Concept Development

SATELLITE
POWER
SYSTEMand

Evaluation Program

DOE/ER-O023(NASA-T,_-go_/S
CONCEPTDEVELOP8ENTREFERENCESYSTE_!SATELLITE
POWER
AND
EVaLUaTIONREPORT(_ational

Aeronautics
HC AI_/_F

Space

and
A01

N79-21538

SISTER:
PROGRAM,

Administration)

321
CSCL

p
10R

G3/4_

US Department

Referenceof Energy
and the

SystemReportNational
Aeronautics
Space
Administration

andUnclas19832

October
1978

Available

from:
National
Technical
U.S. Department
ofInformation
Commerce5285 Port RoyalRoad
Springfield,
Virginia
Price:

Printedcopy:
Microfiche:$ii.75
$3.00

22161

Service(NTIS)

DOE/ER-0023
Dist. Category UC-11,41,60,63,
63a, b,c,e,64,66e,95f,97c

Satellite

Power System

Concept Development
andEvaluation Program

Reference
SystemReportOctober 1978
(Published January 1979)

U.S. Department of Energy
Office of Energy Research
Washington,
D.C. 20545
and theNational Aeronautics andSpace Administration._..____

FOREWORD

The possibility
and transmitting
ing

the

it

followingofgenerating

large

to

earth

satellites

years,

National

Aeronautics

shortagesofment ofEnergyusing

severalstudies

1973 spurredinterest(DOE) (thentheEnergyan SPS Concept

evaluationisprogrammid-1977

studies

electricalconcept(NASA) and industry.
concept

and inrequired

This

which

guide

all

milestones

sub-

are:

June

will

Depart-a period

Final Program Recommendations

It

the

coversJanuary

direction."by thewhich

Updated Program Recommendations

thateffortDur-

Program.

May 1979

evaluation

1968.

Administration)

and Evaluation

DOE-NASA planspace

The energy

1976,

and Development

Development

The key program

early

Preliminary Program Recommendations

statesininwere conducted

October

plan

power

suggested

System Definition

The joint

ation

the

Researchby a joint

activities

Referencetheguidedto mid-1980.

and program

focus

of

in theof

was first

and Space Administration

and NASA initiatedfromquantities

theReferencein whatparticularlyisSystemconsideredemphasize

by DOE to conduct

1980

1980

selectionmilestone

to

be at

technical

environmental,

1978

that

time

and operationalsocioeconomic

"will
the

optimalinform-and comparative

studies.
This
It

reportdefinesissubmitted

a Reference

Systemin response
Concept

based

date.

The concept

presented

is

is

optimum

this

The system

not

at

indicated

technical

continue

to be analyzed

nitionand othe_time.

feasibility

studiesconsideredofand changed
warrant.

totheon thetomajorsystembe in thedefinition

the reference
as thefirst

results

definition

properstudies
conceptofprogrameffortto"direction,"

have,

and theprecedingmilestone.

but

however,
concept

will

systems

defi-

TABLE OF CONTENTS

Page
SUMMARY

1II.

INTRODUCTION

4

Ill.

REFERENCESYSTEM DESCRIPTION

I0

I

,

I0

A

Guidelines

B

System Overview

I0

C

Solar

Cells

and Blankets

14

D

Solar

Array

and Structure

17EPower Distribution

19

F

Rotary22

G

AttitudeControlH.

Microwave

Power Transmission

I,

Mass Statement

47

J.

Space Transportation

47

K.

Natural58

L.

Operations
I.
2.

M.

and Assumptions

Joint

22

SystemSystemResources

25

64

Construction
Commercial

Operation
76

Costs

IV.

TECHNOLOGYSUMMARY79

V.

DOCUMENTATIONSUMMARY83

APPENDICESA.

Systems

Analysis

A-I

Results

References
B.

Contracted

SystemI.

Rockwell2.

BoeingDefinition

StudiesInternational
Aerospaceii

Corporation

B-I
B-_2
B-66

LIST OFFIGURES

Page

Figure

3
4
5
6
7
8
9
I0
II
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26

SPSReferenceSystemConcept
Simplified Diagramof SPSConceptDevelopment
and Evaluation Methodology
Solar PowerSatellite
SPSReferenceSystem- Satellite Configuration
SPSEfficiency Chain (GaAIAsCR2and Si CRI)SPSOperations
Reference SystemCharacteristics
Solar Cell Options for SPS
Laser Annealing Concept
Solar Cell Blanket Support
SPSPowerDistribution
Antenna Yokeand Turntable
Attitude Control SystemCharacteristics
Transmitting AntennaFunctional Description
MicrowavePowerTransmissionSystemParameters
MicrowavePowerTransmission System
MicrowaveArray PowerDistribution
PowerDensity at Rectennaas a Function ofDistance from Boresight
PeakPowerDensity Levels as a Function of Range
from Rectenna
Grating Lobe Characteristics
Grating Lobe Maxima
PeakPowerDensity for Sidelobes and Grating
Lobe as a Function of Rangefrom Rectenna
Near-Field AntennaPattern
RectennaPatterns and PowerLevels
MicrowaveTransmission Efficiency
Noise PowerDensity of Groundfor a l km, 5 GW
SPSAntenna
iii

3
5
7
12
13
15
16
18
18
20
21
23
24
26
27
31
32
33
35
36
38
39
40
42
44
46

Page

Figure27

Heavy Lift Launch Vehicle

5O

28

Launcher/Erector

51

29

SPS Heavy Lift Launch Vehicle Trajectory andExhaust Products Data

52

30

Personnel Launch Vehicle

54

31

LOX/LH 2 Common Stage POTV

55

32

Cargo Orbit Transfer Vehicle _ Gallium Option

56

33

SPS Construction

65

34

Component Packaging Characteristics

67

35

Typical Component Mixing

68

36

Construction

70

37

Scenario for Buildup of Construction

38

Construction

39

Scenario for Construction40

SPS Operational Functions41

SPS Operations Management

Concept

and Commercial Operations

Base Buildup for Silicon System
Bases

Timeline for Two 5 GW Satellites/Yeariv

of Two 5 GW Satellites/Year

71
72
74
77

Concept

78

LIST OFTABLES

Page

Ta blel

SPS Mass Statement - Millions of KG's

48

2

GaAIAs Independent Electric COTV Mass Breakdown57

3

Si Independent Electric COTV Mass Breakdown

58

4

Materials List for Reference System

59

5Total Program Materials List

62

6

SPS Fleet Sizes

75

List of Abbreviations

AC

ALTERNATINGACSATTITUDE CONTROL SYSTEMACSSATTITUDE CONTROL AND STATIONKEEPINGAMO

AIRMASSAPU

AUXILIARY POWER UNIT

Ar

ARGON

BOL

BEGINNING OF LIFE

CMG

CONTROL MOMENT GYROCR

CONCENTRATION

CRM

CREW AND RESUPPLY MODULECWCONTINUOUS

db

DECIBELDC

DIRECT CURRENT

DDT&E

DESIGN, DEVELOPMENT, TEST AND EVALUATIONDOE

DEPARTMENTEMI

ELECTROMAGNETICEOL

END OF LIFEFETFIELD EFFECT TRANSISTOR

GaAlAsGALLIUM ALUMINUM ARSENIDEGEO

GEOSYNCHRONOUS

GFRTP

GRAPHITE FIBER REINFORCED THERMOPLASTICGHzGIGAHERTZ

CURRENTSUBSYSTEM

ZERO

RATIO

WAVE

OF ENERGYINTERFERENCE

EARTH ORBIT(lO 9 CYCLES/SECOND)

vi

GW

GIGAWATT (.lO
9 WATTS)

GWe

GIGAWATT ELECTRIC

HF

HIGH FREQUENCYH2

HYDROGEN

IMPATT

IMPACT AVALANCHE TRANSIT TIME

I

SPECIFIC IMPULSEsp

JPL

JET PROPULSION LABORATORYJSC

JOHNSON SPACE CENTER

KSC

KENNEDY SPACE CENTERkV

KILOVOLTkW

KILOWATTLaRCLANGLEY RESEARCH CENTER

LBF

POUNDS OF FORCELEO

LOW EARTH ORBITLeRCLEWIS RESEARCH CENTER

pm

MICROMETERMPD

MAGNETOPLASMADYNAMICS

MPTS

MICROWAVEPOWERTRANSMISSION SYSTEMMSFCMARSHALL SPACE FLIGHT CENTERMTBF

MEAN TIME BETWEENFAILUREMTMETRIC TONMWMEGAWATT(10 6 WATTS)

N

NEWTONNASA

NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONNMNAUTICAL MILES

vii

NS

NORTH_$OUTH

O&M

OPERATING AND MAINTENANCE

02

OXYGENPOPPERPENDICULAR TO THE ORBIT PLAN

psi

POUNDS PER SQUARE INCH

RP

ROCKET PROPELLANT

RPC

RECTENNA POWER COLLECTION

RFRADIO FREQUENCY

RFIRADIO FREQUENCYSCBSATELLITE CONSTRUCTION

BASE

SECS

SOLAR ENERGY COLLECTIONSYSTEM

SE&I

SYSTEMS ENGINEERING AND INTEGRATIONSGSWITCHGEARSiSILICONSPS

SATELLITE POWER SYSTEMVAB

VERTICAL ASSEMBLY BUILDING (AT KSC)

VHF

VERY HIGH FREQUENCY

VLF

VERY LOW FREQUENCYINTERFERENCE

viii

Transportation

BLOW
CH4

Systems

BOOSTER LIFT OFF WEIGHTMETHANECOTVCARGO ORBITAL TRANSFER VEHICLE

EOTVELECTRIC ORBITAL TRANSFER VEHICLEET

EXTERNAL TANKGCRGAS CORE REACTORGLOWGROSS LIFT OFF WEIGHTHLLVHEAVY LIFT LAUNCH VEHICLEHTO

HORIZONTAL TAKEOFFIOTVINTRA-ORBIT TRANSFER VEHICLE

LCH4

LIQUID METHANE

LH2

LIQUID HYDROGENLOXLIQUID OXYGEN

MOTV

MANNED ORBITAL TRANSFER VEHICLEOTV

ORBITAL TRANSFER VEHICLEPLVPERSONNEL LAUNCH VEHICLEPOTVPERSONNEL ORBITAL TRANSFER VEHICLESRB

SOLID ROCKET BOOSTERSSBESPACE SHUTTLE BOOSTER ENGINESSMESPACE SHUTTLE MAIN ENGINESSTO

SINGLE STAGE TO ORBIT

T/W

THRUST TO WEIGHTULOWUPPER STAGE LIFT OFF WEIGHTAV

DELTA V - CHANGE IN VELOCITY

ix

I.

SUMMARY
The SPS Reference _ystem concept as defined

herein, is largely a product ofsystem definition studies conducted by the Boeing Aerospace

Company under con-tract to the Johnson Space Center (JSC) from December 1976 to December 1977,
and Rockwell International under contract to the Marshall Space Flight Center(MSFC) from March 1977 to March 1978.

The results of these two system definition

studies combined with in-house efforts at both NASA Centers and several smaller
contracted studies provided the data from which the Reference System wasdeveloped.
The two parallel, but independent, system definition
well-integrated

studies resulted insystem concepts in which some major elements were very similar,

but others were markedly different.

To meet the Reference System milestone

with essentially a single concept, a third system was developed by the selectionof system elements of the two individual concepts.
Part III of this report describes the Reference System.
the various systems analyses that have been conducted.

Appendix A describes

Appendix B describes theindependent system definitions developed by Boeing and Rockwell.
The Reference System description emphasizes technical and operational
mation required in support of environmental,
ment studies.

socioeconomic,

infor-and comparative assess-

Supporting information has been developed according to a guide-line of implementing

two 5 GW SPS systems per year for 30 years beginning with an

initial operational date of 2000 and with SPS's being added at the rate of two
per year(lO GW/year) until 2030.

