOMENTUM L REACTION WHEELS (MW) x x

• CONTROL MOMENT GYROS (CMG) x x x

• COMPUTER (COMP.) K x x x

• INSTRUMENT POINTING SYSTEM (MPS) x x

• MISC. ROTARY JOINTS (RJ) x x x K

• INTERFACE ELECTRGNICS UNIT 0 EU) x x

• MAGNETIC TORQUER$ (MT) K x

• MAGNETROMETER (M) x x

• INERTIAL MEASUREMENT UN:T ONO x x

• REAC T• ,N CONTROL SYSTEM (RCS) x x 3 ! x

• STRULIURAL ALIGNMENT MEASUREMENT

DEVICES x A x

CONCLUSION: ALL ACS COMPONENTS ARE LOCATED ON ADAPTERS ON IN CONTROL SYSTEM MODULE.

1-19

ki

0
ORIGINAL PAGE 15

OF POOR QUALfiY.

1.1.5 Structural Temperatures

Peak structural temperatures range from -100' to 80'C for LEO and -200' to

80'C for GEO. Figure 1.1-10 presents the oak temperatures calculated for the

materials shown for an end of life (a/e

i

O_

m

LEO

Gr/Ep

AL

BE

BE

AL

GR/EP

-100

GEO

^_ GR/EPAI

Be

i

u

A

-^ Be

Al

-100 Gr/Ep

0.075 0.150 0.225 0.30 0.075 0.150 0.225 0.30

Thlckneer. (cm) Thickness (cm)

Figure 1.1-10 Maximum and Minimum Temperatures

The materials considered were aluminum, beryllium, magnesium, and several

graphite and metal matrix composites. The peak maximum and minimum orbital

temperatures depend upor the density times specific heat product for a

particular material. For a given material thickness, the smallest density

specific heat product will yield the largest maximum-to-minimum temperature

range. Since a large number of materials is being considered, the materials

were ranked according to this product. Beryllium had the largest value of

this product, and graphite epoxy the smallest. Therefore, the properties of

these materials were- u..., eria used i t the analysis. Aluminum is used currently in

spacecraft construction; therefore, data for aluminum are presented for

comparison.

The member shape; considered were round, square, rectangular tubes and

I-sections. The hottest structural temperatures occur when the member shape

exposes its largest projected area toward the sun. A measure of this peak

heating condition is the ratio of the solar projected area to the radiation

area; therefore, the member shapes were ranked according to this ratio. For

the member shapes considered, this ratio differs only slightly; therefore,

only one or two shapes need to be analyzed. The square tube and the I-section

1-20 
Q

were selected for analysis. The data presented are for the square tube. The

coldest structural temperatures occur during eclipse and are independent of

member shape.

Individual trues members may be oriented at various angles with respect to the

orbit plane and the sun. The orientation which exposes the largest solar

projected area was selected for analysis.

Orbital temperatures were computed with a thermal math model (TMM). This

model, for the square tube, consists of four nodes which are connected

together by conduction and internal radiation. The nodes are connected to the

external environment by radiation to apace and by heat flow rates (QDOT's) for

the incident solar, albedo, and earth emission fluxes.

1.1.6 Servicing

Servicing is assumed to be at one-year intervals, and consists mainly of

replacing consumables or equipment with limited life. Changeout of payloads

may be required from time to time because of "state of the art" or completion

of mission.

The vehicles used for servicing are the orbiter for LEO and a teleoperator !or

GEO. It is expected that the teleoparstor will use the same docking interface

as the orbiter and will have a similar set of RMS.

The design of the platform, therefore, has to include docking provisions at

r' strategic locations, near to payloads and service centers such as the control

module. All items likely to be replaced must be designed for

disconnect/:, emoval/replacement using the RMS and/or EVA. Factors such as crew

visibility, TV coverage, RMS reach, RCS plume effects and orbiter/platform

interference need to be considered when designing for servicing. 6

1.1.7 Orbiter Integratio

The orbiter is obviously an item of major concern in the design of a

deployable platform. The orbiter is called upon to perform several functions:

o Transport the packaged platform to LEO

o Serve as a base for adding payloads, modules, etc.

o Checkout and troubleshooting the platform and systems

C Continued servicing and maintenance for a period of 10-20 years.

In performing these functions, the following factors are of importance:

o Docking/berthing clearances

o EVA capabilities and safety

o RCS plume effects

o RMS reach and capability

o TV & crew visual coverage

o Orbiter power available for platform

o Payload bay volume, c.g. and weight

o Mounting in the payload bay, cradles/pallets, trunnion and keel

fitting location and loads

o Orientation during platform deployment, solar thermal, radiators

o Orientation, free drift, control authority

o Cost

1-21

f 

V 

f.:. _. ' ' ^ i. ... .z. _.r.. . ....'...Kim . .....
fi ORIGINAL PAGE'S'

OF POOR Q UAUTY

1.1.8 Environment

Materials for the deployable platform systems must have a minimum life of ten

years in either LEO or GEO environments. Since transfer time to GEO is

expected to be lees than 24 hours with the use of chemical propulsion, the

more severe LEO-to - GEO environment effects will be negligible.

Space environment effects are due to a combination of environments, some of

which act at the surface and others which act throughout the volume ( mass).

Solar radiation, vacuum, and micrometeoroids are examples of environments

which act primarily on exposed surfaces, while Van Allen Belt particles, solar

flare particles, and the electron-produced Bremeetrahlung are examples of

environments which act throughout the volumes of objects in space. Figure

1.1-11 illustrates the nuclear radiation components at GEO as a function of

aluminum shielding thickness ( the curves are similar for other materials) and

Figure 1.1-12 illustrates the altitude dependence of the natural Van Allan

Belts.

DOSE RATE

RATE (Si)

YEAR

1 ^Natural Van Allen N Belt Electrons

OB — \\

U 

5

SWZr Flare ^

Particles \^

Trapped

\

1` Fission

\ Electrons

Uq ^ 1

\ ; 1

\' 1

1

1

\1

103

I

^ 1

Bremsstrahlung ----r-^-

I 1I i02 1_ I i ...I 1.

in- 2 in-1 1

Aluminum Shield Thickness (gm /cm2)

Figure 1.1-11. Radiation Dose Rates at GEO

— Functions of Shield Thickness

1-22
URIGINAL PAGL !J

OF POOR QUALITY

0* INCLINATION

I00

s.

i

ISS 102

101

101

I

I

I

I

I

0.5 GN/CM2

GM/CNZ

I I

2 c" /Cm

I

I

I

I

I

^

I

I

O NN

(3. 17

I

12

b) (1.IS

NN

N.)

26 NN

(6.6 44)

I

100

10"1

0 4000 8000 12,000 16,000 20,000 21,000 78,000

RO . IABTN MDI US ALTITUDE (NMI)

GM/CN • ALUMINUM

Figure 1.1-12. Natural Van Allen Belt

Dose Rates, 0 0 Inclination Orbit

1-23

r
10-S

10"6

10-7

10-8

e

10-5

10.10

S 10.11

W 10-12

10.11

10-14

10.15

10-16

i

Q

ORIGINAL PAGE 18

OF POOR QUALITY

The solar spectrum and vacuum are well known and are not included herein.

Several materials can be severely impacted by both of these environmental

properties, but thin protective coatings from other materials (such

as thermal control coatings) can and do readily negate any excessive adverse

effects.

Particulate or meteoroid flux can and does get significant as satellites

increase in size and required service life. Figure 1.1-13 shown a

time-averaged meteoroid flux at 1 AU from the sun. This meteoroid flux was

obtained directly from Reference 6 and is sufficiently accurate for both LEO

and CEO platform applications.

PARTICLE MSS III-)

Figure 1.1-13. Natural Time Averaged Meteoroid Flux

at 1 AU from the Sun

1.1.9 Payloads,_ Propulsion Mndules, ACS Modules

Generally speaking, the linear platform is deployed without payloads,

propulsion modules, ACS modules, or other large items which are not part of

the basic platform. It is expected that they are added by the RMS to suitable

interfaces subsequent to deployment of the platform structure. There are,

however, possibilities of building some modules into the deployment system;

depending on the size/shape of the module and the size/shape of the platform.

If the circumstances are favorable, the module may be treated as another

housing similar to a building block, and be incorporated onto the end of a

1-24
trues. Some items may be mounted directly onto an existing control module or

trues housing. If the module is to be added to the platform subsequent to

platform deployment, an interface is provided complete with alignment features

and all the structural/mechanical/electrical/fluid interconnects required for

RM9 berthing.

1.1.10 System Pointing Accuracy

The most stringent pointing requirement is delineated in Table 6 of Reference

4 for the OSP application. A value of 0.05 to 0.10' is listed.

Discussion with electronics specialists at Rockwell indicate pointing

accuracies of 0.05 to 0.10' will be representative of most applications for

the 1990 to 2000 time period, although a small number of applications will

require accuracies of 0.03 to 0.02'.

1.1.11 Requirements Perspective

The design approach and requirements presented in Sectiona 1.1 through 1.10

ere the basis for the generic platform development (Section 1.2); deployable

tivse, utilities folding/deployment, housing, and adapter concepts developed

(Sections 1.3 and 1.4) and concept selection (Section 4). A perspective on

these requirements are summarized below:

o The adopted strength and stiffneris requirements are at the -rpper 

tlimits of the spectrum of requirements for the 1990 to 2000 time

period.

o The generic platform represents the largest size of platform

foreseeable for the 1990 to 2000 time period.

SSti

k o The adopted complement of power utilities is representative of the

P^ maximum number and size of lines consistent with a 50 kW spacecraft

R (30 VDC and 124-164 VDC lines).

o The adopted complement of power utilities is also representative of

a 250 kW spacecraft (460 VAC).

o The adopted complement of data utilities is quite extensive, but can

vary considerably with payloads.

o During the platform design phase, the magnitude of data utilities is

generally driven toward the maximum the structure can accommodate.

o The two 2- cm fluid lines requirement is representative of the

maximum coolant fluid line diameter.

o A pointing accuracy of 0.05 to 0.10 degree is representative of a

majority of systems for 1990 to 2000 time period.

This perspective is applicable to the focus missions (Figure 1.1-1) and to the

spacecraft configurations shown in Figure 1.1-14.

1-25

Af 
r x.

i^

URIGINAL PAGE 19

Of POOR QUALfiY

Figure 1.1-14. Rockwell Comnunications Configurations

and SASP

1-26

s
4

1.2 GENERIC DEPLOYABLE SPACECRAFT CONFIGURATIONS

The generic linear platform (Figure 1.2-1) is based on the requirements

established in this study. The purpose of developing sucL a platform is not

to suggest that this is the beet or only configuration available, but to

display and understand the many problems encountered in designing a deployable

system. Without such a configuration to study, there is a tendency to

concentrate on the trues structure and overlook such important points as

orbiter integration/packaging, deployment sequence and attachment of one

section to another.

The configuration of the deployed platform is similar to the ASASP. The

overall dimensions of the structure are 146 meters by 73 meters. The payloads

shown (i.e., the reference payload system) are the atmospheric gravity-wave

antenna, particle beam infection experiment, astrometric telescope, and IR

telescope. Each of the four payloads in mounted at a "hard point" and not on

the deployable truss itself. The design goal is to avoid mounting equipment

on the deployable truss unless it is absolutely necessary. This policy has

been followed throughout all the drawings. The electrical power for the

reference payloads, -nd spacecraft systems, with allow4nces for line losses is

50 kW. The solar array is suitably sized and has the necessary rotary joints

i for two degrees of fre„dom. A radiator is shown mounted adjacent to the solar

array module.

Although the reference payloads are LEO payloads, an orbit transfer vehicle is

mounted on the aft end of the platform. Four reaction control system (RCS)

modules are shown, of which two are mounted on short booms deployed from the

control module, and two are mounted on structural hard points.

The control module (CM) serves three functions, i.e., it is a cradle or pallet

for mounting equipment in the orbiter, a building platform for deploying the

platform, and it is a part of the spacecraft platform and functions as the

control center and houses equipment such as batteries, communications, data

otorage, and power conversion and control.

Provisions for docking/berthing of the orbiter are shown at the control

module, at too solar array module, and at hard points on the structure,

suitable for servicing payloads or propulsion units.

The structure consists of building blocks, arranged in such a fashion that

they deploy sequentially or in unison to form the configuration shown. There

is no "erection" or EVA involved, and thore is no requirement for a building

fixture or jig.

Is the development of the several concepts for the trusses and building

blocks, the generic platform was modified slightly to incorporate needs as

they arose: for instance, sometimes there are two control modules and

sometimes two trusses are ,joined into one building block.

Subsequent to the development of the generic linear platform, a generic area

platform was studied in whi.ch the structural assembly behaves as a plate

rather than as a number of beams connected together. Ths advantages expected

to be gained of such a structure are:

1-27 
8

m

a:

w°

ri

^ a

t J 4

m

o ai S p c

Na N

p p

7

4

N

ti

N

V

W

v

4

.HU

I' O pY

W ^ O

1 F

n S

`S

r-^

a

i

a

!I

'400"" /

^O ` I '

s 1

Q W

^x

i

\W

O < p 

6y

=yr.:

Y W

6 ^ 8

a

1-28

a

3

d
• A significant increase in stiffness provided payload and equipment

attachments span to three node points.

• A sig-ificant decrease in torsional-type, thermal-induced

deflections.

• Both of the aboie advantages may permit simplification of the

control system, depending on the specifics of the design.

• Significant reduction in GEO orbit transfer-induced member loads,

particularly for thrust in tYe plane of the platform. This is

particularly favorab'.e to joint designs.

• Significant increase in number of paths for routing of power and

data lines.

The disadvantages of an area platform such as the Generic Area Platform

Figure 1.2-2 are:

o A large area platform has drawbacks as compared with a linear

platform when payload servicing/ replacement is considered. The

platform shown is about 1.5 meters deep. If the depth !- increased, 

the accessibility problem is aggravated. This is a problem which is

common to all large area platforms regardless of the type of

structure — deployment method. 

o Deployable area structures which deploy in two directions (i.e.,

length and width) possess certain drawbacks which are not present in

other concepts:

- Difficulty of controlling and holding the platform while it is

deploying.

- All bans deploy simultaneously in both directions, with root

strength not achievable until full deployment.

Rockwell Inc, is not aware of the need for such an area platform application

for the time period 1990-2000 other than solar arrays and antennas being

developed in other study contracts. Therefore, with the concurrence of

NASA/MSFC, the design of a deployable area platform was discontinued.

1-29

F 

P 

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NNVWN

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1-30

no

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1

i
1.3 DEPLOYABLE PLATFOT4 SYSTEMS CONCSPTS

This section describes the individual component concepts developed that, in

total, comprise the basic building block (Figure 1) from which automatically

deployable platform systems can be constructed. The basic components (Figure

1.3-1) are the deployable truss, utilities integration system, deployment

mechanization and rail system, the main housing into which the foregoing

systems are folded during launch, and the end adapter.

The candidate designs for each of the aforementioned components are discussed

in this section. In Section 1.4, the candidate components are integrated into

total building blocks. Also, in Section 1.4, the integrated building-block

designs were used to construct the generic spacecraft configuration, starting

with packaging in the orbiter and ending with the fully deployed basic

platform system.

The following section discusses the candidate designs developed for each of

the building-block components discussed above.

1.3.1 Deployable Trusses

The establishment of the candidate deployable truss concepts gave seripus

consideration to the applicability of existing concepts. The search included

review of in-house documents, the applicable documents listed in the four

Large Space Systems Technology (LSST) bibliographies (Reference 12), and

reviews of reports and discussions with the associated study managers of

concepts recently developed/documented. Figure 1.3-2 illustrates most of the

designs that were compiled as a :exult of that effort. Barring unseen

proprietary designs, there is ,io deployable structure panacea, i.e., a

structure that can be doubly folder, into a very compact configuration (with

utilities integration) that cau ne integrated into an automatically deployable

platform system.

Figure 1.3-2 encompasses flight-proven designs, designs for which

demonstration model© have been made, and proposed concepts. These designs

were reviewed on the basis of their suitability to satisfy the adopted

strength and stiffness requirements and complement of utilities; compatibility

with the total building-block approach; and compatibility with the single bay

at a time deployment approach with maintenance of root strength during

deployment. The designs using X-braced tension cables such as designs A, B,

G, H, and K represent single-folded structures that are not compatible with

the adopted GJ stiffness requirements.

The same comment is applicable to Concept F. A demonstration model of this

concept was observed at the AIAA symposium in Long Beach, California, held on

May 14, 1981. The design is a box truss containing circular longerons and

I-section battens into which the longerons nest during stowage. The diagonal

system uses X-braced tension straps preadjusted on the ground, so the preload

is induced upon extension/locking of the longerons. Strap tension is

approximately 45 N. No evidence of utilities integration was present in the

model cell, which was approximately 4.6 m on its side. Demonstration of the

model was presented by a series of staged slides. The basic deployment

1-31
UTILITIES

INTEGRATION SYSTEM

DEPLOYMENT MECHANIZATION

AND RA ! ' SYSTEM

DEPLOYABLE

TRUSS

^ END

•,r 1 ADAPTER

b^

ORIGINAL PAGE1Y

OF POOR Q

J1 _ ~ V BASIC BUILDING BLOCK

Figure 1.3-1. Deployable Platform System Components

—Basic Building Block

1-32
ORIGINAL PrC- 1-3

OF POOR QCALRY

(A AstroResearch

CONTINUOUS

LUNGEPON

MAST

QE

OOUBLF.-CELL,

DWDLY:-FOLD

TRUSS

Vmht S. G Rockwell

WAOF•N TRUSS- R

TRANSVERSE

PO LD

B AstroResearch

ARTICULATED

L40NOERON

MAST

F

BOX TRUSS

WITH

X-BRACING

Martin

I

Rockwell

CABLE

CROSS - BRACEDTRANSVERSE FOLD

C stroResearch

EXTENDIBLE

SU

SUPPORT

SUPPORTRE J

FOR SEASAT

©

TOWER

SEGMENT

Harris T

I

K Rockwell

CABLE CROSSLONGED

LONGITUDINAL

FOLD

601k

Qp General^namfcs

^l^^(K

TETRAHEDRALi

TRUSS

H

TAPF. RF.D

TURF.

Lockheed L Rockwall

LONGITUD INAL

NIN

FOLD

FOLD

, q

INCORPORATE INTO CANDIDATE CONCEPTS

O TETRAHEDRAL TRUSS O MECHANIZATION

SYSTEM

Figure 1.3-2. Review of Applicability of

Industry Existing and Rockwell IR&D Designs

concept ( obtained from Reference 8), however, for one bay at a time deployment

utilizing stored strain energy (with retention of the remaining structure) was

used initially in the development of Concepts 4 and 6.

Concept E has been successfully tested in the Neutral Buoyancy Tank at MSFC

(References 9 and 11). This design has a limited length storable in the

Shuttle cargo bay, since the longerons are not folded. This concept is

regarded as essentially an erectable concept, and was not considered further,

since no apparent method was foreseeable to use this concept in an

automatically deployable platform system. The Rockwell-proposed designs (I

and J) were not considered further for the same reason despite the attractive

feature of non-folding longerons.

Concepts L and M were not pursued further because of poor packaging

characteristics. Concept C (Reference 10) was not significantly different

from Concept 1 (Figure 1.3-4).

Concept D represents a very efficient double-folded structure. The basic

structure mechanization is quite simple ( in the context of the complexity of

double folding). Examination of the demonstration model ( courtesy of General

Dynamics) revealed well designed joints ( concentric load paths and a minimum

of material). This concept was therefore included (as Concept 3) in the

candidate designs, although difficulty with integration into an automatically

deployable spacecraft configuration was anticipated. Unquestionably, the

design is extremely attractive in an erectable platform system that, through

use of a fixture, joins together numerous deployable truss modules.

1-33
s.

URIWNAL PAGE 15

OF POOR QUALITY

The development of new trues concepts resulted in the matrix of truss designs

Shown iz Figure 1.3-3• This matrix of designs was derived to include the

scope of single and double-folded designs, a design with x-braced tension

cables that could satisfy the GJ stiffness requirements, and designs using

triangular and square cross-sections. The matrix of these design variations

is shown in Table 1.3-1.

OIENTAHEDRAL TRUSS (LONGITUDINAL FOLD( O WARREN TRUSS ILONOITUDINAL FOLD)

nr^W j

^r un

•r, NWW

— Irr,

^

^nr un

umr LrnW w.rn

OGO TETRAHEDRAL TRUSS (DOUBLE FOLD) O TRIANGULAR TRUSS (LONGITUDINAL FOLD)

A-BRACED TO I N BA TRUSS WITH

MNS

O

RECTANGULAR TRUSS & TRIANGULAR TRUSS l / ' WITH

\^^11/

L.. i Im^^

-

a.uu uwnn. ar.^., enoi.0

nuwr .w,o^w n^aw

unwu

nuunw a n

,.m. "^"' rnus^n

urnn

Figure 1.3-3. Candidate Deployable Platform Structure Concepts

All of the designs shown (Figure 1 . 3-3) satisfy the requirements in Section

1.1, with emphasis placed upon:

o One bay at a time deployment to maintain root strength during all

phases of deployment. A rail system is provided on the

building-block main housing for root strength. Upon completion of

deployment, the main load path from truss to truss is through the

adapters and main housing (Section 1.3.4)•

o Satisfaction of the adopted strength and stiffness requirements in

Section 1.1.1.

o Capability to be integrated into the building block concept shown in

Figure 1. For example, the matrix of design variations (Table 1.3-1)

includes no designs that have only a lateral fold despite the significant advantage of such a design. Such a design could have clevis

,joints in the longerons approximately 17 TO apart.

I •

1-34
t

t

R

URIGINAL PAGE 19

OF POOR QUALITY

Table 1.3-1.

Matrix of Truss Variations

0J W O W O

O O

'J N N Y^ W C

W - O N

J J`NN

~ ^C C W

- 4 U D

^JE(

(

~ L

GEOMETRIC

K W

y

_

^

i

W

^ ^`` z

SHAPE J J J U i J .] H

// 

U

^^ 

L

O

© (D O

The eight truss concepts shown in Figure 1.3-3 are discussed in detail in

Section 1.4. The structural drawings are presented in Volume II, The main

features in the development of each of these trusses are as follows:

• Concept 1 (Figure 1.3-4) utilizes the kinematic advantage of the

General Dynamics (,D) tetrahedral truss (for a single-folded

design), the folding advantages of nesting the longerons and

diagonals into the pyramidal members, and provision of clear

space for utilities support trays.

• Concept 2 (Figure 1.3-5) utilizes the high packaging efficiency

of the offset longerons and shear panel as shown. The design

requires only the diagonals shown in the top plane. Redundancy

for meteoroid impact can be provided by addition of lower plane

diagonals. The shear panel was used to provide lateral stiffness

to the hinge line member to minimize deflection due to the offset

longerons.

• Concept 3 (Figure 1.3-6) is included for the reasons discussed

previously.

• Concept 4 (Figure 1.3-7) places emphasis on structural simplicity

(in contrast to the designs of Concepts 2, 3, 5, 7, and 8), both

in regard to member shape and kinematics. All the structural

members are circular tubes, with folding of the longerons and

telescoping of the diagonals as shown.

1-35
ORIGINAL PAGE IS

OF POOR QUALITY

LATERAL

^MI

i

SECTION A-A

GIIOMING WHGERONS NESTING IN

THE PYRAMIDAL MEMBERS

Figure 1.3-4. Concept 1--Pentahedral Truss

PANELS

r

Figure 1.3-5. Concept 2—Warren Truss

s

1-36

A

rI
01 AGORA1

ORIGINAL PAUL 13

OF POOR QUALITY

STOWED

Figure 1.3-6. Concept 3—Tetrahedral Truss

fi

wnuEROR

PRAYE Mwn

9TO^E0

Figure 1.3-7. Concept 4—Truss

1-37
r

END VIEW

ETOEED END VIEW

DEPLOYING THE FIRST Voin

TENSION STRAP

BAT'

FR,

z

•

UsilUINAL PAGE IS

OF POOR QUALrN

o Concepts 5 and 7 (Figure 1.3-8) utilize the advantages of double

folding (increased structure depth, fewer members and Joints, and

increased packaging efficiency) with provision of a rigid main

housing structure. Both concepts utilize telescoping battens to

deploy laterally to the larger crone-section shown. The increased

depth provided by the double-fold permitted consideration of

X-braced tension cables (Concept 5). However, in view of the ever

present concern for maintaining cable tension throughout the thermal

variation spectrum, a design with compression diagonals was

considered (Concept 7).

IONGLRON

Figure 1.3-8. Concepts 5 and 7—Truss

o Concept 6 (Figure 1 . 3-9) establishes a square truss version of

Concept 4, i.e., the Fame emphasis upon structural simplicity.

The square truss can be either statically determinate (battens

braced at end bays only) or redundant (all batten bays braced).

1-38
JGERONS

mLcj^vrnvv vimvwiNAL

ORIGINAt PAGE 19

OF POOR Qk,A. rr'f

Figure 1.3-9. Concept 6—Truss

o Concept 8 (Figure 1.3-10) has the advantages of the square truss

(redundancy, ease of building-block to building-block attachment,

accommodation of payloads), but with the high packaging efficiency

achieved through nesting of the Longerons into the battens, and

the availability of the total area inside the batten for mounting

utilities in trays.

w

i I

J I

DETAIL OF JOINT

Figure 1.3-10. Concept 8—Truss

1-39

• y
1.3.2 Utilities Installatio

1.3.2.1 Requirementa/Ob,jectivee

The requirements/objectives for installation of the utilities into the

deployable structure are as follows!

o Incorporate the adopted requirements for power, data, signal and

fluid lines established in Sections 1.1.2 and 1.1.3.

o Growth capability is desirable.

o Automatic deployment, as part of the building block, in space and on

the ground with manual or ground support equipment (GSE) assistance

if necessary.

• Retraction in space is not necessary, but on the ground retraction

is required with manual or GSE assistance if necessary.

• Minimum number of Joints/connections along a truss with no ,joints

preferred.

• Minimum number of in-space connections with none being preferable.

This depends as much on the method of integrating the building

blocks into the platform, as it does on integrating the utilities

into the building block..

• On the ground end-to-end checkout is highly desirable including all

building blocks and possibly small payloads with no break/remake

connections between ground checkout, loading in the orbiter, and

final deployment.

• Protection from adverse environments (thermal, radiation, vibration

during launch)

o High reliability

• Weight - not a big driver for LEO but important for CEO.

• Separation of power and data/signal lines (minimum electrical

interference)

? o Accessibility - ease of installation, maintenance and replacement,

j and accommodations of design changes

1.3.2.2 Installation Methods

Four general methods for installation of utilities were investigated;

1-40 
sw a opmonan./tn4g(aaon aI Rockwell

S001111b SytNms Dlvlvwn ^ ^ International

0 In trays

• In coils or loops of several configuration.

• On the outside rf structural members; for example, secured

to the outside of longerons

o Inside structural members such as the longerons

The results of a design review are shown in Table 1.3-2. .e conclusions to

be drawn from this review are:

o No one method of utilities installation is best for every

configuration and mission.

• If room is available, the tray method is preferred for

integration of a large quantity of utilities.

• For a limited number of utilities, installation on the outside

of longerons is preferred.

1.3.2.3 Fluid lines

The fluid lines are nominally two centimeter • iameter flexible lines

containing Freon for cooling purposes. Other sizes and shapes are available

to suit the specific requirements of the various configurations investigated.

The two main problems encountered in using such lines are the bend radius and

associated bending moment, and temperature control.

The bend radius required in any truss installation depends on the pitch

distance between batten frames in the packaged configuration. The designer

naturally attempts to keep this distance as small as possible to obtain a high

packaging ratio, hence the importance of bend radius. If it is ner9ssary to

maintain the folded line in one plane the beat bend radius obtainaoia is equal

to the pitch distance less half the diameter (Figure 1.3.-11). If it is

ossible to overlap the lines out of plane, the bend radius can be increased

Figure 1.3-12).

The acceptable bend radius of a 2-centimeter-diameter flexible fluid line is 3

centimeters with an associated bending moment of 7 Nm (courtesy of Metal

Bellows Corp.). In the avant that a 2-centimeter-diameter line cannot be

installed, there is always the alternative of using a greater quantity of

smaller lines i.e. four 1.4 centimeters diameter lines. Another interesting

possibility is the "race-track" shape (Figure 1.3-11) which permits a small

bend radius. Table 1.3.-3 lists some characteristics of utilities installed

on the Rockwell building blocks. In cases where the permissible bend radius

of the installation falls below that which is recommended for a 2 centimeters

diameter line, the system uses smaller lines.

Fluid lines require thermal insulation to prevent the Freon from freezing

during eclipse periods. Six layers of MLI (multi-layer insulation) is

estimated to be sufficient protection (Section 1.1.3). The MLI is retained on

the outside of the flexible line by a loosely woven nylon ,jacket which will

not interfere with the flexing of the line.

1-41

r
ORIGNAL PAGE 19

OF Po©R Q

Table 1.3-2. Comparison of Utilities Installation

METHOD PRO CON

° SIMPLE INSTALLATION —►tRMITS ° EXTRA COST

STRUCTURAL ASSEMBLY, RIGGING, AND ° EXTRA WEIGHT (CEO)

TESTING PRIOR TO UTILITIES INSTALLATION ° EXTRA SPACE

° ACCESSIBLE FOR DESIGN CHANGES

° EASE OF MAINTENANCE B REPLACEMENT

TRAYS ° GOOD GROWTH POTENTIAL

° METEOROID IMPACT PROTECTION

° SEPARATION OF POWER AND DATA

° GOOD SUPPORT FOR LAUNCH

° DISPERSED AGAINST METEOROID IMPACT

° GOOD HEAT DISSIPATION FOR ELECTRIC

POWER LINES

° SIMPLE INSTALLATION —PERMITS ° GROUND CHECKOUT — INCREASED FOLDING

STRUCTURAL ASSEMBLY, RIGGING, AND TIME

TESTING PRIOR TO UTILITIES INSTAL- ° LIMITED GROWTH POTENTIAL

LATION

COILS N ACCESSIBLE FOR DESIGN CHARGES

° NO METEOROID IMPACT PROTECTION

S

° NOT DISPERSED AGAINST METEDROID

LOOPS ° EASE OF MAINTENANCE t REPLACEMENT IMPACT

° LIGHTWEIGHT, LOW-COST INSTALLATION ° LAUNCH SUPPORT COULD BE PROBLEM

° FAIRLY GOOD HEAT DISSIPATION FOR

ELECTRIC POWER LINES

° GOOD SEPARATION OF POWER AND DATA

• LARGE BEND RADII OF LINES

° SIMPLE INSTALLATION ° LIMITED NUMBER OF UTILITY LINES

° ACCESSIBLE FOR DESIGN CHANGES ° MAY NOT BE ABLE TO SEPARATE POWER

AND DATA

OUTSIDE • EASE OF MAINTENANCE B REPLACEMENT

° NO METEOROID IMPACT PROTECTION

OF • GOOD SUPPORT FOR LAUN [M

LONGERONS ° NOT DISSPERSED AGAINST METEOROID

° GOOD NEAT DISSIPATION FOR ELECTRIC IMPACT 

POWER LINES

• LOW-COST LIGHTWEIGHT INSTALLATION

° METEOROID IMPACT PROTECTION ° SMALL BEND RADII OF LINES

(NEW TECHNOLOGY)

° GOOD SUPPORT FOR LAUNCH ° MAY REQUIRE INCREASED LAY ANGLE

° LIGHTWEIGHT INSTALLATION (INCREASED WEIGHT L LOSSES)

° DIFFICULT TO INSTALL—INSTALLATION

INSIDE DURING PIECE-BY-PIECE STRUC. ASSY.

