NASA

Technical

Paper

3088

July 1991

,u; ": ;

%,;-* ,

¢

* t ; •

Packagin_eployment,

and On- Assembl

Options

Geostati

Williaz_T. Davis _ . ....

andCharles B.

-y

ge

Spacecraft

AqINGt '_" V.. I.,.jPM[Nr. , p A'-

• o_ -., i

' (..SC L :_2p, uncles

HIIIQ 0329507
- - "- -" -, - z

_, _ _-k.,- ...... _ _ _ • .T _ ,_ - _k -_,_-_-_ 7 ._-.--

- .,% .--

- j

• =

-5 -

LL

• :....

.... _ -: r;,SD -

7. .L

_. . +..L

-. _-_ -V _ -_._ :_

.:-- - • .i ̧_-: - _

, ;

_ ] _.

_._-,'_',_ .... _ - .__ , -

*'.: ',_.._.-_.-__-#--:.-. _ .,"_ "i. :- • .: •

. _ . : • . _ . __- .._ __- ,

" S ,7' -¢ - -_ _ =; , " _ J i , .

.... -'z..-7,-]£ "" -= L :-i == "_-'-- '." ?"1 .. -"

' -- '" -"-"_.G-- --'"!"" " '

+ ! -- . _'-._{+ £_£.. -_. _, .... _: •
NASA

Technical

Paper

3088

1991

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

Packaging, Deployment,

and On-Orbit Assembly

Options for Large

Geostationary Spacecraft

William T. Davis

Langley Research Center

Hampton, Virginia

Charles B. King

The Bionetics Corporation

Hampton, Virginia

Contents

Tables and Figures .............................. iv

Acronyms .................................. v

1. Introduction ................................ 1

2. Design Processes .............................. 2

3. ESGP Spacecraft .............................. 4

3.1. ESGP Spacecraft Description ....................... 4

3.2. ESGP Assembly Sequence Incorporation .................. 6

3.3. ESGP Assembly Sequence Definition ................... 8

3.4. ESGP Summary Remarks ....................... 13

4. GHRMR Spacecraft ........................... 15

4.1. GHRMR Spacecraft Description .................... 16

4.2. GHRMR Launch Vehicle Mass Considerations .............. 17

4.3. GHRMR-b Deployment/Assembly Concept ............... 17

4.4. GHRMR-a Hybrid Deployable/Erectable Concept ............. 22

4.5. GHRMR Task Summary ........................ 24

4.6. GHRMR Summary Remarks ...................... 26

5. Concluding Remarks ........................... 27

References ................................. 28

Ill
Tables and Figures

Table I. Payload Envelopes

Table II.

Figure 1.

Figure 2.

Figure 3.

Figure 4.

Figure 5.

Figure 6.