Figure l illustrates the Reference System concept, which features galliumaluminum-arsenide(GaAIAs) and silicon solar cell options.

The concept utilizes

a planar solar array (about 55 km 2) built on a graphite fiber reinforced thermoplastic structure.
centration),

The silicon array uses a concentration

whereas the GaAIAs array uses a concentration

kilometer diameter phased array microwave antenna

ratio of one (no conratio of two.

is mounted on one end.

A oneTheantenna uses klystrons as power amplifiers with slotted waveguidesas radiating
elements. The satellite is constructed in geosynchronousorbit in a six-month
period. The ground receiving stations (rectenna) are completed during the sametime period.
The other two major componentsof an SPSprogramare (1) the construction
bases in spaceand launch and mission control bases on earth and (2) fleets ofvarious transportation vehicles that support the construction and maintenance
operations of the satellites.
Thesetransportation vehicles include HeavyLift
LaunchVehicles (HLLV), Personnel LaunchVehicles (PLV), CargoOrbit Transfer
Vehicles (COTV),and Personnel Orbit Transfer Vehicles (POTV). The earth launchsite chosen is the KennedySpaceCenter, pending further study.

2

II.

INTRODUCTION
The DOE and NASA have initiated a joint program which is describeddocument entitled SPSConcept

Development andEvaluationin theProgram Plan

(July 1977 - Auqust 1980), DOE/ET-O034, dated February 1978.

The objective of

the program is to generate information from which a rational decision can bemade regarding the direction of the SPS program after fiscal year 1980.
Briefly, the plan states that NASA will conduct systems definition of the SPS
and the DOE will evaluate health, safety, and environmental factors and will
study SPS economic, international,

and institutional issues.

In addition,

DOE will make comparative assessments of the concept relative to alternative
power sources for the future.
Figure 2 shows a simplified diagram of the development and evaluation methodology.

As indicated, the major milestones of the plan are baseline concept(s)selection-

October 1978; preliminary program recommendations

dated program recommendations
June 1980.

- May 1979; up-- January 1980; and final program recommendations

In this report, the term Reference System is used instead of base-line concept(s) as being more appropriate for the current level of definition
and understanding.

Using the results of this evaluation

program as a basis

and considering other pertinent factors, it will be possible for the Administration to either recommend continuation of the program in accordance witha defined option or terminate the program.
The purpose of this document is to present a description
System.

of the Reference

It is submitted in response to the Baseline Concept(s) Definition

program milestone(October 1978) established in the DOE/NASA plan as indicatedabove.
Section III of the report presents a description
with emphasis on technical and operational
duct environmental,of the Reference System

information required by DOE to con-socioeconomic, and comparative assessment

studies.

recognized that the Reference System lacks maturity as reported herein.
tion work is continuing to develop further understanding of the system.

It isDefini-

-

JULY
1977

t
SYSTEMS
DEFINITION

L
_

[

t

.......
'.....
I

SO¢IOECONOIdD¢

EVALUATION/OCTOBER
1978

COMPARATIVEASSESSMENT

[

I

MAY

JANUARY
Im0

JUNE
1R0

Figure 2.

Simplified Diagram of SPS Concept Development
and Evaluation Methodology

Section

IV provides

a discussion

which

focuseson criticalquestions

Also,

alternateresults

technologies

section,are

A summary ofreaderthat

inlocating

Appendix

over
of

the

A provides

to datein theanalysissummaries

consists

studies

as wellofearly

The Jetmarysourcesas those

information

systems

definition

reports

space

Company underby theSPS feasibility.

in thesummary

of

study

in

V to assist

documentation

summaryforreport.
analysis

results.

The information

thissystemsthekey design

of

study

reports

and otherlate

1960's

by A. D. Little,
and themade significant

in theareaof microwave

for

the

Reference

to

study

studies

The database forthedocuments

fromcurrent

1970's.

Muchand early

Corporation,

NASA Lewis

power

by Rockwell

Spectro-Research

contributions

Systemresults

trade-offRaytheonalso

publishedofand operational

investigation.

in theCorporation

flow

the

This

of

contract

Sectionyears.

areas

B provides
Boeingprovidedmajor

Laboratoryofductedispreparedparticularly

AppendixaffectseveralofGrumman Aerospace

bank,

that

the

past

based on results

Propulsion

requirements

and following

work was performeddata

are delineateda summary of

conducted

Inc.,advancement

be resolved

documents

references

is

lab,

technology

reviewed.

presented

the

which

referenceas a listto

SPS documentation

has occurred

servesofCenter.

tothetransmission.

description

The priareInternational

SPS

the

and Boeing

Aero-

MSFC and JSC, respectively.

systems

Aerospacedescriptions

resultingCompany and Rockwell

from

the

studiescon-International.

BackgroundThis

document

in figure3.

electrical

power

power

to

microwave

earth

It

deals
is

solar

the

solar

electricalenergy)focusedmicrowave

usefulelectricity

1968 ("Powerfrcm

beams into

was suggested

inScience,

162,

Vol.

a primary

(from

via

with

November

22,

power

satellite

powersourceconcept
thatin geosynchronous
beams,

the1968,

and collecting

on theearth'sSun:

Future,"Its

pp 857-861).

as illustrated

involves

orbit,

generatingtransmitting

theand converting

thesurface.
Dr.

This
Peter

concept

E. Glaser,

u')

°p-

Q)
or)
_..

_1

0
r_

0
I-I-0
Z

E_
I,

7

A number of other potential energy system concept involving the use of
space satellites have also been suggested:
I.

Orbiting nuclear reactor power systems (in lieu of solar collector/

converters} with microwave
2.

transmission of power to earth.

The power relay satellite concept (Reference 12} in which powersystems on the earth's surface or in low orbit transmit power by microwave togeostationary

satellites.

These geosynchronous

satellites

then relay (reflect}

the microwave energy to ground stations placed at desired locations on earth.
3.

Solar reflecting satellites(mirrors) in earth orbit which reflectsolar energy directly to earth to augment ground-basedsolar power plants,

allowing night operation or increased output.
4.

Laser power transmission(in lieu of microwave}

from the satellite.

The orbiting nuclear reactor concept has been evaluated to a limited
extent.

While it might offer the advantage of compactness relative to solar

powered systems, its mass and complexity are significantly greater than solar
powered systems.

Significant safety and environmental questions remain to beaddressed.
The power relay satellite is a long distance power transmission
rather than a primary electrical energy source; consequently,

concept

it is not viewed

as a basic alternative to the SPS concept.
The idea of placing large mirrors
erence 14).

in earth orbit has been evaluatedAnalysis indicates that with a mirror in geosynchronous(Ref-orbit, thesmallest illuminated "spot" on the earth would be about 330 km (205 miles) indiameter, governed by optical geometric considerations.

If continuously

illumi-nated at an average level of one sun, this large area would tend to rise intemperature to approximately

150°F, posing severe environmental

problems.

Placinglarge mirrors in lower altitude orbit reduces the size of the illuminated"spot."

However, the mirrors would not be stationary with respect to a point on the earth.
Thus, to achieve continuous illumination at a given location, numerous mirrors
would have to be placed in low orbit.

In addition,

cloud cover and weather con-ditions would have an adverse effect on solar insolationprecluding consideration

of this concept as a primary, baseload electrical

power source.

Laser power transmission has a significantly lower efficiency for longdistance power transmission than is estimated for microwavepower transmission.
Atmospheric attenuation is substantial comparedto microwavefrequency transmission. Therefore, this concept is presently less attractive than the microwaveconcept for transmitting powerfrom geosynchronousorbit.
Alternate system
concepts such as solar collectors and laser transmitters in low earth orbit
with relays in geosynchronousorbit havereceived preliminary consideration.
Another SPSconcept is being evaluated that makesuse of materials derived
from the moonto construct the SPS. The moon's lower gravitational force (onesixth of earth's) would allow muchless propulsion energy to movepayload togeosynchronousearth orbit.
This idea appears to have merit in terms of conserving earth resources and possibly reducing the cost of space transportation;
however, it would require developmentof moon-basedmining, manufacturing andlaunch facilities.
Consequently, the research and developmentrequirements
for such an approachwould be greatly increased.
While the above options offer interesting possibilities,
the present DOE/
NASAprogramfocuses evaluation on the SPSconcept using terrestrial materials
and deployed in geosynchronousorbit as illustrated in figure 3. This evaluation
does not exclude the possibility of future consideration of the alternatives andoptions such as identified above.

III.

REFERENCESYSTEM DESCRIPTION
The purpose
has evolved

of

this

primarily

sectionfromissystemto describe
definitionthe

SPS Reference

studies

conducted

by Boeing

Aero_

concept

presented

hereinspace

Company and Rockwell

International.

The system

is

the

resultnumerous

trade-offand engineeringaresummarizedin Appendix

A ofsystemofdescribed

Evaluation

hereinby bothSystem evolvingA.

iscontinue

to

which

be emphasized

that

the

in

detailinsome areas.

progress

with

the

Referencedetailsaredeveloped.

and assumptions

utilizedinthelistedo

SPS operational

date

2000.

o

Rate ofReference

Systemdefi-

per year;

300 GW total

below.
is year

implementation

costingis

two 5 GW systems

purposes.

oAllgroundrectennaso

SPS operation

sized

for

5 GW.

in geosynchronous

orbit.

Systems

operatingoMicrowave

power

and edge,

respectively,

of

o

All

derivedoSystem lifeis

30 years

o

Zero

rate

failureoTechnologyoNo costo

Cost estimatesinoSystem weight

growthmaterials

launchfrequency
density

is

not

2.45

GHz.

to exceed

23 and 1 mW/cm2 at

centerrectenna.
fromavailability

margins

earth

with

no salvage

value

or disposition

costs.

assumed.
date

will

resources.

is

1990.

be used.

1977 dollars.
factor(25%) to

sized

for

be reflectedin costs.

System Overview
The Reference

ventional

shouldand incomplete

as further

o

B.

It

analyses,and Assumptions

are

for

report.

preliminary

and maturing

Guidelines

capacitythis

DOE and NASA will

The guidelines
nitionstudies

System whichpower grid.

Systemis

The satellite

5 GW DC powerhas one end-mounted

I0

outputantenna

into

a con-

whichtransmits

to a rectenna on the ground.
The configuration

This concept is illustrated in Figure 4.

of the satellite consists of a planar solar arraystructure built from a graphite composite material.
presented.

Two conversion

One is the single-crystal gallium-aluminum-arsenide

with a concentration

ratio of 2 as illustrated in figure 4.

options are(GaAIAs) solar cells

The other energy

conversion option is the use of single-crystal silicon (Si) solar cells with noconcentration.
The size of the solar array is dictated primarily by the efficiency chainof the various elements in the system.

Figure 5 shows the end-to-end efficiency

chain for the GaAIAs and silicon solar cell options.

With the satellite

designed to provide 5 GW of DC power to the utility busbar and an overall
efficiency of approximately 7%, it is necessary to size the solar arrays tointercept approximately
quoted70 GW of solar energy as indicated in figure 5.

efficiency is the minimum efficiency, including the worst-caseThesummersolstice factor (0.9675), the seasonal variation(.91), and the end-of-lifeyear) solar cell efficiency assuming annealing.

For the GaAIAs case, the end-of-life(30 year)concentrator reflectivity

is 0.83.

Since only half of theintercepted solar energy is reflected by the concentrators,
efficiency(30

the equivalent overall

is 0.915.

The GaAIAs option is a five-trough configuration with a solar blanket
area of 26.52 km2, a reflector area of 53.04 km2 and an overall planform area of
55.13 km2.