OF ° NOT ACCESSIBLE FOR DESIGN CHANGES

STRUCTURAL ° DIFFICULTY OF PAINT, S REPLACEMENT

MEMBERS ° GROWTH POTENTIAL POOR

° NAY NOT BE ABLE TO SEPARATE DATA

AND POWER

° NOT DISPERSED AGAINST METEOROID

IMPACT

° POOR HEAT DISSIPATION FOR ELECTRIC

POWER LINES

° NOT SUITABLE FOR SMALL MEMBERS

(SMALL PLATFORMS) 

1-42

C
URIGINAL PAGE 19

M

OF POOR QUALrTY

BEND RADIUS

;H

Figure 1.^-11. Utilities in One Plane

RACETRACK CRUSS-SECTION

SHAPE PERMITS REDUCED

BEND RADII

END RADIUS

Figure 1.3-12. Utilities with Overlap and

Increased Bend P.adius

1-43
Table 1.3-3. Electrical and Fluid Utilities Characteristics

ELECTRICAL FLUID

BEND BEND

RADIUS RADIUS

CONCEPT METHOD

^r

(cm) METHOD (cm)

1 TRAYS 6.3 TRAYS 12

2 TRAYS 3.3 OVERLAP FOLDS 20

3 COILS 24.0 OVERLAP FOLDS 20

4 OVERLAP FOLDS 22.3 OVERLAP FOLDS 14

5 TRAYS 2.9 TRAYS 4

6 OVERLAP FOLDS 22.3 OVERLAP FOLDS 14

7 TRAYS 2.9 TRAYS 4

8 TRAYS 2.5 TRAYS 3.1

8 LONGERONa 20.3 LONGERONS 20

1.3.2.4 Electrical Utilities

The electrical utilities consist of power lines; twisted shielded pairs for both

.signal and data lines; and for data lines, fiber optics and coaxial.

The twisted shielded pairs and coaxial lines are of small diameter and are

flexible and pose no problems pertinent to folding in the trays or into loop'+

as required. The type of overall platform design envisioned by Rockwell

avoids the problems of multiple coaxial connectors. There are no coaxial

connectors or. the trusses where flexin occurs and, since the entire platform

is packaged and deployed as a unit witnout piece-by-piece assembly, there are

very few connectors in a line which may run from one end of the platform to

the other, traversing several building blocks in the process.

Fiber optics are sensitive to radiation degradation and to cracking caused by

thermal cycling. also, at low temperature some fiber optics cease to

transmit. The present drive in the industry is to improve fiber optic

materials and to develop shielding to overcome these problems. This is listed

as a line item in the technology development seition. In the event fiber

optics are not ready for 1986, copper 'lines can be used instead.

Power lines for space platforms do not need the he —•,y insulation commonly

found on electrical cables for earth applications. There is no moisture

problem in space and no requirement for "idiot proi.t' ruggedness".

Consequently, a light, looael'•, v, -^n insulation is recommended. Such

insulation will save weight, f- :d :lexiuility, and avoid cracking when

folded.

M, s acceptable bend D/d rati: cciated with the sizes of power cables

selected are shown in Table

1-44

' ♦ f
I

s^

e g

ORIGINAL PA17E 13

OF POUR QUAL"

Table 1.3-4. Permissible Electrical Conductor Dld Bend Ratios

CABLE SIZE D/d

BEND RADIUS

(cm)

0 6 2.45

2 6 1.96

4 6 1.55

WIRE BUNDLES 10 23

The actual bend D/d ratios (Table 1.3-3) all exceed these acceptable values.

The permissible D/d shown ir. Table 1.3-4 were obtained by review of the parametric data (Figures 1.3-13 through 1.3-16) furnished by Tension Member

Technology (TMT). The review used a maximum permissible strain of 2.5% based

on upward adjustment of the strain data shown in Figure 1.3-18 since the data

shown are for cyclic reversal of strain which does not occur in the bending of

the cables. Further, the value of 2.5% is regarded as reasonable since the

following analysis in(.ludes the strain imposed during fabrication of the cable

which occurs only once. The basis of the 40 cycles of strain shown in Figure

1.3-17 comes from the assumption of 10 cycles of folding (during installation,

checkout, and final deployment) multiplied by a scatter factor of 4.

The analyses are all based upon the use of a 15-degree lay angle (essentially

standard cable fabrication practice).

• For a No. 2 conductor a maximum strain of 1.5% is determined

from Figure 1.3-13 for D/d = 6. The D/d = 6 is conservative.

• For a No. 0 conductor a maximum strain of 1.1% is determined

from Figure 1.3-14 for D/d Q 6. The D/d = 6 is conservative.

• For a 1 x 7 wire bundle comprised of No. 2 conductors, from

Figure 1.3-15, the curvature ratio fox bending to a D/d = 10

is 19, the initial wrap curvature is 32. From Figure 1.3-13

the maximum possible strains are respectively 1.2 and 1.15%,

or a total of 2.35%. A 1 x 7 bundle of No. 0 conductors is

less critical. It is pertinent to note the cyclic strain

is less than half of the total strain.

• A 1 x 19 wire bundle to a D/d = 10 will have less strain.

The bending moments associated with bending these cables, to the D/d ratios

discussed herein, are not significant to the concept development. For example,

the limit bending moment incurred in bending a No. 0 cable to a D/d = 6 is

0.7 ivm.

The background for the foregoing analysis is based upon the additional documentation (and discussions) provided by TMT as follows:

The bending of a helically wound cable produces bending, extensional, and

torsional strains in the individual ciL'_r elements. The bending strain is a

result of the change of curvature of the elements as the cable is bent.

1-45

1 s 
ki

Okl(ANAL PAGE k.

OF POOR QUAD Y

y—D/d - 2 — - -

D/d 4 — -

O- e 0

D - Pend Diameter

d - Cable Diamete r

T

6

—

---- 8

--- _- ^/d 10 -

- - - D/d 1

_ D/d - INFINITY

0 5 10 15 20 25 30 35 40

LAY ANGLE(S), DEGREES

5.0

4.0

3.0

2

2.0

OC

N

C^

_2

UZ

1.0

X 0.9

X 0. 8

0. 7

0.6

0.5

0.4

0.3

Figure 1.3-13. haximum Bending Strain

—No. 2 Conductor

II

1 -46
1.0

z 0.9

0,8

`^ 0. 7

z

0.6

m 0.5

7

7

0.4

I w

0.3

r

S.D

1.0

3.0

2.n

0.2

ORIGINAL P.4re lq

OF F'CX)R QUALITY

— - --

o - 0 .. • 0

D - hand Diameter

d - Cable Diameter

_

— - -

– -

35 40

0.1

0 5 10 15 20 25 30

LAY ANGLE(S). DEGREES

Fi gure 1.3-14. Maximum Bending Strain

—No. 0 Conductor

1-47
ki ORKANAL FACE i7`

OF POOR QUALITY

Inn n
1

D - Pend Disme

d - Cable Dism

---

_-

_

— --- - _

tar

Itor

100.0

90 0

l0.0

0

r 10.0

60.0 i

50.0

A

a

Uu 40.0

V

2

°x 30.0 w

m

S

= 10.0

7

ORt':,ff\^!'i^ r'F1t t

OF POOR QUAL 

C

fn/

LIL 4

15 20 15 3S 40

LAY ANCLE(S), DEGREES

Minimum mending Curvature Ratio 

—1 x 19 Structure

> 10 100 1000

CYCLES TO FAILURE

Figure 1.3-17. Copper Filament

Fatigue Data
•

Extensional strain occ-irs when the elements of the cable are prevented from

axial movement due to friction forces among them, thus requiring the elements

to elongate or foreshorten to comply with changes in their local path length

as the cable 1s bent. This friction - induced extensional strain of the cable

elements is in addition to any strain which may exist due to tensile loading

and elongation of the complete cable assembly. Torsional strains are the

result of twisting of the individual elements due to bending distortions of

their helical paths.

Eoth the bending and torsional strains induced in individual cable elements

become smaller for decreasing element size relative to the total cable size.

Extensional strains due to bending become smaller with reduced values of

internal cable friction.

The analysis described in this report assumes that both the extensional and

torsional strains are negligibl9 and that only bending strains are

significant. This assumption is valid for a frictionless cable wherein the

elements are free to move axially with respect to each other. The known

techniques for approximating this frictionless condition are the utilization

of ample cable lubrication and/or a " loose pack" cable design. For some

applications, low cable friction can be achieved by enclosing the cable

easembly within a loose- fitting tube or hose which contains a lubricant such

sea liquid fluoropolymer. Of course, such a design is not suited for

deployable apace structures. However, a " loose pack" cable design is quite

appropriate. In such a cable, the individual elements ideally are helically

wound and held loosely together with some average spacing among them. The

reeult is a nearly frictionless cable which provides for minimum strain on the

individual elements during cable bending.

The No. 2 (19 x 35) and No. 0 (19 x 55) cable constructions are defined as

"cables" which consist of 19 " strands," each of which consists of a number of

"elements" ( 35 and 55, respectively). The 19 strands are arranged with one

strand in the center of the assembly, a layer of six strands around the center

strand, and a layer of twelve strands around that. The 1 x 7 and 1 x 19

constructions are defined as "cables" ( or "strands") which consist of a number

of "elements" (7 and 19, respectively). In each case, the "elements" are the

amellest subunits of the construction.

The numerical analysis included computation of the bending strains induced in

the individual cable elements both during original manufacture of the cable

assembly and as a result of bending this assembly to various ratios of bending

diameter to cable diameter. Figures 1.3-13 and -14 show the results of this

analysis for the No. 2 and No. 0 cable constructions. For this analysis, it

was assumed that the helical direction of the elements within the individual

strands was opposite that of the stranda in the complete cable assembly. It

was also assumed that the 19 strands were manufactured as a "bunched" strand

configuration. Furthermore, the analysis required that the lay angle of the

elements within the strands be equal to the lay angle of the strands within

the cable assembly. Without this restriction, the analysis would have

required an additional plotting dimension. 

1-50 
The curves in Figures 1.3-13 and -14 are dimensionless and can be applied to

any size cable made with the indicated configurations. On each of these

curves, the reference angle 0 defines the location of the strand within the

cable (0 - 0 corresponds to the strand furthest from the center of curvature

of the entire cable seeembly)• Similarly, the reference angle 0 defines the

location of an individual element within that strand (0 m 0 corresponds to

the element which is furthest from the cable centerline). The combinations of

0 and 0 indicated in these figures correspond to the locations of the

maximum element bending strain (minimum radius of curvature).

The curve for D/d - infinity corresponds to a straight cable assembly and

indicates the maximum bending strain (minimum radius of curvature) in

individual elements as the result of the elements being formed into a double

helix during the cable manufacturing process. The location of this maximum

bending strain corresponds to 0 - 0 for all values of G. In other words,

for a straight cable, this maximum bending strain occures in the cable

elements which are furthest from the cable centerline.

The remaining curves for various values of D/d indicate the maximum bending

strain (minimum radius of curvature) in the individual elements after bending

' the entire cable assembly. In all cases, this maximum bending strain is that

which is produced as a cable element, which is initially straight, and assumes

some final radius of curvature in the bent cable assembly. Not iLcluded in

this analysis is any initial state of strain which the individual elements may

r. have had as the result of wire drawing or heat treating processes occurring

prior to the elements being formed into the cable.

Note that the location of the maximum bonding strain changes as a function of

both the element lay angle and the ratio of the bending diameter to the cable

diameter. The reader is cautioned against attempting to take the difference

between the D/d - infinity curve and any other curve to determine the change

in element bending strain due to bending the entire cable assembly. This

procedure will not yield accurate results in all cases, since the site of

maximum pending strain within a straight cable may be different than the site

of maximum bending strain within a bent cable. Furthermore, even if the sites

of maximum bending strain are identical in both the straight and bent cable,

it is possible that for some cable geometries and bending diameters, the

radius of curvature of an individual element may pees through infinity as the

cable is bent, thereby producing a change in strain due to bending which is

greater than the net bending strain which sxisto either before or after the

cable is bent.

Figures 1.3-15 and -16 for the 1 x 7 and 1 x 19 configurations indicate the

minimum bending curvature ratio, which is the minimum radius of curvature of

an outer layer element divided by the radius of that element. Again, these

curves are dimensionless, and they can be applied to any size cable made with

the indicated configuration. The curve for D/d - infinity can be used to

deacribe the minimum radius of curvature of one of the outer elements in a

straight cable. The remaining curves indicate the minimum radius of curvature

of one of the outer cable elements after the cable is bent.

1-51 .1
All the four curves have been corrected to take into account the fact that the

diameter of the complete cable assembly increases with increasing lay angle of

the individual elements and strands. In other words, for a D/d ratio of 10, a

larger bending diameter is required for a cable assembled with 20-degree lay

angles than is required for a cable assembled with 10-degree lay angles.

The curves in 1.3-13 and -14 may be used directly to determine the maximum

bending strain produced in an individual element of a No. 0 or a No. 2

conductor as the result of the original cable manufacturing p.*ocese or as a

result of bending the cable to a specified bending diameter. 1t is important

to note, however, that when establishing a value for D/d, the "d" applies to

the outside diameter of a bare conductor. Any insulating ,packet which is

applied over the conductor must be ignored for purposes of determining the

maximum strain.

All the nurves require that the lay angle of the cable component be specified

in order to determine the maximum bending strain or the minimum bending

curvature ratio. For a cable component which follows a simple helical path,

the lay angle is defined acs

Lay angle - Arctan 27rr

k

where

r - the pinch radius of the cable component as measured

from the axis of the helix to the center of that

component., and

Z - the lay length of the cable component (the distance

measured along the axis of the helix corresponding

to one helical pitch of the cable component).

1.3.3 DeploZment Meehan_zation Concepts

The requirements for the deployment mechanism are:

o Automatic deployment - in space, on the ground manual or GSE

assistance is permissible.

o Retraction 1s nice but not a firm requirement.

• One bay at a time deployment.

• Controlled rate of deployment. Sometimes it may be necessary to

synchronize several trusass being deployed.

• Root strength of trues maintained throughout deployment.

• Suitable for use with the building block approach for deployment of

a platform.

o Compact or foldable.

1-52

Ir
o Low power consumption.

o High reliability.

o Suitable and safe for EVA operations in the event of a malfunction.

o Able to generate extra force ir. the event of a "hang-up" or Jam.

The following sections describe the deployment techniques considered in this

study.

1.3.3.1 Pressurization

Pressurization eyetsms were investigated but no suitable applications were

discovered.

1.3.3.2 Cable Deployment Systems

Cable systems are sometimes used for deployment and/or retraction of trusses

and similar structures. They may be used as the prime motive source or in

conjunction with deployment springs, in which case the cable may function as a

restraining device. Cable systems tend to deploy all the bays of a truss

simultaneously or in a random fashion. This may be acceptable for a single

trues but is not acceptable for a system with many trusses. A remedy for this

s- uncontrolled deployment is to tie all of the batten frames together by latches

vhich are released sequentially. Alternatively a "hold-back" sequencing 

FF mechanism can be incorporated in the main housing. Another and more complex

F" method is to have a separate cable system for each bay.

Depending on the truss being used, a cable system will probably not develop

root strength of the truss unlesa it is aided by auxiliary guide rails.

Although cable systems were investigated, the truss concepts developed herein

wero not appropriate for the use of cables.

1.3.3.3 3 red Energy Deployment Systems

Of the stored energy devices available, mechanical springs are the most

suitable for truce deployment. Other devices such as gas cylinders are more

complex and bulky and lees relilable. Mechanical springs used may be tension,

torsion, compression, leaf or any of the other forms available.

Spring deployment systems have many of the same characteristics as cable

systems; simultaneous bay deployment, lack of accurate control over deployment

rate, and lack of root strength development. These drawbacks can be overcome

by the addition of other devices, such as restraining cables, sequenced

latches and guide rails.

Concepts 4 and 6 were initially designed for torsion spring deployment using

sequenced mechanisms as shown in Figure 1.3-18. This is a mechanism which

holds adjacent battens together as part of the truss stowed stack. Upon

receipt of a command signal, a fuse wire which holds together the two halves

of a nut is melted. The nut and bolt separate and the batten is free to

deploy under the influence of torsion springs (not shown).

1-53
i3

• LONGERON= OVER CENTER LINKAGE HINGE

0

• LONGERON: CENTER LATCH

r

ORIGiMAL PAGE IS

OF POOR QUALrN

_

All I

r

I latching mechanism

^ ^^ L ^' • —' ( between No. 1 8 No. 2

battens. Shown In the

open configuration.

3 BATTEN NO. 2 BATTEN

NO. 1 BATTEN

DEPLOYED

Figure 1.3-18. Latching Mechanisms

between No. 2 and No. 3 Battens

Spring deployment is usually not as positive as a mechanical drive. It is

sometimes difficult to obtain a reserve of force without inducing unwelcome

accelerations in the item being deployed. For longerons which unfold and

move Lo an over center or "in line" configuration, light springs located at

the longeron center hinge are very useful. Theycan provide a kick over the

last few degrees of movement wt.ich is difficult to obtain by other means.

Such springs were used in the design of longeron latches as shown in

Figure 1.3-19.

EE 4-

• LONGERON: OFFSET HINGE WITH HOOK LATCH A; LONGERON= IN LINE HINGE

Figure 1.3-19. Longeron Latching Mechanisms

1-54
ki

UHI411VNL PAGE 19

OF POOR QUALITY

1.3.3.4 Mechanical Deployment System

The mechanical deployment system used is the same or equivalent to the reciprocating tape and pulley with integral guide rails developed by General Dynamics

for their tetrahedral truss (Figure 1.3-20).

Figure 1.3-20. Deployment tally

M and n'echani,+.5

The system consists of two guide rails (Fig-,re 1.t 11), each carrying a tape

and pulley arrangement which advances to deploy one bay, then returns to

deploy the next bay. To completely control the motion of the truss and to

develop root strength, the guide rails need to extend a distance of two

deployed bays from the front of the stowed truss stack. The mechanism needs

to extend only half way along that distance. The pulleys are motor driven

and are controlled electronically. A method of folding the guide rails and

mechanism is shown on Drawing 42712-016, sheet 4 (Volume II).

DRIVE

DIGITAL

SNAFTENCOOER

UTCN

\

^ SUPPORT RAIL

TRUSSSEAM "DOE

MASTER

MOTOR

/ POWER

CONTROLLER

Figure 1.3-21. Detail of Deployment Rail

1-55
PIL

1

RAILS

STOWED

RAILS IN

POSITION

ING

UNI(iIAVAL PAGE IS

OF POOR QUALrN

For n truss whose profile changes during deployment (concepts 1, 2 and 3) the

two rail system gust described is used. However, for trusses who** profile

does not change during deployment (concepts 4, 5, 6, 7 and 8) a variation is

available. It is possible to use 3 or 4 guide rails, (depending on whether

the truss is triangular or square) the length of the rails being equal to only

ens bay in front of the stowed truss stack. The operation of the rails and

mechanism remains essentially unchaagci.

The reciprocating mechanical deployment system is the design selected for the

deployable platform. It meets all of the requirements and does not suffer

from the drawbacks of the other systems considered.

1.3.3.5 Payload Deployment

One problem associated with the use of guide rails arises when extending a

truss which has a payload or module ao wide that the guide rails cannot

straddle it, (Figure 1.3-22) and therefore cannot be unfolded until the truss

bee extended and moved the large payload out of the way. Obviously, if the

guide rails are not in position when the truss/payload is moving, there is no

root strength developed.

Figure 1.3-22 Deployment Rail Issue

1-56

At
BOOMS RELEASED

L RETRACTED

°""" "' POSITION

2 SPAR BOOMS RAILS STOWED

EXTENSIONS

'M OIZIGINliL f= i. ;E 19

The •e are several possible eol:•tlone.

OF POOR QUALrN

Solutior. J: Use fixed rails ( Figure 1.3-23). This design uses more stowage

length in tle orbiter and the large payload is not supported

directly from the main housing.

Figure 1.3-23. Fixed Rail Method

Solution 2: Penetrate the module (Fi4r -a 1.3-24). It is not likely that the

rails would be allowed 'a penetrate an actual payload but it is

quite possible that t 'aey could penetrate a structural module.

-= 

Figure 1.3-24. Penetration Method

Solution 3: Spar biome ( Figure 1.3-25)• This appears to be the mosrt feasible

solution to the stated p roblem for situations with payloada and

modules an well.

Figure 1.3-25. Spar Boom Method

. F

1-57 17 -3

• ......uwn,........a....>,.txu..+u..Ere...:,^.zw,........,..^ri+
6Y

1.3.4 Main Houainga and Adapters

This section describes the array of main housings and adapter designs (Figure

1.3-26) developed for the candidete concept trusses (discussed in Section

1.3.1) that are compatible with the building-block concept. Additional

information is provided in the drawings listed in Volume II.

Emphasis in :he drawing development was placed upon conce ptual rather than

detailed design. Of primary concern was the overall concept integration with

the truss, rail system, accommodation of docking ports and payloads, and

suitability for inter-building-block attachments. Sufficient detail was

presented to support the weights and meat analyses (Section 1.4).

The primary functions of the housings are as follows:

o Attachment of the deployable truss, mechanization, and rail

system components.

o Attachment of electrical, data, and fluid utilities feed-through

connections (as required).

f

o Provisions for mounting of docking port support rings for orbiter

berthing or payload attachment.

o Provision for structural attachment with adjacent building-block

housings and/or adapters with accommodation of the variations in

the building blocks orientation.

o Provision for str-.ctural attachments for orbiter installation.

The structural concept for all the main housings shown is expected to be built

up from numerically controlled integrally machined aluminum panels (2219-1'6 or

equivalent) to a truss or skin-stiffened construction depending on the

specific design conditions. Thermal gradients can be greatly minimized to

reduce the local thermal distortion (if required). if aluminum is not

adequate, (depending on the pointing accurney requirement) the main housing

can be constructed of composite materials, but with increased cost.

The adapter functions are as follows:

• Provision of attachments to mount onto the extremity of the basic

deployable truss.

• Provision of all hardware to permit subsequent RMS attachment of

payloads, RCS modules, and/or orbiter berthing ports.

• Provision of automatic electrical and fluid line connectors. 

• Provision of structural stability to the main housipg during

orbiter boost.

1-58
OF PCXi 12 QUA^_^

TRUSS

T HOUSING CONCEPT ADAPTER CONCEPT

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1-59 

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UNKAIMl'L PAGE 19

OF PWR QUALM

While three conceits are shown for the adapter, the moat likely construction

is that of a numerically controlled, integrally machined aluminum design

(thermal stability permitting). The machined panel permits the greatest

flexibility for mounting of latches, connectors, etc.

1.4 DEPLOYABLE FL_4TFORM SYSTEMS CONCEPT INTEGRATION

This section duecribee the integration of the specific deployable truss

concepts (1.3.1), utility folding concepts (1.3.2), deployment mechanization

concepts ( 1.3.3) and main housings and adapters ( 1.3.4) into 8 candidate

building block concopts that are subsequently compared in the concept

selection study of Section 4.

Unquestionably, a comparison of these diverse designs for satisfaction of a

specific platforms unique configuration and performance requirements

represents an am - le challenge but is possible within the scope of this study.

Comparison of the 8 concepts for more than one platform of diflerent size,

strength, stiffness, and utilities accommodations needs, while preferable, was

beyond the scope of this study. The beet compromise therefore, was to perform

the comparison in detail ( production of drawings) for one baseline platform

designed to one set of baseline requirements and, to the maximum extent

possible, analyze and/or review the implications associated with departure of

size and requirements from the baseline.

The generic platform described in Section 1.2 was the baseline platform, with

the adopted strength, stiffness, and utilities accommodations representing the

baseline requirements. All 8 concepts were constrained in size to permit

packaging of the generic platform in the orbiter as schematically shown in

Figure 1.4-1. Additional details pertinent to packaging are shown on Drawing

42712-020 (Voliune II). The maulting deployable truss dimensions are shown on

Figure 1.4-2. It is recognized that other options for each concept are

possible. The options shown were the moot likely at that point in the study.

In fact, Concepts 2, 6, end d were subsequently (after completion of

structural and thermal analyses) packaged as shown in Figure 1.4-3, permitting

an increase in the truss width and depth. Hence, the data shown in Section 4

for these concepts are slightly pessimistic.

a el OEED CONCEPT 1 CONCEIT 2 CONCEPT 7

CONCEPTS 4, 5, 7 CONCEPTS 6 r. O

Figure 1.4-1. Concepts for Packaging of

Generic Platform

1-60
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are in meters.

ORIGINAL PAGE 19

OF POOR QUALITY

1.58 FOR

1.47

L T

-

1.27

END VIEW

ST&TDt

Figure 1.4-2. Structural Concept Dimensions

1-61
i

ORIGINAL PAGE 18

OF POOR QUALrrY

S O

CONCEPT 2 CONCEPT 6, 8

W

Figure 1.4-3. Final Concept for Packaging

fconcepts 2, 6, and 8)

Referring to Figure 1.4-2, it is pertinent to note that the largest truss

dimensions compatible with orbiter space available, were used for the

following reusons:

o Packaging efficiency is increased with increase in truss depth

o Best accommodation of utilities

o Fewest structural members and points (reduced cost and ainimum

weight)

The following sections describe the structural sizing method and approach, the

8 integrated candidate building-block Assigns, the packaging of the generic

platform utilizing these designs, and the thermal, mass properties, and

fabrication coat data developed for use in the concept selectio-. Finally, a

discussion of significant miscellaneous issues is presented.

1.4.1 Structural Siring Method and AP4 roach

A summary of the design ultimate compression load, shape, and structural sizes

for the 8 deployable truss individual longeron, batten, diagonal and/or

pyramidal members is presented in Table 1.4-1. The generation of these dada

are described as follows:

o The individual compression loads were obtained by computer analysis

of each truss (dimensions shown in Fig, 1.4-2) for the concurrent

adopted limit bending and torsional moments of respectively,

2.5x104 and 1 x104 Nm. An ultimate safety factor of 1.5 was

used. The computer used equations developed from hand analyses

which are sufficiently accurate for the purposes of this study.

o The individual compression load in conjunction with the member

length and specified material properties was input (automatically)

into a column analyses subroutine. A CRT plot for each member is

obtained such as that shown in Figure 1.4-4 for either longeron,

batten, diagonal, or pyramidal members of circular, square,

rectangular or I shape. The material properties used for the

T300/934 graphite composite design was EL-143,000 and ET -

17,250 mpa and for the P75S/934 EL - 231,000 and ET - 28,000 mpa•

1-62
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ORIGINAL PAGE IS

OF POOR QUALITY

l0 II II b

u 1

8

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b

6 ` Jam}?/`

eb

4 J.,

2

0

0 2 4 6 8 10 12

LONGERON WIDTH, d (cm)

Figure 1.4-4. Example of Column Sizing Data

o For the longerons, data such as shown in Figure 1.4-4 ware compared

with the AE requirement compaAiblg with the adopted flexural (EI)

stiffness requirement of 2x10 am . For example, for Concepts

1, 2, 4, 6 and 8, the AE requirement was more critical than the

column stability requirement. For these cases the higher modulus

material was used to maintain good packaging efficiency and/or avoid

solid members. Note that for an EI requirement equal to half the

adopted value, the T300/934 would be used. The coat penalty with

use of the P75S/934 was accounted for in the fabrication coat

analyses. For all the longerons the sizes were based upon the beet

compromise, through discussion with the designer, in regard to

packaging, suitability of end Joint attachment, and apace for the

folding Joint.

0 For the diagonal, battens, and/or pyramidal members, the data such

as shown in Figure 1.4-4 were developed with sizes determined based

upon the best compromise with the designer as described above. The

sizes obtained rare checked (excepting Concepts 2 and 5) for

satisfaction of the adopted torsional stiffness (GJ). For Concepts

1, 3, 4, 6, 7, ang 8 the desi ne satisfied the adopted torsional

stiffness of 5x10 Nm2 (Concept 3 in fact had a G,T five times

greater than that required). For Coneeot 5 the %-bracing members

AE values were sized to satisfy the GJ requirement. For Concept 2,

NASTRAN analysis was used. The NASTRAN analysis included elastic

stability analysis and shear stiffness analyses of the individual

panel, including the cutouts (reinforcing around cutouts). Further,

1-64

A

.1

i
LATERAL I-SHAPED

MEMBER

CLEVISES

ORIGINAL PAGE IJ

OF POOR QUALITY

since the total Warren truss design included an offset between

the longerons and diagonal braces, a two-bay model of the

structure was made.

The data of Table 1.4-1, as determined by th_ above discussion, are shown on

the structural drawings of the eight candidate concepts. Additional discrete

analyses were performed to support the concept development and mass properties

data tabulated in Section 1.4.5. These analyses encompassed the housings,

adapters, utilities support trays, and launch support cradles.

1.4.2 Eight Candidate Building-Block Concepts

1.4.2.1 Concept 1

The structure (Figure 1.4-5) is a pentahedral truss formed essentially from

one-half of the General Dynamics tetrahedral truss, but with three differences:

• The concept folds only in the axial direction.

• A telescopic diagonal is provided across the base of the

pyramid-shaped bay.

• Members of various cross-sections (e.g., = or I) are used which

nest one within another, thus permitting a higher packaging ratio.

• Dtnon/PO D AXIALLY

• M YQION MAMIDAL M MMId

• 

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O anAr LOADS AT JOINTS

ADAVm

MAIN HONING

avcs wu

Figure 1.4-5. Concept 1

The folded truss is stowed in a rigid triangular housing from which it is

deployed one bay at a time along a pair of guide rails by means of a reciprocating mechanism. To enable the truss to develop root stiffness while deploying, the rails must be a minimum length of two bays plus the stack length. The

guide rails may be folded back alongside the housing for ease of stowing in the

orbiter.

Subsequent to full deployment, root stiffness is developed by attachment of the

truss to the rear face of the housing.

The adapter is a rigid assembly which attaches to, and moves with, the far end

of the truss. It is the interface for payloads or modules and contains mechanical/structural latches, electrical/ fluid interfaces and an alignment system.

1-65

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ORIGMAL PAGE 19

Sp 

Satellitete Sy tems Division ntematlonal

g OF POOR QUAL"

The electrical utilities are mounted in trays which are attached to the truss

pyramidal members and pivot about the tray centerline (Figure 1.4-6). There

is ample room for the full complement of utilities. The fluid lines are

installed on the square face of the trues In a series of pivoting trays.