........................ 18

Packaging Options for GHRMR Geostationary Platform ....... 26

Earth sciences geostationary platform ................ 2

GHRMR spacecraft concept .................... 3

Spacecraft design flows and spacecraft assembly design flows ...... 4

ESGP spacecraft ......................... 5

ESGP cargo pallets for STS orbiter ................. 7

Titan 4 launch packaging for ESGP ................. 8

Figure 7. OFA locations as suggested for Space Station Freedom

enhanced operations capability phase .................... 9

Figure 8. Space Station Freedom box truss section with

proposed add-on shelf structure ..................... 10

Figure 9. Orthogonal construction station at SSF .............. 10

Figure 10. ESGP spacecraft assembled configuration (with tooling) ....... 12

Figure 11. Time-sequence assembly for ESGP spacecraft ........... 14

Figure 12. Principal elements of GHRMR spacecraft ............. 17

Figure 13. Single-fold, double-stowed truss concept .............. 18

Figure 14. Hexagonal-panel solid-reflector deployment concept

developed by Harris Corporation ..................... 19

Figure 15. GHRMR-b stowed in Shuttle C (block 1) ............. 20

Figure 16. GHRMR-b stowed in Shuttle C (block 1) with

prebuilt secondary reflector ....................... 21

Figure 17. Deployment sequence for GHRMR-b spacecraft ........... 22

Figure 18. Reconfigured GHRMR spacecraft bus ............... 23

Figure 19. GHRMR-b stowed in Shuttle C (baseline) ............. 23

Figure 20. GHRMR-b stowed in STS orbiter ................. 24

Figure 21. GHRMR-a hybrid deployable/erectable geostationary platform

stowed in Shuttle orbiter cargo bay (struts and nodes packaged separately) 25

Figure 22. GHRMR-a hybrid deployable/erectable geostationary platform

stowed in two Shuttle orbiter cargo bays (nodes preattached to struts) .... 25

iv
Acronyms

ACCESS

APS/MT

CELV

ESGP

EVA

FF

GCTI

GEO

GHRMR

GPS

ID

LaRC

LCP

LEO

MPESS

MSC

OFA

OMV

OTV

PSR

s/c

SSF

STS

STV

assembly concept for construction of erectable space

structure

astronaut positioning system/mobile transporter

complementary expendable launch vehicle

Earth sciences geostationary platform

extravehicular activity

free flyer

global change technology initiative

geostationary Earth orbit

geostationary high-resolution microwave radiometer

global positioning system

identification

Langley Research Center

launch configured package

low Earth orbit

mission peculiar equipment support structure

mobile servicing center

obstruction free area

orbital maneuvering vehicle

orbital transfer vehicle

precision segmented reflector

spacecraft

Space Station Freedom

Space Transportation System

space transportation vehicle

V

Abstract

Large spacecraft, particularly in geostationary Earth orbit (GEO),

require special attention to the design challenges of launch vehicle

packaging, deployment, and/or on-orbit assembly. Design studies of

two different GEO spacecraft required that packaging, deployment, and

on-orbit assembly analyses be conducted to establish the viability of these

concepts for future NASA missions. This study used these analyses as

"strawman" concepts for an investigation of packaging, deployment, and

on-orbit assembly techniques. It also revealed generic guidelines for inspace assembly and highlighted the importance of early integration of

packaging, deployment, and on-orbit assembly requirements into the

spacecraft design. The first study spacecraft was used to study the

definition and analyses of on-orbit assembly options for large GEO

spacecraft. The second study spacecraft required investigation of the

feasibility of deploying large spacecraft at GEO. The second spacecraft

was also used to examine the packaging requirements of a deployable

spacecraft and the packaging requirements for a hybrid assembled and

deployable version of that spacecraft,. This investigation was done

with attention to minimum volume (and minimum launches) and to

the relationship between packaging and spacecraft deployment and final

configuration.

1. Introduction

The Langley Research Center (LaRC) is conducting engineering studies to identify and

evaluate the role for advanced technologies in proposed future space systems.

The future need for large structures in space, particularly at geostationary orbit, has

frequently been highlighted and examined in the literature. (See ref. 1. Other work has been

done under contract to NASA by Ford Aerospace Corporation, GE Astro-Space Division,

and Lockheed Missiles & Space Company.) Studies concerning in-space construction range

from the construction of spacecraft bound for the Moon and Mars (work done under

contract to NASA by McDonnell Douglas Space Systems Company) to partial telerobotic

or autonomous assembly of the Space Station Freedom (SSF). (See ref. 2.) LaRC has, as

part of its overall spacecraft technology studies, performed packaging, deployment, and onorbit assembly studies relating to two large geostationary platform concepts as an adjunct

to other studies in progress.

The first concept used an LaRC design of a large second-generation Earth sciences geostationary platform (ESGP) (refs. 3 through 6) as shown in figure 1. The ESGP spacecraft

designed by LaRC was based on an earlier concept developed by Ford Aerospace Corporation. Configuration details, subsystem definitions, and large antenna designs necessary to

conduct in-depth performance analyses were developed for the ESGP spacecraft. Assembly

techniques and options necessary for low Earth orbit (LEO) assembly were studied. Packaging of this spacecraft for delivery to LEO is discussed in section 3.2 and in reference 7.

The second concept stemmed from the Global Change Technology Initiative Architecture Trade Study recently conducted at LaRC. This trade study defined an infrastructure

of both LEO and geostationary orbit (GEO) spacecraft. The geostationary Earth science spacecraft from this study (fig. 2) is referred to as the "geostationary high-resolution

microwave radiometer" (GHRMR) spacecraft after its principal instrument.

The study described in this paper serves to supplement the previous technology studies

and provide a mechanism for examining near-term (defined here as prior to 2010) packaging,

deployment, and on-orbit assembly options for this class of large geostationary spacecraft.
Specifically, the ESGP spacecraft was used in this study to examine requirements for LEO

assembly of large GEO spacecraft, and the GHRMR spacecraft was included to investigate

complex packaging and subsequent deployment/assembly activities. This paper addresses

the two concepts separately and then presents some overall concluding remarks.

Item Mass, kg

Instrument & housekeeping modules

15-m radiometer

7.5-m radiometer

Truss structure

OTV interface

Solar arrays

Other scientific instruments

Solar sail

Communication antenna

3456

844

705

377

45

149

1075

10

20

Total 6681

Each

Payload

m/UlHouse::jping

scientific '_ Platform main bus

instruments truss

Figure 1. Earth sciences geostationary platform.

2. Design Processes

As with most engineering design processes, early incorporation of significant requirements enhances the probability of overall success. Similarly, a design for spacecraft requiring on-orbit assembly must incorporate special requirements at the beginning of the design

process. In order that the spacecraft fulfill its on-orbit mission, the designer must carefully

attend to the requirements of in-space assembly as they pertain to ease and reliability of

construction, time and cost of assembly operations, and the safety of the spacecraft and

its assemblers.

2
GHRMRoption GHRMR-aGHRMR-b

GHRMRmass, kg 2417 1947

Otherbay-loaded mass, kg 732 732

OtherS/Cmass, kg 3010 2754

Totalmass, kg 6159 5433

Spacecraft

bus

Figure 2. GHRMR spacecraft concept.

Other instruments I

A process for synthesizing the design of a large geostationary spacecraft that requires

in-space assembly is postulated and shown in figure 3. The normal or usual design process

is indicated by the solid lines. A new group of design flows which are necessary for on-orbit

spacecraft assembly are indicated by the dashed lines.