The silicon option has the solar blanket with no concentration

resultingin a blanket area of 52.34 km2 and a planform area of 54.08 km2.

The satellite in either option is oriented so that the antenna main
rotational axis remains perpendicular to the orbital plane.
The end-mounted microwave antenna is a one kilometer diameter phasedarray transmitter.

The phase control system utilizes an active, retrQdirectivearray with a pilot beam reference for phase conjugation.

Klystrons are used as

the baseline power amplifier with slotted waveguides as the radiating element.
The ground rectenna has subarray panels with an active element area of 78.5 km2.
The satellite is constructed in geosynchronoustime being six months.

orbit with construction

The initial estimates of construction

IIcrew size are 555

C
0

0

°_,,.

m

a,)
U
e-"
I.-

%..

°¢..,.

12for the silicon option(480 in GEO and 75 in LEO) and 715 for the GaAIAs

option (680 in GEO and 35 in LEOn.
The transportation

system is made up of four major items.

clude: (1) the Heavy Lift Launch Vehicle
VehicleThese in-

(HLLV), the Cargo Orbit Transfer(COTV), the Personnel Launch Vehicle (PLV), and the Personnel Orbit

Transfer Vehicle (POTV).

The HLLV is a two-stage, vertical launch, winged,

horizontal land-landing, reusable vehicle with 424 metric ton payload to low
earth orbit.

The earth launch site was chosen as Kennedy Space Center pending

further study.

The COTV is an independent, reusable electric engine-powered

vehicle which transports cargo from(LEO) to the geosynchronous

the HLLV delivery site in low earth orbit

earth orbit (GEO).

For the GaAIAs SPS option,

the COTV is powered by GaAIAs solar cells, whereas a silicon solar cell power
supply is used for the silicon SPS option.
Personnel for the orbital construction

and support functions are trans-

ported to LEO via the PLV which is a modified space Shuttle Orbiter with apassenger module.

The POTV, a two-stage reusable, chemical fuel vehicle isused to transfer personnel from LEO to GEO and return to LEO.
The satellite construction
trated in figure 6.

scenario for the Reference System is illus-

The HLLV is shown transportingcargo to LEO while the

COTV and the POTV are illustrated transporting cargo and personnel,respectively,from LEO to GEO.

A LEO operations

supplies and propellant.

base is used for temporarystorage of

One satellite is shown in GEO during the con-struction phase while another satellite is shown in the conventional
phase transmittingenergy to ground rectenna.

Figure 7 summarizes

operational

the character-istics of the Reference System.
C.

Solar Cells and Blankets
Both GaAIAs and single-crystal

reference energy conversion devices.

silicon solar cells are considered
Figure 8 shows a cross-section

GaAIAs and Si solar cells and blankets.

of theThe basic GaAIAs solar cell consistsof a 5 _m thick GaAs-P,N cell with a 0.03 to O.05)qm thick GaAIAs front-side
window.

The solar cell efficiency

is 20% at AMO, 28°C.

14

The design operating

REPRODUC131TJITY OF THE
ORIGINAL PAG2 IS POOR

15

SPS generation

capabiliW

Overalldimensions

Power

converzion-photovoltaic

Satellite

Mass (Kg)

Structurematerial

Construction(utility

interface)

5 GW

5.3

(Kin)

x 10.4GaAIAs34(CR=2I

X 106

51

Graphitelocation

Silicon(CR = 1 )X 106

composite

GEO

Transportation••

Easth-to-LEO-CargoVertical(payload)(424,000Kg)-Personnel

Modified(Number)(75)

4;argoDedicated

LEO-to-GEO

Microwave-Personnel

2-stage(Number)

(75)

power

take-off,

shuttleelect.

LOX/LNtransmission•

No.

of antennas

•

DC-RF

1

•

Frequency

(GHZ)

•

Rectenna

dimensions•

Rectenne

power

conwerter

Klystron
2.45(Km)

density

10 x 13(mw/cmCenterFiqure 7.

2)
23

Edge

1

Reference System Characteristics

16

wingedOTV

22-stagetemperature

is 125°C, which produces an 18.2% cell efficiency.

self-annealing of radiation-induced

At 125°C,

damage occurs in the GaAIAs cells.

As

indicated in figure 8, the solar cell blanket material is Kapton (25_m
The 20 _m

thick).

synthetic sapphire (Al203) substrate, used in an inverse orientation,

also acts as the cell cover.

The blanket weight is 0.252 kg/m 2.

The projected

cost of the GaAs solar cell blanket is $71/m 2.
Solar reflectors for the GaAIAs CR2 concept consist of thin reflective membranes tIz.5 Mm)of aluminized Kapton mounted

in a 60e Vee trough configuration.

reflector membrane has a reflectivity of 0.9 BOL and 0.83 EOL.
life efficiency is 0.915 as previously stated.

thick single crystal siliconsolar cells with borosilicate cover glass electrostaticallybonded to the cells

The cells are designed with both P and N terminals brought tothe back of the cells.

This feature makes it possible to use 12.5_m

plated copper interconnections

interconnections.

silver-which are formed on the substrate glass.

panels are assembled electricallyA_design

The effective end-of-

The membrane mass is 0.018 kg/m 2.

The silicon solar blanket consists of 50_m

front and back.

The

Completeby welding together the module-to-module

The cell efficiency is 17.3% (AMO, 28°C) at beginning of life.

operating temperature of 36.5%,

the efficiency

is 16.5%.

A key

feature of the blanket design is the ability to perform in-situ annealing of thesolar cells using a laser annealing concept.
gimbal-mounted

CO 2 lasers.

The laser annealing concept utilizes

The gimbals would be mounted on an overhead gantry

structure to permit transversing of the entire solar array by several lasers
as illustrated
temperature
Annealingin figure 9.

The laser beam heats

the solar cells to annealing(500°C) without damaging cell interconnect

and substrate materials.

is required to recover radiation induced degradationof the cells.

The

projected cost of the silicon solar blankets is $35/m 2.
D.

Solar Array and Structure
The solar array consists of the deployed solar cell blankets attachedto solar array structure.

In the case of the CR2 GaAIAs option, the array in-cludes the solar reflectors mounted on each side of the solar cell blankets.
The solar cell blankets would be transported to orbit in modular packages

17inGroovesaroundrefract
grid fin(ligh

0 /._m Sapphire
Substrate,u,m GaAIAs.03--.05
5 /_m GaAs

75 ,u.m Borosilic
glass

13 ,u.m FEP
25 ,u,m Kapton(blanket)

50 /.Lm Silicon(textured)
50 ,u.m Borosilicate
glass substrate

6 /._m Polymer
thermalcoating
Gallium
Arsenide

SiliconAluminum
Solar Cell

Figure 8.

Satellite

Module,, , __,__,____I I I I II I I I i I'_ , I _7

Solar Cell Options

Solar

Cell

for SPS

Laser Scanner
44 per_,_Gantry_H_FpA_LF,_141,,,,-.-'"'
I_liiiiiiiiii_,-T(I

t-I

I I I I I

5,348m

/.

660m

I

-

_2,678m

/

_-"_7
/19

_/__

_Gantry

Typica1667.5m
Bay_

TypicalAnnealer

Track///, --_15m:J/_//_

667.5mGimballed

II

./1_

.....

Scanning

II

_

/

Laser

,":" "_--._
Figure 9.

Laser Annealing

18

Laser

Concept

_.,._
__

__/

eithera rolledor folded

configuration.

nectedtogetherin a parallel/series

outputof

40 to

45 kv.

solar

cell

blanketsarrangement

FiguretotheThe blanket

I0 illustrates

supporttostructure.

fashion.

The reflector

panelspleatedattoan accordianfoldshipment

to orbit.

The reflector

structureby cables(catenaries).

isThe basic

elements

fabricationin space.

but

would

solar

includeThe CRI siliconconstructedwithconstruction

additionalforprimary

truss-type

rein-

designedfor

optionwould

beams.

The CR2 GaAIAs

elementsfortheto formthestructureelements

are

solar

array

and

antennaa graphite-fiber(beams)

cell

the

are

storagetransmitting

composite,berolledand microwave
issiliconGaAIAs option

arraythermoplastic

similar

would

solar

material

utilizes

the

the

to

The structure

optionforvoltage

attaching

be attacheddesign.

configuration

desiredwould

an open truss-type

rectangularofbe con-panelsstructure

automatictheand then

The primaryforcedfor

would

The GaAIAs blankets

in a similar

produceobtainthe method

attached

intervals

modulesutilizeastructure,forattachingistoreflectors.
E.

Power Distribution
The primeand control

function

prime

of

power from

regulatethequantity

microwave

generators;thepower distribution

the

solar
of

the electricalprovidetherequiredforocculationor system

maintenance

fault

detection,and fault

isolationIIand distribution

panels;and quality

energyFigurecellsystemcontrol,powerenergyshutdown;

power

sectors.

dividedswitchable

and can be isolated

from

Solar

array

annealing

the

solar

power

is

cell

during

array

array

solar

storage

for

solar

The solar

tatinggeneratedmonitoring

disconnects.

system.
is

and

for

diagramsectorcondition,and provide

shows a schematic

power

accumulateofoperations(forcontrolled

by highthesilicon

into
the

main

Each

bus,

facili-otherservic-circuitbreakers

nearstrings

on or off,

depend-cells)

power

collection

and/oring.

thebuses.

Voltage

power
is

is

controlled

by turning

ing on loadrequirements.

Two sections

age atsliprings

therequiredthevoltage

using
at

the

sheet

of

groupstheconductor

sliprings.

19

voltageofarrayvoltageprovide
drop

the
to

requiredachievevoltthe/TENSIONING
n

m

660m
i

-

--I

14.9

I--- lo.o.,,Figure I0.

Solar Cell Blanket Support

2_

m

--

J=

=__

,..__

.'-'

--_- o__=•" _ 0 0_2: O_.-m _u3 _..-_'ID
LrJc_i _ m

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The microwave power transmitting antenna includes a power distribution
system which transmits DC power from the slipring System to the DC-RF generators.
Conductors from the slipring brushes are tied to DC/DC converters
switchgear to allow isolation when performing maintenance.

through

Conductors

are thentied between voltage summing buses through other switchgear for transmitting
the required power to the klystrons.
An electrical energy storage power system is located on the antenna
structure with a bus routed along the regular network for operation during
powered-down periods such as may occur during solar eclipse periods.

A

battery energy storage system would be utilized for storage of about 12 MWhours of electrical energy.
F.

Rotary Joint
Power transfer from the solar array section to the microwave

antennais accomplished via a rotary joint (figure 12) using a slipring/brush assembly.
Mechanical rotation and drive is provided by a mechnicaldiameter.

turntable 350m

in

The antenna is suspended in the yoke by a compliant mechanicalto isolate the antenna from turntable vibrations.
aimed by control moment gyroscopes
feedback with a low frequency

joint

The antenna is mechanically(CMG's) installed on its structure.

band-pass allows the mechanical

A positive

turntable to drive

the yoke to follow the antenna and also provide sufficient torque through thejoint to keep the CMG's desaturated.
G.

Attitude Control System (ACS)The attitude control system for the reference 5 GW system is describedfor a CR2 GaAIAs option.

The ACS for the CRI Si option will be similar, although

the number of thrusters and propellant requirements will differ.
Preliminary Baseline ACS Description - GaAIAs-CR2- The baseline ACS

features an argon ion bombartment thruster reaction control system (RCS) whosecharacteristics are given in figure 13.