The packaging ratio is 15 to 1.

w +A w H

N d `1

# d + `;; 1,-

6

Figure 1.4-6. Installation of Utility Trays

(Concept 1)

h w 1.4.2.2 Concept 2

This concept is a deployable Warren truss, as shown in Figure 1.4-7. The

truss is supported in a housing during orbiter launch and is deployed from

it along a pair of tails by a reciprocating mechanism. The length of the

housing is dependent on the number and size of the bays to be packaged. For

the truss to be under control during deployment, the minimum length of the

tails is equivalent to the length of two bays plus stack length. This will 

usually provide fairly long rai:s which reed to be folded for convenient

packaging in the orbiter. During deployment, the root strength of the truss

is developed by the truss attachments rolling in the deployment rails.

Subsequent to full deployment, the truss root strength is developed by the

attachment of the truss to the housing structure. The truss consists of

folding longerons, telescoping diagonals and rigid sandwich panels that form

the shear panels. Because the panels are rigid, there is a choice of design

such as machined waffle or honeycomb panel.

Each longeron is hinged at the middle, and the two half-longerons fold one

inside the other. There is an offset between the axes of the diagonals and

the longerons. 

The adapter is a square, rigid asesmbly which also acts as the last shear

panel in the truss. It is the interface for payloads and other modules and

contains all of the elements necessary for alignment, berthing and utility

interfacing.

The design of the shear panels is particularly suitable for the installation

of utilities in trays (Figure 1.4-8). The electrical power and the

I signal/data cables are in two separate trays, pivoting in Blots cut in the

'j shear panels. The fluid lines which use elbow fittings, to avoid small bend

.j radii, are mounted in similar fashion in other slots in the shear panels.

The packaging ratio is 21.6 to 1.

1-66

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..^Ya
FLUID

ORS IK'1: r"A": I'i

L. !')(if, I' ALrN

*DEPLOYS FOLDS AXIALLY

• PLATE TYPE SHEAR MEMEERS

• ROUND SECTION TELESCOPIC DIAGONALS

o t9i SECTION LONGEkONS

• LONGERONS ARE OFFSET

MAIN

HOUS ING

ADAPTER I

Figure 1.4-7. Concept 1

DETAIL OIF JOINT 

rvncn uivo

Figure 1.4-8. Installation of Utilities

(Concept 1)

1-67

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Uhlh;riAL 0'.L,. I'"

OF POOR QUALfrf

1.4.2.3 Concept 3

Concept 3, the tetrahedral trues, is the 3sneral Dynamics design used in its

entirety (Figure 1.4-91• It is a good, practical design of a deployable trues

which has beai demonstrated as a working model on many occasions. The truss

is a double-f.ld system which packages into a small volume. All of the

members are round tubes which oonverge without offset at the Pointe. Spring*

are used at the Pointe to aseirt in the final deployment of struts and to lock

them in position.

ADAPTER/

Figure 1.4-9. Concept 3

The deployment mechanism is the same as in Concepts 1 and 2, i.e., a

reciprocating mechanism with folded guide rails.

During deployment the truss develops root strength through Its interface with

the guide rails. Subsequent to deployment, root strength is developed through

the attachements between the trues longerons and the main housing, or between

the truss longerons and a subsidiary structure (Section 1.4.3.3) which is

deployed for that purpose.

The adapter consists of folded plates attached to the truss. It is

automatically erected as the first fold of the truss is deployed. The erected

adapter contains the same type of alignment, berthing and interface devices as

do the rigid adapters previously described for other concepts.

The double folding of the trues causes some problems with the utilities

installation (Figure 1.4-10 ). To install the full complement of utilities,

the electrical cables are mounted on the exterior of the pyramidal members in

a looped configuration. They will extend easily but must be retracted and

r3-loopel manually. The fluid lines are mounted on the lateral members in a

series of folds.

1-68
^y

s

The packaging ratio is 20 to 1 (alone the length).

•ALL MEM62RS Ala ROUN_ TUEES

• DOUELE -FOLD SYSTEM

• NO OFFSET LOADS AT JOWS

Figure 1.4-10. Installation of Utilities

(Concept ?;

1.4.2.4 Concept 4

This concept (Figure 1.4-11) is a triangular to es which deploys/retracts

axially iut changing its end profile. Each bay consists of a rigid

triangL. came, three folding longerons, and three telescopic diagonals.

'ne .r.oe loagerons folded into the stowed configuration are shown in

Figut, 1.4-12. It is recognized that there will have to be a support in the

middle cf the triangular frame to brace the longerons during launch. All of

the true members are round tubes which converge at the corner fittings

without oihet. The folding of the longerons toward the middle instead of in

iine with the batten tubes and diagonals is necessary to achieve a high

packaging ratio.

The main housing is a conventional rigid framework which contains the folded

trues and deployment mechanism, and carries the loads from the truss into the

rest of the deployable platform. The selected deployment mechanism is a

reciprocating device which deploys and rigidizes the truss one bay at a time.

Initially, torsion springs were used but were replaced by a reciprocating

mechanism for better packaging. During truss deployment, root strength is

developed by guide rails which are stowed alongside the housing when not in

use.

Because the truss does not change its triangular end profile dimensions while

deploying, there is a choice of using either two or three guide rails, i.e.

with two-guide rails a rail length equal to 2 bays plus the stack length is

required; with three guide rails a rail length equal to 1 bay plus the stack

length is required.

1-69

f

f 

_....... . r.l r-...' . V.. •... ,1 L. ^ . .,..,b 
ORIGUNAL PAGE !,

OF POOR QUALITY

i

ADAPTER

Figure 1.4-11. Concept 4

A-^^

FOLDED

LONGERONS

Figure 1.4-12. End view of

Folded Configuration

The adapter is a rigid triangular ea3embly which can be identical to that

described for Concept No. 1.

Utilities cannot be mounted in t:ay3 beca•. ee the longerons are folded into the

center. Therefore, they are mounted directly onto the batten frames in a

series of folds ( Figure 1 . 4-13 ) . For the f0 l complement of utilities they

are installed on both the inside and outside of the battens.

With the use of the reciprocating deployment system, a packaging ratio of 20.2

to 1 was achieved.

1-70

i
IBIE IS SHOWN.

)LITER sLoACES

OF ALL THREE SIDES ARE

L'SED TO MOUNT UTILITIES

1^^0

JOINT DETAIL

Figure 1.4-14. concepts 5 and 7

ADAPTER

3

i)

Y 1-71

u 0AIL*NAL

OF POOR Q' IALITY

Figure 1.4-13. Utilities Installation

(concept 4)

1.4.2.5 Concepts 5 and 7

Concepts 5 and 7 are the same except that Concept 5 is a tension X•-braced

structure, whils Concept 7 uses compression diagonals for shear and torsion

capability. Concept 5Till be described (Figure 1.4-14).

Concept 5 is a triangular section truss which deploys along two axes. First,

the triangular section expands to approximately 170% of its original size;

then, the truss deploys along its longitudinal axis. A high ratio of deployed

length/stowed length is achievea by all of the cross-bracing and longerons

nesting inside the batten frames when stowed.

-CONCEPT $CROSS BRACED WITH TENSION STRAPS AS SHOWN

-CONCEPT 7 BRACED WITH COMPRESSION DIAGONAL

• TRIANGULAR BATTEN FRAME

• EXPANDS 70%

•

 CM

• H SECTION MEMBERS

LONGERON SECTION

• LONGERONS 6 DIAGONALS

NEST THE BATTENS

• NO OFFSET LOAD$ AT JOINTS

NOISING

RAILS
BATTEN

FRAME

OCOMPRI

TELES(

e OTHER4

COICEP

ORIGINAL PAGE IS

OF POOR QUALMY

The expanding triangular batten frame is of H-section construction. The

long,irons are a channel crone-section with a hinge and a latch at the mid

length. The crone-braces are rectangular section tension straps, constructed

of a graphite/rubber composite or similar material which has a low CTE and is

flexible enough to benave as a strap instead of a rigid bar , The crone-braces

are rretensioned by deploying the truss bays. The longerona serve as natural

over-center tensioning devices to apply load to the cross-br p cee which are

accurately fabricated to a predetermined 'ength.

The expanding triangular trued .a stowed in an expanding triangular housing.

An the housing expands it pulls the trues with it. The housing which f.s made

in three sections is expended/contracted by a number of powered screw jacks.

The expanding force from the housing to the trues is via the guide rails

attached to the housing and the guide wheels which are attached to the truss

battens. As in Concept 4 the number of guide rails can be either two or

three. The longitudinal deployment mechanism is a "standard" reciprocating

system using the guide rails.

Concept 7 is shown in Figure 1.4-15. Becausa of the double fold nature of the

truss, the compression diagonal has two folding joints in addition to a

telescoping point which is considered to be a significant disadvantage.

STOWED END VIEW DIAGONAL

DEPLOYING THE

FIRST FOLD

Figure 1.4-15. Concel,

1-72
kr»

r^

Y'

The full complement of utilities can be integrated iuto Concepts 5 and 7. The

electrical cables are installed inside the truss triangle in trays which are

pivoted at their ends from clevisee mounted on the battene ( Figure 1.4-16).

The fluid 1'nes are outside the trues triangle on shallow trays which are

pivoted at their and from the battens.

A packaging ratio of 27.5 to 1 was achieved.

1.4.2.6 Concept 6

Concept 6 ( Figure 1 . 4-17) is a lines -ly deployable trues with a square

cross-section consisting of rigid batten frames joined to each other by

folding longerons and telescoping diagonals pin-connected at corner fittings

integral with the frames.

The longerons ( Figure 1.4 - 18) fold toward the middle of the truss at a

45-degree angle, have self- aligning spherical ball end fittings, and

„ hinge/ lock fittings in the center of the longeron length. V i e telescoping

diagonals have a lock mechanism and self - aligning end fittings.

E-braced tension cables can be providad for all the interior batten frames for

redundancy.

All of the trues members (round tubes for the longerona, battens, and

4lagonals, and rectangular straps for the lateral bracing) are loaded along

their centroids and converge without offsets at the batten frame corner

;fittings.

The stowed truss is contained in a square rigid housing to which the guide

rails and deployment mechanism are attached. The number of guide rails may be

2, 3, or 4. For 2 guide rails, the rail length is equal to 2 bays plus the

stack length. For 3 or 4 guide rails, the rail length is equal to 1 bay plus

the stack length.

The folding guide rails and the reciprocating deployment mechanism are the

same as in the concepts described previously. Initially, torsion springs were

considered. The method of deployment is "one bay at a time," with root

stiffness developed by the guide rails during deployment. The longerons

attachment to the rear of the main housing provides the load capability

subsequent to deployment.

The adapter is a square rigid assembly at the far end of the truss. It

contains all of the electromechanical latches, alignment features and

connections required for a payload/module interface.

The full complement of utilities is attached to the batten frames in figure-8

lt•-n ps ( Figure 1.4-19). Both the inside and the outside of the batten frames

can be used ( Figure 1.4-18).

With the use of the reciprocating mechanism, a packaging ratio of 20.4 to 1

was achieved.

1-13
LLLATION

ORIGINAL PAGE 19

OF POOR QUALITY

^I

Figure 1.4-16. Utilities Installation

(Concepts 5 and 7)

a

i

11,`7

^r Y

5

• All members are round tubes MAIN HOUSING

• Inngerons di center hinge joints e

• Telescoping diagonals ^ !J^-

• No offset loads at joints

• Battens braced by tension straps 'I

TRUS

i

ADAPTER

O

0 ^ DEPLOYMENT RAILS

0

90

Figure 1.4-17, Concept 6 

1-74
ORIGINAL PAGE 19

OF POOR QUALITY

l

Figure 1.4-18. End View Stowed

FOR CLARITY, ONLY TWO LINES ARE SHOWN

Figure 1.4-19. Utilities Installation

(Concept 6)
1-76

r

ORIGINAL PAGE 11

OF POOR QUALM

1.4.2.7 Concept 8

Concept 8 ( Figure 1.4-20) consists of square rigid batten frames connected

together by folding longerone and telescoping diagonals. The batten frames

are H-section, cross-braced by thin tension straps ( as required in the

interior bay$ for redundancy). Each longeron consists of two rectangular

section memb6re with locking hinges at their centers. The diagonals are

rectangular sections which telescope within each other. The diagonals and

longerons stow within the confines of the H-section battens. In the deployed

configuration all the members converge at a common point.

The houHing, rails, deployment system and and adapter are the same as in the

square truss of Concept 6.

Concept 8 has the advantage of utility trays which can extend the whole width

of the interior of the batten frames if desired, permitting a large separation

between power and data lines plus the advantage of growth in the number of

lines ( Figure 1.4-21).

A packaging ratio of 22 to 1 was achieved.

• ARO"OtM AKULLY MAN

• H Stao,. Nn7N FFAMB

.O wCDCH NUKOINO M400HAU

• g wC,ON FOLDNO LONO.Fde

• "Mm w.aD n

nNSION Srun

qADAFUR

atA^ of ^oNr

'F

DM LONGINONS

DuooNU

Figure 1.4-20 Concept 8

Figure 1.4-21 Utilities Installation

(Concept 8)

• d
' 44 ORIGINAL PACE 13

OF POOR QUALITY

1.4.3 Orbiter Packaging

This section describes the studies and results applicable to the packaging of

the generic platform ( Section 1 . 2), constructed from the eight candidate

building blocks, into the orbiter. The packaging is based upon the packaging

efficiencies shown in the preceding section.

1.4.3.1 Concepts 1 d 4

Concepts 1 and 4 are axially folding trusses of triangular cross - section.

Because of its better packaging ratio, Concept 4 uses leas spe;e (7•5 "I) is

the Payload Bay than does Concept 1 (9.0 K), Figure 1.4-22. Ap rt f^om this,

the two concepts are identical in packaging and deployment. The figUris and

description apply only to the generic platform, although daay otbar

configurations can be built/deployed by .following the approaoh outl. nud Si

this report.

AYLOAD l__ OM5

BAY KIT

9M, CONCEPT 1

7.5M CONCEPT 4

DOCKING MODULE

Figure 1.4-22. Orbiter Packaging

—Concepts 1 and 4

The platform system deployment is 100% automatic without resort to EVA, RMS

assembly or to building fixtures. All of the necessary modules, orbiter

supports and cradles are incorporated in the design.

There are eight triangular building blocks and two cradles. The cradles are

structures which interface with the orbiter trunnion and keel fittings, join

some of the building blocks together at their hinge points, and provide

mounting surfaces for the spacecraft systems.

Each of the building blocks 1s a single truss design, i.e., only one truss is

deployed from each housing.

The platform is assembled and checked out on the ground before it is folded

and placed in the orbiter. All of the utility connections from end to and of

the platform are made on the ground and are not broken during stowage or

deployment.

1-77 

.... dw..:.....n Yw...aa `•. .. G.:J'..a.. c. w ^..... ...axs^1{
• REORIENT THE

REMAINDER OF THE

BUILDING BLOCKS

• EXTEND THE TRUSSES

TO COMPLETE THE

CONFIGURATION

REORIENT THE

BUILDING BLOCKS AS

SHOWN AND EXTEND

THE TRUSSES

l

I

C

I

Nit

i

ORIGINAL PAGC is

OF POOR QUALMY

The stored platform in removed from the orbiter by using the RMS, and remains

in the grasp of the RMS and/or in attached to the NAPA during the initial

stages of deployment. If continuous operation and a low deployment rate of

3/cm per second are assumed, the platform could be deployed in about 72

minutes. Some trusses deploy simultaneously. There are three basic

operations involved in deploying the platforms, i.e., reorienting the building

blocks, extending the trusses, and latching.

One sequence of operations to deploy the platform is shown in Figure 1.4-23.

Several other sequences aro possible in achieving the same configuration. The

solar array is not shown as an integrated deployable item in this design,

although it is quite possible that it could be incorporated. Items such as

large payloads, modules, etc., are added to the deployed platform by the RMS,

using the interfaces provided.

ORBITER_`

BAY

ODD

^+^^,,y7^^ 1\\ l

l

I/ • USE RMS TO

REMOVE STOWED

PLATFORM FROM

ORBITER

Figure 1.4-23. Platform Deployment —Concepts 1 and 4

1.4.3.2 Concepts 2, 6 and 8

This section discusses the packaging and deployment sequence for axially

deploying building blocks, of square cross-section of which Concepts 2, 6, and

8 are typical examples. The three concepts are identical in packaging and

deployment with the exception of the alight difference of their lengths in the

payload bay (Figure 1.4-24.).

1-78

•

.!r._ ._Y. _.... ...^ .. r..
ORIGINAL PAC: —, i.,

OF POOR QUALrN

13.562 M

PAYLOAD OMs

BAY KIT

8.2 M, CONCEPT 2 •DOCKBVG MODUIf MoOUIf 7

MODULE 8.8 M, CONCEPT b

8.2 M, CONCEPT 8

Fi gure 1.4-24. Orbiter Packaging

—Concepts 2, 6, and 8

There are five building blocks and two modules in the assembly

(Figure/1.4-25). Of the five building blocks two are "singles", i.e., one

trues from a main housing, and three are "doubles", i.e., two trusses, one

from each and of a main housing.

o Module 1 mounts three of the building blocks and is a cradle which

supports them in the orbiter. It forme part of the deployed

platform and is used to mount spacecraft systems and equipment.

o Module 2 mounts the remaining two building blocks and in the cradle

which supports them in the orbiter. It too is deployed with the

rest of the platform and is used to mount spacecraft systems.

The complete package is removed as a unit from the payload bay using the RMS

and is retained on the RMS (or mounted onto the NAPA) for the initial stages

of deployment. The stages of the deployment sequence are:

• Reorient four of the building blocks by rotating them until all five

building blocks are in the same plane.

• Extend the trusses as shown to form the vertical of the T shape of

the platform.

• Rotate the building blocks which form the diagonals and extend tho

trusses to complete the T shape of the platform.

The platform f.s automatically deloyable without the assistance of EVA,

building/assembly fixtures or piece-by-piece assembly using the RMS. There is

no requirement to make electrical or fluid connections in flight except where

separate unite such as payloads are added subsequent to deployment by the RMS.

1.4.3.3 Concept 3

Although this concept was successfully packaged (Figure 1.4-26) and deployed,

this study does point up the difficulties with the use of a double-fold

structure. The advantages gained from the dense volumetric packaging of the

structure are largely illusory when one turns from a simple truss to a

complete deployable platform system. Some of the problems which contribute

are:

1-79

k4
uRIG11-AAL PAGE IS

OF POOR QUAL"

• ORBITER

PA (LOAD

BAY

REMOVE

WS ST

RE STOWED

PLATFORM FROM

THE ORBITER

• DIAGONAL TRUSSES

'r ROTATE ABOUT

THE HINGE POKE`S

ON THE HOUSINGS

*EXTEND

THESE

..••//// TRUSSES

} •REORENT THE

1 BUILDING BLOCXS

• EXTEND THE TRUSSES

SHOWN

Figure 1.4-25. Platform Deployment —Concepts 2, 6, and 8

562 M

J IQAYLQA - OMS

BAY ,-.- KIT

• DOCKING ^"— 9.1 M

MODULE

Figure 1.4-26. Orbiter Packaging

—Concept 3

1-80

e
o>norn)AL r-act f:^ 1 OF PU()R QUF,LrrY

• The stowed double-fold truss is difficult to support in the

orbiter with provision of stiffness compatible with frequency

separation (10 Hz) from the orbiter.

• The methods of joining truss to truss, or truss t.o module,

are more complex. it is sometimes necessary to erect a subsidiary structure for structural continuity.

• The end adapter cannot be a simple plate to support latcheq,

interfaces, etc.

• Mounting of the adopted complement of utilities is difficult

because of limited apace.

In spite of these difficulties, t^Ie generic platform can be built using

Concept 3. It is an automatic deployable platform which requires no RVA, no

building/assembly fixture and no part-by-part erection. The electric/fluid

utilities are installed and checked out on the ground from end to end of the

platform. There are no connections broken/reconnected between stowage inL-T

the orbiter and final deployment.

The platform consists of five building blocks and one module. Of the five

building blocks, three are "double building blocks," i.e., a truss deploys out

of each end of the housing. The module is an integral part of the platform,

is deployed along with the building blocks, is intended to mount spacecraft

systems, join three of the truss together, and serves as a base for deploying

subsidiary structure for reacting truss loads.

The packaged platform (Figure 1.4-27) is removed ftcru the orbiter by means of

the RMS. While it is in the grasp of the RMS (or HAPA) the initial stages of

platform deployment are made. There are several stages in the deployment

sequence.

V/ORBITER PAYLOAD RAY

/ 1 . USE RMS To

REMOVE STOWED

PLATFORM FROM

ORBITER

• REORIENT THE

WILDING BLOCKS

-EXTEND

THESE

/// TRUSSES

-NOTATE THE

OUTER TWO

BUILDING BLOCKS

•UNFOLD TRUSSES &

}^ SUBSIOWIY STRURUR! 

4 •EXTEND TRUSSES IN

J' CIRECTIONS SHOWN

Figure 1.4-27. Platform DeployMent —Concept 3

1-81

Aff..•:...... .. 4 a ., .. x..,,.. .....:

::a^..-v ^r ...... .. m+
1

}.

ORIGINAL PAGE 19

OF POOR QUALITY

• Unfold the building blocks.

• Deploy the subsidiary structu.-:- on the module. This includes e

docking port and attach points for reacting truce loada.

• Erect the first fold of the trusses.

• Exterd the trusses part way.

• Make the points at the module attach points.

• Extend the trusses completely.

This completes the deployment of the platform. Payload&/solar array/RCS pods

may be added, as required, by a'second launch or by using the empty portions

of the payload bay of the first launch.

1.4.3.4 Concepts 5 and 7

Concepts 5 and 7 are stowed in the orbiter ( Figure 1.4-28) and deploy to the

final configuration in identical fashion. The stowage and deployment shown

are for the generic platform only. It is a completely automatic deployable

system which requires r EVA and no assembly or erection by RMS. The

deployable platform inc den six buiding blocks and two modules.

13.562M _^ 140DULE 2

PAY LOAD I OMS

BAY KR

MODULE I

DOCKING 6.73M —^-1

MODULE 

Figure 1.4-28. Orbiter Packaging

—Concepts 5 and 7

o Module 1 mounts three of the building blocks and is a cradle which

supports the whole of the stowed platform in the orbiter. It forms

part of the depioyad platform and is used to contain spacecraft

systems.

o Module 2 mounts the remaining three building blocks. It does not

interfrce directly with the orbiter, but is attached to Module 1.

It too forme part of the deployed platform and is used to mount

spacecraft systems.

The complete package is removed from the orbiter in its stowed form by using

the RMS. While it is in the grasp of RMS (or RAPA) automatic deployment is

initiated. There is no requirem• ^ t for a building/asaemhly fixture to hold

the platform while the RMS atsa : ne, trusses or modules section by section. As

in all of the Rockwell platform de p loyment concepts, the utilities are

installed and checked out on the gound and not disturbed thereafter.

1-82
Y

i1 ORIGINAL PAGE 13

OF POOR QUAfflY

Because of the high packaging •fflciancy, it ',s poaa s to stow some of the

building blocks with their lengths normal to the major axis of the payload

bay, i.e., stow them "across" the ship instead of lengthwise. This permits

arrangements such that the building-blocks do not have to be reoriented to

achieve the deployed configuration.

There are only two stages in the deployment se q uence (Figure 1.4-29),i.e.,

expand the triangle shapes of the building blocks and extend the trusses.

Of all the concepts which were studied, Concepts 5 and 7 proved to be the moat

suitable for packaging and deploying the Generic Platform. They atowed in the

shortest length of the payload bay and required the least reorientation of the

building blocks.

ORBITER

PAYLOAD\

BAY \^

i

• EXTEND THE TRUSSES

• NO REORIENTATION OF

BUILDING BLOCKS IS

REQUIRED

• USE TO REMOVE

STOWED PLATFORM

FROM ORBITER

'AND THE

TRIANGLES OF

ALL THE BUILDING

BLOCKS

• WAGONAL TRUSSES

ROTATE ABOUT THEIR

HINGE POINTS A:

THE TRUSSES EXTEND

Figure 1.4-29. Platform Deployment —Concepts 5 and 7

1-83 

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VRILANAL PAGE IS

Of POOR QUALffY

1.4.4 Thermal Analysis

A therm..l analysis provided a prediction of the thermal gradients across the

candidate deployable trusses (Table 1.4-2). The temperatures shown are the

average over the length of each member. These gradients were determined for

tho structural arrangement/geometry shown in Figure 1.4-30 and for GEO

applications. LEO applications have lower gradients due to earth and albedo

Lr.a . ing .

Table 1.4-2. Thermal Gradients between Longerons

LONGERON

d DIAGONAL TEMPERATURE (°C)

AT

FIGURE OF

MERIT

CONCEPT I (m) PRESENT SUN SIDE SHADE SIDE °C) d/ AT

1 1.6 YES 21.6 -13.5 35.1 0.046

2 1.3 YES 20.7 -11.6 32.3 0.040

NO 24.0 4.8 19.2 0.067

3 3.0 NO 24.0 2.7 21.3 0.140

4 1.7 YES 21.7 -7.4 29.1 0.058

5 2.5 YES 23.8 -36.0 59.8 0,041

(TS)*

6 1.3 YES 22.5 -22.6 45.1 0.029

7 2.5 YES 23.6 -56.7 80.3 0.031

8 1.3 YES 22.5 -17.5 40.0 0.031

*(TS) TENSION STRAP

The thermal gradients are due to shadowing when one or more structural members

pass between another member and the sun during orbit.

For analysis, each configuration can be represented as a Z-section (Figure

1.4-31) or as two parallel members if a diagonal is not present in the plane.

A thermal math model of the Z-section is used for the analysis. It consists

of 24 nodes and provides for conduction betxaen nodes. The conor,ruction

material is graphite composite. The surface radiation properties were assumed

to be nearly black. Emmittance and absorptivity values of 0.85 were used.

The thermal model for the two parallel member ca ges is a single node heat

balance.

Two importaa: assumptions in the analysis are the estimate of the shadow time

and the view factor for solar impingement during shadow. The shadow time is

(minutes) is estimated by the following expression. For two parallel members,

is •. 8(0.2566 + sin -1 D/S)

where D - member diameter

S - distance between members

1-84
ORIGINAL PACE ray

OF POOR QUAL17Y

Concept Arrangement Concept Arrangement

O O

5

DOi O

5

!J30

`-'

O

4

^— --^ 1 .68

O O

a

1.26

4-30. Structure Arrangemen t for Thermal Gradient Analysis

(Dimensions in Meters)

SUN-EXPOSED

LONGERON

SHADED

LONGERON

4AL

r

Figure 1.4-31. Thermal Analysis

Configuration

1-85
This expression accounts for the size of the shadowing member, the distance

between members, and the included angle of the sun. For the section, a

modified form of this expression was used. The distance between members was

allowed to very with nodal position.

The view fs.ctor for solar impingement of the shadowed member was approximated

by the following cosine function,

F - cos 2 (180 t/ts)

This function has a value of unity at the extremes and a value of zero at the

csnter of the time interval.

Table 1.4-2 also illustrates the parameters to evaluate the re'itive merit of

each concept for thermal stability. The end rotation of a beam 3 is

determined by

P a(AT)R

d

For beams cf equal length and the same material, then P is proportional to

(AT/d) or the figure of merit is proportional to (VAT) with high values

being best. The foregoing data were based upon the structural design sizes as

determined by the strength/stiffness characteristics, packaging, minimum

weight, and joint design considerations. The design was not adjusted for

reduction of the thermal gradients. For example, for Concept 6 the thermal

gradient can be reduced by reduction of the diagonal diameter (reduced shadow)

with a small weight impact (of no consequence to LEO platforms). Also, since

the data are based upon a/E - 1.0 with E and a . 0.85, if reduction of

the thermal gradients is required, initial values of a - 0.10 can be

realized by wrapping the graphite composite structure with silver teflon

tape. Though a will increase over the life of the structure, particularly

in the CEO environment, the end of life value will be less than 0.85.

A perspective on these thermal gradients is appropriate. For a 40-meter

cantilever using concept 6, the end will have s thermal induced rotation of

0.028' (a - 0.36 x 10- m/mC) due to the specified gradient of 45'C across

a design with a 1.26 m depth. (The 40 m length is used since precision

antennas would be mounted closest to the center of the configuration).

Section 1.1-10 indicates the desired pointing accuracies should be 0.05 to

0.10 degree (based on statistical analyses). If necessary, a value of 0.28°

can be reduced by many of the techniques delineated above. In addition,

placement of the square cross-section as shown in Figure 1.4-32 can reduce the

0.028' rotation to approximately 0.020'. If necessary, larger trusses

(Figure 6) can further reduce the end rotation. The use of composites with

negative coefficients of expansion and metal fittings can also reduce these

rotations.

In the light of the foregoing, it is also appropriate to note that a 22.2 N

shear on the end of a 40 m cantiliver will induce an end rotation of 0.005'

for the adopted flexural stiffness, or for 4.45 N induce an end rotation of

0.001'. Hence, RCS induced distortions are small by comparison.

T

.w

1-86

s
I

R

ORIGINAL PAGE 19

OF POOR QUALITY

AV

Oft

e

(NO DIAGONAL

SHADOW FOR

SUN AT A)

Figure 1.9-3'. Approach to Redure

Thermal- Induced Truss Rotation

1.4.5 Mean Properties Analysis

The results of the deployable platform systems mass properties analysis of the

generic platform contructed from Concepts 1 through 7 are shown in Table

1.4-3. These masses are determined by a combination of detailed calculations

where specific structure sizing was avail a ble (deployable trusses, for

example), and standard preliminary design mass estimation techniques

(housings, for example). The generic platform mass is based upon the adopted

strength, stiffness, and complement of utilities delineated in Section 1.1.

Table 1.4-4 illustrates the mayor differences in the masses for Concepts 1

through 8. The data for Concept 8 are estimated from the data provided for

Concept 6 in Table 1.4-3. The data shown are used in the orbit transfer cost

data described in Section 4. The data in Table 1.4-4 were also used as a

baeepoint for extrapolation of the weight differences for a generic platform

designed to 1/10 the adopted strength/stiffness requirement (Table 4.3-5)•

A review of Table 1.4-3 illustrates all of the candidate deployable platform

systems, including a 20% growth allowance, are launchable to a 210 nmi orbit

28.5 0 inclination (Figure 1.4 . 33) without use of an OMS kit. The balance of

mass between the values shown and an allowable 20,000 kg is available for

payloads, RCS propulsion modules, etc. An addditional 2500 kg is launchable

if an OMS kit is provided. The 210 Nmi orbit is ooneldered suitable, since

for the generic platform, decay to 150 nautical miles would not occur until 60

days after insertion into orbit. This is considered ample time for

deployment, installation of payloads, checkout, etc.