The normal inputs to a spacecraft design are shown as mission requirements, instrument

requirements, spacecraft subsystem hardware needs, checkout requirements, and launch

vehicle constraints. Additional items to be considered for on-orbit assembly include new

techniques and hardware specifically for in-space construction, such as EVA techniques,

assembly-specific jigs and tooling, and unique deployment schemes. For this new group

of design flows, the ESGP and the GHRMR spacecraft provide "strawman" vehicles with

a wide-ranging technology representation, which provide an elementary understanding of

considerations pertinent to on-orbit assembly/deployment requirements.

3
Spacecraft i ,- -_

! instrument 1 | Launch |

f Mission1 / uniqueI / vehicle/ Oes,gnfiow

requirements|_ k reqmrements Jl [parameters | \ Assembly design flow

On-orbit t / ,..f Final

/ check-out I--------_ Spacecraft i / _ |spacecraft design I

( requirementsJI x,, ]design options| / ]including assembly I

\, _ _ / I,, sequence •

I Spacecraft | _\ ,I ,/ \_ _ [

I suosystem !// \ I / ( _ I ]

I .hard.ware | \ \ ] / I Assembly node 1

k.aescnpuons • \ "a _ _/ _ requiren.'ents | I [

\ "_ • V _/_ [. &constraints J I

r , "_\\. [ Assembly/ t I [

[ :x_:m:/Ye c t _'[ inftoe_nhtologbYas edepl°yment I d.._ nd_

Past m-spa e __

t, _ • , . (Grouna test a Ill c--'------_

/_ _ ] verification _ Mission |

/ [ L "actuals" J--

[ ewm-space|JI 1 1

assembly New in-space

methods assembly

technology

Figure 3. Spacecraft design flows and spacecraft assembly design flows,

3. ESGP Spacecraft

The ESGP spacecraft as shown in figure 4, with part identification (ID) numbers, is a

large Earth science geostationary platform concept. It may be delivered to low Earth orbit

by either the Space Transportation System (STS) orbiter or the Titan 4 complementary

expendable launch vehicle (CELV). (See ref. 7.) After assembly in LEO, it is transferred

to GEO by a space transportation vehicle (STV).

3.1. ESGP Spacecraft Description

The overall dimensions of the ESGP spacecraft are approximately 40 by 40 by 10 m.

It uses a 3-m box truss structure (ID 1, fig. 4), which acts as a strongback for all

spacecraft subsystems and for the payload complement of 18 instruments including two

large microwave radiometer antennas (15- and 7.5-m-diameter solid reflectors). This

erectable truss provides a rigid and stable base to which the various components of the

spacecraft are connected. The individual truss element design (unpublished data from

Ford Aerospace Corporation) is derived from a quick-connect type used on the ACCESS

experiment conducted on STS mission 61-B. Truss elements are composed of graphite-epoxy

tubes of approximately 5 cm diameter.

Attached to the second full bay (from left to right) of the truss is a stack consisting of the

payload module, the housekeeping module, and the orbital transfer vehicle (OTV) thrust

interface tube (ID 2). At the top of the stack, the payload module contains additional

science instruments. Connecting the payload module to the truss is the housekeeping

module which contains the major subsystems of the spacecraft such as power and attitude

control. Both modules are of aluminum honeycomb construction. The OTV interface tube

attaches to the bottom of the housekeeping module and extends through the second bay

of the truss to a standard OTV docking ring.
ID Item

1

2

3

4

5,6

7

8

9

10

11

3-mboxtrussstructure

Modulestack(payload&housekeeping modules, thrustube)

15-m-diameter trussmicrowaveradiometer reflector

7.5-m-diameter trussmicrowaveradiometer reflector

Subreflector masts

Undeployed15-mantenna

Solarsail

Solarcellarrays

Solararraydeployablebooms

Otherscientificinstruments

Earth

A

I

E (+x)

\

9

\

S (+y)

3

\

!0

1

Figure 4. ESGP spacecraft.

The 15-m-diameter microwave radiometer reflector (ID 3) and the 7.5-m-diameter

microwave radiometer reflector (ID 4) are the largest instruments on the platform and are

the primary platform configuration drivers. Each uses offset-fed Cassegrain-type geometry,

which makes use of folded optics to enhance the scanning performance of the radiometer.

In this configuration, the large primary reflector focuses the radiation upon the smaller

subreflector which, in turn, focuses the radiation upon the feed array. The scanning

of the radiometers is accomplished by pivoting their subreflectors through the use of

electromechanical actuators. The subreflectors are mounted at the top of the subreflector

masts (ID's 5 and 6). These masts are deployable and are ofa Minimast design. (See ref. 8.)

This type of boom requires the use of a deployment canister mechanism which extends

the triangular-cross-section boom two bays at a time. The 7.5-m antenna is similar in

concept to the precision segmented reflector (PSR) being developed at the Jet Propulsion

Laboratory. (See ref. 9.) It consists of 12 precision surface hexagonal-shaped reflector
segments mounted atop a strongback truss structure. The 15-m antenna strongback is

similar in concept to a General Dynamics deployable concept (ref. 10) and includes a

deployable reflector membrane. The undeployed 15-m antenna is represented in figure 4

by the cylinder (ID 7) at the left end of the spacecraft.

An adjustable solar sail (ID 8) is attached to the right end of the box truss structure

to reduce spacecraft control torque requirements.