A total of 64 thrusters is includedin the system to provide the required redundancy

assuming: an annual maintenanceinterval, 5000 hour thruster grid lifetime and a 5-year thruster MTBF,

22

Approxi-

_I,_PIIOOUC!B!T,!TY

OP TItt_

(1)
e-tO

tO
(1)
v
o
>-

c-(1)
c--

r--"

\\

0_.LL.
ill

mO

z
"_a
zm
u_9
_-z

23

\

ATTITUDE REFERENCE DETERMINATION (7 LOCATIONS)• CCD SUN SENSOR (l PER SYSTEM)• CCD STAR SENSORS (2 PER SYSTEM)• ELECTROSTATIC OR LASER GYROS (3 PER SYSTEM)• DEDICATED MINIPROCESSOR

\
\,

\

\
\\

./
J
jj,J

_J

REACTION CONTROL SYSTEM (4 LOCATIONS)• ARGON ION BOMBARDMENT THRUSTERS• CRYOGENIC PROPELLANT STORAGE - ELECTRICREFRIGERATION FOR HEAT LOSS MAKEUP• HEMISPHERICAL PLUME CLEARANCE• SERVICEABLE IN-PLACE
THRUSTER CHARACTERISTICS: EACH
• THRUST - 13N• SPECIFIC IMPULSE - 13,000 SEC.
• POWER - 1,275 kW• APERTURE - IM
• MASS (INCL. SUPPORTS & CABLING) - 120 kg• RESTART TIME - 15 SEC.
• OPERATING LIFE (GRIDS & CATHODES)>5000 HR.

Figure 13.

Attitude Control System Characteristics

24

mately 36 operating thrusters are required. Sixteen thrusters are locatedon the lower portion of each corner of the collector.
Eachthruster is gimballed individually to facilitate thruster servicing using a servicing cab_ topermit operation of adjacent thrusters during servicing, and to provide
redundancy.
The system is nominally designed for X-POPoperation (.long axis per-L
pendicular-to-orbit-plane).
The pertinent features and locations of the Attitude Reference Determination System(ARDS)are also given in figure 13. The average
power required for the system is 34 megawatts.
H.

Microwave

Power Transmission

The reference microwave
sidering environmental
heating limitations,

Systempower transmission

system was developed con-effects, antenna size tradeoffs, antenna thermal

and ionospheric heating effects.

The present microwave

system has DC-RF power converters feeding a l km diameter phased array antenna with a lO-decibel (dB) Gaussian taper illumination

across the array sur-face.

This antenna

matelylO meters X lO meters on a side, having slotted waveguides

radiatingshown in figure 14, is composed of 7220 subarrays, approxi-

surface with DC-RF power tubes mounted

subarrays.

The antenna structure is a graphite

as theupon the backside of thecomposite material while theslotted waveguides are aluminum.
Each subarray has its own RF receiver and phasing electronics to process a pilot beam phasing signal from the ground.

The subarrays are phased

together to form a single coherent beam focused at the center of the groundantenna/rectifying

system (rectenna).

This power beam has approximately 88%

of its energy within a 5 km radius perpendicular

to boresight, with a resultant

beam width of 1.2 arc-minutes.
Microwave System Parameters. and sizing Considerations- Some of the

key parameters of the SPS microwave system are presented in figure 15.

The

power capability of the SPS system was sized by: (1) thermal limitations of
22 kW/m2 in the transmit array due to waste heat from the DC to RF power converter tubes; (2) a peak power density limitation

25

in the ionosphere of 23"_mW/cm2

(.0°l-*e-iI.)
i.¢i
a)
t-_

%
z'0
*r.4-)
0"-i
I,.Ic'-

c'-

c°_

I#'l
el.

,dl-,!
°I-LL

26

v-

E

E

to

.rE
cIt

I_l

o

C_
0
U
o)B
,_===

_4
%.=
-=_
°_
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R,.I_T:.:. _i :tli('.: q! ,WY OP THB

27

°t-

O

E

E

0
,i,m
-rE

0

t_
0
(_)

LL.

28

to prevent non-.linear heating interactions betweenthe ionosphere and themicrowavebeam;and (3) microwavetransmission efficiency considerations(particularly the RF levels into the rectenna).
I

Studies into the microwave

beam/ionosphereinteractions

indicatethat non-linear thermal self-focusing instabilities in the F-region (200 to
300 kilometers altitude) and thermal runaway conditions
kilometers)

may

approximately

in the D-region

limit the maximum power density at the center of the beam to

23 mW/cm 2 at the 2450 MHz operating frequency

29, 30 and 31).

(lO0(_references 26,

Above this threshold power density level (which is a theo-

retical number, not yet verified by experiments), non-linear
between the power beam and the ionosphere begin to occur.

interactions

These non-linear

heating interactions are of concern because of possible degradationsing HF and VHF communication and VLF navigationand multipath degradations.
jitter and/or differential
ionosphere/microwave

to exist-systems due to RFI effects

The heated ionosphere may also introduce phase
phase delays on the uplink pilot beam signal.

Thesebeam interactions are now being studied, both theoreticallyand experimentally.
Antenna Characteristics , The microwave

antenna has both a primary anda secondary structure composed of a graphite composite material.

The primarystructure is an open truss, 130 meters deep, with an octagonal shape over lO00
meters in width and length.

The secondary structure is a deployable

cubic

truss, 9.93 meters in depth, which provides support for installation of the micro_
wave subarrays.
The aperture illumination

function across the l-kilometer transmit array

was optimized to provide the maximum amount of RF power intercepted by the groundrectenna and to minimize the sidelobe levels.
nation functions, operating
element

A number of different

illumi-in the presence of phase and amplitude errors and(subarray) failures, have been studied (see Appendix A).

The lO-

decibel Gaussian taper has the best overall performance of the optimized

illumi-nation functions after considering the maximum power density constraints

in thetransmit array and rectenna.

29

A I0 step, I0 dB Gaussiantaper for the transmit array is given in
figure 16. There are 36 DC_RF
pQwerconverter tubes per subarray at the centerof the antenna_decreasing in quantized steps downto four tubes per subarrayat the edge to provide the lO dB Gaussiantaper. There will be a total of
I01,552 tubes in the antenna, integrated into the subarrays as shownin
figure 17. This particular configuration uses a 70 kw klystron tube_ an alters.
native concept has a 50 kw klystron, which requires approximately 140,000 tubes.
The numberof tubes per subarray would then vary from 50 tubes at the center of
the antenna to six tubes at the edge in order to provide the lO dB illumination
taper.
The radiation pattern for the lO dB taper, l kmarray (and o-= lO° RMS
phase error; _l dB amplitude error, and 2%randomfailures) at the ground rectenna is shownin figure 18. The effect of the antenna errors is to produce awider, lower intensity main beamwith higher sidelobes. For the SPSsytemconcept, only a portion of the main lobe will be collected; the sidelobe energy
occupies a very large area at very low density levels and is not economicallyfeasible to collect.
The peakpower densities are 23 milliwatts per square centimeter at
the rectenna boresight, l milliwatt per square centimeter at the edge of therectenna, and 0.08 milliwatt per square centimeter for the first sidelobe,
which is two orders of magnitude below the U.S. radiation standard of lO mW/cm
2.
If there is a total failure within the phase control system (for example,
the uplink pilot beamtransmitter is shut off), the subarrays will no longer
be phased together and the total beamwill be defocused. As shownin figure 18,
the peak intensity of the beamdrops to 0.003 mW/cm
2 and the beamwidth greatly
increases. This peak powerdensity is significantly less than the USSRguideline indicated on figure 18. Consequently, this is a fail-safe feature of thephasing system. In addition, there are sensors near the rectenna to detect any
large changesin incident powerdensity; this information would immediately betransmitted to the antenna to cease operations.
In addition to the sidelobe patterns near the rectenna, the far-sidelobe
patterns havebeen calculated. There had been someconcern about the radio
30

_o__

REPE?:;.::.;_gTTJTYORIGP.:A;__" .......

o_

OF THt_

el

o

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co

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.,_-

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31

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32

5O

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-

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No pilot beam present| ||...............

t"1"..........

|--,'. .... ...I. .... &-4- ..........

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I

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I(t°tal phase c°ntr°l failure)i |

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0

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5000

10 000

i_l
15

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_

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20 000Ground radius, mFigure 18.

Power Density at Rectenna as a Function of Distance from Boresight33

interference levels at large distances from any given rectenna becauseof
frequency allocation problems. The SpSdownlink power beamlies in the
2400-2500MHzfrequency band which has been reserved for Governmentand
non-Governmentindustrial, medical_ and scientific (IMS) usage. By definition,anyoneoperating in an IMSband must accept interference from any other user
within this band. However,this lO0 MHzband is not recognized by someof
the eastern Europeancountries which reserve 2375MHz+ 50 MHzfor the IMS
band. The far sidelobe levels as shownin figure 19 indicate the peaklevels for one 5 gigawatt SPSsystemare three to four orders of magnitude
below O.Ol mW/cm
2. For simultaneous operation of 200, 5 gigawatt SPSsystems,
the average peak level is still
O.Ol mW/cm
2.

one to two orders of magnitude lower than

Grating lobes, which occur at 440 kmintervals from the rectenna, arefunctions of subarray size and mechanical misalinement of the subarrays within
the l kmphase array. The grating lobes occur at spatial distances corresponding to angular directions off axis of the antenna array where the signals fromeach of the subarrays add in-phase.
Whenthe boresights of the subarrays are not alined with the uplinkpilot beamtransmitter at the rectenna, the unwantedcontributions of thearray factor of the antenna do not lie in the null-points of the subarray pattern as shownin figure 20. Eventhough the phase control systemwill still
point the composite beamat the rectenna, someenergy will be transformed fromthe main beaminto the grating lobes. The amountof energy in the grating
lobes dependsupon the misalinement (or howfar the array factor is displaced
from the null points of the subarray pattern). These grating lobes are somewhat unique in that they do not spatially movewith misalinement changes, rather
they are stationary with an amplitude dependenceupon the mechanical misalinements. This behavior is due to the operating characteristics of the retrodirective phasing system. Basedon environmental considerations, the grating
lobes are constrained to be less than O.Ol mW/cm
2. The total mechanical alinement requirements for both the subarrays and the total array can be determinedfrom this constraint. The I0.4 meter X I0.4 meter subarray which is considered34

22
Conditions:....

=10", +1 dB, 2%
10 dB. Gaussian Taper Antenna

1.0

.

++,

,_

A

_

e+ e_"

EE,>.01 _-e

_

__

i

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==

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.._"-_ _,_

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._1
10"4

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-

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l"

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0

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2000

Radius from 8oresight

•

I
3000

•

•

!
4000

.-

•

e°

!

5000

•

eee

•

e e

•

•

!
6000

I
7000

8000

9000

(KM)

t

I

I

I

I

I

I

I

I

I

t

I

0

500

1000

1500

2000

2500

3000

3500

4000

4500

5000

5800

Radius from Boresight (Miles)

Figure 19.

Peak Power Density Levels as a Function of Range From Rectenna
35

%
or-

or-

U

±
T \\

0
o

°r-

\
41-eft
o-G_

\

U

%
%
O

%

I
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36

°i--

to be the smallest entity for phase control, has the peaks of the grating lobe
patterns at the ground as shown _n figure 21.

Since the distance between maxima

for the grating lobes is inversely proportional to the spacings between subarrays, a I0.4 meter square subarray has peaks every 440 km,
control system is extended down to the power module

If the phaselevel, the grating lobes

will be spatially smeared and the peaks greatly reduced in amplitude.
improvementin grating lobe pattern

would

be due to differences

between the power tubes within the antenna.
misalinements:

in spacings

There are two types of mechanical(1) a systematic tilt of the entire antenna structure, and (2) arandom tilt of the individual subarrays.