1.4.6 Cost Analysis

Consideration was devoted, in support of the Concepts Selection Trade, to the

establishment of the relative coats associated with the design, development,

testing, special technology needs, fabrication, shuttle launch, and orbit

transfer of GEO platforms. The design, development, and testing coats were

not included since they are dependent on the number, types of application and

1-87
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1-88

t'

Ai
ORIGINAL PAGE 19

OF POOR QUALITY

• DATA SHOWN FOR 28.5 DEG INCLINATION ONLY

74 INTEGRAL

0 SHADED AREAS REPRESENT PPIMARY REGIONS AFFECTED

BY PROGRAM VARIABLES. PAYLOADS REQUIRING PERFORMANCE OIAS IN THESE REGIONS SHOULD VERIFY CAPABILITY WITH

^i JSC AS EARLY AS POSSIBLE 0

2 FIRST OMS

r KIT ADDED

N

SECOND OMS

o

/ KIT ADDED

a 1

THIRD OMS KIT

ADDED

400 500 600 700 800

ULAR ORBIT ALTITUDE, N MI

DESIGN

iP ao• ASCENT

Y CARGO

0

0 20

DESIGN

.n -- LANDED

1 CARGO

0

0

a

u

0 10 20 30 40 50 60

CARGO L.G. POSITIONI(FT)

LANDING CAPABILITY - ^--—

22,500 KG

Figure 1.4-33. Orbiter Payload Ma_ —Launch /Landing Constraints

1-89
ORIGMAL PAGE 13

Of POOR QUALRY

Table 1.4-4. Comparative Mass (kg x 10 -3 ) -Generic Platform

ITEM 1 2 3 4 5 1 6 1 7 8

BASIC TRUSS ELEMENTS 1.9 1 3.2 2.0 1.8 1.6 1.8 1 2.2 1.8

ALL JOINTS 1.3 0,9 1.0 1,6 0.6 1.9 1,3 1.9

UTILITIES INSTALLATION

SYSTEM

6.8 0.2 0,1 0,3 0,1 0,2 0.7 0.1

HOUSING AND ADAFTERS,

HOUSING-TO-MOUSING ATTACH, 0.9 0.9 0,6 0,8 0,8 0.9 018 0,6

DEPLOYABLE STRUCTURES FOR

DOCKING PORTS

0.1 0.1 0.1

TOTAL 4.9 5,2 3.8 4,5 3.8 1 4.8 5.1 5,0

o MASS 1.1 1 1.4 1 0.0 1 0,7 1 0.0 1 1.0 1.3 1.2

ALL 6ESIGRS. IRCWOIRG OEROTIKHT IHCIWIISHS. UTILITIES LIKES, DOCKING PORTS, CONTROL HDORI

EQUIRHEHT, ARO CRADLE WIGHT (WITH 701 GROWTH(-LESS TI 16,600 KG.

requirements of future platforms which are not defined. The special

technology development coats (as subsequently shown in Section 3) are

primarily the same eeross all the candidate concepts, with the few exceptions

resulting in negligible coat impacts (particularly if spread across

several platforms).

The analysis, therefore, determined the recurring fabrication cost, Shuttle

launch and for GEO platforms, the cost of orbit transfer. All the coat

analysis data are expressed in terms of FY 1981 dollars and are relative,

i.e., applicable only to items which are different. The coats of items that

are the same across all concepts are not included. For example, the coat of

the basic power, data and fluid utilities themselves, docking ports, etc., is

not included.

1.4.6.1 Fabrication Coat Data

The component fabrication costa to construct the generic platform are shown in

Table 1.4-5. These data are incorporated directly into Table 4-4-1. These

data were derived by a combination of material and fixture costa (as

appropriate) and manufacturing hours to which a composite coat rate was

applied. The composite cost rate includes direct labor, overhead, and general

and administrative (G&A).

The materials, fixture, and manufacturing hours data presented were determined

from estimates by advanced manufacturing personnel using drawing details.

Where detailed information was not available, in-house Rockwell parametric

cost model techniques were used.

1-90

j
0R cmn rro' 11

Op ; I,Ljh QUALfTI(

Table 1.4-5. Fabrication Cost Data ($ Million)

CONCEPT

COST ITEM 7

9ASIC TAUS$

MCA$ENf (COMPOSITE) 2.2 5.0 1.0 1.7 2.0 1.9 713

ANDS JOINT$ y

AMC TI TTINGS 1.7 1.5 2.1 1.4 3•0 1,9

TRUSS ASSEM9LY AND

CMECKruT 0.4 0.6 0,0 0.6 0.7 0.6 0.6

UTILITIES SUPPORT

SYSTEM VARRICAIION 0.4 0,2 O.1 0,1 0.9 0,1 0.9

AND INSTALLATION)

INSTALLATION OF

UTILITIES ANO 0.1 0.1 0.1 0.1 0.1 0.1 0,1

CHECKOUT

END ADAPTERS AMC 0.4 0.0 0.0 0.0 0.4 0.0 O,S MAIM HOUSINGS

NIILDING 9LOCA TO -

9UILDIHG $LOCK 7.0 2.9 0.5 2.2 1.7 2.9 1.7

ATTACH MECHANISMS

TOTAL (SM) 9.0 10.9 6.9 7.0 6.5 6.6 7.6

0 IfMI 1.5 4,0 0.0 0.0 O 2,0 1.7

NOTE: CONCEPT 9 ESTIMATED FROM Ef ISTING DATA FOR CONCEPT 6

It is pertinent to note that Table 1.4-5 does not include the costa of

deployment mechanisms since the building blocks can utilize either a

reciprocating device (GD design or equivalent) or a stored strain energy

system. The differences between the same type mechanism across the single

folded designs were regarded as negligible. In retrospect, the additional

coat of the mechanisms for the double folding of Concepts 3, 5, and 7 should

be included into the last cost item in the table. To do so would only enhance

the conclusions and deciaions made in Section 4. and require extensive

numerical change throughout the tables for the sake of consistency. Hence,

this item is not included.

1.4.6.2 Shuttle Launch Cost

The Shuttle launch coat data for use in Table 4.4-1 are based upon a cost of

$2.6M/meter of Shuttle cargo bay length. It is derived from an FY 1982 total

cost of $48M for 18.3 m of bay length. The $48M is the FY 1981 cost

extrapolated from the baseline FY 1975 launch coat of $32M and is for a medium

traffic model (40 launches per year). A significant reservation on the use of

this total value is that for a dedicated mission the system coat is incurred

regardless of the usage of the bay. This reservation is reflected in the

allocation of points in the totaling of the major criteria.

1-91

do
r

1.4.6.3 GEO Orbit Transfer Cost

The Rockwell space station studies have identified a cost (in FY 1981 dollars)

of $8,800 per kg of mass delivered to GEO orbit including the lEunch to LEO by

the shuttle. This coat is also based on a medium mission model.

1.4.7 Miscellaneous Design Issues

During the concept development discussed throughout this section , three

design concerns surfaced as follows:

• Extension of a truss which has a large payload while maintaining the

deployable truss root strength with the guide rails

• Implicatiun of building block to building - block structural

attachments that are not fully fixed about all three axes

• Potential significance of "Joint alop"

The first issue has been resolved as discussed in Section 1.3.3.5. The second

issue has been investigated and resolved as follows:

A configuration such as the generic or ASASP platform can have Joints without

full structural continuity at places, such as Joints 1 and 2 of Figure 1.4-34,

with acceptable reductiona in effective stiffness. 

A sensitivity analysis of platform overall stiffness variations associated 

with different end Joint design characteristics was performed. The analysis

was performed with a RASTRAR model that simulates the ASASP platform

configuration and mass distribution. The construction platform was not

included. The ASASP was used since the mass distribution was more readily

available and the generic platform was fashioned after the ASASP. The adopted

stiffness values were used.

Since the overall variation in platform stiffness is most easily defined by

the resulting modal frequencies, modal analyses are performed for the expected

diagonal member end Joint variations. At all other Joints, flexural and

torsional moment continuity is maintained.

The table on Figure 1.4-34 illustrates the first four modal frequency

variations relative to that of the baseline design. For all of the four eases

shown, the stiffness reductions were acceptable.

For other possible configurations:

o Moment and torsion capability of cantilevered members can be

rovided by fixed Joints or appropriate self-locking latches

p{ Figure 6)

o In worst case, EVA attachment of strut or pair of struts is

feasible, if necessary.

1-92 

# 

6
I

_H

i

UHIGINAI. Fr• .,L 13

OF POOR QUALITYY

MODEL,

• JOINT 1 i 2 LENGTHS

ARE EACH 2.5

*EI & GJ ARE

ADOPTED VALUES

CASE

FREQUENCY IHa) COMMENTS

FIRST SECOND THIRD FOURTH FIR JOINTS 1 6 2

EASE

LINE 0.0471 0.0200 0.007 0.1181 FULLY FIXED

1 0.0061 1 0.0656 0.06% 0.1127 ZdXIS PINS

2 0.0651 0.060 0.1120 Z-AXIS PINS AND

1/10 OF EI, UJ, AE

3- 0.0061 0.0650 0.0686 0.1152 Z-AXIS PINS WITH

OFFSET FROM DIAGONAL CENTERLINE

0.0131 0.0628 0.030 0,100 BALL JOINTS

`OFFSET - 1.3 a

Figure 1.4-34. ASASP Modal Frequency Variation

with Diagonal End Joint Design

The third design concern has not been resolved in this study and is applicable

to the concepts studied which all utilize clevis ,joints, folding and /or

telescoping points.

Consider the following illustrative example (Refer to Figure 1.4-35):

o Consider a cantiliver member 60 m 'long (38 bays)

o Suppose ACS thruster shear - 4.5 N

o Limit longeron load - 2.8 N (one bay from RCS thruster)

o Limit longeron load - 107 N (38 bays from RCS thruster)

D For the adopted EI - 2 x 108 Nm2 , and trues depth of 1.26 m the

longeron AE - 1.26 x 108 N

o Elongation of longeron for load of 2.8 N - 0.0000035 cm

o Elongation of longeron for load of 107 N - 0.00013 cm

o For a design to twice the depth (same EI), the above values of

0.0000035 and .00013 cm are multiplied by 4, i.e., elongation is

0.000014 and 0.00052 cm

1-93 
GjUGjNAL PAGE !S

of pooR QUALrN

V .

f 1.126 m

Bay

1.26m—^ length

• 1.58 m

Figure 1.4 - 35. Model for " Implication of

Joint Slop" Analysis

• For both depths, the elongations are very small compared to

conventional olevie / pin clearances

• problem is increasingly more severe with reduced RCS thrust

Hence, for future resolution: ( 1) To what extent does friction preclude

"joint slop?" and (2) What is statistical implication of point slop in actual

design with numerous points?

The potential implications of "point slop" to the control system design are:

o It will be of increasing concern in figure control applications where" "3oint slop alone results in structural deflections

approaching total allowable deflection

o "Joint slop" is highly non-linear phenomenon with strong potential

for destabilization and limit cycling

c It will be of increasing concern if " equivalent frequency" (fe)

(Figure 1.4-36) approaches control system bandwidth

The resulting potential control system design implications are:

o Develop accurate " Joint slop" model to facilitate controller design

o Typical controller enhancement - nonl_near gain scheduling,

compensation and limiters to enhance performance

1-94
1-95

is UHIGINAL PAGE 19

OF POOR QUALITY

W

Vl

2

O

CL

W

W

GC

T FE

Figure 1.4-36. Equivalent Structure

Frequency due to "Joint Slop"

The primary problem is that joint slop introduces a series of distributed nonlinearities in the spacecraft dynamics. In general, this will degrade control

system performance. It is destabilizing and a potential cause of limit cycling.

Special control enhancements can be used to minimize the detrimental effects of

joint slop, as indicated above. In general, the problem is tractable if the

magnitude of the joint slop can be eliminated or made small relative to the

control system pointing accuracy requirement. Also, the control problems are

eased with increasing friction levels at the joints.

The preferred solution is to eliminate or negate joint slop by mechanical means

(eccentric clevis pin or bolt with an expanding sleeve) or with thermosetting

materials at the joints (one bay at a time deployment enhances this possibility).

Another alternative is to increase joint friction providing it is not detrimental

to deployment.
1

2. MATERIAL DATA RASE

The material requirements for the manufacture of structural members for large

deployable platforms include those needed to produce tension/compression

members, tension members only, fittings, hinges, bearings and springs. The

materials selected for t 'ne data base to satisfy these requirements are those

currently available in present technology and those which could be available

in the technology for point design by 1986. This philosophy permits the

consideration of not only metals and organic composites, but metal matrix

composite materials.

This material data base in not intended to peruit the detailed optimization

needed for the ultimate point designs to be made by 1986. It is intended to

supply a sufficiency of general information to allow the material trades

required for the generation of preliminary designs. These trades will specify

the direction of the ultimate designs for deployable space structures.

All materials selected are suitable for the temperature fluctuations

(-200 to 80C) that will be encountered in service in either a low earth or

r geosynchronous orbit. These temperature excursions can be reduced by the

judicious appllca . 4 on of thermal control coatings and insulations.

Temperature rssiat.;;:-e of the materials is, therefore, not a critical.

consideration. The radiation resistance of the material is important only in

the casef geosynchronoua orbit. Using a radiation criterion of a does rate

_ of 6 x 10^ rade ( Si) per day an the maxim4m anticipated In geodynchronbue

orbit, the materials have life expectancies in exceec or 5 to 500 years

(Reference 13). Since _ sly a few of the materials have low damage

thresholds and these are ueieally restricted to low radiation exposures in the

interior of spacecraft, a minimum of thirty-year life is expected for any

platform system designed using the materials in the data base.

The data base is contained in Tables 2.2-0 through 2.0-11. Table 2.0-1

provides a glosaary of e•:^rma. Tables 2.0-2 through 2.0-11 present candidate

materials and their pertinent material properties.

Tension and Compression Members

Some typical metals such as aluminum 2219 and beryllium (Table 2.0-2) are

included in the data base for comparison with the more advanced materials such

as the organic and metal matrix composite materials ( Table 2.0-3).

Considering the design parameters of low thermal distortion, low coat, and

high modulus, the most attractive candidate materials within the state of the

art are the epoxy/graphite composites. The metals have a high coefficient of

thermal expaneio- which contributes to excessive thermal distortion of the

structure. The polyimide/graphite in state of the art, but it is used only

for high temperature applications because it is excessively expensive for

general use. The metal matrix composites have all the right properties, but

are at present very expensive and the technology is in the experimental stage

and may not be available for production Ly 1986. Therefore, they may have to

be relegated to specialty applications where no other material can reasonably

suffice. Some new materials that will very probably be available for general

application to point designs in 1986 are the very high modulus graphite

fibers. These fibers permit the production of structures with a lower

2-1

Ar
Q URIGJNAL PAGE 19

OF POOR QUALrN

Table 2.0-1. Glossary of Terms for Data Base

A ACC'EPT'ABLE r K MY

a ACCEPTABLE WITH SPECIFIC COUPONS L I4003 FACTOR

9, BEARING STRENGTH N/A NOT APPLICABLE

C ACCEPTABILITY MUST BE DEMONSTRA- Na NOT RATED

TED BY TEST OR ANALYSIS a RIYUCTANCE

D DUCTILITY SERFAO RESISTIVITY

da, DI"UBION Of SELIUM

3

S BET PO

I

N':

D I• DIELECTRIC CONSTANT SP SPECIFIC HEAT

Di, DISSIPATION FACTOR TML TOTAL MALI$ LOSS

E MODULUS D UNAVAILABLE

W aLON(ULTION VCR VOLATILE OONOEBSIDIJ MATERIAL

F

STRESS I/7 VOLUME RESISTIVITY

I FATIGUE PROPIRTTES I UNACCEPTABLE

O $BELL MODULUS Q ABSORBTIVtTT

B SANDINESS E KMIESIVITY

K COEFFICIENT N POISSON ' W RATto

k THERMAL COIIDUCTANCE P DENSITY

C

[MUM=

COMPRESSION C

I

TEMPERATURF

E MODULUS a THERMAL ETFARS:OR

t TINSI E f FRICTION

u ULTIMATE L LONGITUDINAL

• SNIAN T TRANSVERSE

DIMENSIONS

MPa _ MEGA PASCALS

®a - GIOA PASCALS

J/kKK - JOULES /91 LOGRAY - DEGREE KELVIN

N/K _ WATTS/METER-DEGaII KELVIN

M/K - WITR/MRTKR-DEGREE KELVIN

kkB^/m3 _ KILOGRAM/CUBIC METER

sE /sec-ATM - $QUARK MITERS SECOND-ATMOSPHERE

coefficient of thermal expansion, a higher thermal conductivity and a high

specific rigidity. The two more prominent contenders in the field of high

modulus fiber composites are the P753/934 and P1003/934 graphite/epoxy

composites listed in Table 2.0-3,

Fittings

In addition to the conventional materials used in the manufacture of fittings

for spacecraft, such as the metals in Table 2.0-4, more unconventional

materials (Table 2.0-5) are being considered to obtain both a closer thermal

expansion match to the basic structure and a closer match to the modulus.

Another major advantage for the use of composite fittings is that it permits

the molding of fittings dire.:tly into structural members. The type of organic

materials moat anenable to this type of processing are the thermoplastic

composites such as the polyoulfone/graphites. There is some development work,

however, that needs to be accomplished between now and PY 1906 to reduce the

direct molding to structural member process to common everyday practice.

2-2

i 
r

r

f

ORIGnve t. Fr- f ,

OF PO( t QUA,"

Table 1.0-2. Metallic Materials For Structural Members

MATERIALS ® ALUMINUM

BERYLLIUM 2219

PROPERTIES m m

MECHANICAL

Ftu

(MPa) 427.8 276.0

FC C(MPa) 331.0 207.0

Fsu

(MPa) 248.4 172.0

E t (GPa) 72.5 289.0

EC (GPa) 74.5

i

289.0

G L

!GPa) 27.6 138.0

N 0.33 0.10

9 (MPa 

) e/1 . 1.5 655.0 N/A

e e/D . 2.0 835.0

PHYG I k ^!.L

Sp (J/kg K) 848.0 1825.0

k (W/mK) 12.1 8.4

Ke (m/mKx iu' 6 ) 22.0 12.3

e 0.05 0.10

a 0.10 0 55

R 0.90 0.45

p (kg/m3) 2830.0 1850.0

APPLICATIONS TENSION & TENSION &

COMPRESSION COMPRESSION

BEARINGS

CD ALL PROPERTIES MCASURED UNDER AMBIENT CONDITIONS

Q ALL MATERIALS LISTED MEET SP-R-0022 FOR VCM AND TML

n,

{

2-3

.. . ,. ... .. - .. m..r ... ^. A^ "w Y ... (W l .^ uda. .^Iilatiawtuu iu..r--^•iMry^4.r.`.uu.uNYa
ORIGONAL 7f - 7 1'

OF POOR Q'JC,--' Y

Table 2.0-3. Structural Member Candidate Composite Materials

- a

^

r

F

r J^

d

4II e

NA TKRIAI TWO ORA►NIT[ EPPRY POLY I MI DE/ ALUMINUM/ MAGNESIUM/

GRAPHITE GRAPHITE OPAPNITE

PROPERTIES 7700/554 ► 752/970 PIW3/554 m m (D

MECHANICAL

F tu (NA) 1481.0 999.0 1140.0 865.0 493.0 67S d

Ft0

(MPA) 4d.0 52.0 17.0 20.7 48.0 55.0

/L !* o) 1276.0 728.0 711.0 578.0 642.0 621.0

F" (me) 186.0 119.0 U 89.0 104.0 104.0

FIN

(MPS) 69.0 4410 48.0 41.0 48.0 48.0

FTN(MPS) 77.0 21.' 27.0 20.0 55.0 U

E L, (ap e) 145.0 768.0 418.0 504.0 201.0 724.0

El

(GPR) IC,0 6.0 U 6.2 35.0 21.0

[^ (OPA) 173.0 146,0 428.0 I5410 U U

M (E) 5.0 4.9 U 4.85 24.1 17.3

PLAYS I CAL

S P (S/kOR) 880.0 U U 875.0 96x.9 962.9

k (N/MK) N/A N/A N/A I N/A 257,0 1030.0

kL (M/MR) 9.0 157.0 520.0 77.87 N/A N/A

kT (N/MR) 0.7 1.8 U 1,21 N/A M/A

R.I. I M/MRN 10- 6 ) 0.27 -I.Oh -1.6 -0.99 1.77 0.55

,N7 (M/MRR 10- 6 ) 25.2 25.6 26.9 28.8 26.E 28.2

c GAS 0.85 0.85 0.85 0.45 0.45

0 0.80 0.80 0.80 0.80 0.45 0.45

A 0.20 0.20 0.I0 0.20 0,55 0.55

P kg, 1600.0 1/65.0 1

1820.0 1580.0 1 2410.0 1880.0

ALL COMPOSITE PROPERTIES ME EASED ON 0' FIBER ORIENTATION

® ALL PROPERTIES MEASURED UNDER MEINET CONDITIONS

(TALL MATERIALS LISTED MEET SP-4-0022 FOR VCM AND TML

i

2-4
i

J0 y`e 4)'

ES ~a' ' y ^a y

ORIGINAL PAG[ !'

OF POOR QUALITY

Table 0-4. Metallic Materials for Fittings and Springs

MAT[OT61c(2)

w,

MECHANICAL

Ft0 (Mod) 518.0 235.0 1790.0 427.8 276.0 1000.0 587.0 1830.0

Fc0 (WA) )70.0 110,0 1655.0 ))1.0 207.0 10)4.0 414.0 1107.0

FL 0 (MP4) 345.0 159.0 1080.0 248.4 172.0 621.0 366.0 793.0

Et (GPI) 200.0 448.0 200,0 72.5 289.0 116.0 193.0 200.0

EC (GPO 200.0 448,0 200.0 74.5 289.0 113.0 193.0 206.0

GL (GPI) 79.4 16,6 75.9 27.6 1)0.0 42.8 79. 4 75.9

p 0.28 0.35 1

0.72 0.)) 1

0.10 0.311 0.28 0,28

PRVSICAL

SO (J/k9K) 502.0 1004.8 477.0 840.0 1825.0 544.0 I 502.0 460.5

k (N/MK) 176.0 76.0 38.0 12.1 8,4 ].0 16.25 16.1

KO

(m/mK . 10'6) 17.2 25.3 Ii.) 2I.0 IZ.) 5.4 1' 10.4

p !49/1,7) 8030.0 1770.0 7840.0 2830.0 1850,0 44)0.0 8030.0 7760.0

APPLICATIONS FITTINGS FITTINGS FITTINGS FITTIm S FITTINGS FITTINGS SP SPRINGS SPRINGS

RINGS

NOTES: (1) ALL PROPERTIES MEASURED UNDER AMBIENT CONDITIONS.

(2) ALL MATERIALS LISTED MEET SP -1 !-0022 FOR VCM AND TML.

2-5
2-6

Y

ORIGINAL. PAGE 19

OF POOR QUALITY

Table 2.0-5. Non-Metallic Materials for Fittings and Springs

(71

MATERIALS Ow"ITE/HO%V (t) POLYUAIDE/ POLVSULFONE/

PROPERTIES X74 "S"34 1 PI00S/174 (2) GRAPHITE (1) GRAPHITE (1)

MIC14ANICAL

FL' (Mh) I42.0 999.0 1140.0 867.0 1317.0(4)

FT (ANN,) 46.0 12.0 (4) 77.0 (4) 10.7 42.0(4)

F^ (ANN,) 17!6.0 720.0 211.0 S71.0 1139.0(4)

FT (MN) 156.0 119.0 U 59,0 166.0(4)

F^ (ANN,) 69.0 44.0(4) 46.0 41.0 54.0(4)

FT (M►N,) 33.0(4) 21.0(4) 23.0(4) 20.0(4) 26.0 (4)

E, (ON) 145.0 765.0 425.0 704.0 119.0(4)

E T (ON) 10.0 6.0 U 6.2 4.3(4)

E -L

(ON) 177.0 246.0 425.0 254.0(4) U

OLT (Gh) 3.7 4.9 U 4.62 U

NLT

0.29 0.765 U 0.70 U

PHYSICAL

Sp (JAU K) 150.0 U U 079.0 U

Y L(W/eK) 9.0 19.0 520.0 77.57 91.0(4)

A T (W/TK) 0.7 1.0 J 1.21 U

K N,E (m/mK a 10- 6 ) 0.27 -1.04 -1.6 -0.99 0.23(4)

KN,T (m1mK w 10-6 ) 25.2 25.6 16.9 28.0 U

0.15 0.115 0.05 0.85 0.65

a 0.60 0.00 0.00 0.80 0.80

R 0.20 0.20 0.20 0.20 0.20

P (kp/m7 ) 1600.0 1765.0 1820.0 1550.0 1490.0

APPLICATIONS FITTINGS FITTINGS SPRINGS

SPRINGS SPRINGS

NOTESs (1) ALL COMPOSITE PROPERTIES ARE BASED ON 00 FIBER ORIENTATION

(2) ALL PROPERTIES MEASURED UNDER AMWENT CONDITIONS

(7) ALL MATERIALS LISTED MEET SP-R-0022 FOR VCM AND TALL

(4) CALCULATED VALUE FROM BASIC MA(ERIAL PROPERTIES

swk+^+zvnc#:L$^.=e.amo-e::...wo..--r--. _ ^ ^ .•,.. . .. _.... ^..x^...^e+.,....-,.,... s.. ..

4 x., IFS,`._ ^ ... .,.:.^r.:.n..

... .... 

4 •.._ - _
6i

Springs

Table 2.0-4 presents the conventional spring materials normal,y used in

spacecraft. These materials are metallic and have the same shortcomings of

excessive thermal ezpaneion as metals proposed for other applications on the

deployable platform systems. To obtain a close match of the coefficient of

thermal expansion to a graphite composite structure, it is desirable to

manufacture the spring out of the same basic materials. Some of the new

organic composite materials being considered for the production of springs are

presented in Table 2.0-5. These composite springs may not be commercially

available by 1986. However, titanium springs can be considered as the closest

compatible compromise in the event commercial availability of composites lags

the need date.

Flexible Tension Members

Flexible tension members usually consist of metal cables such as the 17-7PH

stainless steel presented in Table 2.0-6. This type of const ruction has two

mayor deficiencies for this application. One, the high coefficient of thermal

expansion of the metals makes these cables incompatible with the contemplated

graphite/epoxy structure. To further aggravate this condition, the twist in

the material can exaggerate the dimensional mismatches that occur with thermal

excursions. To reduce, if not eliminate this problem, graphite fiber tapes of

monofilamente with flexible organic binders are proposed. A sample of one of

these materials is RTV566/graphite and is presented in Table 2.0-6. Most of

the properties of the individual components of the composites are known but

significant work remains to confirm the potentials of these materials in a

combined form as flexible cables.

Table 2.0-6. Materials for Tension Cables

MATERIALS (2) RTV 566/ 17-7PH STAINPERTIES (7) GRAPHITE (1) LESS STEEL

MECHANICAL

FP(MPG) U 1870.0

41 (O►-) U 200.0

PHYSICAL

SP (JAW U 460.5

It (W/mK) U 16.1

K, (-/-Ka 10-6) U 10.4

[ 0.90 0.70

a 0.85 0.10-0.25

R 0.15 0.90-0.75

P (6a/m7) V 7760.0

APPLICATIONS I TENSION TENSION

NOTES, (1) ALL COMPOSITE PROPERTIES ARE BASED ON 0° FIBER ORIENTATION

(2) ALL PROPERTIES MEASURED UNDER AMBIENT CONDITIONS

(7) ALL MATERIALS LISTED MEET SP-R-=2 FOR VCM AND TML

2-7
I

4

ORICjNnL PA(3'

Of POOL QUA" rrY

Electrical Conductors

The primary criterion for electrical conductors, apart from the obvious

properties such as specific electrical conductivity, is the ability to coil

and uncoil the conductors several times through very tight radii for stowage

and deployment in a deployable structure. To this end, copper with its smaller

radius of wire for equivalent electrical conductivity, surpasses aluminum.

Its superior mechanical properties for like vire gauges surpass silver for

this application (Table 2.0-7).

Table 2.0-7. Materials for Electrical Conductors

® MATERIALS (2)

PROPERTIES ALUMINUM

1100 COPPER SILVER

MECHANICAL

Ftu

(MPa) 165.0 207.0 138.0

Et (MPa) 68.2 117.0 73.0 .

PHYSICAL

Sp (.1/kg K) 958.8 385.2 238.6

k (W/mK) 222.0 7.11x 10 5 427.0

Ke

(m/mK x 10 -6 ) 23.86 16.74 19.0

p (kg/0) 2710.0 8920.0 10,500.0

1/7 (ohm-cm x10-6 ) 2.6548 1.6730 1.50

APPLICATIONS FLUID LINES m ELECTRICAL ELECTRICAL

ELECTRICAL

CONDUCTOR CONDUCTOR

CONDUCTOR

O ALL PROPERTIES MEASURED UNDER AMBIENT CONDITIONS

O ALL MATERIALS LISTED MEET SP-R-0022 FOR VCM AND TML

2-8

_... ._. .Aff
i OF POOR QUALRY

Diffusion Barriers for Fluid Lines

Fluid diffusion bcrrier s ( Table 2.0-8) are one application where only a metal

can perform effectively for long exposure times. There are a number of

candidate materials with little to choose between. The possible exception is

the use of the metal matrix composites which technically provide the beet of

all worlds, low coefficient of thermal expansion, high resistance to gas

diffusion, high modulus and low weight. The drawbacks are extreme high cost

and the commercial availability by 1986.

Table 2.0-8. Diffusion Barriers for Fluid Lines

MATERIALS 17-7PB ALUMINUM/ (4)

(^) STAINLESS GRAPHITE ALUMINUM ALUMINUM ALUMINUM TITANIUM

PROPERTIES (3) STEEL (3) 5052 5056 1100 (GAL-4V)

MECHANICAL

Ft0 (MPs) 1830.0 483.0 289.8 414.0 165.0 1100.0

EL (OPS) 200.0 207.0 70.4 71.1 68.2 118.0

PHYSICAL

SP (J/kRx) 460.5 962.8 962.0 921.0 858.8 544.0

k (R/mk) 16.1 233.0 138.0 116.8 222.0 7.0

Xm (m/m[ x 30'6 ) 10.4

T 29.6 3

24.1 24.1 23.86 9.4

dHm (m2 /men-ATM) NEGLIGIBLE NEGLIGIBLE NEGLIGIBLE NEGLIGIBLE NEGLIGIBLE NEGLIGIBLE

P (k/m3 ) 7760.0 X410.0 2880.0 2640.0 2710.0 4430.0

CHEMICAL

HYDRAZINE A A A A A A

HELIUM A A A A A A

FREONS AS A CLASS A A A A A X (5)

NITROGEN TETROXIDE A A A A A A

APPLICATIONS FLUID LINER FLUID LINES FLUID LINES FLUID LINES FLUID LINES FLUID LINES

(1) ( U (1) (1) (1) (1)

NOTW

(1) THE FLUID LINES MAY BE OVERRRAPPED HYBRIDS OR SINGLE MATERIALS

(2) ALL COMPOSITE PROPERTIES ARE BASED ON 0 0 FIBER ORIENTATION

(3) .ILL PROPERTIES MEASURED UNDER AMBIENT CONDITIONS

(4) ALL MATERIALS LISTED MEET SP-R-0022 FOR VCY AND TML

(5) CHLORIDES AND CHLORINATED HYDROCARBONS OTHER THAN FREON TF ARE UNACCEPTABLE

Diffusion Barriers-Bellows

Considering the extended exposure times anticipated for flexible portions of

the fluid lines the only materials reasonable to consider for this application

are metal bellows. The two primary materials currently in use are presented

in Table 2 . 0-9. It is not anticipated that any significant new material will

be available for point design by FY 1986.