The foldable solar arrays (ID 9) are attached to the housekeeping module by deployable

booms (ID 10) which place them at sufficient distance from the spacecraft so that they

are not shadowed by the 15-m radiometer reflector. These deployable booms are 0.5 m

in diameter with flexible longerons and are deployed from the housekeeping module by

electric motors. Each solar panel boom rotates to maintain solar pointing.

Attached to the third, fourth, fifth, and sixth bays of the box truss, between the housekeeping module and the 7.5-m reflector, are other elements of the scientific instrumentation

complement (ID 11) of the spacecraft.

3.2. ESGP Assembly Sequence Incorporation

A number of factors influenced the incorporation of the on-orbit assembly sequence into

the ESGP spacecraft design. First, other studies relating to this spacecraft were reviewed

to acquire spacecraft functional understanding (i.e., the design of the mission, instruments,

checkout requirements, launch vehicle constraints, subsystem hardware design). This, in

effect, was a review of the spacecraft design process, as shown in figure 3 by the solid lines.

The launch packaging study (ref. 7) identified three viable scenarios for delivery of

the ESGP spacecraft to Space Station Freedom (SSF) in LEO. Two scenarios use the STS

orbiter while the third utilizes the Titan 4 complementary expendable launch vehicle. In the

shuttle scenarios, the ESGP components are packaged in three specialized pallets as shown

in figure 5. A similar arrangement of three modified payload modules (fig. 6) is integrated

into the Titan 4 launch vehicle in a No Upper Stage configuration. The launch configured

payload pallets or modules are henceforth referred to as "launch configured packages"

(LCP's). Mass and center-of-gravity launch constraints for the three configurations were

verified as being met. The Titan 4 scenario and the shuttle single launch scenario both

require orbital maneuvering vehicle (OMV) retrieval to SSF for assembly, whereas the

shuttle two-flight scenario is launched directly to SSF. This paper uses these launch vehicle

assumptions and the launch packaging findings of the launch packaging study (ref. 7) as a

starting point for assembly definition.

The design flows represented by the dashed lines in figure 3 were next undertaken to

define a strawman example of an assembly sequence. An investigation of previous inspace assembly experience (ref. 11), new in-space assembly methods (ref. 12), and new

in-space assembly technology (ref. 13), along with previously gathered spacecraft hardware

and design data, became the basis for an assembly/deployment technology information

base. This information base, along with the spacecraft design options chosen, allowed the

definition of an assembly node (an assembly place and process) for the ESGP spacecraft.

The specific design form of the assembly node can have a number of acceptable solutions

and is normally based on the engineering criteria of simplicity, safety, cost, and so forth,

and their respective priorities.

6
GeOdynamic laser ranging experiment

Solar panel

High-ga n/ pins

_ Innff_:dometer Instrument module 7--_

spectrometer

antenna /- Multichannel microwave

/ radiometer

F 3-m truss elements

I I ""- 4.24-m truss

;'ae2e ':

1 '- Housekeeping module

Keel pin

(a) Pallet 1.

15-m passive

rm_tlr°Wet:e feed array____,._,_x_r_ _

M_cnihm2S_boom deployment---x\ //_, _ _ (

.,----15-m passive microwave

radiometer membrane surface

" _15-m passive microwave radiometer

x_ >i strongback truss

Longeron pins

"_ L '"*'-_ Launch support structure

"//A

' " _ -------- 15-m passive

[ microwave

radiometer boom

7.5-m passive microwave

radiometer boom

- Keel pin

(b) Pallet 2.

_j-- 7.5-m passive microwave

Grapple fixture 7 J _ radiometer truss

/ _----_ " / structure

Longeron (, _j_" _ _ /-- 7.5-m passive

pins _ v ..,¢'_,._.__ / m_crowave

__1____ radiometer surface

]II / /-3 \ )"_l_..._-- Longeron

Launch \-x_ J II ,,_¢_¢"_"

support \ _

structure __.x _ _}_..---_Keel pin

t-Solar sail

(c) Pallet 3.

Figure 5. ESGP cargo pallets for STS orbiter.

7
Payload fairing

OMV grapple

fixture

Launch support structure

Figure 6. Titan 4 launch packaging for ESGP.

3.3. ESGP Assembly Sequence Definition

Assembly sequence design decisions are strongly influenced, if not dictated, by the

specific spacecraft design under consideration. In any event, certain issues must be

addressed (with mutually compatible conclusions) prior to determining an acceptable

assembly sequence. These are as follows:

Where is the assembly to occur? The options considered were the Space

Transportation System orbiter, a free flyer (FF) arrangement where the spacecraft

is its own assembly node, and the SSF. Although assembly at the STS does not

depend on other elements such as the SSF or an FF, it does have limited on-orbit

mission duration, leaving little, if any, contingency assembly time. The FF requires

EVA support and/or astronaut transfer from either the SSF or the STS and active

spacecraft control during at least part of the assembly period to point the large

reflectors away from the Sun. The analysis assumes that the SSF has obstruction

free area (OFA) available for both the assembly and the storage of the spacecraft in

its launch-configured packages. EVA support is assumed to be available from either

the orbiter or the SSF, and there is no limited-time constraint (before the shuttle's

return to Earth) on the allowable time to build. Other factors, not quantifiable

in a strawman scenario, such as disturbance of microgravity science experiments,

should also be considered. Based on these considerations, the SSF was chosen as

the assembly node.
Where and how large is the OFA? Sufficient OFA for assembly of the ESGP

spacecraft must be found and allocated for both the assembly process and the storage

of LCP's (until required for assembly). Although availability may change as the SSF

design evolves, sufficient OFA for assembly was found along the back of the lower

boom in the enhanced operations capability configuration of the evolutionary SSF

(unpublished data from the Space Station Freedom, Office at LaRC). (See fig. 7.)