The systematic

tilts have a greater

impact than the random subarray tilts on the grating lobe peaks,of the first grating lobe peak for a total antenna/subarray
minutesThis

is shown in figure 22.

alinement requirements

An example

tilt of 3.0 arc-

Other simulations have established mechanicalof less than l arc_minute for the antenna tilt and lessthan 3-arc-minutes for the random subarray tilts.
The near-field antenna pattern for distances close to the transmit
array is shown in figure 23.

A peak density of approximately

32 k_m 2 occurs

at a distance of 1600 km.
Rectenna Characteristics- The present ground receiving

antenna(rectenna) configuration, which receives and rectifies the downlink power beam,
has half-wave dipoles feeding Schottky barrier diodes.

Two-stage low-pass filters

between the dipoles and diodes suppress harmonic generation
matching.

and provide impedance

Eor economic reasons, the rectenna is a series of serrated panels

perpendicular

to the incident beam rather than a continuous

structure.

panel has a steel mesh ground plane with 75-80% optical transparency.

Each
This

mesh is mounted on a steel framing structure, supported by steel columns in
concrete footings.

Aluminum conductors are used for the electrical

power col-lection system.
The rectenna will produce

RFI effects due to rescatteredation and harmonic generation within the diodes.

incident radi-

There will also be a small

amount of RF energy leakage through the ground screen as well as knife edge
diffraction

patterns at the top edge of each rectenna section.

37

The general

0
_-

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X

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Q

_0

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o
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_D_
t.t--

0

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ZlU/M){ 'X.LISN30_l]MOd4OO[0rectenna parameters may be summarized as follows_
Total Active Panel Area:
Configuration;

78.5 km2
panels (multiple antenna elements feedinga single diode.

Panels should be open.

faced to reduce wind loading, with a maximum of I% leakage energy.
Receive Elements:

Half-wave rectifying diodes.

To provide an estimate of the power levels around the rectenna,studies

indicate that the dipoles have 98% collection efficiency under normal loading
conditions(reference 9).

The 2% reflected microwave

ward and towards the southern horizon.

power is directed up-

This reflected energy at 2.45 GHz is

only partially coherent since the regions of coherence for the incident beam
are limited due to phase irregularities

in the heated ionosphere and atmosphere.

Harmonics of 2.45 GHz will be generated within the half-wave rectify_
ing diodes and will be reradiated back through the low-pass filters and dipoles.
Initial measurementsof the harmonic levels relative to the fundamentalindi-cate the second, third, and fourth harmonics are down by _25 dB, -40 dB, andKTO dB, respectively,

for the normal dipole/dioderectifier configuration(reference g).
There will also be leakage power through the rectenna.
plane transparency

For a groundof 80%, approximately I% of the incident power appears asleakage through the wire mesh.
Since the rectenna's receiving surface appears serrated with individual
panels perpendicular

to the incident radiation,at the top edge of each section.

there will be diffraction

An analysis of the knife edge diffraction pat_tern has been made to determine the variation in power density
adjacent rectenna section (reference 9).

incident upon the

The power density will vary in the

shadow region (area behind the rectenna section) as shown in_gurenecessarylosses24.

It may be

to extend the size of the rectenna section to intercept part of the

shadow region.

There will also be energy lost as heat in the rectenna due to

41

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12R losses in the receiving elements and in the diodes.

POOR

Approximately 7% of

the incident energy is expected to be lost as heat.
The expected power levels around the rectenna for reflected energy,
leakage energy, harmonics, diffractions, and heat losses are summarizedinfigure 24.
Microwave System Efficiency - The total microwave

system efficiency

from the DC power output at the rotary joint to the collectedof the rectenna is 63%.
systems

A breakdown of the efficiencies

DC power outputof the microwave sub-is given in figure 25.
Phase Control System - The phase control system has an active, retro-

directive array with a pilot beam reference for providing phase conjugation.
Each subarray or possibly each power module(that portion of a subarray fed by

one klystron tube), has its own RF receiver which processes the uplink pilot
beam reference and inserts the proper phasing signal to form a single coherent
beam at the ground rectenna.

Tradeoff studies are now being conducted to deter-mine if the phase conjugation

should be at the subarray level or the power

module level.

Conjugation

at the power module level improves main beam effi-

ciency and microwave environmental

effects, but increases costs and complexity.

The Reference System includes:(1)

phase lock loop around each power tube for phase stability andnoise suppression.
(2)

double sideband, suppression carrier modulation

about the downlink power beam frequency fdl' as shown below.

fdl
43

which is symmetrical

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The two sidebands are demodulated in the RF receivers in the subarrays(or power modules) and the carrier is reconstructed.

This prevents beamsquint problems arising from different uplink and downlink frequencies, andit allows thelproper phase conjugation to be made.
the frequency separation_f

The ionosphere constrains

between the sideb.ands and the downlink carrier tobe greater than lO MHz, the maximum plasma resonance frequency.
is to prevent intermodulation

products between the uplink pilot beam andthe downlink power beam from creating parametric
overdense ionosphericThis limitation

instabilities associatedwithheating,

C3) coding of the pilot beam for security and pilot discrimination.
Since multiple SPS satellites will be _lluminatedmitter,by a single pilot beam trans_.

each satellite has to recognize whilch pilot beam signal it should

respond to.

In addition, coding will prevent power drain from any intentional

interfering signals.
(4) ground safety control system (ground sensors for interpretingbeam

shape) with a command link capability to the satellite.
RFI Characteristics_ The radio frequency interference comes primarily

from the DC-RF power converter tubes,
three main categories:

This interference can be divided into

(1) interference from the high power downlink beam due

to sidelobe and grating lobe radiations,the carrier frequency(2) spurious noise generated

by the tubes, and (3) harmonic generationnear

within the tubes.

The sidelobe and grating lobe levels were previously examined.

Withinthe phase control system, the phase lock loop around each power tube will reduce
the spurious noise close to the carrier frequency.

A representative

have a 5 MHz bandwidth with a second or third order filter.
affect the tube noise characteristics

loop mightThis loop will not

outside the 2450 ± 50 MHz band.

However,the klystron tubes will have a multiple cavity design which provides additional
filtering(24 dB/octave) to reduce the out_of_band noise.

characteristics

The SPS noise density

are summarized in figure 26 (_reference 31),

The CCIR (International Radio Consultative

Committee)

requirement for

power flux density at the earth'_s surface is _180 dBW/m2/Hz for S_band frequencies with an angle of arrival above 25°.
45

As shown in figure 26, the RFI

effects due to spurious noise will be below the CCIRrequirements, provided
the klystron tubes are phase-locked for noise reduction and a multiple cavity
design is used.
-l6O
CRITICAL RADIO ASTRONOMYFREQUENCIES-180

1400-1427 MHz (Hydrogen
Resonance Line)-200

1660-1670
MHz (Hydroxyl
Resonance Line)"_=

2690-2700
MHz_
4990-5000 MHzJ__

Astronomy
Bands

-220

v

-240
Z

o

0
Z

-260

-280
-300

-320

NOTE:

+l

+lO

2450 MHZFREQUENCY

+lO00

+lO0

2690 MHZCCIR REQUIREMENTS ARE - 180 dBW/m2/Hz FOR ARRIVAL ANGLES GREATER
THAN 250 •Figure 26.

Noise Power Density at Ground for a l km, 5 GW SPS Antenna46

RIGINA!_I.

_, : ,....

Mass Statement
A summary of the satellite mass properties

is presented in Table I.

The masses are separated into three major segments: solar array, microwave
antenna, and array antenna interfaces (the section of satellite between thearray and antenna which includes the rotary joint, sliprings, and antenna yoke).
The GaAIAs configuration utilizes a concentration

ratio of 2 which

reduces the required blanket area and in turn the blanket mass.
The antenna section mass properties are the same for both options.
The antenna mass is dominatedklystrons and waveguides.

by the transmittersubarray which includes the

The other large items are the power distribution

system and the thermal control system for the klystrons and DC/DC converters.
The total mass for the two options, including a 25% contingency factor,
is 34 and 51 million kilograms for the GaAIAs and silicon options, respectively.
J.

Space Transportation
This section provides descriptions of the reference space transportation

system vehicles.

The alternative

selected are described

concepts from which the Reference System wasin Appendix A.

The vehicles are distinguishedby theirprimary payload, either cargo or personnel, and their area of operations
earth and low earth orbit (LEO) or between LEO and geosynchronous(GEO).

betweenearth orbit

Cargo is transported from the earth's surface to LEO by the HLLV and per-sonnel (and priority cargo) are transported from earth to LEO and back by the PLV.
Transportation

between LEO and GEO is provided by the COTV and the POTV.

The general groundrules followed in the development
the transportation•and evaluation

system are:

The SPS transportation

system elements, with the possible exceptionof shuttle derived PLV's, are dedicated and optimized for the installation,
operation, and maintenance•ofof the SPS.

The SPS transportation

system will be designed for minimum

program cost.

47

total

TableI.

SPS Mass

Statement

STRUCTURE

SECONDARYSOLAR

STRUCTUREBLANKETSSILICONOPTION

DISTRIBUTIONINFORMATION
ATTITUDE& CONDITIONINGMANAGEMENTCONTROL27.258

4.172

3.388

0.581

0.436

6.696

22.051& CONTROL& STATIONKEEPING

1.1_

1.134

0.0_

0.050

0.200

0.200
13.382ANTENNA
PRIMARY

STRUCTURE

SECONDARYSTRUCTURE

TRANSMITTERSUBARRAYS

POWER

DISTRIBUTION

THERMALCONTROL

INFORMATION
ATTITUDE& CONDITIONINGMANAGEMENT& CONTROLCONTROL

ARRAY/ANTENNA

13.382

0.250

0.250

0.786

0.786

7.178

7.178

2.189

2.189

2.222

2.222

0.630

0,630

0.128

0,128
0.147

INTERFACES

0.147

0.094

0.094

0.003

0.003

MECHANISMS

0.033

0.033

POWER

0.017

0.017

PRIMARY

STRUCTURE

SECONDARYSTRUCTURE

DISTRIBUTIONSUBTOTALCONTINGENCY

(25%)

TOTAL

Rotary joint,

CR = 1

0.955

CONCENTRATORS
POWER

KGs

13.798ARRAY

PRIMARYof

GaAIAs CR = 2
OPTIONSUBSYSTEM
SOLAR- Millions

slip rings, antenna yoke

48

27.32740.7876.832

10,19734.15950.984

_- .....

•

Energy requirements will be minimized•

Environmental"y OF TKB

consistent with minimum cost.

impact will be minimized and, so far as possible,

protective measures needed will be factored into cost analyses.
•

The use of critical materials will be minimized consistent withcost, energy and environmental

impact requirements.

Heavy Lift Launch Vehicle
vertical take-off,vehicle.

horizontal landing (VTOHL), fully reusable winged launch

The launch configuration

27 and the launch/erector
uses 16(HLLV) - The reference HLLV is a two-stage,

CH4/02on the orbiter

and overall geometry are detailed in figure

concept is illustrated in figure 28.

The vehicle

engines on the booster (first stage) and 14 standard SSME's(second stage).

The booster engines employ a gas generator

cycle and provide a vacuum thrust of 9.79 X lO6 newtons each.

The orbiter

SSME's provide a vacuum thrust of 2.09 X lO6 newtons each at I00% power level.
The gross lift-off weight of the HLLV is ll,040 metric tons with a payload to
LEO of 424 metric tons.
An airbreather

propulsion system (aircraft jet engine)is provided

on the booster to provide flyback capability and simplify the booster operations.
Its landing weight is 934 metric tons. The orbiter deorbits and performs a glideback landing maneuver.