2-9 
i ORIGINAL 

PRO: "f OF POOR Q

Table 2.0-9. Diffusion Barriers for Bellows

MATERIALS

PROPERTIES* 321 STAINLESS STEEL INCONEL 718

MECHANICAL

Fl u(MPa) 586.0 1240.0

FTu

(MPa) N/A N/A

F cu (MPa) N/A N/A

Et (GPa) 193.0 200.0

ET (GPa) N/A N/A

u 0.28 0.31

PHYSICAL

Sp (J/kg K) )02.0 418.7

dHe (sec/m2 s) NEGLIGIBLE NEGLIGIBLE

P (k9/m3)

117 (ohm-cm x 10- 6 ) N/A N/A

CHEMICAL

HYDRAZINE A A

HELIUM A A

FREONS AS A CLASS A A

NITROGEN TETROXIDE A A

APPLICATIONS BELLOWS BELLOWS

*ALL PROPERTIES MEASURED UNDER AMBIENT CONDITIONS

2-10 

-
ORIGINAL Pf.•'' .

y OF POOR QL'ALI i'^

Vibration Damping Materials

Currently, there are primarily two significant space - rated materials for

vibration damping. Theme materials are presented in Table 2.0-10. It is not

anticipated that any unusual developments will require any unique materials

development in this area.

Table 2.0-10. Vibration

Damping Materiels

MA RUUS SMIEW SYNTACTIC

R RT G.E.

PO TI

AIRCRAFT

MECHANICAL

P^ (KP.) 2760.0 }20.0

(x) 100.0 55.0

Ta.M,J UAW We) U 1025.0

G (00 U 41,400.0

T«, S"Ih (N-M) U 11.2

H (Shwa A) 50.0 win. 70.0445.0

01 . 0.162 0.12

(i KHa) (i MHa nw..)

PHYSICAL

VCM (%) 0.02 0.10

TML M 0.19 0.22

Gim TnwWHan(OK) U 2V.0

APPLICATIONS VIYATION VIMATION

DAMPING DAMPING

Thermal Control Coatings

Some typical apace - rated thermal control coatings currently available are

presented in Table 2.0-11. This area requires some extended development not

only for deployable apace structure, but for all spacecraft. Very few

materials serve in this capacity and meet the outgassing contamination

requirements of NASA specification SP-R-0022. These materials have some

problems with adhesion and general in-apace deterioration from causes not

fully understood.

2-11

1

j
uwuil AL PALL N

OF POOR QUALITY

Table 2.0-11. Thermal Control Coatings

i

C

s .4

11ALf HACK CHEM-

'• GLAZE x701 f I3GAO11111111

MECHANIC L

A&mkn (1) NO EVIDGICE NO EVIDENCE

OF LIFTING Of LIFTING

FIMSICAL

d 0.91 0.22

0.86 0.85

VCM (%) 0.01(3) 0.03(2)

UAL (%) 0.56(3) 0.0 (2)

Af9lICATIONS TNERMAL TI91MAL

CONTROL

j

CONTROL

COATING COATING

NOM (1) 1ESTED FER ASTM D.=

M flaLY CUED (30 DAYS O/ LONGER) AT ROOM TEMFERATURE

(3) CUED 30 DAYS AT ROOM MWERATURE 

Lubricants

Several apace-rated lubricants exist for most of the typical applications that

are anticipated. For solid film lubricant applications, there are the

molybdenum diaulfides such an Mclykote 3402 which conforms to MIL-L-3937 and

Adrecolube 13 which conforms to MIL-L-46010. There is also a tungsten

diaulfide solid film lubricant Microseal. Where greases are needed there is a

fluorinated material, Braycoat 3L-38RP; and for an oil, a fluorinated

material, Brayco 815Z. All theme materials meet the requirements for

outgassing in SP-R-0022. The materials listed are typical of the many

available and are presented for illustration only.

p

4ii

N

ag 2-12

A'

d
OF P(j, , 1.1 L(1loLI*TY

3. IDENTIFICATON OF TECHNOLOGY DEVELOPMENT NEEDS

Table 3.0-1 presents the special technology needs summary. It eummarizee the

estimated cost of resolving each new technology need, as well as the estimated

calendar time.

Table 3.0-1. Special Technology Needs Summary

POTENTIAL TECHNOLOGY REQUIREMENT ESTIMATED SCHEDULE TO

(LISTED BY RATING) TOTAL COST (f) ACCOMPLISH (YR) PRIORITY

• DAMPING CHARACTERISTICS PREDICTION TECHNIQUE 350-400,000 4.0

• ORBITAL TRANSFER THRUSTER (1335-2225 N THRUST)' 780-960,000 3.5

• SPACE CHARGE DISSIPATION TECHNIQUES 124-150,000 4.0 HIGHEST

• JOINTS WITHOUT •SLOP' 190-220,000 1.5

• MINIMIZE STRUCTURAL LOAD AMPLIFICATION 295-340,000 3.5

• RADIATION-RESISTANT FIBER OPTICS 120-150,000 1.0

• LIGHTWEIGHT, LOW CTE, HIGH-STRENGTH CLUSTER FITTING 210-240,000 1.7

• PASSIVE STRUCTURAL DAMPING TECHNIQUES 100-125,000 2.0 MQDlUM

• MICROMETEOROID STRUCTURE DAMAGE 100-150,00^ 1.5

• RADIATION-RESISTANT THERMAL CONTROL COATING 95-115,000 018

• HIGH-CAPACITY HEAT PIPE 600-800,000 3.5

• ACTIVE STRUCTURAL CONTROL TECHNIQUES 300-350,000 3.3

• ADAPTIVE CONTROL TECHNIQUES 340-385,000 2.7 LOWEST

• LIGHTWEIGHT, HIGH-STRENGTH MEMBER (MALE) FITTING 155-U5,000 1.7

• INSULATED FLEXIBLE COOLANT LINES 190-240,000 2.5

• LOW CTE, HIGH MODULUS, FLEXIBLE TENSION MEMBER 100-130,000 1.0

TOTAL 4,930,000 4.0

(HIGHEST) (LONGEST)

THIS TECHNOLOGY REQUIREMENT IS INVALID IF EARTH STORABLE PROPELLANTS ARE CHOSEN FOR THE

ORBIT TRANSFER FUNCTION,

The technology needs are presented in three priority groups. The groups are

based upon consideration of net effect on performance, hardware level of the

problem, type of logic for resolution, level of teat simulation required,

development teat approach, and required hardware interfaces. There is no

significance to the order of the items listed within each priority group.

It is pertinent to note that numerous development testa need to be performed

pertaining to electrical cable bending, fatigue data, suitability of folding

points, telescoping joints, etc. These were not considered new technology

items since the technical approach is based on )mown methodologies.

The logic diagram shown in Figure 3.0-1 indicates the approach Rockwell used

to identify, validate, and estimate the coat of the special technology needs.

+c

3-1
sww opgMlom/mNefNlun a Rockwell

ORIGINAL Wl^ t s.I*mq sr.bms olvubn I, International

OF POOR QUALn-d

R

GATE 1

EJECT

SPECIAL TECHNICAL PASS PRELIM NARY

flEVIEW

TECHNOLOGY TECHNOLOGY BOARD'

FORM I •^LIU^I Iun FORMS 1 L 3

QUESTIONS

(1) 121131 141

TECHNICAL PASS SYSTEM

MATE

M

L ESTI COST/

REVIEW IMPACT 

SCHEDULE

R

AND

 RANKING 

U REQMTS

• Dr. Jim Haffner Domino Peebles

TECHNOLOGY Norm Scully Dave Pankopl

DATA OASE Dr. Ed French L Nishlmoto

Lee Smith Fred Etherldp

Figure 3.0-1. Logic Diagram—Special Technology Needs

Table 3.0-2 summarizes the 23 potential technology items identified, along

with the rationale used to reject seven of the items. As might be expected,

the greatest number of items are identified for structures and materials, as

well as dynamics and control.

Each of these potential technology requirements was validated by asking the

four questions shown in Table 3.0-3. The Technical Review Board (TRB),

representing in-depth knowledge in each of the disciplines overed, was able

to identify solutions/ongoing R&D activities for seven of the 23 items discussed. The remaining 16 items, which passed Gate 1, were scrutinized further

as the TRB collected confirming data. The data were reviewed at the second

gate,with confirmation of the 16 primary technology candidates.

Most of the requirements apply to all the structural concepts generated, not

just to individual concepts. As an example, radiation-resistant fiber optics

apply to all the concepts generated—not just one or two. For this reason,

the output of this task is deemed to be inappropriate for inclusion in the

selection criteria, but will be a direct input to the Preliminary Test Plan

of the Part II study.

A fis.al year schedule for each identified task, including major milestones,

was completed for each technology deficiency. Estimated costs were established through discussions with each responsible engineer in each discipline,

using a checklist which includes the costs for design and analysis, manufacturing, laboratory testing, major ground tests, flight tests, and (outside)

consultation fees.

3-2
ki ORIGINAL F,', • .

OF POOR Ql %: i V

Table J.0-2. Summary of Potential Technology Requirements

POTENTIAL I"NS RAiIONA.1 FOR SELECTION

STRIE7URES AND RAr 111 S

• MICIINETIOROIN IMPACT STRUCTURE DAM"

• OAOIATIOII-R91

FIBER OPTICS

• SPACE CHARGE DISSIPATION TIEMS114M

• RADIATION-ARSISTART THERMAL COATING

• LW CTS, HIM AMAUS, MIME VERSION MEMBER

• LW CTS, HIM MODULUS, COMPOCI If WIRIAL FDA CURRENTLY UNDER DEVELOPMENT IV G.E. AN COWAIR DE►LOVAIMI STRUCTURAL ELEMENTS

• LIWTNEIGNT, LW CTS NIW-IMNGTH CLUSTER

FITTINE

• LIWIIRIWT, HIG11-STRENGT11 rAMA (MALE) FITTING

• COOSITE MATERIAL CARPENTER'S MINIS MCARPENTER'S SHORT-TIRO DEVELOPMENT EFFORT

• ZERO-GAEF.A1* JOIN"

RSIRNAL CONTROL

• 10164"1IN FL911SU COOLANT LINES

• LONE-LIFT FLUID PUMP ONNITER'S RADIATOR FREON PEP

• HIGN-CAPACITY NEAT PIPE

• NUT PIPE INTERCONNECTION TECHNIQME CAN 04:01 V1 SV DESIGN TECHNIQUES

DESIGN APPRUACII DEMONSTRATED WAIN FLYING LUNAR

• ROTATING FLUID JOINT ION ARTICULATING RADIATOR EXCURSION ERSAIMENTAL PLATFORM CONTRACT

(NASIasIG, 1570)

UTILITIES

• TECHNIQUES TO REDUCE DENT RADIUS OF ELEC. CARllS DESIGN SOLUTIONS ARE WITHIN aMA/NT TSOROLOOV • TECHNIQUES TO MOM DERO RADIUS OF NNI- ►wER REALM

PROPULSION —_

• ORBITAL TRANSFER THROSTEA (I3J5-ZZ15MTMRIST)

--

DTNARICS AM CONFRW

• DAPPING CHARACTERISTICS PREDICTION TECHNIQME

• ADAPTIVE CONTROL TECHNIQUES

• ACTIVE STwRURAI CONTROL TECHNIQUES

• PASSIVE STRUCTURAL DAMPING TECHNIQUES

• MIMIMIZI STRUCTURAL LOAD AMPLIFICATION

Table 3.0-3. Validation Questions

)MPARABLE WORK BEING CONDUCTED NOW (OR CONTEMPLATED)

4SA, DOD, OR INDUSTRY?

D THE REQUIRED NEED DATE BE SATISFIED BY THE

)ING TECHNOLOGY RATE/TREND LINE?

(HERE VIABLE ALTERNATIVES IF THE TECHNOLOGY NEED

)T SATISFIED?

1E SOLUTION TO THE PROBLEM PRIMARILY A SHORT-TERM

IT (LESS THAN ONE YEAR)?

3-3
i The final products of this task are provided at the end of this section. The

individual tasks also lend themselves to more accurate milestone schedules and

time-phased cost eatimatoe.

The 16 technology items were then prioritized according to the criteria

described in Table 3.0-4.

The use of criteria is a method of measuring the impact of a new technology

need on deployable platform system performance. The first criterion is a

direct measure, while the remaining five are indirect meaaurements based on

the relative difficulty of solving a particular technology deficiency.

The results of the rating are summarized in Table 3.0-5, which indicates that

the point count ranges from a high of 4.75, down to 2.00. In those cases

where two or more technology deficiencies receive tbi same numerical total,

the same priority rating is assigned. Although this technique is rather

arbitrary, the ratings are relative to each other, and a difference of one or

two positions in not critical.

Because these rating criteria have only one direct measurement of the impact

of a new technology need on system performance, a sensitivity analysis is

conducted. The results of this analysis are shown in Table 3.0-6. The upper

limit of the first cri*arinn is raised to 3.00 (rather than 1.00) and each

priority rating valvt in thin column is multi p lied by 3.00 to intensify the

first criterion. The results (Table 3.0-6) indicate that roughly half the

technology require2 - its de not change position and the ether half moves only

one or two positions, i.e., the change is negligible.

3-4
OF POQ ^, (.- F 6

Table 3.0-4, Priority Rating Cciteri

g

 IZ

1 , N M H AN OVL FALL MEASURE OF THE IMPACT THAT A TECHNOLOGY DEFICIENCY HAS ON AN

OPERATIONAL DEFLOVABLE PLATFORM 0 SIGN IS TN! N(T EFFECT ON PERFORMANCE: AND, BECAUSE THE NET EFFECT INCLUDES ALL

TYPES OF TECHNOLOGY OEFICIENCIES, THE TERMS USED TO DESCRIBE THE IMPACT ON THE SYSTEM ARE RATHER GENERAL. OUR

NUMERICAL RATING ASSIGNS A VALUE OF 1.00 FOR SRVfOR PlareftumcS CIMITIATTON, WHILE AO IMPACT It ASSIGNED A VALUE

OF ZERO.

• SEV(C; PERFORMANCE LIMITATION 1.00

• CONSIDERABLE DEGRADATION 0.75

• MILD IMPACT 0.50

• MEμURABL. DEGRADATION 3.25

• NO IMPACT ON PERFORMANCE 0

r7 HARDWARE -REpE^Ep-^ ON[ INDI AI[T M[μURE Of TN( SYSTEM IWAGE OF A TECHNOLOGY DEFICIENCY IS THE LEVEL

^SRt/CUML XITTI Of INE HAR6IIARI REQUIRED TO RESOLVE THE PROBLEM. GENERALLY SPEARING, A COMPONENT ( SUCH AS A

PUMP) ACPAESENTS A MUCH LOWER COST IMPACT TO RESOLVE A PROBLEM, COMPARED TO A COMBINED SUBSYSTEMS M INTEGRATED

SYSTEM TEST PROGRAM, THE LATTER NORMALLY REQUIRES MANY ITCH$ Of GROUND SUPPORT EQUIPMENT AND KNOWLEDGEABLE TECHNICAE PERSONNEL IN A VARIETY Of DISCIPLINES TO SUCCESSFULLY DIAGNOSE AND COP RECT A SYSTEMS-LEVEL. DEVELOPMENT TASK,

A COMPONENT, ON THE OTHER NAND, USUALLY I[MILRES A MUCH SMALLEN TEAM WITH MORE IN-DEPTH TECHNICAL TRAINING AND A

FEW ITEMS OF GROUND SUPPORT EQUIPMENT. OUR NUMERICAL RATING ASSIGNS A VALUE Of 1.00 FORA SYSTEM-LEVEL PAOSLEM

BECAUSE Of THE RELATIVELY GREATER COMPLEXITY Of A SYSTEM VERSUS A COMPONENT. A COMPONENT-LEVEL PROBLEM HAS A

VALUE Of 0.25.

• COMBINED SUBSYSTEMS/INTEGRATED SYSTEMS 1.00

• SYSTEM/SYSTEM CRITICAL INTERFACE 0.75

• SUBSYSTEA 0.50

• CNPONEIT 0.25

• MATERIAL OR METHODS 0

^. L JC S_ U OM THE Tiff Of (COMPUTATIONAL) LOGIC REQUIRED TO RESOLVE A TECHNOLOGY DEFICIENCY

ALSO AN IMDIRZCT MEASUREE 0! SYSTEM IMPACT. CUT-AID-TRY LOGIC IS A FORM Of CUAV[ FITTING IN A REGION WHERE

RELA'IONSHIPS ARE NOT WELL DEFINED. THIS METHOD IS TIME-CONSUMING AND REQUIRES EXPERIMENTATION. AN EXAMPLE WOULD

SE THE KNOWLEDGE Of FLUID BEHAVIOR IN ZERO GRAVITY IN MO. AT THE OTHER EXTREME, ANALYSIS Of A PROBLEM USING

PROVEN FORMULAS/RELATIONSHIPS AND/011 COMPUTER-STGN(D PROGRAMS IS RELATIVELY SIMPLE (LEAST COST IMPACT).

• CUT AMC TRY 1.00

• DATA EXTRAPOLATION 0.75

• STATISTICAL/EMPIRICAL 0.50

• ANALYSIS/SYNTHESIS 0.25

LEVEL OF TEST/Sim-u-MrS-M790-1 TEST LEVELS REQUIRED TO FIND SOLUTIONS TO A PROBLEM ARE ALSO INDICATIVE

OF RELATIVE SYS12M IHFACT. A IVUOI FLIGHT YiST, FOR INSTANC:, IS VERY COSTLY BUT 15 SOMETIMES JUSTIFIED AS THE

ONLY TEST LEVEL WHICH WILL CREATE ACTUAL CORNERED TEST ENVIRONMENTS WHICH ARE NECESSARY FOR HIGH-CONFIDENCE RESOLUTION. LABORATORY TESTING, ON THE OTHER H IND, TENDS TO BE LESS COSTLY DECAUSE THE TEST TEAM IS GENERALLY SMALL

AND THE TEST ENVIRONMENT, EVEN IF CONSIDERED COSTLY, DOES NOT LAST VERY LONG (APPLIED INTERMITTENTLY).

• MAJOR FLIGHT TEST 1. DO

• MANOR GROUND TEST 0.75

• FLIGHT/GAWND PASSENGER 0.50

• LABORATORY 0.25

DEVELOPMCN T[ APPROACH THE DEVELOPMENT TEST APPROACH ACFCRf TO THE PRIMARY METHOD i0 BE USED TO OBTAIN

H 31RRLDTES SUL 1. AS AN EXAMPLE, COMBINED TEST ENVIROIM[NTS AN( GENERALLY PREFERRED, FROM A TEST POINT

OF VIEW, SECAUSE THEY STRESS THE TEST ARTICLE SIMILAR TO ACTUAL USE CONDITIONS AND THEREBY ELIMINATE SOME OF THE

UNKNOWNS (THE ADDITIVE EFFECT OF INTERNAL STRESS LEVELS CAUSED BY DIFFERENT ENVIRONMENTS), COMBINED TEST ENVIRONMENTS ARE DIFFICULT TO ACHIEVE μ'D ARE, THEREFORE, COSTLY TO SIMULATE, A DEMONSTRATION TEST, HOWEVER, IS

USUALLY STRAIGHTFORWARD AND, FROM A COST IMPACT STANDPOINT, REPRESENTS RELATIVELY LOW TEST COSTS.

• COMBINED SNVIROMMENTS TESTING 1.00

• DETERMINE SAFETY MARGINS 0.75

• DETERMINE FAILUAF "DOES 0.50

• DEMONSTRATION 0.25

• GATHER DATA 0

OTHER HARDWARE IE R ACES R QU RED OTHER HARDWARE INTERFACES ACQUIRED TO CONTROL, MONITOR, SIMULATE, RECORD,

COOL. SU^NT, OR T^T^ ARTICLE ARE ALSO AN INDIRECT MEASURE OF THE IMPACT OF A TECHNOLOGY OEFICIEM CY. OTHE E HunuuF IWWRFLEES REPR(SFNT CONSIOFRAULE COST TO OBTAIN AND OPERATE. ON TO SIMULATE THE EFFECT OF

TVA INTERFACE. FOIL THESE REASONS, MNBISRD SURRYSTfMf/TNTfGXA= SYSTICAT (μ A HARDWARE INTERFACE) IS THE MOST

MMTIY ANM THKRFPMF HAS THE NNGNEST PRIORITY RATING VALUE OF I.W . M M SUBSYSTEM INTERPACBs REPRESENTS THE

I

LEAST COST IMPACT AND HAS THE LOWEST PRIORITY RATING VALUE OF 0.25.

• iurdlneu SuvaviTEASii•I[GF.AT0 G.iTEiw

• OTHER COMPLETE SYSTEMS 0.75

• OTHER PARTIAL SYSTEMS 0.50

• OTHER SUBSYSTEM INTERFACES 0.25

3-5
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3-7

EiNvMO i
R

ojr IGIWAL PAGE 13

OF POOR QUALr"

soKll opMsosahsawaam a,^^ Rockwell 5"1114 s scam. Division Intematlonel

r ..

I NEW TECHNOLOGY NEED I

TASK TITLE: I

Damping Characteristics Prediction Technique

JUSTIFICATION:

A highly damped structural system will minimize any requirement to augment motion with the attitude control function. A low damped structural

system may require augmented attitude control ( to meet pointing accuracy)

as well as a variable gain function. The latter situation would be

more complicated (less reliable) and weigh more than the first situation.

The ability to accurately predict the degree of damping may alleviate

a complicated attitude control function and avoid an overkill with passive

damping techniques.

TECHNICAL OBJECTIVES:

Develop a method of analysis (verified by test data) which will predict

the damping characteristics of deployable structures containing extensive

utility lines and cables in zero gravity.

TASK DESCRIPTIONS:

1. Develop a modeling technique for utility lines and cables which

are secured to basic structure in a variety of ways (various

damping coefficients).

2. Verify the above modeling characteristics by free-free modal

testing of representative structures and utilities.

3. Verify the above modeling characteristics by space flight testing

representative structures and utilities ( zero-gravity mode).

4. Update the analysis technique as required.

r .; 3-8
ORIGINAL PAGE

OF pOOR UALITY

spats OpNMlons/Imo rtelon a 

o1, 

Rockwell

s 111w Systems DN1sten 1 International

I NEW TECHNOLOGY NEED I

TASK TITLE: ORBITAL TRANSFER THRUSTER (1335 TO 2225 N THRUST)*

Justification

Many light-weight structure concepts will be constrained to very low

acceleration forces (thrust to weight ratio) during transfer to

geosynchronous orbit. A new thruster, in the 1335 to 2225 N thrust

range, utilizing liquid oxygen and liquid hydrogen, needs to be

developed for this application.

Technical Objective

Design and demonstrate reliable operation of an orbital transfer

thruster in the 1335 to 2225N thrust range, pump fed.

Task Lescrintion

Conduct detail design, analysis, manufacture, and development tests to

verify performance and reliable operation of an orbital

transfer thruster in the 1335 to 2225 N thrust range. A soft thrust

buildup transient is a key requirement (5 to 15 second rise time), which

may require a variable area injector, a sequenced injector design, or

variable-speed, motor-driven pumps. 

*This technology requirement is invalid if earth storable propellants

are chosen for the orbital transfer function.

3-9
'ttFF

Y

Space Operations/ Integration &

I^, 

Rockwell

Satellite Systems Division International

TECHNOLOGY

ITASK TITLE: SPACE CHARGE DISSIPATION TECHNIQUES I

IJUSTIFICATION

The space environment, particularly at geosynchronous altitudes, is known to

cause a spacecraft to accumulate a differential charge as high as 20 kilovolts when a magnetic disturbance (called a substorm) occurs in its vicinity.

Large arcs, or the corona discharge so produced, radiate large-amplitude

fast-rise-time electromagnetic pulses that can be detrimental to circuits and

circuit piece-parts.

While there are active mechanisms to reduce surface charging (electron gun,

heated filament, or a plasma source), long-term, high-reliability requirement

suggest a high secondar y electron emission material. This area has not been

fully explored to date and holds the promise of a passive, long-life, highly

reliable method to minimize surface charging of the system.

TECHN?CAL OBJECTIVES

• Investigate and recommend, based on simulated environmental testing, a

surface coating which will produce secondary electron emissions (or

luminescence) at relatively low-voltage spacecraft charging levels.

• Develop a suitable application process for the above recommended coating,

based on laboratory testing and verification methods.

• Verify, by space testing, that the basic theory of reducing charge voltage

levels, through secondary yield or photon emission of the coating, actually

works.

TASK DESCRIPTION

Part I will investigate, by means of laboratory testing, various coating

materials which electroluminesce or produce high secondary yields at

relatively low-voltage platform charging levels.

The purpose of this task is to demonstrate the above effect by subjecting

a test structure, coated with a suitable material, to a simulated electrostatic charge.

Part II will concentrate on developing a suitable method of applying the

recommended coating. Laboratory application tests will be conducted, and

methods will be developed which verify uniformity and adhesion.

Prrt III will verify the basic theory of the new coating by space testing.

A spacecraft (an ATS geosynchronous vehicle) will be crated with the new

material and be suitably instrumented (a harness noise monitor and highvoltage-charge-accumulation monitor) to detect possible static charge

buildup and arcing.

3-10

6ttu..-.. , ._...._, •,^...... _..,. ^.._^ . _..,. u;. „,,; .^,._. ._^_.Aic.J_—^.._.......,,
Sp. Oa...aon. /:atoWauoaa ,I^ Rockwell sna snwns Division11 Intemaponal .w

NEW TECHNOLOGY NEED

TASK TITLE: Joints without Slop

JUSTIFICATION:

Stiffness of the deployable platform structure will be lost if backlash

is present in the joints. Also, the predictability of structural

performance characteristics will be imperiled.

TECHNICAL OBJECTIVES:

Develop a zero—backlash joint, suitable for deployable platform designs

and operational environments.

TASK DESCRIPTIONS:

1. Perform design and analysis to investigate and select several promising

approaches, both passive and active concepts.

2. Manufacture several joints of each design selected above.

3. Perform laboratory test program to verify zero backlash for all

operational (load) conditions and environments. Select the design

approach which best meets the performance requirements.

3-11
Q

Space Opmaoona/Integraoon&^^ Rockwell

Satellite Systems Division International

NOLOGY NEED

TASK TITLE: Minimize Structural Load Amplification

JUSTIFICATION

Several approaches are possible which will reduce load amplification

factors for lightweight structures, namely: (1) soft transient thrust

buildup of RCS thrusters (0.1 to 0.25 lb thrust, 10-20 second rise time);

and (2) optimum pulsing of RCS thrusters to deadbeat the structural

response of the system. The most cost-effective approach must be defined

and demonstrated because a structural load amplification of two

will require structural members with greater stiffness (more weight) for

the deployable platform system.

TECHNICAL OBJECTIVES

Select the most cost-effective approach to minimize structural load amplification for lightweight deployable space platform structures.

TASK DESCRIPTIONS

Two approaches are p esented. The optimum pulsing approach is shown to

be the most cost-effective method and will be adopted as the baseline

method to minimize. load amplification.

1. Develop and demonstrate an RCS thruster (0.1 to 0.25 lb

thrust level) which has a 10-20 second rise time.

2. Define a suitable method (hardware and mechanization) for

optimum pulsing of current design RCS thrusters, such that

the structural response of the system does not lead to

appreciable load amplification factors. Include verification testing using representative structure and utilities

in a space flight mode.

3-12

d.. .,..'. _. .,...,......._.....:,.^..:.^ 

.^^_.... _.......
6i sa.oa'o •/wn a 0 International

NEW TECHNOLOGY NEED I

TASK TITLE: RADIATION RESISTANT FIBER OPTICS

JUSTIFICATION:

Quartz fiber optics are subject to color center formations, called

opaquing (degrading effect to light transmission), caused by radiation

impingement in a vacuum. The Quartz fiber apace lattice is composed of

silicon and oxygen atoms which have a one-to-one correspondence (single

bond) between each element. Double bonds are formed within the lattice

structure (color is added) when radiation (electrons, protons, or

photons) displaces an oxygen molecule, thereby destroying one of the

single bonds.

Shielding would add considerable mass. The high density material

required would be a source of secondary emissions (Bremsstrahlung

radiation, produced by the impact of electrons, protons, or photons)

which can be more damaging (because these are harder to shield against)

than the original radiation.

Two approaches to resolve this problem are presented. The approach

which shows the most promise, at the end of the development period, will

be adopted as the baseline method for fiber optic applications.

TECHNICAL OBJECTIVES:

1. Determine the possibility that organic materials (such as Acrylic),

with higher energy linkage between bonds, are more resistant to

galactic radiation and will maintain high transmissibility of light.

2. Determine a suitable method of annealing the Quartz fibers with

heat, applied intermittently.

TASK DESCRIPTIONS:

1. Select the most promising organic materials which possess suitable

light transmission characteristics.

Conduct a laboratory testing program to determine the amount of

resistance to simulated galactic radiation of each material selectod

above.

Based on the above test results recommend an organic material

suitable for 10-year life in geosynchronous orbit.

2. Investigate possible methods of annealing Quartz fibers with heat.

Conduct laboratory tests to verify that annealing restores the

single bonds between elements (no degradation).

3-13

de
F

sass. opus" a/InwPMbn s 

^l, 

Rockwell

semmft sr.ns w oNrion 1 Intematlonel

I NEW TECHNOLOGY NEED I

I TASK TITLE: I

Lightweight, low CTE, High-Strength Cluster Fitting

JUSTIFICATION:

Without a low CTE cluster fitting, the effective CTE for the overall

deployable platform design could be appreciably higher and result in

increased thermal deformation.

TECHNICAL OBJECTIVES:

Develop a technique for producing multi -directional cluster fittings

which have the following characteristics: ( 1) low CTE in each -f the

projected ( longitudinal) directions; ( 2) close - tolerance clevie pin holes;

(3) high bearing allowables at the clevie pin holes; ( 4) low overall

weight; and (5) high strength across the fitting.

TASK DESCRIPTIONS:

1. Manufacture several cluster fittings using different ratios of resin to

fiber, different fibers, mixes of fibers (glass / carbon), as well as

combination techniques ( partial hand layup, partial injection mold).

Consider metal inserts for clevie pin holes.

2. Perform a laboratory test program to evaluate longitudinal CTE,

bearing stress limits, weight, load capacity across the fitting.

Select the manufacturing methods and material which best meet the

characteristics desired.

3. Develop a process specification to be used tc produce multi -directional

cluster fittings.