This configuration is scheduled to be within the time frame used for this study

(before 2010). OFA used in assembly must also allow room for any special tooling

required. OFA for storage of the spacecraft LCP's should be in near proximity to tile

assembly OFA and could be on either the port or the starboard keel near tile lower

boom. Any environmental protection needed for the LCP's must also be included.

For this study, it is included in the launch packaging. Alternative solutions to the

OFA include (1) along the inside of the starboard lower keel (fig. 7), (2) a hanger

structure built on the evolutionary SSF dual beam truss structure (particularly if

a protective enclosure is needed), (3) a temporary shelf-type truss structure (fig. 8)

specifically designed for this spacecraft assembly, and (4) a special construction

station capable of efficient assembly with the spacecraft oriented orthogonally to the

SSF truss structure (fig. 9).

What will be the assembly sequence? Several options to this major question

are (1) build the spacecraft truss structure first and then add the other components,

(2) build the three-piece module stack first and expand from there, or (3) build

the spacecraft in an end-to-end sequence with major deployments last. Although

the first and second options may also provide acceptable assembly sequences, the

third option was chosen to complement the special tooling discussed later and as

a method of minimizing the EVA translational movements and resulting astronaut

physical exhaustion as recommended in reference 14.

.<

\

\

\

\

N

x.

x N

• [

_, Alternate obstruction free area

Obstruction free area along back of

lower boom

Figure 7. OFA locations as suggested for Space Station Freedom enhanced operations capability phase.

9
5m

Existing truss Add-on truss

structure5m _'-y--,_k'_ structure5m

/

\

\

Figure 8. Space Station Freedom box truss section with proposed add-on shelf structure.

Figure 9. Orthogonal construction station at SSF.

How is the assembly implemented? The choices are EVA, telerobotics, automation, or combinations thereof. EVA requires much of a limited resource and provides significantly less astronaut safety than the other methods. Telerobotics may

require as much or more assembly time than EVA but provides the desired safety.

10
Telerobotics is applicable to many but not all assembly tasks. Robotics may be

faster than either EVA or telerobotics and also maintains astronaut safety but is

usually best applied to repetitive tasks. In addition, its application to this type of

assembly task is not a near-term option. (See ref. 13.) Therefore, for the near-term

assembly techniques that were set up as ground rules for this study, a combination

of EVA and telerobotics was chosen as a baseline assembly method. The assembly process, in whole or in part, could be automated as an enhancement when the

necessary automation technology becomes available.

What assembly tooling is required? The unpacking and temporary stowage of

the LCP items would require the use of a translatable remote manipulator arm

device such as the SSF mobile servicing center (MSC). These one-time tasks would

appropriately be done telerobotically to avoid EVA. For spacecraft truss assembly

and instrument/appendage attachment via EVA, a build-extend turntable-type

structure was devised. This special tooling was designed with similarity to ACCESS

(ref. 11) and to a proposed SSF truss structure assembly tooling (ref. 12) to minimize

astronaut EVA physical exertion. The build-extend descriptor refers to a truss

section or bay being built on the tooling fixture, then extended outward but still

held as the next bay is assembled. As each bay is being assembled, the turntable is

partially rotated or rocked to give the astronauts access to all assembly nodes. This

process is repeated until the complete truss is assembled. The same basic tooling

apparatus that has been used to build the spacecraft truss can then be reconfigured

to support the assembly of the 7.5-m solid reflector and its truss strongback. Another

possible option, although not considered here because of unknown availability, would

be to use the reconfigured SSF astronaut positioning system/mobile transporter

(APS/MT) proposed for SSF assembly. (This system was discussed by D. R. Barron

of McDonnell Douglas Mechanical Systems at a planning review meeting at Johnson

Space Center.) This tooling configuration would have been used in SSF truss

assembly and possibly could be downsized to accommodate the smaller truss of

the ESGP spacecraft. A drawing of the assembled ESGP spacecraft and its tooling

is shown in figure 10.

Restating from the above discussion, the strawman assembly sequence was defined

within the following guidelines:

1. The assembly is to occur at SSF

2. The assembly is to use OFA along the back of the lower boom

3. The spacecraft is to be built in an end-to-end order

4. The minimum technology approach (EVA, telerobotics) was chosen as a baseline

instead of an advanced technology approach (automation)

5. The basic tooling required was defined as a build-extend fixture incorporating a

turntable

Once the basic approach has been selected, a more detailed review of the assembly

scenario should be conducted. Examples of this more comprehensive review of the assembly

design are now discussed.