Its landing weight is 453 metric tons which includes

an assumed returned payload of 63.5 metric tons or 15% of the payload delivered
to LEO.
The HLLV trajectory and exhaust products data are provided in figure 29.
This figure shows the propellant expended and the exhaust product components
by weight for intervals of altitude versus range from lift-off to orbiter engine
cut-off.
Personnel Launch Vehicle(PLV_-The PLV provides for the transportationof personnel and priority cargo between earth and low earth orbit.

The reference

vehicleis derived from the current space shuttle system.

wingedliquid propellant fly-back booster instead of the Solid Rocket Boosters

and has a personnel compartment

It incorporates

ain the Orbiter payload bay capable of transporting

49

I

I

79.9m
262 ft_8.5m
60.7

Engines(16)ft

PayloadOrbiterBay

Booster
41.2m
135 ftJ80.6 m
264 ft

73.8
242 m
ft-

Figure 27.

1

Heavy Lift Launch Vehicle

5O

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52

-_,"
REPRODUC_-:"-2TY 0P THE
ORIGINAi. i." ', ,? PO(_R

75 passengers.

The overall configuration and vehicle characteristics

are shown in figure 30.

The passenger module is illustrated in fiqure 30

also.
The booster employs four 02/CH4 engines similar to those on the
HLLV booster.

A series burn ascent mode is utilized and the external tank(ET} is a resized, smaller version of the space shuttle tank, carrying 546
metric tons of propellant versus 715 metric tons for the current STS.
Personnel Orbital Transfer Vehicle {POTV) - The functions of the POTV
are to deliver personnel and priority cargo from LEO to GEO and to return
personnel from GEO to LEO.
The reference vehicle is a two-stage (common stage) LO2/LH 2 configuration as illustrated

in figure 31.

The start burn weight is 890 tons with an up payload of 151 tons and adown payload of 55 tons.

The up payload consists of 160 personnel in a passenger

module, 480 man-months of consumables in a resupply module, and a flight control
module piloted by a crew of two.

The down payload is identical except the

resupply module returns empty to LEO.
Cargo Orbital Transfer Vehicle (COTV} - The function of the COTV is todeliver SPS cargo to GEO from the LEO staging area.
the construction

The basic concept involvesof a fleet of reusable electric powered round trip vehicles

and their dedicated solar arrayin LEO.

The vehicle uses ion bombardment

thrusters with cryogenic argon as the propellant.

The ion thruster propellant

was selected on the basis of availability, storability,
mental impacts, cost, demonstrated

absence of serious environ-performance, and technicalsuitability.

conversion options are GaAIAs and Si photovoltaic array systems illustrated

Power
in

figure 32.
The first option utilizes a self-annealing GaAIAs array with a concentration ratio of 2 and provides a LEO-GEO trip time of 133 days and a total roundtrip time of less than 180 days.

Ion bombardment thrusters of lO0 cm diameterare used with an Isp of 13,000 seconds and argon as the working fluid.

53

The primary75- Passenger

ShuttleTransfer

ModuleOrbiterHigh Pc HLLV
engines
O2/ CH4
thrust-9.6
x 106NIsp (SL/VAC)

Reducedsize External
Tank

Flyback

Booster

55.7 m37.9 mFigure 30.

Personnel Launch Vehicle

54363Â455sec--• _O_UOyT]
REPRi....... ...... _TY OF THE:: Pi)ORORIGD,:A!__>_ _" Y_e_ ca_.=_._

g

k

=

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I

II

II

o

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_"

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_,

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,,,_

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oo

oo

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-_i

_-_

THRUST
MODULES,,,,o,oI! i<__

I
1140 M _-_PAYLOAD UP
4000 MT
VEHICLE DRY MASS ]206 MT
PROPELLANT UP
835 MT
PROPELLANT DOWN
150 MT

164 M

Silicon Solar Cells
PAYLOAD ATTACH POINTS

'\",

/
¢
i

I

THRUST
MODULE I
\

/
J

'\
2250M

PAYLOAD UP
VEHICLE DRY IvLASSPROPELLANT UP
PROPELLANT DOWN

t

3,469 MT715 MT
185 MT
27 MT

GaAIAs Solar Cells

Figure 32.

SOLAR
ARRAY

Cargo Orbit Transfer Vehicles Options

56

REPRODUCI2!r.H,yORIGINAL
PA(];]OF THEfS POORthruster array of 259 thrusters is suspended by cables and located at thevehicle center of gravity.

Additional attitude thruster control packages arelocated at the structural extremities.

The component mass breakdown is given in

Table 2.
The second option utilizes a silicon photovoltaic
planar configuration with no concentration reflectors.

solar array in aRound trip time from

LEO-GEO-LEO

is approximately 160 days which also allows two trips per year foreach COTV.

Ion bombardment thrusters of 120 cm diameter are used with an Ispof 7000 seconds and argon as the working fluid.

Thruster modules of 296

electric thrusters each and an appropriate number of chemical thrusters arelocated at the four corners of the COTV.

The component mass breakdown isgiven in Table 3.
Transportation

Fleet Requirements - Section L of this report provides

a summary of the total transportation
orbital bases and of two 5 GW SPS.

requirements for the installationof theRequirements are expressed in terms offlights per year for each vehicle and fleet size for both the GaAIAs SPS andsilicon SPS.
Table 2GaAIAs Independent Electric COTV Mass Breakdown

Vehicle

(MT)

(Dry)

Power Generation/Distribution

249

Thrusters

26

Propellant

Tanks

and Lines

39

Control

229

Structure/Thermal
Rotary7

Joint

AttitudeControl/IMS

Primary

Power

22

Unit

Total (Dry)572

25% Growth Margin

143

Payload3469

Propellant Up

185

Propel_own

27
4396

Total in LEO

57

Table 3

Si Independent Electric COTV Mass Breakdown

Vehicle

(Dry)

(MT)

Power Generation/Distribution

570

Thrusters

70

Propellant Tanks and Lines

60

Structure/Thermal

80

Control

Rotary Joint
Attitude Control/IMS

K.

Primary Power Unit

185

Total (Dry)965

25% Growth Margin

241

Payload4000

Propellant Up

835

Propellant Down

150

Total in LEO

6191

Natural Resources
The materials for the Reference System are listed in Table 4.

The

specific materials are identified in column l under individual program components
such as the satellite, rectenna and various transport vehicles.

The second col-umn indicates the mass of each material in metric tons for each component
specific to a silicon solar cell satellite program.

In those instances wherecomponents can be used in both the silicon and gallium arsenide solar cell programs such as several transport vehicles, the masses are listed in the thirdcolumn.

Finally, the masses of those materials

for components

specific to thegallium arsenide program are listed in the fourth and last column.

In the caseof the low earth orbit staging and orbital transfer vehicle construction
form and the geosynchronous

orbital construction

system only, the estimated total mass is given.

plat-base for the gallium arsenide
These components

studied in depth at this time to identify the materials

are being

and better estimate

component mass.
Table 5 compares the total materials masses for all componentsofthe program through the first satellite in columns 2 and 3 and for each year

58

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63}TI.I_ _- OF THE

thereafter with two satellites launched each year in columns 4 and 5.
first satellite includes not only transportation
orbital staging and construction
orbital vehicles.

The

and rectenna, but also thebases, plus the constructionof the cargo

It should be noted that because the gallium arsenide system

has both construction bases as total mass estimates at this time, (see Table5) the mass of those components

is indicated under miscellaneousand organics

heading.
L.

Operations
SPS operations include those activities

then to operate and maintain them.
as illustrated

by figure 33.

required to build SPS's andThis requires a wide variety of activities

Because studies of most of these activities

arenot yet complete, this section will briefly outline the overall SPS operation
as it appears at this time.
I.

Construction Operations
Construction of an SPS starts with two supporting operations.

First the necessary raw materials are mined and manufactured
components and propellants.

A significant mass production

to be developed to produce the high number of components

capability will needneeded per satellite(e.g., about IO II solar cells, lO5 klystrons, lO lO dipoles).
requirements

for large amounts of propellants

into launch-ready

Similarly,(oxygen, hydrogen, argon) demand

expanded processing capabilities.
The other supporting function involves ground transportation of raw
materials, fabricated components,

and assemblies

to the launch site.

Also,

about lO6 MT of hydrocarbon propellants are used per year.

Among the possi-

bilities for propellants under investigation are processing

coal at mines andusing a gas transmission system, transporting
rectenna-supplied

coal to the launch site, using

electricity to electrolyze water, and others.

At the launch site, principal activities
and processing of material and propellants;
ing and checking out returning vehicles.

64

involve receiving,storing,launching vehicles; and refurbish-

,-,_ vVv oF

q_gtS

01%101.:,
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65

Incoming material (via rail, air, etc.) is off.loaded,inventoried, and stored in warehouses.

Component packaging

inspected,(for construction

material, consumables, spares) is very significant for construction as well
as space transportation.

Packages must meet dimensional and weight con-straints of the launch vehicle and have appropriate mass density for costeffective

space transportation.

Figure 34 illustrates the dimensions, densityand part counts of various SPS components.

As indicated, densities vary froma low of 12 kg/m 3 for antenna subarray elements to about 2500 kg/m3 for power
conductors.

To obtain desired densities,

components must be packaged inappropriate mixes as indicated in figure 35.

Such packaging minimizes

number of launches, thereby reducing transportation

thecosts.

The silicon option requires 375 HLLV flights and the GaAIAs
option requires 225 HLLV flights to transport construction(two 5 GW units) capability.

material for lO GW

Construction personnel are launched in an up-

dated shuttle Personnel Launch Vehicle.
Operations in LEO include COTV construction and maintenance,

payloadtransfers between HLLV's and COTV's, POTV stage mating, crew transfers, vehicle
and base maintenance, and propellant

storage and transfer.

After payload transfers, COTV's travel to GEO over a period of several
months.

At GEO, a small interorbital transfer vehicle docks to and moves the

cargo to the construction base.

After off-loading,

the COTV returns to LEO

with packing materials, damaged or defective equipment, and parts and consumables containers.

At LEO, argon tanks and thruster grids are replaced, thevehicles refurbished and readied for the next transit.
Personnel arriving at LEO from earth, transfer to POTV's
for the trip to GEO, which takes a few hours.

Personnel returning from GEO

transfer to Personnel Launch Vehicles for the trip back to earth.
Because detailed construction

techniques,both for COTV's and SPS's,

for the Reference System have not yet been developed, the reader is referredto the Appendices.

Appendix A discusses construction

issues while Appendix B

presents techniques that were developed for the Rockwell and Boeing independent
systems.

The techniques for the Reference System would be similar in general

concept.
66

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68

g., },
OgiOE<a_t . :i, ]_; I_ pOOR
' +Figure 36 presents a typical timeline for the silicon option for
constructing

the initial LEO and GEO bases and the COTV's required to thenconstruct SPS's.

Six months are required to construct the LEO base.

Threemonths are required to then construct the first COTV, three additional months
to construct the second and third COTV's, and a year to complete the rest ofthefleet needed to transport SPS materials from LEO to GEO at the rate of lO GW
per year.

Once the first COTV is completed, it begins to transport materials

to GEO needed for the GEO construction

base.

port the remainder of this material.

Nine months are required to construct the

GEO base.

The second and third COTV's trans-

After two years, all of the major elements are available to begin

production of the first SPS.
For the gallium option, it has been assumed that the GEO base would
be built first in LEO where it would construct the COTV's.

Then two COTV's

would transfer the base to GEO and leave only staging facilities

in LEO.

Figure 37 shows estimates of the number of flights required, payload
characteristics,

launch vehicle packaging factors assumed, and numbers ofpeople associated with building the LEO base, the COTV's and the GEO base overthe initial two-year period.

Data is presented for both silicon and GaAIAs

options.
Figure
SPS's.