3-14
q

spWa00waWns/Inlpnl, a

salaslla syaa"na Ob410n 01

Rockwell

Intematlonal

I NEW TECHNOLOGY NEED—CONTINUATION SHEET

TASK TITLE: PASSIVE STRUCTURAL DAMPING TECHNIQUES

JUSTIFICATION. The visco-elastic material currently used to attenuate

structural response is very heavy and, therefore, WOU12 ^JL be suitable

for application to the entire deployable platform. A judicious application to each pinned joint, to cause the joint to respond more like a

cantilevered beam, needs to be investigated and design solutions verified

by testing.

TECHNICAL OBJECTIVES. Develop passive structural damping techniques

suitable for the pinned joints of large deployable apace platform

structures.

TASK DESCRIPTIONS

1. Design studies to determine best means of applying visco-elastic

material to pinned joints.

2. Laboratory test program to verify attenuation of structural response

(pinned joints respond more like cantilevered beam).

3. Establish a design standard for applying visco-elastic material to

pinned joints for deployable space platforms.

3-15
Space operations/ ln"imlbn aQ% Rockwell

Satellite Systems Ohleton International

TASK TITLE: Micrometeoroid Impact Structure Damage

JUSTIFICATION

The probability of micrometeoroid impact upon the structural members In a

deployabla truss is sufficiently signlfl.cant to be of concern to the

structural integrity. Presently, there is virtually no data descriptive

of the impact damage upon structural tubes of circ-ilar, square, or rectangular cross sections fabricated from composite materials.

TECHNICAL OBJECTIVE

To develop analytical method (confirmed by testing) for prediction of

micrometeoroid impact damage to circul • r, square, and rectangular tubes

of graphite epoxy and aluminum construction.

(1) Develop analytical methodology, using existing techniques to

maximum extent possible, to predict structural damage of candidate constructions.

(2) Manufacture nine representative sections of graphite epoxy tubing

and nine representative sections of aluminum tubing. Two tubing

shapes will be made (nine round, and nine square).

(3) Characterize each different tubing shape (round, square) for

compressive strength.

(4) Test six sections of each different tubing shape for hyper!elocl.ty

impact, using glass beads with a mass of one gram (or less) to

produce velocities of 7 km/sec or more. Three sections are to be

hit on centerline, and three sections hit one inch off centerline.

(5) Retest each tubing section (which was penetrated) for strength

characteristi.:s as described in item (2) above.

(6) Correlate analytical predictions with test data and mobility

prediction techniques, as required.

3-16
space Ope rations / Integration •,^, Rockwell

satellite system$ Division International

Y NEED

(TASK TITLE: Radiation-Resistant Thermal Control Coating

JUSTIFICATION

The white pigments used in thermal control coatings (such as zinc oxide

and titanium dioxide) tend to increase in solar absorptivity in the space

environment. The cause is believed to be a combination of ultra violet

and entrapped radiation, causing the loss of a small percentage of oxygen

or water in the pigment. Most white paints will tulowly dlscolcr because

of color centers formed (caused by the loss of oxygen or water). This

phenomenon causes spacecraft temperature to increase and could imperil the

operation of temperature-sensitive components using structure as a heat

sink.

TECHNICAL OBJECTIVE

Verify that a new pigment (such as zinc orthotitanate), with an appropriate

binder, will not degrade significantly over the ten-year life of the space

platform in geosynchronous orbit.

TASK DESCRIPTION

Conduct laboratory tests on new pigments (such as zinc orthotitanate), with

an appropriate inorganic binder, to determine the amount of degradation to

be expected in a space environment over a ten-yeat period. Recommend a new

pigment and binder combination for long-term space applications.

3-17

_ c J ,. ..
sPa oowuae "w rwon aV%flockwell

smM apWn• otsNbn .a Internabonel

NEW TECHNOLOGY NEED

TASK TITLE: HICH CAPACITY HEAT PIPE

Justificat?on

High capacity heat pipes, in the range of 20 KW-meters and above, have

not been developed and demonstrated.

Technical Objective

Develop t. , d demonstrate a high capacity heat pipe.

Task Description

Develop and demonstrate a high ca pacity heat pipe design ( 20 KW-meters

and above), suitable for apace platform application.

3-16
spec. Operations/ InYq.ann• Rot.:;well

3auaay.wn.oti:.ion Interrmtlonel

I NEW TECHNOLOGY NEED

TASK TITLE: A,tive Structural Control Techniques I

JUSTIFICATION:

This technology will permit arstem performance requirements to be achieved

in the presence of relaxed structural stiffness, frequency, damping, and

alignment requirements.

TECHNICAL OBJECTIVES:

Develop an active structural control technique for vibration suppression

and shape control ( including use of distributed sensors and actuators)

in order to reduce requirements for structural stiffness, frequencies,

and damping.

TASK DESCRIPTIONS:

1. Develop meth^dology.

2. Synthesize active structural control approach for space platform.

3. Simulate evaluate sensitivity to "real world" errors and

develop hardware requirements.

4. Develop sensor and actuator requirements.

5. Verify sensor and actuator I.arformance by space flighttesting representative structure and utilities.

3-19

..., ....emu ....., ...,:.... .... , rwl uJn}..,L.,. ..._....1rd^.,l.^.Y.'......n..,u.^,-..::,..t,^xw.;eh^eflewmueti:L'ra"n^,. n''.^_^.'•^•^
e,9

trs^.^

Soate Opersnons/ Integration &„ Rockwell

Satellite Systems Division I nternational

NEW TECHNOLOGY NEED

TASK TITLE: Adaptive Control Techniques

JUSTIFICATION

Large tolerances on structural dynamic parameters, from preflight analytical estimates, will not permit achievement of a high-performance

controller. Ground testing of very large structures is difficult and

expensive. Classical frequency separation criteria impose unnecessarily

severe structural dynamic requirements.

TECHNICAL OBJECTIVES

Development of an in-flight dynamic mode identification technique in order

to reduce tolerances of structural dynamic parameters. These data will

permit adaptive readjustment of the controller to achieve higher stability/

performance and/or reduction in structural bending stiffness and frequency

requirements.

TASK DESCRIPTIONS

(1) Develop methodology.

(2) Synthesize adaptive approach for space platform application.

(3) Simulate system and evaluate sensitivity to "real world” errors.

(4) Formulate new criteria for structural dynamic design requirements.

(5) Verify in-flight dynamic mode identification technique by ground

testing representative structure and utilities.

3-20

I

1<
Spme 0pwsWns/1n6*W*don a ,I, Rockwell

SM"He SVM m Motion jj Intematfonal

NEW TECHNOLOGY NEED

TASK TITLE:

Lightweight, High-Strength Member (1lalei fitting

JUSTIFICATION:

Separate (male) fittings will weigh more (and require more volume) than a

fitting which is integral with the structural member. Packaging efficiency

of the deployable platform may be degraded.

TECHNICAL OBJECTIVES:

Develop a technique for producing lightweight, high-strength member (male)

fittings as an integral part (extension) of the structural compression

member.

TASK DESCRIPTIONS:

1. Manufacture several member (male) fittings by a layup technique,

transitioning from a circular (tube) member to a flat tongue.

Consider the addition of different fibers in the male fitting,

as well as metal inserts for the clevis pin hole.

2. Laboratory test program to evaluate compression/tension limits

and bearing stress limits. Select the manufacturing methods and

materials which best meet the desired characteristics.

3. Develop a process specification to be used to produce integral

(male) member fittings.

c

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sue. opwadw./1n Wa w s,I^ Rockwell

sn.mc.sy.nso^. W ^ 1 Intematlonel

NEW TECHNOLOGY NEED

TASK TITLE: INSULATED FLEXIBLE LINES

Justification

Insulated fluid lines, currently on the market, are not flexible enough

to bend to a small radius. Those designs available which will bend to

a small radius are quite bulky and require too much room when stowed

in a deployable structure.

Technical Objective

Develop a new concept for insulated flexible lines.

Task Description

Design and develop a new concept for an insulated flexible line which is

not bulky and will permit bending to a small radius.

t

i

P

4

3-22

..r : ♦ F ..

....

1

_ ..L. .6. Y^ na..._ ice. • ._^ .- 1 .^.... - ..casnu...^... u._. .iur.^.. L-.....v......_. .nws
Specie OpwallonslIntograilon & Rockwell

SiAtille Sysisms Division Intemational

I NEW TECHNOLOGY NEED

I

TASK TITLE: Low CTE, High Modulus, Flexible Tension Member

JUSTIFICATION:

The only high-strength tension member which is quite flexible and currentl

available is stranded wire cable. Wire cable, however, has a substantiall

higher coefficient of thermal expansiun compared to the anticipated

composite structure. Net result would be a loss of tension in sunlight

and excessive tension in the earth's shadow. The structure's shape,

pointing accuracy, and load-carrying capacity would be severely degraded.

TECHNICAL OBJECTIVES:

Develop a high-strength tension member which will bend to a small radius,

have an extremely small coefficient of thermal expansion, and not

degrade in the space envirormien^..

TASK DESCRIPTIONS:

Laboratory test program to determine the best composite material (such as

P100S graphite fibers, imbedded in an elastomeric matrix material) which

will bend to a small radius, have an extremely small coefficient of

thermal expansion, and will not degrade in a space environment.

3-23
t

4. CONCEPT SELECTION

This section describes the process used to select the moot suitable concept

from the e + ght candidate building block designs, described in Section 1. This

process identified Concepts 6 (Figure 1.4-17) and 8 (Figure 1.4-20) as

together being the designs that beet satisfy the mayor criteria tabulated in

Table 4.1-1. The major reasons for identification of Concepts 6 and 8 are

anu®erated as follows:

o The open-square shape of Concept 8 has the beet growth potential for

utilities and can accommodate (in trays) up to twice the adopted

study requirements (increased utilities requirements are typical

with program maturity).

o For significantly reduced utilities requirements (mountable on

longerons) Concept 6 is adequate and is a simpler structural design

than Concept 8.

G, o The two designs %.ogether can satisfy the LEO/CEO platform

requirements with common concepts of housing, adapter, deployment

F mechanization and building-block to building-block attachment

I. designs.

o Redundancy for meteoroid impact survival (if necesoary) is available.

o The square housing is most suitable for mounting of payloads,

subsystems, propulsion modules, and mounting ports.

o The square housing is most amenable to support in the orbiter cargo

bay.

The identification of Concepts 6 and 8 was accomplished through review of

summary tables (4.10-1 through 4.10-5). These tables encompass the major

criteria of Table 4.1-1.

At the conclusion of this selection process, the beat features of Concepts 6

and 8 were configured into a new concept called 6A, (Figure 4.11-2). Concept

6A is the same design as Concept 6 except the longerons are folded at 30'

rather than 45 * (Drawing 42712-29, Volume II), clearing the center of the

square frame for installation of utility trays such as are shown for Concept 6

(Drawing 42712-025, Volume II). The overall features of thi9 design are

summarized as follows:

o Building-block approach for automatic deployment of platform systems.

o The square-shaped truss is moot suitable for inter-building-block

attachments; mounting of payloads, docking ports, propulsion

modules, etc.

o Circular tubes for all trues members provide minimum cost

construction with use of graphite composite construction.

• 4-1

X ! T 7

. ....:_. .in: 

•S,.w... rY•.J_:.4.._^ ... ..... ......c •. .~...hwv.:.....^a 1.. '.a^..,aTss...cArw..... rm u
o Trays for mounting of adopted complement of utilities provide ease

of initial installation, repair, and replacement during total ground

fabrication period with minimum trues structural design constraints

imposed by utilities integration

• Small complements of utilities can be mounted directly onto the

longerona (design reduces to Concept 6).

• Square-shaped housing with reciprocating deployment mechanism

• Bey-by-bay deployment to facilitate identification of deployment

problem (in the event this occurs).

o Rail system for root strength during deployment permits orbiter

berthing and orbiter VRCS firing, if necessary.

o Adapters for mounting of payloads with automatic electrical

connector interface.

o Payloads and propulsion modules attached using RMS.

Further detail concerning the selection process is provided in the remainder

of this section.

04

4-2

, ..-•. d. -. •__. .rte. _. ._x..,
ORIGINAL PAGE 19

OF POOR QUALITY

4.1 MAJOR CRITERIA

The criteria used in the W action are listed in Table 4.1-1. Many other

criteria were included and then rejected, since there was no difference among

the concepts insofar as these criteria are concerned. For example, one early

criterion was the ability to deploy in a straight line. All of the concepts

have this characteristic, hence that particular criterion was eliminated.

An explanation of each criterion together with the approach to grading the

concepts in provided in Sections 4.3 through 4.9.

Table 4.1-1. Major Criteria Used in

the Selection Process

1. DESIGN VERSATILITY (WITH DISTINCTIONS BETWEEN LEO AND CEO)

OF STRUCTURAL CONCEPT

• Accommodation of adopted power and data utilities requirements

• Accommodation of reduced power and data utilities requirements

• Accomodation of fluid utilities: Two 2-cm lines (or equivalent)

• Satisfaction of adopted strength and stiffness requirements

• Satisfaction of strength and stiffness requirements that are each

1/10 of the adopted values

• Satisfaction of the adopted strength requirement and 10 times

the adopted stiffness requirement

• Platform construction

• Accommodation of aluminum and graphite composite materials

2. COST OF TOTAL BUILDING BLOCK IN GENERIC PLITFOYA

• Launch cost

• Fabrication cost

• Orbit transfer to GEO

• Technology development differential (negligible)

3. THERMAL STABILITY OF STRUCTURAL CONCEPT

4. METEOROID 1HPACT SUITABILITY OF STRUCTURAL CONCEPT

5. RELIABILITY OF DEPLOYMn (BUILDING BLOCK)

• Basic truss structure • Docking port structure

• Housing • Materials variation

• Adapter • Mechanization

6. PREDICTABILITY OF PERPORHANCE OF STRUCTURAL CONCEPT

7. INTEGRATION SUITABILITY OF BUILDING BLACK

4-3
ORIGINAL PAGE 19

OF POOR QUALr"

4.2 lIETHODOLOG7

The eight concepts are graded by an allocation of points. Points are

allocated to the concepts in two ways:

o Qualitative data art converted to points judgmentally.

o Quantitative data are converted to points using a linear systea as

shown in Figure 4.2-1. Regarding the line marked "baseline

evaluation" the most desirable concept is awarded 100% points and

the least desirable concept is awarded 50% points. The other

concepts are graded on a linear basis between the two extremes.

This method was used for all the tables shown up to and including

Tables 4.10-1 and -2.

QUANTITATIVE DATA

MAP. _

001N

Nl RVALUATIM

STUDY

u EXAMPLE

rolNrs

MINIMUM DESIGN :ORETMINT!-

MNIMUM OIlIWL[ VALUE LEAST 20

OF PERFORMNCE PARMET[R DElI MOLE

VALUE

/[AROR

MA

MARCI

PARAMETER C

10

P[AFORI WI PARAMETER Y

0

0 2.0 A 0 6.0

MIL LEMOTM (.)

Figure 4.2-1. Point Evaluation Methodology

Another approach is shown on Figure 4.2-1, but uses the line marked

"sensitivity study". The only difference is that zero % points are awarded to

the least desirable concept instead of 50 %. This method was used in compiling

aummary tables ( 4.10-3 and -4). Both the " baseline study" and the

"sensitivity study" are incorporated into this selection process.

There is only one important distinction between LEO and GEO in terms of 

concept selection, and that is mass. Other differences between LEO and GEO

are listed below, but do not affect concept selection:

• Tranafer to GEO introduced a somewhat higher loading regime, but not

sufficient for distinction in trade

• No significant differsncea observed in flexural and torsional

stiffness requirements

4-4
ORIGINAL PAGE i5 OF POOR QUAL."

o No differences (significant to concept selection) observed in

- Utilities requirements (function of payloads)

- Meteoroid environment

- Thermal environment

- Servicing and maintenance

4.3 DESIGN VERSATILITY OF STRUCTURAL CONCEPT

The grading of the eight buiding-block concepts for this criterion is shown on

seven tables:

o Table 4.3-1 - Electrical Accomodations (GEO)

o Table 4.3-2 - Electrical Accomodations (LEO)

o Table 4.3-3 - Fluid Utilities Accommodations (LEO and GEO)

o Table 4.3-4 - Structural Materials Variation (LEO and GEO)

o Table 4.3-5 - Strength 3 Stiffness Accommodations (GEO)

o Table 4.3-6 - Strength 8 Stiffness Accommodations (LEO)

o Table 4.3-7 - Platform Construction (LEO and GEO)

The following notre are intended to explain the rationale behind the grading

and points alloc^,ion of Tables 4.3-1 through 4.3-7.

For Tables ? .3-1 through 4.3-3:

0 Mass implications apply to the mass of trays, clips, and shoes,

etc., necessary to support utilities.

0 The high electrical requirement is the baseline adopted requirement.

0 The low electrical requirement consists of eight No. 4 lines and 1/8

the complement of adopted data lines.

0 The values assigned are based on accommodation of utilities by the

utilities installation designs tabulated as follows:

CONCEPT LOW REQUIREMENT NIGH REQUIREMENT FLUIDS

I TRAYS TRAYS TRAYS

2 TRAYS TRAYS LOOPS

3 COILED COILED LOOPS

4 LONGERONS LOOPS LOOPS

5 TRAYS TRAYS TRAYS

6 LONGERONS LOOPS LOOPS

7 TRAYS TRAYS TRAYS

8 TRAYS TRAYS TRAYS

4-S 
ORIGINAL PALL 17,

1 Of POOR QUAMY

Tables 4.3-1 and 4.3-2 address the comparative capability of the designs to

accommodate the electrical utilities, respectively, for GEO and LEO

applications. The foll ,)wing clarification of the evaluation parameters is

presented:

o Reliability for Space Deployment

This parameter addressee the

electrical lines from the at

Retraction is not a factor.

the longerone are thought to

similarly reliable for small

reliable for large bundles.

reliable.

likelihood of successfully moving the

owed to the deployed position.

Linea which are mounted on trays or on

be the moat reliable. Looped lines are

quantities of lines, but slightly lace

Coiled lines are fudged to be the least

o Ease of Ground Deployment and Checkout

r'

This section addressee the ease/difficulty of extending and

I retracting the electrical utilities several times in a 1-g

environment. Factors included in the assessment are accessibility

and the necessity for manual resetting of the lines for retraction.

Table 4.3-1. Electrical Accommodations (GEO)

e1 MICA:. Witlrin K r

fall 644a1e

4

IM w fat W

6

elllMilln Sure "leash:" 1411.1111110" 4114m, anwnrn

e Ieea eletn011" n tam wleeert rWA Ipei V•4 wnenen h tr lu r.eN 1 L) ew4 to 1101 •loll t1 MAL to M •IM 4 sees m •.11 m er m r•a m nw,eeel sew h1 me, M ere M WHIM lamina" ^wlt,4w

eT

11N e w w nom .el. —M-1-19-111K ""' .elm Wrolm NU nl

20 10 70 20 70 20 to 20 20 70 140

1 1 20 2 9 2 1e .14 10 57 1 20 1 10 1 20 .69 10 60 117

2 1 20 2 9 0 1e 0 20 67 1 20 1 10 1 20 0.1 19 69 129

2 9 16 9 4 9 10 0 20 49 a 16 9 4 6 12 0 20 61 100

4 1 20 1 10 1 20 0 20 70 6 19 6 6 7 e .1e 19 46 Ile

6 1 20 7 0 le .14 10 e6 1 20 4 e 1 20 .64 11 69 114

6 1 20 1 10 1 20 0 20 70 6 6 7 6 0.1 19 49 119

7 1 20 6 7 2 16 ,14 10 66 1

J80

6 1 20 .64

1693

114

9 '1 20 e 6 9 19 0.0 20 66 6 9 1 20 0.2 129

.ones -

• ko it lo-1

4-6
ORIGINAL PW. - 7 9

OF POOR Q,:,;LIVY

Table 4.3-1. Electrical Accommodations (LEO)

EEE CTOICAL UYILIfIES ACCOMMODATION

REQUIREMENT N ON REQUIIEPIENT

Fo

N

C

E

RELIABILITY

SPACE

OALOTNNT

EASE Of

GROUND

OEPLOTHENT

S CMECROUT

SUITABILITY

TO LAUNDI

ENVIRONMENT

TOTAL

LOW

REQMT

Nlll ABllln

S ►ACE

O[rIOTMlIIi

IASE a

6ROUMO

OE► LOTMMT

L CNECSOUT

SUITABILITY

TO LAAAPI

ENVIRONMENT

TOTAL

NICK

NEW

TOTAL

LOW

AND

Him

RANK ►n RANK Pn O MR ►n ABNR ►n RANK ►ri RANI m NOW

P

T HAB POINTS

70

HAB POINTS

30

HAS POINTS

70

MAX ►TS

e0

MAR POINTS

7

MAB POINTY HAS POINTS

70

HAS ►n

50

MAB PTS

_

100

1 1 70 7 0 7 19 47 1 70 1 10 1 70 60 97

7 1 70 D 7 is 47 1 70 1 10 1 70 e0 97

7 a 15 e 4 e 10 79 a le e R 9 17 91 DO

1 70 1 10 1 70 eo a 1e e e 7 e 70 DO

D 1 70 a 7 7 le Re 1 70 t a 1 70 A e]

a 1 10 1 70 e0 a le a a 7

7 1 70 e

101

 le 60 1 70 e 1

e 1 70 le 441 .s 19 0 e 1 70 ^e w

o Suitability to Launch Environment

This parameter subjectively accounts for the degree of support

provided to the utilities during launch by the utilities

installation systems ( longerons, trays, clips, shoes). Electrical

utilities mounted on the longerons are beet, with trays second, and

the remaining concepts last. The points allocated are a judgmental

estimate of the relative difference betweeL the designs.

o Mass Implication

This parameter quantitatively accounts for the GEO transfer coat

implication of the utilities support system mass. The masses shown

represent the delta mass above that of the minimum value.

4-7

-
ORIGINAL PAGE P3

OF POOR QUA.M..

V

Table 4.3-3 compares the comparative capability of the designs to accommodate

the fluid utilities. The following clarification of the evaluation parameters

is presented.

o Minimum Rend Radius

A large 'And radius in the fluid lines is fudged to be better than e

small bend radius. A small bend radius has higher stress and takes

more force to fold.

o The notea for Tables 4.3-1 and 4.3-2 apply to the other parameters

shown.

Table 0.3-3. Fluid Utilities Accom^_odation

(GEO and CEO)

C

0

N

CE

P

T

FLUID UTILMIS ACCOMNIODATIOW

RELIABILITY

SPACE

DEr,,'/NENT

-TIAS1 OF

CROWD

DEPLOYMENT

A CHECKOUT

SUITABILITY

TO WINCH

ENVIRONMENT

NININUN

REND RADIUS

RANK ►TS ROHR I PTS ARK T

RADIUS

AS P

NM POINTS

20

NAB POINTS

!O

" POINTS MAX POINTS

20

MAX ►TS

70

1 1 70 I 1 10 1 ^ 90 12 10 e0

9 0 19 D a 3 1a 20 20 as

3 a 10 a 7 a Is 20 20 53

4 S 19 a 7 7 e 14 Aa

0 1 7o 3 9 1 70 4 10 DJ

e s 19 e 7 7 a 14 is 50

7 1 70 9 9 t 20 4 10 as

e 1 20 1 10 1 1 20 3.1 10 e0

Table 4.3-4 compares the versatility of the candidate structural designs to

use either aluminum or composite materials (graphite epoxy or metal matrix).

The judgemental evaluation is based upon the degree of static determinacy o

the structure. The maximum points are assigned to statically determinate

structures. A pure statically determinate structure can experience thermal

gradients between members with no loads incurred and no resistance to closure

dust prior to looking at the end of the deployment phase.

4-8
e

4-9

i

w ORIGINAL P.1'' r '

OF PO4 17^ (' „L'l'Y

Table 4.3-4. Structural Materials

Variation (LEO and GEO)

STRUCTURAL RRTIRIR1

VAR:ATI OP — to ro:NT1

caRan am RO:wn

1 1 70

9 1 90

e 0

7 1 10

e e e

a 1 70

7 1 70

e e e

Rerts t

K0W011TC MATERIALS OR

AIRRIWO

Tables 4 . 3-5 and 4.3-6 compare, for CEO and LEO platforms, respectively, the

candidate concept trusses accommodation of varied strength and stiffness

requirements. The three ranges of strength and stiffness requirements are

defined in Table 4 . 1-1, (Mayor Criteria used in the Selection Process). The

accommodation of strength and stiffness is described by packaging efficiency

and structure maser hence, these parameters are the basis for this

assessment. Clarification of each of the parameters is as follows:

o Packaging Efficiency

The packaging efficiency in the ratio of deployed length to stowed

length. This is a quantitative evaluation with linear distribution

of points between the maximum and minimum values. Consideration was

devoted to use of a volumetric efficiency term. The linee,r

efficiency is used because the study of packaging ( Section 1.4.3)

the concepts indicates that it is the most significant factor.

o Mass Implication

This parameter represents the estimated weight difference between

the concepts as used in the generic platform. A detailed breakdown

of the designs for the adopted stren gth and stiffness requirements

Is presented in Table 1.4-3. The concern here is the implication on

CEO transfer coat.

hiL= _ - ..
4-10

V

ORIGINAL PAGE IR

OF POOR QUALITY

Table 4.3-5. Strength and Stiffness Accommodation (CEO)

STRENGTH N T FFN SACC DATION

C

0

N

C

E

P

T

RE000TO STRENGTH

AND STIFFNESS

SUBTOTAL

30

ADOPTED

AND STIFFNESS

STNENGTN

SUBTOTAL

50

INCREASED

STI!!NES S

SUBTOTAL

20

TOTAL

ACROSS

ALL

RIGHTS.

100

PACKAGING

EFFICIENCY

(PI)

MASS

IMPLICATION

PACKAGING

EFFICIENCY

(PI)

MASS

IMPLICATION

PACKAGIIN

EFFICIENCY

(PI)

MASS

IIII.I CAT ION

►E rif

e

IASf PTS ►[ ►TS

0

NABS PTS ►[ ITf

e

NA5 ►ri

MX POINTS

10

MAX P016TS

20

MM POINTS

20

MAX POINTS

30

MAX POINTS

10

MAX POINTS

10

1 22 5 1.0 17 1 17 15 10 1.1 17 27 6 O 10 0 5 40

2 32 LO 1.2 10 20 21.6 17 1.3 10 32 B D 14 0 O

3 26 7 0,0 20 27 20 15 0.0 30 45 20 9 0 10 19 BE

4 20 7 0.2 19 20 20 15 0.7 22 37 10 6 10 0 6 00

5 25 0 0.9 13 19 25 20 0.1 29 49 25 10 3 3 13 BE

6 31 9 0.0 16 24 20 1 16 11.0 16 33 30 0 12 0 6 63

7 26 6 1.3 '^. 16 25 1 20 1.3 15 35 25 10 3 3 13 04

6 10 a O,g 1S 23 22 17 0.9 16 35 14 , la 0 7 05

MOTES: MASS DATA (KO M 10'J) BASED ON GENERIC PLATFORM (WITHOUT MECHANIZATION SYSTEM).

Table 4.3-6. Strength and Stiffness Accommodation (LEO)

C

STRENGTH AND STIFFNESS ACCOMMODATION

REDUCED STREW 4

AND STIFFNESS

ADOPTED STRENGTH

AND STIFFNESS

INCREASED

STIFFNESS

0

N

C

E

P

T

PACKAGINA

EFFICIENCY

V[

PACKAGING

EFFICIXNCT

►[

PACKAGING

EFFICIENCY

►[ TOTAL

ACROSS

ALL

RIGHTS

40

PE POINTS; PC MINTS P POINTS

MAXIMUM POINTS

10

NAXIMIM F71NTA

20

MAXIMUM POINTS

10

1 A2 5 Is 10 S 5 20

2 32 10 21.6 17 S 5

32

3 26 7 20 15 20 9 31

4 26 7 20 15 10 6 26

U25 6 25 20 25 10 36

e 31 0 20 13 10 0

30

7 25 6 25 20 25 10 30

6 28 a 22 17 14 7 J/

....:.^ ..1 „uw4:^a^awm-.tom' —
ORIGINAL PAGE 19

OF POOR QUALITY

Table 4.3-7 compares the candidate concepts' buiding-block platform

construction versatility. The following discussion clarifies the comparison

parameters.

Table 4.3-7. Platform Construction (LEO and GEO)

PLATFORM CONSTRUCTION

C EASE OF BEST MINIMUM

0 JOINING ACCOMMODATION DESIGN

N BUILDING OF PAYLOADS CONSTRAINTS—

C BLOCKS AND DOCKING RAIL

E DEPLOYMENT

P RAN_ POINTS RANR POIITIS POINTS TOTAL

T iUR. POINTS NAR. POINTS PAR. POINTS NAR. PTS

20 20 20 60

1 4 19 1 4 18 3.2 17 1 49

2 1 20 1 .20 2.5 18 59

3 a 6 a 6 6.0 10 22

4 4 14 4 Is 1.0 20 52

5 6 IQ 8 10 2.8 17 37

6 1 20 1 1 20 1 1.6 20 1 00

7 6 10 8 :O 2.8 17 37

a 1 20 1 20 1.8 10 58

o Ease of Joining Building-Blocks

This section ranks the ease and versatility of joining building blocks

together to form platforms of many configurations. Building blocks are joined

to each other or to other modules via the main housing or adapter. Rigid

square housings/adapters are the best, followed by rigid triangular shaper

Expanding triangular shapes, such as Concepts 5 and 7, pose difficulties

because of the change in dimension of the housings/adapters. Concept 3

offers the most difficulty because of the flat shape of the housing/

adapter, and because only two of the longerons are tied into the main

housing. The other two longerons must be tied into another structure

(subsequent to truss deployment) to maintain structural integrity.

o Beet Accommodation of Payloads r,nd Docking

Paloade/docking accommodations are prov-ided by the main housings and the

adapters. Adapters which do not change shape are judged superior to those

which expand (Concepts 5 and 7) or to those which unfold (Concept 3). Rigid

square main housings are beat for mounting interfaces, closely followed by

rigid triangular housings. Concepts 5 and 7 have expanding triangular

housings which pose obvious difficulties. The mounting of an interface on the

main housing of Concept 3 requires a deployable substructure.

4-11
C

ORIGINAL PAC.• i3

OF POOR QJALIl-f

o Minimum Design Constraints -Rail Deployment

The length of the deployment/guide rails poses some design

constraints. If the rail is to be folded for atowing, a longer rail

may require multiple folds. If the rail is to be moved into

operating position subsequent to a partial deployment of the

trues/payload, a longer rail requires a longer "partial deployment"

wY.ich in turn implies a longer 'partial deployment mechanism".

Finally, if a fixed rail system is to be used, a longer rail implies

that more of the orbiter payload bay is required for stowage.

4.4 COSTS FOR GENERIC PLATFORM (LEO AND GEO)

The grading of the eight building-block concepts in shown on Table 4.4-1.