This review should investigate the need for identifying other resources and their impact

on the assembly node. This would include electrical power, propellant handling scenarios,

other fluids usage, special checkout requirements, and any special instrumentation required

for either assembly or checkout. The appropriate selection of either EVA or (tele) robotics

11
fortheimplementationofparticularassemblytasksmustbeconsidered.Forinstance,some

tasksmaybesignificantlyquickerto implementviaEVAalthoughtherisksassociatedwith

othertasksmaybetoo high for EVA.Useand schedulingof otherSSFresourcesuchas

the MSC,the SSFcrew,if available,or an auxiliary crew must betakeninto account.

The requirementsfor environmentalprotectionof the LCP whilestoredat SSFmust be

addressed.

I _ Tooling

Lower boom in

enhanced operations capability configuration

(Space Station Freedom Office at LaRC)

Figure 10. ESGP spacecraft assembled configuration (with tooling),

Furthermore, the minimization of assembly sequence times (or other selected parameters) can be significantly enhanced by attention to the following generic design

considerations:

1. Design of the vehicle/node interface with possible incorporation of assembly requirements into the node design

2. Design for minimum spacecraft assembly (maximum prebuild on Earth)

3. Design for quick assembly (e.g., snap-lock fittings versus nuts and bolts)

4. Incorporation of self-alignment features

5. Incorporation of self-verifying locking features

6. The innovative design and use of assembly aids

7. Examination of the assembly sequence to minimize the need for time and other

resource-intensive operations such as MSC translation and plane changes

8. The use of hybrid deployable/erectable designs that can yield significant resource

savings or may give other unrelated benefits

9. The generation and use of a lessons-learned compilation to avoid previous mistakes (a

study made by Rick Vargo, Fred Mitchell, Ken Flemming, and Maurice Willis of McDonnell

Douglas Space Systems Company for Kennedy Space Center)

12
Eachspacecraftelementmust be kept compatiblewith the designof the assembly

sequence.As an illustration of this necessity,an incompatibilitywas found with the

proposeddesignof the spacecraftpropulsionsystem. The hardwarethat wasselected

in the spacecraftdesign(not theassemblydesign)wasanall-titanium, all-weldeddesign

for both propellantstoragetanks and propellanttransferlines. Thrusterswereplaced

on the instrument moduleand its supportingtruss bay. To minimize tile potential

complexityof in-spaceoperations(the assemblyandweldingof this hardware),the option

of prebuildingthe modulestackandits trussbayasa singleassemblybeforelaunchwas

selected.Therefore,for a small Earth to LEO transportationvolumepenalty (i.e., tile

truss baylaunchedassembledvs launchedasstrutsand nodes),the needfor complexinspaceassemblyandweldingof the propulsionsystemplumbingwasavoided.This typeof

potential incompatibilityillustratesthe desirabilityof integratingthe assemblysequence,

asearlyaspossible,into the total spacecraftdesignprocess.

A time-sequencedepictionof the constructionsequenceis shownin figure 11. The

assemblybeginswith the +x end of the ESGPspacecraft(left end in fig. 4) and grows

towardthe oppositeend. In all drawings,the build-extendtoolingwouldbeat the right

endof the so-far-assembledspacecraft.In general,attachmentsaresecuredincrementally

as constructionprogresses.The payloadand housekeepingmodules,thrust tube, and

associatedtruss bay are constructedon Earth beforelaunch and assembledinto the

spacecraftas a singleentity. The 7.5-m reflector is constructed after all truss bays are

completed. This requires reconfiguration of the tooling turntable. After construction,

the 7.5-m reflector is attached to the spacecraft. Next, using the Minima,st deployment

mechanism, both Minimast subreflector booms are deployed and attached to the spacecraft

with the aid of the MSC. The 15-m reflector canister (ID 7 in fig. 4), attached to the left

end of the spacecraft, is then deployed. Individual cl_eckouts occur as appropriate during

spacecraft construction with full integrated checkout occurring last. A rough order-ofmagnitude estimate for assembly time was desired. Based on previous experimental inspace construction timeline studies (refs. 12 and 14), a preliminary estimate for assembly

of the ESGP spacecraft, using the specified baseline technology (EVA, telerobotics), is at

least three and probably four standard 6- to 8-hour EVA periods. Contingency assembly

time was not estimated for this study. Integrated checkout, being spacecraft and payload

dependent, was not included in this estimate.

3.4. ESGP Summary Remarks

A spacecraft design process to accommodate on-orbit assembly design for an Earth

Sciences Geostationary Platform was postulated. With this process, an information base

was collected utilizing relatively near-term technology parameters. The information base

was next applied in the specific definition of an assembly sequence for an already defined

strawman vehicle (the ESGP spacecraft). The use of a strawman test case for an on-orbit

assembly-required spacecraft design helped to identify and highlight the significant steps

needed for this design process.

We recommend that the assembly tasks be planned as an integral part of the original

design process and considered second in importance only to the spacecraft functional requirements. When this is not possible, they must be integrated as soon as feasible within

the overall design process. The on-orbit assembly tasks must be verified ms compatible

13
I 2 3 4 5 6 7 8

9 10 11 12 13 14

15 16

(_z)

(+x_"__ N (-y) W (-x)

E

20

Figure 11. Time-sequence assembly for ESGP spacecraft.