38 presents a typical timeline for constructionof two 5 GW

All the material for one SPS is taken to LEO by HLLV flights in a six-month period.

It is transferred to COTV's for six-month trips to GEO, withall materials arriving at GEO over a six-month period.
during this six-month arrival period.

Construction takes place

While the constructionof the first SPS

takes place, material for the second is being taken to LEO, transferred to
COTV's, and is enroute to GEO.

When the first satellite

material for the second begins arriving at GEO.

is complete, the initial

Although the entire sequence

to build any two 5 GW SPS's takes 18 months, once the process is underway, two
SPS's are being produced every year.

69

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Figure 39 estimates the number of flights required, payload characteristics, packing factors assumed, and numbers of people associated with building two SPS's over a one-year period.

Data is presented for both silicon andgallium options.
Table 6 shows the fleet sizes of HLLV's, PLV's, COTV's, and POTV's
needed (1) for the buildup period prior to SPS construction and (2) for theconstruction of two SPS's per year.
gallium options.

Data is presented for both silicon andFewer COTV's are needed for the gallium option due to thefollowing combination of factors: different COTV design and flight times, different satellite weights, and different packing factors.
Besides space construction, there is the task of constructing
ground rectenna.

theTechniques for accomplishing this have not yet been developed;

however, some of the relevant issues are present under the constructiondis-cussion in Appendix A.
2.

Commercial Operations
Once the SPS and rectenna are constructed,

produce commercial power.

the SPS begins to

The main tasks are to operate the interface withthe grid and maintain both the SPS and the rectenna.
With regard to the grid interface,power would remain uniform at all times.

it would be ideal if the SPS

In reality, however, there will bevariations from a number of seasonal, daily and orbital path causes.
periodic shutdowns due to earth eclipses will occur.
power generation systems must have a throttling

Also,Thus, ground-based

capability to smooth out thedemand load.
With regard to maintenance,

the large number of components

lO5 klystrons) will result in random failures.

Also, it is probable that

sections of solar blankets will have to occasionallyoroid damage or part failures.

It is estimated(e.g.,

be replaced due to mete-that between 5 and 20 people

would be required in GEO per SPS, probably stationed at the construction

base.

Parts and personnel are transported from LEO to GEO and between SPS_s in GEO
using OTV stages.
73

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3.

Integrated Operations Management
All of the above construction

simultaneously.

and commercial operations

must occur

Thus, there needs to be an integrated operations management

scheme that assures they occur in a coordinated manner.

Figure 40 illustrates

the various functions involved in SPS operations as they have been discussed.

Figure 41 illustrates an operations

coordinate,

management

concept to control,and integrate these functions.
The program headquartersfunction would manage the flow of ground-based resources through manufacturing and transport to the launch site.
would coordinate the communicationsItsystems necessary to carrying out SPS

operations and, in general, assure that overall SPS operations

are conductedsatisfactorily.
The launch and landing control function manages the preparation,
launch, and landing of cargo and personnel to and from LEO.

It maintains

and refurbishes launch vehicles.
The LEO and GEO bases manage on-site their respective
involved with construction,

logistics, transportation,

activities

checkout, maintenance,etc.
SPS ground control units at each rectenna would operate and control
the functioning of their assigned SPS, rectenna and grid interface.
M.

Costs
Cost

System.
theThis

Comparative

informationhas

costing

information

Assessment

been

Study.

76

estimated
will

be

for

the

included

SPS
asReferencea partofGE 00pc.r,_tional

I

Unit
SPS
Production

andConstruction|

Facilities
M'anufactu'ing

_
1

Base

1

Supportl

Logistics

P-'asel

Facilities

_

and

|

Refurbishmentt/"_inlcnance1.t----

Unit
._

SPS
Ope_al;onv_sits J
SPS

Operatlor,al Elements

•

Solar

Power

Satellites

•

Space

Transpo_b_tlon

Caigo

and

Orbit

Transfer

Per_otlnel!

|

SystemVehicles

•
•

l
I

LEO
an# G_O
Operational
Ground Support System

Bases

J

I

t

!

Base

/Facilities|

Facilities

illllllllllil
r._al(.,ria

Ft_uipment
Supplies

Lau,_chy/J_//Jg/_
/_

JS

_

z

L og'stics
Facilities/

Facihhes

Vehicles_//b_IL/_/_,,,o_,ao_ V/*_7/7/////A_e,,o.ne'

PlupeIIanls
Personnel

S_:;;orl_.

/t_e(.e,,ery

Facilitiesj.

[artllOperatior,

LaunchFigure 40.

and

P_c

liliesa', l_.se

Recover)' SiteSPS Operational Functions

R}:P.:(!DUCIB!LT.Ty
OF THE
ORIG.iNAL
P,_....G i_ l_,o POrt R

77

1

SPS

m_mbe, I_,J

Bese ConSrof£,E00p{_alional

IS'S°°m"_'']........F°°"'"_"""'!
_'_'S"'""I[_-_,.,,o.0..
lOps
j
OpS
.I
L
ISPS

t

numbs" Nr

T ,an_ p c_'ia
ffo_[. EO Operalior,

Cunst
Ops

tuber t_l SPS

number

NI

na

lo"s

c°? trOI

l--""_[_line,ace

Progr,,mand MainlJ

al EaseControl

LO_s-----_

14,,,dcll,arlers

:r

]

L

Ir..a,er,aehi_.,,,_--,l,('c,,-/r;a_,on[///////j_u ch ,=nd R_.cove_

I'L......

Ops

I[_l'.;_,n_,_,,_,,,,i,,g
Ops _.__jr---_T_C:--.l- :. .... -- -I

L7_
.....
LJ_
---_
I v_',_clet.'_.i.:_.:,,'.n_eo_s-I

r-''__,":....-'"_"_:"-:,oo_I'....................... -'Figure 41.

SPS Operations Management

78

ConceptIV.

TECHNOLOGYSUMMARY
A.

Critical questions
Beginning with the earliest systems definition studies, it has beenrecognized that SPS will require significant technologicaleral disciplines.

advancements

Continued study has indicated that establishment

designs, performance levels, cost expectations, development
environmental
tions.

acceptability,

in sev-of firm

requirements, anddepends on the resolution of several critical ques-

Although overall success of SPS development

is possible over a range ofperformance and design, establishment of specific attainable

performancelevelsis important to establishment of designs and system specifications.
The critical questions involve engineering,

economic, and environmental

issues, each of which require laboratory test effort to obtain the necessary
information for full evaluation of the SPS concept.

This summary will confine

itself to critical questions primarily related to questions of engineering
feasibility.

Many of these questions and the required test activity are interlacedwith answering critical environmental and economic questions.
The critical questions may be categorized according to the following
general subsystem areas.
I.

Microwave power transmission

2.

Solar arrays

3.

Power distribution

4.

Structures and control

5.

Materials

6.

Construction

7.

Space transportation

Critical questions under each of the above categories are summarized
below.

79

I.

Microwave

Power TransmissionoMicrowave

power amplifier(DC-RF converters)

design andoMicrowave power transmission phase control system development.

oMicrowave power transmitting antennadevelopment.

development,(subarray) design andintegrated systems test.

2.

oRectenna electronic component development.

oRectenna structures and construction

techniques.

Solar ArraysoSolar cell array (blanket) design and materials selection

for automated blanket fabrication/productionoat low cost.

Long-term environmental effects on candidate solar cell arraymaterials to determine radiation(electron/proton)degradation characteristics,

UV susceptibility, and plasma interaction with high voltage arrays.
o

Determine thermal annealing characteristics

and, where

necessary, develop in-situ annealing techniques such as laser heating devices
and self-annealing methods.
o3.

Solar reflectors(for concentrated

systems).

Power DistributionoInvestigate spacecraft charging and high voltage interaction.

oHigh voltage (40 to 50 kv) DC switchgear development.

oHigh voltage (40 to 50 kv) electricaloRotary joint/slipring

cable insulation

development.
development

for 30 year life (electro-

mechanical components).
oPower processing

sion systems.

80

technology development for electric propul-o Powerprocessing technology developmentfor radiation
resistant control electronics, EMIsuppression, and long-life, high efficiency
transformers with weight constraints.
4.

Structures and ControlsoDevelopment of integrated structure/control

systems forlarge lightweight systems.
5.

MaterialsoDetermine outgassing/UV/particle

radiation effects on com-posite materials and coatings.
oLongterm stability of composites

and other materials

in

GEO and LEO environments.
6.

Constructionodevelopment,

Space fabrication techniques and equipment

including the accrual of space operational(beam builder)

experience with such

equipment.
odevelopment

Manned remote work station and large-scale crane-manipulator

including docking and berthing techniques.
oHardware deployment, handling, and installation techniques

for solar array blankets, power conductors, structural elements, and antenna
subarrays.
7.

Space TransportationoElectric propulsion technology development;

thrusters andoLarge liquid-fueled rocket engine development

control s.
for low cost,long-life HLLV operations.
oFluids handling and transfer technology

81

for use in orbit.

B.

Alternate Technologies
In consideration of the above technology advancement

is appropriate to identify alternative
program risk.

requirements,ittechnology options available to reduce

The major alternatives and options are summarized

below according

to system element.
I.

MicrowaveoPower Transmission

Use of amplitrons, magnetrons, or solid state RF generators

in lieu of klytrons.
2.

Solar Arrayso

Use of amorphous silicon, high efficiency cadmium sulfide,or other potentially low cost solar cells.
oDevelopment of thermal engine systems such as the potassium

Rankine cycle and Brayton cycle for use in lieu of solar photovoltaics.

Focus

on lightweight radiators and leak-tight joints for fluid system.
3.

Materialso

Use of aluminum structure with thermal distortion

stabilizing

design features for transmit antenna and/or solar array.
4.

Space TransportationoSingle-stage-to-orbit

booster (HLLV).

82

booster in lieu of two-stage vertical

T_
T'_
REPRODE, T C_i_!
_Y 0F THE
0RIGINA{. ?' .....
TS POORV,

DOCUMENTATION SUI_4ARY
The following

is a partial listing of the reports of studies dealing withsystems definition of the Solar Power Satellite and related subjects.
are grouped by general subject.

Reports

Listings under each subject heading are chrono-

logical.
Most of the reports listed are available through the National Technical
Information Service.

NTIS numbers are listed in parentheses after each refer-ence if applicable.

Where a series of numbers is given, the report consistsof more than one volume.

An "X" prefix indicates restricted distribution.

These restrictions are listed in this document as follows:
N

_

G

-

GC-

NASA only
Government agencies only
Government agencies and contractors only

System Definition
I.

Glaser, P.E., "Power From the Sun: Its Future," Science, Vol. 162,

November 22, 1968.
2.

"Initial Technical, Environmental and Economic Evaluation of Space

Solar Power Concepts," Johnson Space Center, JSC-ll568, August 1976 (N77-16442
and 16443).
3.

"Satellite Power System Engineering and Economic Analysis," MarshallSpace Flight Center, TMX-73344, November 1976 (N77-15486).
4.

"Satellite Power Systems (SPS) Feasibility Study," Rockwell International

Corporation,
5.

SD76-SA-0239, Contract NAS8-32161, December 1976.

"Systems Definition of Space-Based Power Conversion Systems," Boeing

Aerospace Corporation,

D180-20309, Contract NAS8-31628, February 1977 (X77-lOlOl and lOl02 (GC)).
6.

"Space-Based Solar Power Conversion and Delivery Systems," Econ, Inc.

and Grumman Aerospace Corporation, Contract NAS8-31308, Second Interim Report,
June 1976.

83

7.

"Solar Power Satellite System Definition

Study - Part I," Boeing

Aerospace Corporation, D180-20689, Contract NAS9-15196, June 1977 (N7813099 through 13103).
8.