Additional information is provided on Tables 1.4-4, and 1.4-5

The following notes are intended to explain the rationale behind the grading,

points allocation of Tables 4.4-1.

Table 4.4-1 presents the comparative costs determined for the candidate

concepts as used in the generic platform and sized for the adopted

atrength/atlffness requirements and adopted complement of utilities. The

Table 4.4-1. Costs for Generic Platform (LEO and GEO)

C

0

N

C

E

P

T

6 LAUNCH COST d)

A FAB.

COST

(SM)

TOTAL

A COST

FOR LEO

ISM)

EOUIV.

POINTS

FOR

A COST

!0R LEO

A OTV COST FOR

CEO TRANSFER ID

EVUIV.

POINTS

EON

A COST

A PRO. FOR GEO

LENGTH

IMF TCRS)

A

($M)

A

,u55

A

ISM1

MAX PTS

20

MAX PTS

zn

1 2.25 5,9 1.5 7." 11 1.1 9.7 12

2 1.5 3,9 4,3 8,2 10 `1.3 11.5 10

3 2.3 6.0 0.3 6.3 12 0.0 o, r) 20

4 0.75 1.8 0,5 2.3 17 0.7 6.2 14

5 0.0 0.0 0.0 0.0 20 0.0 0,0 20

6 2.1 5.5 2.3 7.8 11 1.0 8.8 12

7 0.0 0.0 1.3 1.3 18 l.J I1.5 10

8 1.5 3.9 3,0 6.8 12 1,2 10.9 11

NOTES:

m BASED ON SI.611 PER METER OF PACKAGED LCNGTN

USED ON $8.8K PER KG

4-12

i

r

7

l

F

0
^G

ORIGINAL PAGE iz^

OF POOR QUALITY

difference in packaged lengths is determined from the packaging arrangements

shown in Section 1.4.3. The recurring fabrication costs are extracted from

Table 1.4-5. The use of the unit Shuttle launch cost of $2.614/meter is oased

upon a FY 1981 launch cost of $48M divided by the 18.3 m bay length. A

significant reservation on the use of this value is that for a dedicated

mission the $48M cost is incurred regardless of the length of the bay actually

used.

The differences in platform mesa are obtained from the data shown on Table

1.4-4. A significant reservation pertaining to the OTV transfer costs is that

it is representative of only the adopted etrength/stiffness requirements and

for the generic platform. For this reason, and the reservations noted above,

the point allotment of the coat criteria on the final summation charts is no

more than 40. Further, the maximum coat differences shown are very small

compared to the total system cost.

4.5 TNEHMAL STABILITY OF STRUCTURAL CONCEPT

The grading of the eight concepts for thermal stability is shown in Table

4.5-1. This table is a summary of the thermal data shown on Table 1.4-2.

Table 4.5-1. Thermal Stability of Structural Ccn.:ept

PT

d

W

DIAGONAL

PRESENT

LONGERON TENF. ('G)

AT

('t)

FIGURE OF

SUN SIDE SHADE SIDE NERIT,d/(AT1 POINTS

1 1.6 YES 21.6 -13.5 35.1 o.o46 II

2 1.3 YES 20.7 -11.6 32.3 0.)53 12

No 24.0 4.8 19.2

3 3.0 NO 24.0 2.% 21.3 0.140 20

4 1.7 YES 21.7 -7.4 29.1 0.058 12

5 2.5 YES 23.8 -36.o 59.8 o.o4o II

(TENSION

STRAP)

6 1.3 YES 22.5 -22.6 45.1 0.029 10

7 2.5 YES 23,6 -56.7 80.3 0.031 10

8 1.1 YES 22.5 -17.5 40.0 O.G31 10

4.6 METEOROID IMPACT SUITABILITY OF STRUCTURAL CONCEPT

This section discusses the issue of potential meteoroid impact and structural

survival. The data in Table 4 . 6-1 were derived using the meteoroid model

stipulated in Reference 6. The model is sufficiently accurate for the GEO

environment. Man-made debris was less critical (Reference 15),

The size of the meteoroid particles and associated probabilities are shown for

two sizes of platform and for a 10-year exposure. The projected area applies

to the totality of trues members. Little to no recent information exists

pertinent to the size of holes resulting from the meteoroid strike,

particularly for graphite composites. From discussions with Materials

personnel and reviews of meteoroid impact damage ( Reference 7), it is estimated that the hole size may be 2 to 4 times the diameter of the particle.

Considering the low levels of stress and low number of cycles associated

4-13
w

C

ORIGINAL P^.GE IY

OF POOR QUALITY

with RCS systems, and the negligible impact of the hole on local or Euler

stability, it is likely that the structural damage will be acceptable.

However, since this needs to be verified, redundancy in the structural

design is an advantage; hence the grading of the eight concepts is based

on that consideration and is shown in Table 4.6-2.

Table 4.6-1. Probability of Meteoroid Damage

GENERIC PLATFORM (220 m2)

PROBABILITY

OF MIT IN METEOROID POTENTIAL

10 YR DIAMETER HOLE DIA.

(2) (Cm) (cm)

I 0.60 1-1/4 TO 2-1/2

2 0.48 1 TO 2

5 0.38 3/4 TO 1-1/2

SMALLER PLATFORM (70 m2)

PROBABILITY

OF HIT IN METEOROID POTENTIAL

10 YR DIAMETER HOLE DIA,

(21 Cm) Cm

1 0.42 3/4 TO 1-1/2

2 0.36 3/4 TO 1-1/2

5 0.30 1/2 TO 1-1/4 

Table 4.6-2. Meteoroid Impact Suitability

CONCEPT

METEOROID IMPACT

SUITABILITY-20 POINTS

RANK POINTS

1 5 10

2 4 16

3 1 20

4 5 10

5 5 to

6 1 20

7 5 10

8 1 20

4-14

._._^w.. .. ^.. ..:u.........•..r^ ^-:.^.: y.Ee^t.... - .alt.a^..^7 . L. ., _acfamm. u^i -"- .! 
4 ORIGINAL PA-E: KJ

OF POOR QOAL"

4.7 RELIABILITY OF DEPLOYMENT OF BUILDING BLOCK

This section compares the candidate building- blocks for reliability of

deployment based on the parameters which are explained in detail below. The

grading of the eight concepts is listed in Table 4.7-1.

o Basic Trues Based on Number of Joints.

This evaluation is based on the number of Joints in the length of trues

required to build the generic platform. The Joints included are eliding

Joints in diagonals and battens, and folding / rotating Joints in the longerone

and pyramidal members. There is an inverse linear relationship between the

number of points and the number of points awarded. Table 4.7•-2 describes in

detail the numbers of Joints for each of the eight concepts.

o Basic Truss Based on Complexity.

This is an assessment based on the type of Joints/ movements used in deploying

the basic truss structure. Sliding ,joints in diagonals are fudged to be more

complex than folding Joints. The I-section eliding battens of Concepts 5 and

t 7 are ,judged to be the most complex.

Table 4.7-1. Reliability of Building-Block Deployment

THERMAL

C BASIC TRIMS EFFECTS— THERMAL DOCKING

0 BASED ON BASIC TRUSS GRAPHITE EFFECTS— PONT

N NUMBER OF BASED ON COMPOSITE ALUMINUM HOUSI146 ADAPTER SUPPORT

C JOINTS COMLEXI TY TRUSS TRUSS STRUCTURE STRUCTURE STRUCTURE

P

0. Oi

T DINT ►i1 RANK PTS RANK PTS RANK I PTS RANK PTS RINK PTS RANK PTS TOTAL

NAB POINTS PAX POINTS MAX POINTS MBX POINTS MAX POINTS MAX POINTS MAX POINTS MAX POINTS

10 10 15 5 20 20 10 90

9 2 9 1 15 1 5 1 20 1 20 1 10 97

 2 9 0 12 6 4 1 20 1 20 1 10 82

 1 10 7 12 7 3 1 20 8 12 8 4 69

 2 8 1 15 1 5 1 20 1 20 1 10 88 15 g 5 7 1.11, 7 3 '1 Q 8 10 0 8 80 6660 2 8 1 15 1 5 1 20 1 20 1 10 85

6 5 1 15 1 5 7 8 B 18 6 6 63

8 S 1 15 1 S 1 20 1 20 1 10 80

ra

4-15

^ 

^fdlroaa5._-

jj

w.».....i , .•w. Att..:.x .x-..^: 

.......j.. - ^iL r'_..u-._v ..J. ^:.^ r.xr...

_.;. .»e.. 

.^. 

i-m se ...t.+lLtJ:T.%
ORIGINAL PA ." t .

OF POOR QUALri'T

Table 4.7-2. Number of Joints in Basic Structure for Generic Platform

`M

CONCEPT

NUMBER

Of

BAYS

NUMBER Of

LOMGERON

FOLDED

JOINTS

NUMBER Of

DIAGONAL

TELESCOPING

JOINTS

NJASER Of

DIACONAL

FOLDED

JOINTS

NUMBER Of

BATTEN

TELESCOPING

JOINTS

NUMBER Of

PYRAMIDAL

JOINTS

TOTAL

JOINTS

1 212 636 212 _ —. _ 648

2 270 1080 270 — _ _ 1360

3 114 466 _ _ _ 612 1388

4 207 621 621 — ._ _ 1242

6 122 366 — — 366 732

6 216 830 630 — — _. 1880

7 122 308 386 732 366 — 1464

8 104 1472 736 — — 2208

NOTES: COUNT DOES NOT INCLUDE BASIC CLEVIS JOINTS BECAUSE THEY ARE SIGNIFICANTLY LESS COMPLEX

THAN FOLDING AND TELESCOPING JOINTS. THE NUNEMS ARE BASED ON APPROXIMATELY 140 METERS

Of TRUSS. WHICH IS THE LENGTH REQUIRED TO BUILD THE GENERIC PLATFORM.

o Thermal Effects, Graphite Composite Truss or Aluminum Truss

These parameters account for the reduced reliability of deployment inherent in

the structures that are indeterminate for the materials shown. The values are

judgmental between the maximums and minimum shown.

o Housing Structure

Concepts 1, 2, 3, 4, 6, and 8 are ranked equal because they are rigid with no

mechanisms required. Concepts 5 and 7 are ranked last, because they require

a mechanism for lateral extension.

o Adapter Structure

Concept 1, 2, 4, 6, and 8 are ranked first because they are rigid structures

which require no mechanisms. Concepts 5 and 7 are judged as having less reliability because they are expanded by a mechanism. Concept 3 is ranked last

because it requires a separate mechani3m to unfoli it.

o Docking Port Structure 

A docking port interface mounted to a rigid main housing is the most reliable. 

Concepts 5 and 7 have expanding main housings which degrade the reliability of

the interface. Concept 3 requires a separate substructure to be deployed to 

obtain a docking port interface. This requires additional mechanization, 

which is the reason for its being ranked 8th.

4-16

..

., ^^..;.Y.Naxsn.3iNrr:as,-L..we ..,...ms
OF POOR QUA;.fry

4.8 PREDICTABILITY OF PERFORMANCE OF STRUCTURAL CONCEPTS

The grading for the eight concepts is listed in Table 4.8-1.

The features of the eight structural concepts which affect the accurate

prediction of structural performance are:

• A determinate structure is better than an indeterminate structure

for analytical purposes

o The difficulty of maintaining the tensioi in "E" braced structures

which is essential to performance predictability

• The disadvantages of designs with offset load paths at the points

While NASTRAN analysis and development testing during a program can deal with

these effects, Table 4.8-1 Judgmentally and qualitativaly account¢ for these

additional requirements.

Table 4.8-1. Predictability of Performance

COMPOSITE MATERIALS ALUMINUM

C

0

N

C

E

PREDICTION

OF

INTERNAL

LOADS

PREOICTI ON

OF

EFFECTIVE

STIFFNESS

TOTAL

PREDICTION

OF

INTERNAL

DADS

PREDICTION

OF

EFFECTIVE

STIfFN S

TOTAL

TOTAL

FOR BOTH

RANK POINri MANN T POINTS RANK I POINTS RANK POINTS NATIRIALS P

T PAR POINTS

10

HAS POINTS

10

AAR PTS

20

PAR POINTS

10

AAR POINTS

to

PAR PTS

20

1 1 10 1 10 20 1 10 1 10 20 40

2 O 9 5 5 14 8 9 7 5 14 28

0 8 0 1 10 1 le 7 4 1 10 14 20

4 1 10 1 10 20 1 10 1 10 20 40

5 8 8 7 8 14 8 4 8 2 8 20

0 1 10 1 10 20 1 10 1 10 20 40

7 1 10 1 10 20 1 10 1 10 20 40

8 1 IO 1 10 20 1 8 1 10 18 00

4-17
ORiwNAL f AG_ v,-

OF POOR QUAL;'IY

4.9 ORBITER INTEGRATION SUITABILITY

This section compares the candidate designs in regard to their orbiter

integration suitability.

The grading of the eight concepts is shown on Table 4.9-1.

Table 4.9-1. Orbiter Integration Suitability

C IT HOUSING m CASE OF m EASE OE ul CRAOLE COIe LEXITY,

0 LAUNCH PACKAGING VACRAGIiW STIIUCTUR[ PACINGEO

N ENVIRONMENT INTO INTO A MASS

I

CONFIGURATION SPARE

C SUITABILITY (,AEITTR ORBITER (KC • 10' 1 ) TO DEPLOYS? VOLUME

V

RANR POINTS RANR POINTS RANK PhfNTS POINTS RANK POINTS PWr POINTS TOTAL

T MAX POINTS MAX POINTS MAX POINTS MAXPOINTS MAX POINTS MAX POINTS MAX ITS

10 10 10 0 SO 10 60

1 1 10 a fl 1 10

F2.3

a 3 a 7 s 48

2 1 10 O 6 1 10 7 s 7 4 7 47

3 a 3 4 7 1 10 O 6 9 1 10 41

4 1 10 3 a 1 10 1.0 a 3 a 7 s 49

6 1 10 1 10 6 a 0 10 1 10 3 a so

6 1 10 a a 1 10 1.6 a s 7 4 7 46

7 1 10 1 10 0 a 0 10 1 10 R a sa

a 1 10 6 a 1 10 1.4 7 S 7 4 7 47

NOTES; 'm APPLICABLE TO GENERIC PLATFORM

m APPLICABLE TO PLATFORM SMALLER THAN aEMERIC

m BASIC BUILDING BLOCK CONCEPT

A discussion of each of the design parameters, listed above, is as follows:

o Housing Launch Environment Suitability

This parameter judgmentally accounts for the relative capability of the

candidate housing designs to sustain the launch /inertia induced loads and also

to provide ( in conjunction with the cradle structure) a minimum natural

frequency above that of the orbiter ( say, 10 Hz).

o Ease of Packaging into Orbiter (Generic Platforms)

The rankings of this section are based on the studies and drawings made for

i the generic platform only. Factors which influence the package are:

t

o Packaging ratio o expanded length of a truss

stowed length

o The shape of the truss section

0 The size of the truss section

4-18

V

f 

t1M... a .. ^ • - ^ ^ x^.. a_....._....m.+W ., ..., __.. .. ,m
Q

Concepts 5 and 7 gain first due to their high packaging ratio.

Triangular-shaped trusses generally fit together in a circle (such as the

orbiter payload bay) better than square-shaped trusses. If the packaging can

be arranged so that the building blocks fit across the payload bay instead of 

along the longitudinal axis, it is an advantage. This is reflected in the

ranking of Concept 3.

o Ease of Packaging into Orbiter (Smaller than Generic)

No drawings of packaging "smaller than generic" platforms into the orbiter

were made. As the platforms become smaller, so does the differential between

them, until the point is reached for verT small platforms when there is

generally little significant advantage oT one concept over another. It is

fudged that small platforms can probably be packaged across the width r.° the

payload bay. Concepts 5 and 7 are ranked slightly lower because of the ne,d

for clearance around the expanding main hrusings.

o Cradle Structure Delta Weight.

This quantitative assessment describes the delta weight additional to each

concept as packaged in the orbiter. The delta weight represents cradle

structure, trunnions and, where required, reinforcement of the housing

structure. No distinction is made between LEO and GEO designs at this stage

of the investigation.

o Complexity of Configuration—Packaged to Depluyed 

This evaluation addresses the difficulty and complexity of moving the building

blocks from the packaged configuration to the final deployed configurations.

For each concept, a count was made of the number of mechanisms/movements

required to unlock, rotate, relock each building-block housing and adapter,

and the unfolding of the deployment/guide rai'_o. The concept with the least

number of mechanisms/movements, etc., was awarded first place with the other

concepts graded accordingly. This assessment is applicable only to the

generic platform.

o Spare Volume

In the length of the orbiter payload bay which is occupied by the packaged

generic platform, a certain percentage of t:ie volume is available for other

purposes. The concept with the maximun space is awarded first place, and the

other concepts are ranked accordingly. This assessment applies only to the

generic platform.

4-19 

e 
Q ORIGINAL PAGE I$

OF POOR QUALITY

4.10 SUMMARY OF POINTS AND GRADING

The results of the points allocation and grading of Sections 4.3 through 4.9

are presented in five tables:

• Table 4.10-1, Total of Normalized Points (LEO)

• Table 4.10-2, Total of Normalized Pointe (GEO)

• Table 4.10-3, Total of Normalised Points (LEO)-Sensitivity Trade

• Table 4.10-4, Total of Normalized Pointe (GEO)-Sensitivity Trade

• Table 4.10-5, Summary of Pointe (LEO and GEO)

Table 4.10-1 (LEO) presents the total points in each of the seven mayor

selection criteria for each building-block concept. The point values shown

for each concept are obtained from the total point value determined in the

individual preceding criteria ahests multiplied by the appropriate factor

to be compatible with the maximum point allocations shown on this chart for

each criterion. For example, for Concept 1, the total points for "Reliability

of Building-Block Deployment" (obtained from Table 4.7-1) is 87 out of a total

poasible 90 points. The 97 points shown on Table 4.10-1 is obtained by 100/90

x 87 - 96.66 or 97. The "Design Versatility" criterion includes the points

obtained from Tables 4.3-2, 4.3- 3, 4.3-4, 4.3- 6 and 4.3-7, normalised in the

same manner as shown above.

Table 4.10-1. Total of Normalized Points (LEO)

CONCEPT ]POINTS

(2)

 COST

(3)

THERMAL

STABILITY

(A)

METEOROID

IMPACT

SUITABILITY

(S)

RELIABILITY

(6)

PREDICTABILITY

(7)

ORBITER

INTEGRATION TOTAL

NAB POINTS

40

MAX POINTS

70

NAZI POINTS

40

PAX POINTS

100

MAX POINTS

20

MAX POINTS

80

NAB POINTS

380

I 87 ^ 77

20

11

17

70 •

1 32

07 3

91

20 3

14

46

47 J

303

2 94 310

3 60 74 20 3 40 3 77 15 41 277

4 78 34 3 12 20 96 3 20 3 49 309 m

5 80 40 3 11 20 97 10 BB 3 784

6 83 77 30 40 J 94 3 20 ./ 46 315

7 84 3e 3 30 20 70 20 3 56 / 2B0

8 Be It 24 10 40 3 89 18 47 3 316 G)

NOTESI

CIRCLED NUMBERS IN TOTALS COLUMN DENOTE RANKING.

J FOR TOP 3 VALUES IN EACH CATEGORY

S' M

5

, 4-20

.j
^r

ORIGINAL PAGL 15

OF POOR QUALITY

The assignment of maximum points of 100 to "design versatility" and

"reliability" represents the greater importance of these criteria. Phe

rationale for reduced emphasis on cost in discussed in Section 4.4. It is

important to note that this generic requirements point allocation would be

different for a very specific mission, depending on constraints. for example,

suppose the available orbiter length is a constraint, or suppose subsequent

test data indicate meteoroid impact to be more critical than expected. These

parameters and/or criteria take on an increased significance. The total

points are shown to the right for each concept, with Concepts 8, 6, 2, and 4

representing the most suitable design for a LEO platform.

Table 4.10-2 ( GEO) is compiled in the same fashion as Table 4.10-1. The table

indicates Concepts 6, 8, 4. 3, and 2 to be most suitable.

Table 4.10-2. Total of Normalized Points (GE01

CONCEPT

11)

DES ION

VERSATILITY

U)

COST

(3)

THERMAL

STABILITY

(L)

METEOR0ID

IMACT

SUITABILITY

(5)

NELIAl I L TV

(6)

PREDICTAEI L TT

(7)

ORBI TEA

INTEGRATION TOTAL

MAR POINTS

100

PAR POINTS

40

MAR POINTS

70

111E POINTS

40

MAX POINTS

100

MAR POINTS

70

MAX POINTS

e0

MAR POINTS

380

1 77 74 11 30 97 3 70 3 4e 790

90 3 30 12 37 91 14 47 V 307

3 09 40 3

_

70 3 40 3 77 10 41 307

4 78 78 3 17 70 90 3 70 3 49 303

0 70 40 3 11 70 87 10 0e ! 779

9

RO 3 74 10 40 3 94 ! 70 3 40 314 m

7 1 70 70 1 10 1

70 170 70 3 Be 3

771

6 83 ! 77 10 40 3 89 18 47 3 307

NOTES:

CIRCLCO NUMBERS IN TOTALS COLUMN DENDTE RANKING.

! FOR TOP 3 VALUES IN EACH CATEGORY

Tables 4 . 10-3 and 4 describe the corresponding sensitivity data derived as

described in Section 4.2. The results are summarized in Table 4.10-5. The

sensitivity study data, however, is not considered on au OlL31 basis with that

of the baseline.

4-21
fll

DESIGN

VERSATILITY

(2)

COST

(1)

TNERNAL

STABILITY

(4)

METEOROID

IWACr

SUITABILITY

(5)

RELIABILITY

(6)

PREDICT.

ABILITY

171

ORBITER

INTEGRATION TOTAL

CONCEPT MAX POINTS

100

MAX POINTS

40

MAX POINTS

20

MAX POINTS

40

MAN POINTS

1

100

MAX POINTS

20

MAX POINTS

00

MAX POINTS

380

1 58 6 2 20 93 ! 20 4 44 243

2 76 ! 0 4 32 BB 14 44 255

3 05 40 ! 20 ! 40 ! 71 15 36 287 m

4 69 ! is ! 5 20 91 ! 20 ! 47 ! 270

5 88 40 ! 1 20 62 10 56 ! 255

6 73 ! 10 0 40 ! 89 ! 20 ! 43 275

7 61 0 0 20 64 1 20 ! 56 ! 221

8 89 ! 4 0 40 ! 83 18 44 258

NOTES:

Circled numbers in totals column denote ranLing.

! For top 3 values fO cach category.

4-22

ORIGINAL F . Mr'-' V

OF POOR QUALIT3

Table 4.10-3. Total of Normalized Points (LEO) —Sensitivity Trade

(11

DESIGN

VERSATILITY

Iii

COST

111

THERMAL

STAULI*Y

(4)

ME TC ORO10

IMPACT

SUI TANI LI"

(S)

Pi U -

All L TT

161

PAID.,

All L TS'

lil

ORII TEN

INTEGRATION TOTAL

CONCEPT MAX POINTS

100

MR POINTS

40

NAX POINTS

20

NAR POINTS

40

MX POINTS

100

MAX POINTS

20

MAX POINTS

80

NAR POINTS

380

1 75 4 ., 20 93 ! 20 ! 44 258

2 90 ! 0 4 32 as 14 259

3 53 a 20 ! 40 ! 71 15

_

36 243

4 70 20 ! 5 20

20

91 ! 20 !

10

47 ! 281

5 78 40 ! 1 52 56 ! 265

6 79 0 0 40 ! 89 4 20 ! 43 271

7 80 ! 34 ! n 20 64 ' 20 ! 56 ! 274

8 82 ! 6 0 40 ! 53 le 44 273

worts,

Circled numbers 10 total• coluae denote rant IO;,,

! For top 3 values in each category.

Table 4.10-4. Total of Normalized Points (GEO) —Sensitivity Trade

n
r`

ORIGINAL I ,;',_ iI

a

OF f)OoR QIIAI_ITY

Table 4.10-5 summarizes the data obtained from the precedin g tables. The

number of criteria for each concept that are within the top three is indicated.

The major result is that there in little difference in the total points

obtained for the top four designs. Hence, the decision between these is

judgemental, but based on the knowledge gained from the details of the study

and tabulation of the selection process data.

Table 4.10-5. Summary of Points and Grading,

LEO d GEO

LEO PLATFORM GEO PLATFORM

BAS{LINE

DATA

SENSITIVITY

DATA

NO. OF CRITERIA

IN TOPI

VASELINE

:}TA_

SENS IT IV ITV

DATA

NO. OF CRITERIA

IN TOP

CONCEPT PACE POINTS PLACE POINTS EASE SINE. PIJ.CE PORTS PLACE POINTS RASE SENS.

a 1 310 3 773 3 1 2 2 307 4 Tee 3 2

0 7 315 4 271 3 3 1 314 2 270 4 4

4 4 30N 1 291 3 4 3 303 3 270 3 S

2 3 J10 0 269 1 4 1302 0 200 2 1

4.11 CONCEPT SELECTION

A review of the selection process of Sections 4.1 through 4.10 and Table 4.10-5

leads to the following general conclusions:

• Concept selection between the first four designs is judgmental, but

based on knowledge obtained f om selection process study.

• Specific mission : rf zica 1 r%qui rem .. • t 4 wt; i impose special emphasis

on particular pm...!, mter.3 1.L/ oncept selection.

• No clear concept choice distinction for LEO or CEO application is

represented by the data.

• The major difference in concept sole^tion will result from extent of

payload power and data requirements.

• No one design is beet on the basis of satisfaction of every

requirement - improvement toward one requirement rlmost always

rcaults in degradation uf another requirement. 

• The thrust of the program for the next 4 years should be to resolve

the major design & technology issues pertinent to ueployable

platform systems.

4-23
CONCEPT

OJMNAI. FAG[ 19

OF POOR QUALITY

o The major design and technology issues are essentially the same

across the candidate iesigns.

o The open-square shape of Concept 8 has best growth potential for

.cilities since it can accommodate in trays up to twice the baseline

study requirements (increased utilities requirements typical with

program maturity).

o For significantly reduced utilities requirements (mountable on

Longerons), Concept 6 is adequate and is a simpler structural

design than Conc^pt 8.

o Concepts 8 and 6 together present the most promising combinations

of top four designs (Figure 4.11-1).

Figure 4.11-1. Concepts 6 and 8

Subsequont to the identification of Concepts 6 and 8 as the two designs best

suited for LEO/GEO platforms, it: became evident that the best features of

both cau be combined into one. This concept, known as 6A, i.s shown in

Figure 4.il-2, and further discussed as follows:

o Concept 6A is the best utilization of the advantages of Concept 8

(utilities accommodation) and Concept 6 (structural simplicity).

o Con,aj.c 6A permits full accommodation of the adopted utilities

requirement with a 0.5-meter-wide • ray. The longerons are

folded at 30°.

o Concept 6A has an acceptable packaging efficiency of 18 for the

adopted requirements, and a 1.26-m-deep truss. lncreased packaging .,fficiency is achievabL2 with increased truss depth and

reduced loads.

4-24

-I ^:.
• LOPIGERONS WRH CENTER HINGE JOINTS

-TELESCOPING DIAGONAU

*NO OFFSET LOADS AT JOINT

.BATTENS CROSS BRACED BY TENSION STRAPS III

TRUSS.

ADAPTER

OA

0 0 1

O

^^ 9o;

e

END VFW FOLDED

DEPLOYMENT RAILS

LKnLTES INSTALLATION

DETAIL OF JOINT

f

i

ORIGINAL PACE I3

OF PGOR QUALITY

i

o The technology development of Concept 6A is applicable to a

family of designs ranging from Concept 6A to Concept 6.

• ALL MEMKRS ARE ROUND TUBES RUIN HCUSING

Figure 4.11-2. Concept 6A

l

4-25
SPACE STATION

CONCEPT

Q ORIGINAL r,10I^ 09

OF POOL QUALnY

5. DEPLOYABLE VOLUME ENCLOSURES

This section describes several applications of deployable volumes for a

typical Space Station (Figure 5.0-1). The three specific applications are

habitat modules, an orbit transfer vehicle (OTV) hangar, and crew transfer

tunnels.

HABITATS

CREW TRANSFER

TUNNEL

im

Figure 6.0-1

w

.y

OTV

HANGAR

Deplo^;able Volume Enclosures

Current designs for Space Station modules typically visualize rigid bodies

which are approximately the same size as the orbiter payload bay. With the

advent of deployable volumes there arises the poasiblity of launching much

larger modules without the problems associated with assembly in apace. This

section examines these possibilities and demonstrates that in many cases they

are more than dust possibilities - they are emi ; aently feasible.

5.1 HABITAT MODULES

5.1.1 Habitat Requirements

The requirements for the design of the habitat module are:

o A life of 10 to 20 years in LEO.

o Compatiblity with the orbiter for transportation to LEO.

o Compatiblity with existing designs and concepts for manned space

stations.

5-1
• Accommodation of tree of up to 8 for 90 days + 90 days emergency.

• Normal pressurization of 14.7 psi (limit); 8 psi for an emergency.

• Provision of two separate prrasurized compartments.

o Provision two routes for ingresa/agrees.

• Automatic deployment to maximum extent feasible, EVA assist if

advantageous.

• Incorporation of all the equipment normally associated with habitat

modules, including lift support equipment, command control center etc.

• Provision of two docking systems.

o Minimum of one airlock.

o Exterior mounting of equipment such as fuel cells, toxic items etc.

o Radiation protection of 0.50 gram/cm 2 (1.0 lb/ft2).

o Adequate meteoroid protection (discussed suLaequently)•

5.1.2 Radiation/Meteoroid Shie l ding Review

The suitability of a potential habitat wall design concept (gratuitously

furnished by Goodyear Aircraft Corp.) for radiation shielding has been evaluated (Figure 5.1-1). The basic 2.5-cm-thick design with a foam density of

0.032 gm/cc 3 is adequate for precluding skin damage, but not for protection

of the astronauts' eyes. An additional 0.24 gm/cc 2 needs to be provided by

either increasing the bladder thickness, foam thickness or density, a combination of the foregoing, or other means. The implication on packaging needs

to be assessed (in subsequent studies).

The associated probabilities of meteoroid puncture of the inner wall using

two different foam densities are shown. These data are based upon the

meteoroid model specified in Reference 6 and the man-made debris model of

Reference 15. The analyses assume an effective stopping power of 15 for

the foam, i.e., the foam is as effective as 15 times the same mass per unit

area of a sheet of aluminum. This information was obtained from documented

tests performed by Goodyear, and is consistent with predictions by Rockwell

researchers.