14
with thespacecraftdesignandwith theEarth to LEO packaging.Thisis donenotonly to

assurecompatibilitybetweenspacecraftdesignandassemblyrequirementsbut to increase

theprobabilityofoptimalengineeringdesign,costandtimesavings,andsafety.In addition,

to assurecompatibilitybetweenspacecraftdesignandassemblydesign,certainmajorissues

shouldalwaysbeaddressed.Theyareasfollows:

1. Whereis the assemblynode?

2. Howmuchassemblyspaceis needed?

3. What is the basicbuildsequence?

4. What arethe crew/EVArequirements?

5. Are theretelerobotics/automationrequirements?

6. What assemblytoolingmayberequired?

7. How much overall time is required for the assembly?

A review of the assembly design followed an assessment of the major issues. Recommendations for assembly time minimization are made in the form of suggested generic design

considerations.

Initial feasibility of on-orbit assembly for a strawman spacecraft was assessed. An

assembly sequence for this spacecraft was designed and EVA requirements were estimated.

Small spacecraft design modifications (in this instance, prebuilding the module/bay 2 stack)

may significantly enhance (or even allow) the assembly task.

Analysis of alternative assembly sequences for the ESGP spacecraft could similarly be

conducted by selecting any combination of available alternatives of the major questions

discussed in section 3.3. The examination of resultant timeline scenarios from alternate

assembly sequences can be an effective assessment tool for assembly sequence timeline

definition and overall sequence selection based on comparative assembly times.

4. GHRMR Spacecraft

The geostationary high-resolution microwave radiometer spacecraft (fig. 2), which is

named for its principal instrument, was conceptualized as a geostationary Earth science

platform for geographic regional process environmental studies. A conceptual design of

the GHRMR spacecraft was done as part of the Global Change Technology Initiative

Architecture Trade Study. The GHRMR spacecraft is described in section 4.1.

The deployment, packaging, and assembly design concepts chosen for the GHRMR

are discussed in section 4.3 and section 4.4. They were selected in parallel with and

in conjunction with the spacecraft design. This follows the guideline established while

studying the in-space assembly of the ESGP spacecraft, that is, to incorporate the in-space

assembly requirement into the overall design process as early as possible.

15
4.1. GHRMR SpacecraftDescription

Althoughthe GHRMRis only oneof seveninstrumentsonthis platform,it dominates

theconfigurationof thespacecraftill that its largesize,offsetparabolicantennadesign,and

viewingrequirementsgreatlylimit theplacementof otherinstrumentsaswellasplacement

of the spacecraftbus. The GHRMR instrumentwas conceptualizeduring the GCTI

tradestudyby LaRCresearchersto meetthe requirementsof bothhighspatialresolution

and high accuracy.Its Cassegrain multiple reflector antenna design provides wide-angle

scanning to cover large portions of a hemisphere of the Earth from a geostationary orbit

position. The concept is composed of a 15-m-diameter primary reflector, a 7.5-m-diameter

secondary reflector, a moving tertiary reflector, and a phased-array feed system. The large

scanning angle requirement necessitates long focal lengths for the GHRMR reflectors, the

longest being around 30 m for the primary reflector. The overall spacecraft dimensions are

approximately 40 by 24 by 15 m.

Spacecraft designers developed two options for the structure of the GHRMR. Option a

(GHRMR-a) was a hybrid deployable/erectable concept which provided the greater capability of the two with a surface sufficiently smooth to operate up to 220 GHz, whereas

option b (GHRMR-b) was a completely autonomously deployable concept which provided

a surface accuracy sufficient for operation up to 90 GHz. Option b, however, completely

eliminated the need for in-space construction. The first option was based on precision

segmented reflector technology (ref. 9), which includes solid surface reflector panels and a

stiff lightweight supporting truss. This option would be deployed/assembled into its flight

configuration at low Earth orbit by using both EVA assembly and autonomous deployment steps. It subsequently uses a space transfer vehicle for transfer and insertion into its

operational geostationary orbit. The total mass of the GHRMR-a spacecraft is 6159 kg

(13 581 Ib).

The second option is based on a solid hexagonal panel concept developed by the Harris

Corporation (ref. 15) and was originally designed to operate up to 40 GHz with the

technology extrapolated to 90 GHz for the GHRMR application. This option is designed for

a single launch to geostationary orbit (GEO). It requires a propulsion stage (Centaur G r)

for geostationary transfer and orbit circularization. After GEO placement, it will be fully

deployed and made operational. The total mass of the GHRMR-b spacecraft is 5433 kg

(11 980 lb).

The complexities of packaging and autonomous deployment of either of these very large

spacecraft have a significant influence on the overall spacecraft design. The packaging

and subsequent deployment at GEO of the GHRMR-b spacecraft was chosen as the

primary strawman design task for investigating a large geostationary spacecraft having

complex packaging and autonomous deployment schemes. A preliminary packaging-options

examination for the partially erectable GHRMR-a spacecraft was also conducted.

Principal elements of the GHRMR spacecraft (whether a or b) are identified in figure 12.

At the lower end of the spacecraft is the 15-m-diameter primary reflector and its supporting

truss structure. It is connected via truss sections A and B to the feed and tertiary reflector

support bay. This bay in turn is connected by truss section C to the spacecraft bus assembly.