The probabilities shown indicate that with development of a leak-detection

system and repair capability, the inflatable wall design shown can be suitable

insofar as meteoroid impact is concerned (propagation of the puncture is not

expected to occur).

i

5-2
ORIGINAL PA(,k Q3

ki

OF POOR QUALITY

H

ELASTIC RECOVERY

/OLYUAETNANE aTIA COVEA AND

IOW TNEMIL COATING

ETWCTUMI

LAVIR

rAISSU"

MADDER

A !LANE

PAR 10

DETAIL WEIGHTS OF EXPANDABLE MATERIALS•

CONSTRUCTION ( •/u

• ALUMINUM INNER LAYER 0.001

• ADHESIVE 0.008

• PRESSURE BLADDER 0.080

• ACRES Iva 0.008

• STRUCTURAL LAYER 0.031

• TABLAN INTERLOCKING LAYER

AND ADHESIVE 0.094

• WLYURETRANE IOW 0.080

• OADHESIVE R 0.006

• OUTER COVER AND COATING 0 0

TOTAL 0.783

FOR PROJECTED AREA OF IOD m2

0.016 GM/CM3 FOAM (2.5 cm THICK)

•1% PROBABILITY OF PUNCTURE 

IN 3 DAYS

•ON AVERAGE, ONE PUNCTURE

EVERY 9 MONTHS TO BE REPAIRED

0.032 GM/CM3 FOAM (2.5 cm THICK)

*I% PROBABILITY OF PUNCTURE IN

27 DAYS

•ON AVERAGE, ONE PUNCTURE IN

7.5 YEARS TO BE REPAIRED

ADDITIONAL

ERIAL REQUI

TO SATISFY

u+

*Courtesy of Goodyear Aerospace Ifor Expandable Airlock Technology Experlment)

Figure 5.1-1. Habitat Radiation/Meteoroid Implications

5.1.3 Habitat Design Approach

A review of the habitat requirements and logistics associated with delivery

of habitats to LEO resulted in the following design approach.

• The habitat module should be large as compared to conventional

modules to significantly offset the reduced cost and higher

reliability of a conventional design.

• For habitats with inflatables, use combination of hard (fixed)

structure and inflatables.

• Design to accommodate equipment in its correct locations (on

the hard structure) during Shuttle launch to minimize work/

rearrangement or orbit.

• Separate the c:Iev quarters from the equipment not in regular

use with placement of heavy equipment on hard structure and

crew quarters it the deployable structure.

• Divide the crew quarters into large volumes for communal use

and into smaller private volumes.

• Build radiators into the exterior of the hard structure.

• Provide capability to repair inflatables from the inside.

• A meteoroid bumper is desirable for inflatables.

5-3
O^ 3

IIJ/^'/^f/

MI DU

12.7M !- T

HARRAT OTV

HANGAR ^.J

—11.JM--^

10.7M

J_

—^EE

0 V

DU

Oi{G{NAl PAGE I5

OF PpOR QUAL"

o Utilize the crew inside the Habitat Module for relocation of minor

equipment, final stages of deployment such as locking points and

removal of temporary deployment mecnanisms.

5.1.4 Candidate concepts for Habitat Modules

Eight concepts for deployable Habitat Modules were considered ( Figure 5.1-2).

Three were selected for further study ( Section 5.1.5).

•DE nVA LTIONS SHOWN SHADED

-i 0

) 4MDIA

^17.RM DEPLOYED i

10.9M STOWED

1

© 4M DIA

MET

RIGID

FLOOR

11. Im STOWED T

3.05M

O

II

I S4 DLA O

SECTION 4.-II.6M DIA^^—

FUQO LENGTH

TORUS

IIAML I UA DLA

^17.

-^{ f^-3.09A D1U

AM DIA

T__—.

^Iom— f

Figure 5.1-2. Candidate Deployable Habitat Module Concepts

Concept 1 expands axially from a central airlock/docking port and two hard

decks on which equipment is mounted. The stowed module occupies most of the

orbiter bay and the deployed volume is minimal. The inflatable sections are

supported by light deployable internal structures.

Concept 2 is a derivative of Concept 1 but has spherical end bulkheads.

Concept 3 has a rigid central floor which acts also as a strongback for

launch. There are two docking ports and an airlock. The two inflatable

eectioLS are shown as approximately 4.5 meters in diameter but can be larger.

They can be subdivided into wardroom, sleeping quarters, etc., as desired.

Concept. 4 consists of a nearly " standard" rigid habitat module with an

inflatable torus attached. It has excellent capability for mounting

equipment, good radia`or area and a large inflatable volume for crew

quarters. The arranr;ement for airlocks and docking is the same as for

"standard" rigid habitat modules.

5-4
0 e,niaiu,^ rw+

66^^

/ r

ORIGINAL FACE 19

OF POOR QUALfTY

Concept 5 is a "etaudard" rigid habitat module with two deployable hard

sections. The deployable sections can be sealed either by large internal bags

or by sealing around individual points ( or both). The volumes available for

equipment and crew quarters are excellent. This concept has the further

advantages of all- hard structure, large radiator area, standard decking, and

no unusual structures.

Concept 6 consists of Concept 4 with the addition of a hard deployable shell

which serves as a meteoroid bumper for the habitat module and as an OTV

hanger. The system is deployable and is packaged for a single launch.

Concept 7 is a rigid module with two inflatable sections deployed from it.

It is somawLat similar to Concept 4 but the deployed volume is probably not

as useful.

Concept 8 is a variation of Concept 7, as shown.

5.1.5 preferred Concepts

The three preferred concepts are shown in Figure 5.1-3. The drawings are

contained in Volume II. These concepts were selected for further investigation

for the following reasons;

o Representation of a wide variety of designs

o A good ratio of deployed volume / stowed volume is provided

O+mrcnau sKw .mr vuw uo+w+u arnn. ^'O'0j ^ aROreu

sna,cnr

i^S4X,

n

Qi rv.o rKwr.uu veus

Figure 5.1-3. Preferred

Habitat Module Concepts

5-5
• INFLATABLE WITH A

STRONG LACK

• AUTOMATIC DEPLOYMENT

• VOLUME STOWED 176 M.

• VOLUME DEPLOYED

LIVING 216

STOWAGE

UNPRESSURIZED 16

PRESSURIZED 12

AIRLOCKS , a

212 M3 1

12.6TM

RADIATORS a COVERS

FOR STOWED INF ATABLES

AIRLOCK

...., .....^ STRONGLACK

i

ORMINAL PA^.}" r,

OF POOR QI)ALfN

• High growth potential

• Adequate hard structure for mounting equipment

• Safety (redundant volumes available)

• The crew quarters can be subdivided into large and small rooms as

desired

• No technology development of an unusually difficult nature is required

5.1.5.1 Habitat Concept 3

Concept 3 ( Figure 5.1-4) consists essentially of an inflatable shell mounted on

each aide of a strongback. The 11trongback serves as a launch cradle,

(contains trunnions and keel fitting), a mounting platform for equipment, and

as a structural support for the airlocks and docking systems.

Figure 5.1-4. Habitat Module (Concept 3)

5-6
OPTIONAL

METEOROID

SHIELD

LIVING SPACE,

INFLATABLE

TUNNEL RADIATOR

AREA

F^-4.6M __ '

ORIGINAL PA6. 1,

OF POOR QUALrry

The inflatable shells are shown as 4.5 meters diameter but they can be

lP.rger. When stowed for launch, they are contained behind covers which

conform to the shape of the orbiter payload bay. The covers are deployed to

allow the shells to inflate and are then used as radiators. The deployed

module forme two separate sections for safety/'redundancy, with the strongback

providing a long central floor. Equipment is mounted both inside and outside

of the pressurized volumes. The two docking systems conform to standard

habitat module practice for design and placement.

5.1.5.2 Habitat Concept 4

This design ( Figure 5.1-5) is a combijation of an inflatable section and a

rigid body. When stowed, the module occupies all of the orbiter payload bay

with the exception of space for the docking module and the 0145 kit. The

module is supported in the payload bay by standard truanions and a keel

fitting and requires no separate cradle.

The rigid body is used mainly for equipment, arranged around a central tunnel

which rcus the length of the module. The outside diameter of the rigid body

is used as a radiator for the whole of the Habitat Module. The inflatable

section is in the form of a torus which surrounds one end of the rigid body.

The torus provides a large living volume which can be used in many ways. One

requirement is probably to divide it into two spaces for safety / redundancy in

the event of lose of pressure. It should be noted that there is space inside

the rigid body which can be used in ar emergency. A docking mechanism is

provided at one and of the module while at the other end an airlock with a

docking mechanism is provided.

To,

-----------

4.42M

— ' DIA

12.82M

• INFLATABLE a HARD SHELL

• OPT;ONAL HUD METEOROID

BUMPER

• AUTOMATIC DEPLOYMENT

• VOLUME STOWED 181 MS

• VOLUME DEPLOYED

LIVING 9D2

STOWAGE 95

AIRLOCK 11

TUNNEL 12

420 M3

Figure 5.I-5, habitat Module (Concept 4)

5-7
SCREWJACI

OR AST M

AIRY

2 PU

n,7 w^: .:.,, PI

Thin module can be used with or without a hard meteoroid cover over the

inflatable torus. The cover in folded and stowed around the outside of the

collapsed torus. When the to n:a inflates, the cover unfolde and expands along

with it. Deployment of the cover is completed when the panels which form its

end wells are released and awing into position under the influence of springs.

The shape and size of the torus can be chosen to suit requirements. it can be

much larger and it does not have to be c circular croas section. The shape,

as drawn, is a large circular room (10 m diameter) with a column in the

middle. The column is a portion of the tunnel which can be used to stow

equipment for launch. The equipment can be subsequently moved into the torus

living space.

5.1.5.3 Habitat Concept 5

When stowed, Concept 5 (Figure 5.1-6) resembles a standard habitat module

(i.e., as currently designed for the Space Station) which is roughly the same

size and shape as the orbiter payload bay. There are two sections which

deploy from opposite sides of the nodule and form the living volumes. They

are of excellent shape and can be subdivided to provide neparate sleeping

quarters, wardroom, working areas, etc. The whole module divides naturally

into two spaces for safety/ redundancy. At each and of the module is a large

airlock and docking mechanism. Some of the space currently allocated for

I.irlocks can be directed to stowage of items outside of the, pressurized volume.

6i

• HARD SHELL THROUGHOUT

• AUTOMATIC DEPLOYMENT

VOLUME ST-WED lei M2

• VOLUME DCPLOYED

LMNG 247

STOWAGE 78

AIRLOCKS al

265 M2

-- 4.I2M

DIA

12.03M ^_...,

0

RADIATOR AREA

Figure 5.1-6. Habitat Module (Concept 5)

5-8
M

The ntructure is metallic, including the deployable seatiunb and the cloee-out

panels. The whole structure is supported in the orbiter r+y azandard trunnion@

and u keel fitting. Expansion of the deployable sections can be obtained by

either internal pressure or by mechanical means such as astromasto or

serewjacks. Sealing between the deployable and the fixed sections can ^o

realized either by a large internal pressure-tight bag and by sealing around

individual points. The area of the module available for use as a radiator is

unusually large.

The ability to carry pressure-induced loads across the deplo;• .ble points is

appreciated. These loads may be bending moments is addition to axial loads.

EVA-installed, moment-carrying fittings are possibly needed with this concept.

5.2 TUNNELS

No study effort was devuted to tunnels since examination of Goodyear Aerospace

Corporation accomplishments indicated ample work has been done relative to

this component. The study time was better used in relation to hangars and

habitats, for ,hick no readily applicable effort is available.

5.3 OTV HANGAR

5.3.1 OTV Hangar Requirements

There are few really firm requiruments for an OTV hangar beyond the obvious;

i.e., it must remain in apace for the same duration as the Space Station and

it is launched using the orbiter.

Previous studies have indicated the infeasibility of pressurizing the hanger.

The problems associated with opening/closing such a large pressurized volume

are:

• Sealing the hangar

• The huge quantities of air lost each time the hangar is opened

• Or alternatively, the very large power requirements if a pump is

installed to recover the air before opening the hangar

The other hangar requirements listed are uncertain. They can be questioned:

o Does the OTV require protection against meteoroids? Would it be more

cost effective to accept a small risk of OTV damage than to build an

expensive hangar?

o If work platforms are required, is it reaso able to build a hangar for

that purpose only; or wnuld it be better to design the OTV so that

astronauts attach themselves directly to its outside shall, rather

than to hangar-mounted work platforms?

However, it was necessary to establish a basis for design studies, hence in

spite of the lack of maturity in hangar and Space Station definition the

following requirements were tentatively assumed:

i

5-9
i

• The UCVH size shall be suitable for t l, e servicing/maintenance of an

OTV. Assume an 0';' size.. of 4.3 m diameLer x 10 m long as a maximuT,

and work platform widths up Lo 2 meters.

• The OTVH will be unpressurized.

• The OTVH shall provide for a method of controlling the egress/ingreec

and stabilization of the OTV.

• Provisions shall be made for the following equipment:

• Work platforms

• Lighting

• Electric power

• OTV replacement items

• OTV refueling

• Minimum life support system (emergency)

o The OTVH shall provide radiation and micrometeoroid protection to the

crew, OTV and equipment. ( The and of the hangar pointing to earth is

not subject to micrometeoroid bombardment and can be left open)

o Life in LEO for 10-20 years

o The OTVH shall be compatible with an orbiter launch

5.3.2 OTV Hangar Design Approach

There are two basic methods of designing the OTV hangar, i.e., inflatable or

hard shell. The hard shell is a more conventional approach (( and provides a

firmer base for work platforms (reaction of astronaut loadZ. Its mayor

problem is packaging in the orbiter and deploying from it. If the Hangar is

to be packaged into a small portion of the critter instead of using the whole

of the payload bay, the problem of deploying it becomes even more formidable.

While an inflatable hangar 1s relatively easy to deploy, present designs may

not provide adequate stiffness for the work platforms. There are various

methods of hardening an inflatable structure such as foaming between walls,

use of a skin material which hardens on exposure to radiation, and an erectile

shell. In the erectile shell, a film of aluminum is added over the inflatable

which is then pressurized to stress the aluminum past its yield stress. After

the gas escapes the structure maintains its shape.

Regardless of which structural concept is used, all of the auxiliary structure

and equipment are to be built into the assembly and deployed with it. These

items are the crew work stations, lighting provisions, OTV ingress /egress

provisions, OTV stabilization provisirns, crew ingress /egreee provisions, OTV

refueling, and attachment to space station.

An alternative to automatic deployment of the hangar is erection or, more

likely, a hybrid arrangeuent. The complications of automation must be

balanced against the difficulties of EVA erection. If EVA creation is

considered, it is better to use a method that requires a minimum of orbiter

support.

5-10
UkKaNK 1 1 AGC 19

[7f POOR QUALfTY

5.3.3 Meteurold Impact Analysis

Since one of the purposes of a hangar is to provide protection for the OTV

against the impact of micrometeorioda, it is pertinent to examine the

potential damage to the OTV if an OTV hangar is not present.

The critical parts of the OTV are the LH2 and L0 2 tanks which are

contained insido an unpressurized structural shell as shown (Figure 5.3-1).

Present designs (ground-based and space-based) utilize a graphite composite

faced sandwich construction with a 0.035 gm/cm 3 (2.2 lb/ft 3 ) aluminum

honeycomb core. A rigid polyurethane foam core is expected to be structurally

adequate for a apace-based OTV design and provides superior meteoroid

shielding capability over that of the honeycomb core. The foam core does not

have sufficient shear and compression stiffness to stabilize the face sheets

against wrinkling duriag Shuttle launch of a ground-based OTV.

$PACE-BASED DESIGN CONSTRUCTION

0.020 M

GIWMI T[

COWOSI T[

I TO 2 u !AC[ WITS

OUTER --

JMDM[YCOIO COME

OUTERCT RIGID POLY- iWELL UN[TNM[ !OM

AWF. TAMM YELL

0.050 TO 0110 M

Figure 5.3-1. OTV Construction

The Rockwell Space Station studies are pursuing a space-based OTV which will

be returned to earth after 16 missions (6 to 18 months). The micrometeoroid

analysis for this design indicates there is a 1% probability the space-based

OTV tank wall will be punctured in six months, and a 32 probability it will

be punctured in an 18-month lifetime. This analysis is regarded as conservative since most of the tank wall is located a significant distance from the

outer shell. Furthermore, even if a tank containing LH2 or L0 2 is punctured

by a micrometeoroid there will not be a catastrophic failure. The tank wall

materials (2219-T87 or 2014-T6) are thin gauges and have a "leak before

5-11
ORIGINAL PP-GE IS

OF pOOR QUALITY

failure" characteristic tna' precludes explosive decompression. More than

100 Lycles of pressure are required for a crack 0.04 cm deep by 0.38 cm wide

to pro,a;ate to a through-crack.

The additional protection of a hangar wall fo^ the OTV is illustrated in

Figure 5.3-2. Thin figure also shows the cost implications of a hangar versus

replacement of an OTV. Clearly, the coat of a conventional hangar cannot be

,justified on the basis of meteoroid protection alone. Only if the - oat of

launching + hanger can be drastically reduced by tight packaging in the

orbiter or by other means, does the project become reasonable (if meteoroid

protection is th- only significant requirement).

ADDITION OF HANGAR STRUCTURE

(WITHOUT FOAM)

HANGAR

A LL

0.050 a

50^ ^ISo

It .RaeAenlTT

METtoeolo

1 TO : •

rARTiCI[S vn.L

Ory

GCT IAfT IMN(R

$HILL I i !III DNGER °MH[o.R

O OS TO 0. 10 .

ALAI. TANK WALL

S HANGAR REALLY COST EFFECTIVE If M[TEON010 PROTECTION

IS PRIMARY JUSTIFICATION}

ONLY WITH RELIABLE TIGHTLY PACKAGED CONFIGURATION

( REQUIRING 7 TO 5 M OF PAYLOAD SAY.

t

COSTS (SM)

OTY REPLACEMENT . 25

OTV LAUNCH AB

TOTAL 73

HANGAR LAUNCH AD

HANGM COST 10

•SPAChSASEL DESIGN I

Figure 5.3-2. OTV Meteoroid Protection/Hangar Cost

To date the studies of the Space Station have not produced hard requirements

for an OTV hangar. However, recognizirlq the possibility that requirements may

be developed, tho OTV hangar concept development was continued.

5-12

V
O RIGID DEPLOYED

HANGAR

--13.7M --+^

O RIGID DEPLOYED

HANGAR

OINRATARLE HANGAR

hA

—1— 14.414

STOWED s 4.514 DIA

RIGID DEPLOYED

STOWED

6M r 4.5M DIA

HANGAR

Q

C

ORIGINAL ?ACC is

OF POOR QUALr(Y

5.3.4 Candidate Concepts for OTV Hangnr

Seven OTV hangar concepts are prese lated (Figure 5.3-3). The concepts may be

clsesified into three categories, i.e., deployable Hangors, rigid

nondeployable hangars, and deployable/erectable hang<re.

The four deployable hangar concepts, i.e., Concepts 1, 2, 3 and 7 are

discussed below.

Concept 1 is a rigid shell which packages from a shape which le essentially

cylindrical (8.2 m dia.) to a star shape which fits into the orbiter bay. An

extendable trues which is used for ingress/agrees of the OTV is part of the

deployable shell. The stowed hangar occupies all of the orbiter payload bay.

Concept 2 is a rigid shell which consists of four half cylinders stowed in

the form of one cylinder insiie another. It, too, occupies all of the orbiter

payload bay when stowed.

Concept 3 is an inflatable shell which expands from a module which contains

the interface for docking to the Space Station, and the mechanism/structure

for OTV ingress/egress There are various methods which may be used to harden

or rigidise an inflatable structure which otherwise might be too soft for a

working interface. The stowed hangar occupies approximately 1/4 of the

orbiter payload bay. 

0

SPACE LAUNCHED

STATION WP.TI Otv

RIGID 

HANGAR RIGID HANGAR it 714

_+•yj x 4.5IA DIA

O

RIGID PANELS

E.V.A WORK STATION,

rSTOWED VOLUME

O

6.7M „ 4.314 DIA

RIGID HANGAR

ROTATING

,YORY. STATION AND PO

PANEL STOWAGE f'^ ePACE

STATION

Figure 5.5-3. OTV Hangar Concepts

5-13
Concept 7 is similar to Concept 1 except that it has a double fold (i.e.,

longitudinal as well as lateral) which enables it to occupy only a short

length of the) orbiter paylc.md bay. (This concept is, subsequently, discussed

in greater detail.)

Concept 4 is a rigid nondeployable hangar in the shape of a clam-aholl which

fits closely around the OTV. The OTV is launct,:d in the orbiter inside the

hangar, and exits from the hangar by %sane of an extending trues. Normal

esrvicing/fueling of the OTV is performed through the aft end of the Space

Station which is docked to the OTV. Access to other areas of the OTV is by

doors and foot holds strategically located in tha shell of the Hangar.

Concept 5 is a deployable/erectable system built out of the orbiter. A

deployable section with an interface for the Space Station is raised

incrementally by an extendable trues wnich also rotates about its major axis.

There is an EVA astronaut work station in the orbiter bey with storage for a

number of rigid panels which are attached manually to build up the hangar as

it rotates and raises.

Concept 6 is similar to Concept 5 except that the hangar is constructed

from the Space Station instead of from the orbiter. For both Concepts 5 and

6, it is estimated that approximately 1/3 of the orbiter bay is used for

stowage.

None of these concepts represents design that Rockwell favors. In an 

additional study effort superior configurations can be developed. However, to

provide further understanding of the design problems to be encountered, three

of the foregoing concepts are discussed in further detail in the next section.

(The drawings are provided in Volume II).

5.3.4.1 Hangar Concept 1

This concept (Figure 5.3-4) is a hard shell structure which folds laterally

for stowage. When folded it occupies all of the orbiter rayload bay. The

operational sequence is:

• Use the RNS to remove the hangar from the orbiter

• Attach to the Space Station using the docking interface

• Unfold the panels of the main structure using the actuators built into

the hargar

• The working platforms on the insido wall of the hangar may be deployed

either by springs or by actuators

The extendable/retractable astromaut is used for ingress/egress of the OTV,

for which purpose it is equipped with a docking mechanism tt interface with

the OTV. Lighting power outlets and similar items may *w built '.nto the wall

of the hangar.

5-14

l
2-7YA

ORIGINAL i ,,,, r^

OF POOR QUAU-Cy

• NAID ST%UCTM, NO INFLATALLfS

• AUTOMATIC DlM.OVMINT

• OCCVRS ALL Of +Nt PAYLOAD MV

L

DEN YING

►Ari0A0 MV

STOWED -

DEPLOYED

WOU MATFOW

^11.

i

ASTtOMAW

Figure 5.3- 4. OTV Hangar ( Concept 1)

Although the basic deployment is automatic, it is quite possible that EVA

astronauts could profitably be used to lock the wall panels together for

increased structural integrity, to install equipment, spares, etc.

5.3.4.2 Hangar Concept 6

This concept (Figure 5.3-5) uses EVA astronauts to erect the hangar walls

around a deployable bass. When stowed in the orbiter, approximately 6.0

meters of the payload bay is used. The stowed package consists of a

deployable base or hub and cradle. The deployable base contains a docking

interface for the Space Station, an astromast with a docking interface at its

free and. deployable "umbrella type" ribs, and a hub structure. The cradle

contains 40 separate panels which form the main structure of the hangar, a

docking mechanism to interface with the astromast, a mechanism to dispense the

panels as required during the hangar erection, and work stations/foot

restraints, etc., suitable for EVA astronauts.

The deployment sequence is as follows:

• Remove the two units from the orbiter using the RMS (the two units are

coupled together by the astromast and temporary fasteners)

• Attach the hub end of the hangar to the Space Station.

o Automatically deploy the umbrella ribs.

5-15
r

UrdUiNAL FAGS (3

OF POOR QUALITY

PAYLOAD NV

• MM STRUCTURE, NO INPLATARLES

• EVA CONSTRUCTION L 17

STOWED ^.--.6M__A

b PANELS

•

i

Figure 5.3-5

RECTION

EVA ASTRONAUT—.-_.

\ STATION

PANEL

STOWAGE

a.

I^----11.6eM 1.7YA

OTV Hangar (Concept 6)

o EVA astronauts begin to remove panels from the cradle and farm the

hangar structure. The panels attach to the umbrella ribs to form the

and wall of the hangar. The wall panels attach to the end wall and to

each other.

o The cradle rotates on the astromast and the astromast advances as

required to enable the EVA astronauts to attach successive panels

o The cradle is removed and returned to the orbiter (alternatively it

may remain with the Space Station and be used for logistics stowage).

There are many methods which may be used for Joining the panels together such

as spring loaded " click-in" ,joints, hand-held power tools for tightening

,joints, over-center locks, or gang latches.

Y' 5.3.4.3 Hangar Concept 7

n

This design (Figure 5.3-6) is a development of Concept 1. The difference is

the incorporation of a double-fold system (instead of a single fold) to reduce

the space occupied in the orbiter payload bay. In this version, the stowed

y length is reduced from 14.4 meters to 5.6 meters with a significant cost

saving. The installation procedure is:

o Use the RMS to remove the stowed package from the orbiter

o Attach the hangar to the space station using the docking interface

5-16

. 
Lop

S.6M

ASTROMAST

, 11.66M --- --. r-4 2.74M

STOWED s

DEPLOYED

B. seM

^r ORIGINAL PACE 19

OF POOR QUALMY

• Deploy the first fold in the same manner as described for Concept 1

• Deploy the second fold as shown in Figure 5.3-6

PAYLOAD

. HAW $TRUCfURE,

NO INKATARLES

• AUTOMATIC DEPLOYMENT

• DOURLE FOLD

SYSTEM

Figure 5.3-6. oTv Hangar (Concept 7)

The second fold is a series n° celescoping sections of the main hangar wall.

Extension of the telescoping oections can be accomplished by a series of stem

actuators at each panel to provide aushing action, or by the use of the

astromast (with a suitable attachment) to draw out the telescoping sections.

It is recognized that the double folding of such a large volume poses

formidable problems, and that much investigation remains to be done before

such a system could be rated as reasonably feasible.

5.4 MAJOR DESIGN ISSUES

The major design issues that have surfaced during this study applicable to

habitats and OTV hangars are listed below. The issue of OTV hangar

justification is dependent on what probability of no puncture is desired, and

the relative costs of OTV replacement versos the coat of a hangar.

Habitats

o Means for installation of equipment in inflatable structures

o Capability and frequency of repair of inflatable inner wall due to

micrometeoroid puncture

5-17
W

• Rigidity of unpressurized inflatable structures

• Inflatable materials suitability in space environment for 10 to

20 years and compatibility with crew (non-flammability, non-toxic)

• Provision of radiation shielding for crew protection

• For hard metallic shelle, the capability to sustain pressureinduced loads across de?loyable joints

OTV Hangars

o Requirements favoring tht need for an OTV hangar have not been

developed in Rockwell Space Station studies

o Justification for the hangar as a meteoroid bumper is dependent

on the extent of design conservatism desired by NASA.

o Cost of OTV possible replacement vs. the cost of hangar development and launch

o Rigidity of inflatable OTV hangars relative to crew EVA imposed

loads

5.5 P!tt'ENTIAL TECHNOLOGY DF - 7ELOPMENT NEEDS

The potential new technology development needs that have surfaced during the

course of this study are shown below. Rockwell is aware, through much of the

design information gratuitously furnished by Goodyear Aerospace Corporation,

that some of these items have been addressed by Goodyear to various stages

of completion but to different requirements (in general, for smaller volumes,

reduced pressures, shorter lifetimes). The time devoted to this activity

did not permit a screening of the applicability and status of these needs,

but this should be done in future studies.

o Habitats Utilizing Inflatables

- Assurance of no serious propagation of meteoroid

puncture hole size

- Repair of pressure containing wall

- Micrometeoroid puncture resistance

- Material suitability to crew and 10 to 20 years exposure

in space environment

- Maintainence of structural rigidity despite loss of

pressure

yy", o Fabitats Using Metallic Structures 

E

- None applicable at this time

,r

y

5-18
o OTV Hangars

- For inflatables —maintenance of structural rigidity

compatible with crew-induced foot loads, equipment

attachment, space platform inertia loads

- On-orbit foam in-place techniques

5.6 CONCLUSIONS

The following major conclusions pertinent to the use of deployable •, nluueu aia

evident from the foregoing discussions:

• Application to habitats appears attractive—large useful

volumes achievable.

• Metallic structures (can be sealed) provide ample meteoroid

radiation protection and equipment mounting surfaces, but

are constrained by pressure loads/packaging requirements.

• Metallic structures may require subsequent EVA for momentcarrying capability to withstand pressure.

• Inflatables alone are not sufficient —hard structure required

for mounting of consoles, orbiter and space station integration, heat rejection.

o Inflatables in conjunction with rigid core module provide

variety of feasible large volume designs, provided:

- Materials are suitable to crew safety/space environment

- Foam micrometeoroid stopping power is comparable to existing

data

- Repair of puncture or use of meteoroid bumper is feasible

- Additional 0.24 g yms/cc2 for protection to crewman eyes

is provided

o Hangar requirements are ill-defined —OTV metenroid protection

alone is not sufficient justification.

o Most attractive OTV hangar concept appears to be metallic

deployable/erectable or inflatable with foam core (provided

stiffness is adequate).

i

5-19

a

r

a
f

I

6. REFERENCES

1. Satellite Power Systems (SPS), Final Report (Exhibit E), LEST Systems

Analysis and Integration Task for SPS Flight Test Article, Rockwell

Report, SSD 80-0102, dated August, 1980.

2._ LSSr System Analysis and Integration Task for an Advanced Science and

Application Space Platform (Contract NAS8-33592), Final Report,

McDonnell Douglas, MDC G8533, dated July, 1980.

3. Geostationary Platform Systems Concepts Definition Follow-On Study, Final

Report, Vol. IIA-Technical, Task 11, LSST Special Emphasis, General

Dynamics, GDC-GPP-79-010 (IIA), dated Septenber, 1980.

t. Space Platform Design Requirements Summary, NASA /MSFC report, dated

December 31, 1980-

5. Geostationary Platform Systems, Concepts Definition Study, Final Report,

General Dynamics, GDC-GPP-79-006 (I) (Contract NAS9-33527), June, 1980.

6. Meteoroid Environment Model-Near Earth to Lunar Surface, NASA SP-8013,

March, 1969.

7. Review of Physical Processes in Hypervelocity Impact and Penetration,

Rand Report RM-3529-PR, July, 1963.

8. Development and Application of Space Deployable Box Truss Structures,

Martin Marietta MMA Report No. 78-52411-01, December, 1978.

9. Neutral Buoyancy Test Results of a Deployable Space Beam, AIAA-81-0437.

10. Efficient Structures For Geosynchronous Spacecraft Solar Arrays, Phase I,

II, and III Final Report, Astro Research Corp., Sept. 14, 1981.

11. Erectable Concepts for Large Space Systems Technology, Vought Technical

Report, 2-51400/OR-LSS-02, dated March 3, 1980.

12. Technology for Large Space Systems (NAS SP-7046).

13. "Nuclear and Space Radiation Effects on Materials", NASA SP-8053,

June, 1970.

14. Space Construction Systems Analysis, Part 2, Final Report, Platform

Definition, NAS9-15718, Rockwell International, April, 1980.

15. An Analysis of the Space Debris Problem - Present and Future, SSD82-0011,

January 29, 1982, Dr. J. K. Haffner.

6-1