Each truss bay in a section is 2 m in length along each orthogonal direction. The bus

assembly houses spacecraft subsystems (e.g., power switching, communications equipment)

and also serves as a mounting platform for the other spacecraft science instruments. A onepiece hinged boom structure attaches the secondary reflector to the spacecraft bus.

16
Spacecraft bus and

other science instruments

7.5-m-diameter /

solid secondary

reflector and

support truss

Deployable truss+

section C

15-m-diameter

solid primary

reflector and

SU

Preassembled feed

and tertiary reflector bay

Deployable truss,

section B

Deployable truss,

section A

Figure 12. Principal elements of GHRMR spacecraft.

4.2. GHRMR Launch Vehicle Mass Considerations

The Titan 4/Centaur G' mass capability to GEO is 4536 kg (10000 lb) not quite

enough for the deployable GHRMR-b whose mass is 5433 kg (11980 lb). The mass

capability of the STS/Centaur G _ to GEO (before this program was canceled) was 6350 kg

(14000 lb). For this study, it was assumed that a Shuttle C (block 1)/Centaur G' and a

Shuttle C (baseline)/Centaur G _ would have at least the same or more mass capability to

GEO than the already sufficient, though canceled, STS/Centaur G/. The STS, however, can

easily carry either the GHRMR-a or GHRMR-b to LEO, where an STV could be utilized

for GEO transfer. The GHRMR packaging task, therefore, becomes a volume-constrained

problem, that is, to determine if GHRMR-b can be packaged in Shuttle C (block 1) or even

in Shuttle C (baseline) and to also determine if GHRMR-a can utilize the STS orbiter for

transfer of the assembly parts to LEO.

4.3. GHRMR-b Deployment/Assembly Concept

The deployment design concept for the GHRMR-b was undertaken first because portions

of it would be applicable to GHRMR-a. The spacecraft design was examined, and the major

functional elements were identified as the 15-m primary reflector, the 7.5-m secondary

reflector, the tertiary reflector and feed assembly, the spacecraft bus including the other

science instruments, and the supporting and connecting truss structures. Not shown in

figure 12 is the final major element, the geostationary propulsion stage (the Centaur Gt).

17
The descriptions of current and near-term launch vehicle payload envelope capabilities

(table I), including the Titan 4, the Shuttle orbiter, the proposed Shuttle C (baseline)

and Shuttle C (block 1), gave volume and shape bounds for the GHRMR-b packaging

envelope. These envelope restrictions indicated that accommodation of this spacecraft by

any current or near-term launch vehicle of reasonable size would require that deployable

technology be applied, as a minimum, to the truss structure and the large primary and

secondary reflectors.

Table I. Payload Envelopes

Payload to Payload a to

Launch vehicle Length, m Diameter, m to LEO, kg GEO, kg

Shuttle orbiter

Shuttle C (baseline)

Shuttle C (block 1)

Titan 4

18.3

25.0

27.0

18.9

4.6

4.6

7.6

4.6

25850

70000

60000

18600

_6000

>6000

>6000

4536

a In combination with Centaur G t.

4m

Stowed

truss

I

Unfolding

tl'tlSS

Figure 13. Single-fold, double-stowed truss concept.

After review of a number of deployable truss options, a deployable truss system

described as a single-fold, double-stowed design (fig. 13) was selected for the spacecraft truss

sections. Since this truss design is not self-deploying, it requires a deploying mechanism.

Although this mechanism was not specifically designed in this study, it was verified

that sufficient space was available for several different acceptable designs of a motorized

deployer. For the 1.25-cm-diameter struts chosen for this spacecraft, a stowed truss length

18
compression of approximately 98 percent was calculated (ref. 16) as achievable (i.e., a

deployed 100-m-long truss section would have a length of 2 m in its stowed configuration).

For the GHRMR, having truss bay lengths of 2 m, the 14-m truss section A (7 bays)

becomes only 0.28 m long when stowed. Likewise, the 20-m truss section B (10 bays)

becomes 0.4 m, and the 10-m truss section C (5 bays) becomes 0.2 m.

Figure 14. Hexagonal-panel solid-reflector deployment concept developed by Harris Corporation.

The truss bay containing the tertiary reflector and the feed assembly was used as a

starting point for the design of the deployment sequence (fig. 12). This bay, whether it is a

truss structure or a solid structure, would be preassembled, aligned, and checked out before

launch. The deployable truss sections B and C would be attached to opposite longitudinal

sides of the feed assembly bay. The spacecraft bus and secondary reflector assembly would

be attached to the opposite end of truss section C. At the opposite end of truss section B,

truss section A must be attached in an orthogonal direction. Since the truss bays are 2 m

in length on all sides, judicious selection of the truss section fold placements makes the

direct attachment of truss section A orthogonally to section B possible with deployment

in the required direction. The primary reflector assembly would then be attached to the

opposite end of truss section A.

A hexagonal-solid-panel concept developed by the Harris Corporation (ref. 15) for

extreme precision antenna structures was chosen as the method of assembly-compatible

implementation for the large primary and secondary reflectors. This concept is presented

19
an appropriateSTV. Options2a and 2b meetthe GHRMR-aconceptrequirementsof