"Solar Power Satellite System Definition Study - Part II," Boeing

Aerospace Corporation, D180-22876, Contract NAS9-15196, December 1977 (N7820156 through 20163).
9.

"Solar Power Satellite System Definition Study - Part Ill," Boeing

Aerospace Corporation, D180-24071, Contract NAS9-15196, March 1978 (N78-20164).
IO.

"Solar Power Satellite Concept Evaluation," johnson Space Center,

JSC-12973, July 1977 (N78-12116).
II.

"Satellite Power Systems (SPS) Concept Definition Study," Rockwell

International Corporation, SD78-AP-O023, Contract NAS8-32475, April 1978.
Other Concepts
12.

Ehricke, Krafft A., "The Power Relay Satellite: A Means of Global Energy

Transmission Through Space," Rockwell International Corporation,

Report E74-

3-I, March 1974.
13.

"Application of Stationkept Array Concepts to Satellite Solar Power

Station Design," Aerospace Corporation, ATR-76(7575)-l,Contract NAS8-31842,

November 1976 (X77-10236 through I0238 (GC)).
14.

Billman, K. W., W. P. Gilbreath and S. W. Bowen, "Introductory Assess-ment of Orbiting Reflectors for Terrestrial
Center, TMX-73230, April 1977 (X77-73872

Power Generation," Ames Research(N)).

Orbit Characteristics
15.

"Orbital Motion of the Solar Power Satellite," Otis F. Graf, Jr.,

Analytical and Computational Mathematics,

Inc., ACM-TR-I05,

Contract NAS9-15171,

May 1977 (N78-15148).
Power Conversion
16.

"To Design Variable Orientation

Solar Concentrators for Space Applications,"

University of Georgia, Contract NAS8-31565, October 1975 (X76-72006 through72009 (N)).
84

17. "Design of ForwardLighted Solar Concentrators for SpaceApplications,"
University of Georgia, Contract NAS8-32149.
18. "Evaluation of Solar Cells for Potential SpaceSatellite PowerApplications," Arthur D. Little, Inc., Contract NAS9-15294,June 1977 (N77-30612).
19. "Developmentof a Directed EnergyAnnealable Solar Cell System," Spire
Corporation, FR-20066,Subcontract to NAS9-15196,July 1978.
Microwave
20.

Power Transmission"Experiments in the Transportation of Energy by Microwave Beams," Brown,

W. C., 1964 IEEE Intersociety Conference Record, Vol. 12, Part 2, 1964, pp 8-17.
21.

"Free-Space Microwave Power Transmission Study, Combined Phase III and

Final Report," W. C. Brown, Raytheon Report No. PT4601, Contract NAS8-25374,
September 1975 (N77-16619).
22.

"Reception-Conversion

Subsystem (RXCV) for Microwave Power Transmission

System," Raytheon Report No. ER75-4386, JPL Contract No. 9533968, September 1975
(N76-15598).
23.
4368,

"Microwave

Contract24.

PT-4964,

J.,

26.

SystemResearch

AugustStudies

performed

27.

"A SolarPower Satellite

Beam Forming,

Steering

NAS9-15237,

June 1978.

28.
Dynamics,"Achievable
Convair

Contract

Div.,

Contract

through

Company,

ER75-

15597).

Receiving

W. C. Brown,

Terminal

RaytheonofaCompany,

1977 (N77-31613).

Affecting

SPS Rectenna

Performance,"

Study,"

L.

NAS9-15212,

September

1977

Transmission

SystemIncorporating

LinCom,in a Large Microwave
NAS9-15423,

85

Gutman,1977.

Beam Interaction

and Phase Control,"

Flatness

Raytheonof theSystem,"at JSC, August"Ionosphere/Microwave
Rice University,

Studies,"

Improvement

Power Transmission

NAS3-19722,"Microwave

SystemDecember 1975 (N76-15594

and MechanicalMicrowave
Contract25.

Gordon,NAS3-17835,"Electronic

Free-Space

RonaldPower Transmission

Inc.,

M. Duncan and W. E.
(N77-33389).

TR-7806-0977,

Power Antenna,"

September

Automatic
Contract

General1978.

: _oDrrc-7._rLTT;::'Y_::d POOR29.

Perkins, F. W. and R. G. Roble, "Ionospheric Heating By Radiowaves:

Predictions for Arecibo and the Satellite Power Station," J. GeophysicalResearch30.

(to be published).
Holway, L. H., A. H. Katz, G. Meltz, "Ionospheric Effects of a HighPower Space-Borne Microwave Beam," Raytheon Technical Memorandum

T-I028,Waltham, M.A., Nov. 1977.
31.

Arndt, G. D. and L. Leopold, "Environmental

Considerations for theMicrowave Beam from a Solar Power Satellite," 13th Intersociety Energy
Conversion Engineering Conference, San Diego, CA, August 1978.

Transportation
32.

"Future Space Transportation

Systems Analysis Study," Boeing Aero-space Corporation, D180-20242, Contract NAS9-14323,
N77-31235; Vol 2:X77-78969
33.

(N); Vol 3:X77-79819

December 1976 (Vol. l:(G)).

"Orbital Propellant Handling and Storage Systems for Large Space

Programs," General Dynamics Convair Div., CASD-ASP-78-OOl,

Contract NAS9-15305,

April 1978.
Orbital Construction
34.

"Orbital Assembly and Maintenance Study," Martin Marietta Corporation,

MCR-75-319, Contract NAS9-14319, August 1975 (N75_32144).
35.

"Orbital Construction Demonstration Study," Grumman Aerospace Corp.,

NSS-OC-RP-O08, Contract NAS9-14916, December 1976 (N77-23136).
36.

"Orbital Construction

Support Equiprlent," Martin Marietta Corporation,

MCR-77-234, Contract NASg-15120, June 1977 (N77-27157).
37.

"Space Construction

Automated Fabrication Experiment Definition

General Dynamics Convair Div., CASD-ASP-77-OI7,

Study,"

Contract NAS9-15310, May 1978

(N78-25111 through 25113).
Rectenna Construction
38.

"Feasibility Study for Various Approaches

to the Structural

Design and

Arrangement of the Ground Rectenna for the Proposed Satellite," Bovay Engineers,
Inc., Contract NAS9-15280, May 1977.
86

Technology Advancement,,p.. .
reilmnary

39.

Assessment of Technology Advancement Requirements

for Space

Solar Power," Johnson Space Center, JSC-12702, March 1977.
Economic and Political40.

"An Initial Comparative Assessment of Orbital and Terrestrial Central

Power Systems," R. Caputo, Jet Propulsion Lab., JPL-DOC-900-780,

March 1977

(N77-22612).
41.
Utility
July"Impacts

and Benefits

Industry,"

Arthur1977(N78-24255).

42.

"Political

Power System,"August1977 (N78-25003).
"A Study

Implementation
ContractofSome Economic

C-80020,ofContract

Developing

ECON-77-195-1,

Factors

Power Systems,on the954639to954652Development

ECON, Inc.,(JPL),and Operating

Contract

Relating

Electrica(JPL),

andNASA-CR-150602,

March 1978.

"Research

Plan

Power Satellite

Contract

NAS2-9655,

45.

"Compilation

Laboratory,
46.

ment,"

Inc.,

Power System

and Environmental

44.

west

Inc.,

Implications

Econ,

Satellite

NAS2-9655,

Biological

Solar

of

a Satellite

D. Little,

and Legal

Satellite

43.

of

Study

of

Transmission

Biological

System,"and Ecological
Bernard

Effectsofthe

D. Newsom and Associates,

March 1978.
and Assessment

AO-O2-OI/EA81028,"Satellite
Floyd

for

Power System

R. Livingston,Jetof Microwave

Bioeffects,"

Contract

EY-76-C-06-1830

Environmental

Impacts

Propulsion

Laboratory,

Pacific

North-(DOE),

May 1978.

- Preliminary

Assess-900-822,A.

Rev.

May 1978.
Work Breakdown
47.

Structure"Satellite

Space Flight

Center,Power System
TM-78155,

Work Breakdown

January

Structure

Dictionary,"

1978.
REPRODUCIBILITY_)I':,qINAL
PAGE

87

Marshall

OF TH_IS POORAPPENDIX ASYSTEM ANALYSIS RESULTS

This appendix provides discussions of the key trade-off study
results and analyses that have been conductedto date.

The approachtaken in documenting the study results is to describe the evolutionof the design work from initial concepts to the present.
The following topics are addressed in this appendix in the orderindicated:A.

Design Considerations

B.

Solar Energy Collection, Conversion, and Power Distribution

C.

Power Transmission,

D.

Structures and Materials

E.

Space Transportation

F.

Construction

G.

Natural Resources

Collection and Conversion

A-l

A.

Design Considerations
This section summarizes several trade studies dealing with orbit selec-tion, orientation,system sizing and constructability.

direct, major influence on the overall configuration
Orbit Selection - Geosynchronous
all studies.

These

trades have aof the satellite.

altitude (35800 km) has been used in

Near uninterrupted transmission

is possible, antenna steering

accelerations are low, and the satellite is stationary with respect to a pointon earth.
Three inclinations have been considered(figure A-l).

gives a stationary satellite, simplifying the rectenna design.

Zero inclination
A 7.3° inclination

eliminates the lunar and solar perturbation which would otherwise cause theinclination to vary between zero and 15° over a period of years, if uncorrected.
A 23.4 ° inclination places the orbit in the ecliptic plane, permitting
lite to be oriented simultaneously

the satel-

toward the sun for maximum output and perpen-

dicular to the orbit plane (POP) for minimum gravity gradient torque.

Non-zero

inclinations cause a daily variation in the angle of incidence on the rectennaof about twice the inclination, increasing

the rectenna area required and imposing

constraints on the rectenna design for which solutions have not yet been identified.

The zero-inclination orbit can be maintained

propellant budget, however.

with a relatively

All studies have concluded

smallthat zero inclination ispreferred.
Solar radiation pressure acting on the large, low density solar array
induces a small, variable eccentricity

in the orbit.

lation of several degrees of longitude.
is relatively minor:

This causes a daily oscil-

The design impact of this oscillation

the resulting angular accelerationof the transmitting

antenna is manageable and the moderate east-west motion of the microwave
can be easily accommodated

by most proposed rectenna designs.

propellant penalty for maintaining

Moreover, theis substantial,

amountingto roughly half of the total attitude control and orbit maintenancerequirement

for typical configurations.

zero eccentricity

beam

Consequently,

early studies tended to tolerate theeccentricity in order to reduce propellant resupply requirements.

A-2

SubsequentFigure A-l.

SPSOrbit Inclinations

A-3

results, however, indicate that orbit maintenancecan be combined,at least partially, with attitude control to reduce total propellant consumption. In an
eccentric orbit (uncorrected for radiation pressure), the large size of the SPS
results in an unacceptable probability of collision with other synchronoussatellites.
Holding the eccentricity to zero will greatly reduce this probability.
In summary,the preferred orbit for the SPSis geosynchronouswith zero
inclination and eccentricity; i.e., geostationary.
Orientation and Attitude Control

The orientation of the SPS mustsatisfy two conditions:the solar energy collection system (SECS) must point at

the sun and the microwave

power transmission

ground station.

system (MPTS) must point at the

The first consequence of this is that the MPTS must rotate con-tinuously at one revolution per day relative to the SECS.

This rotation is aboutan axis perpendicular to the orbit plane (POP), requiring that the MPTS be locatedat the north or south end of the SECS, mounted centrally on a rotary joint ofsufficient size to span the SECS carry-throughstructure, or mounted centrally

on a small rotary joint within dielectric structure for minimum
with the microwave beam.

interference

Early SPS concepts used the last of these.

Subsequentstudies have avoided this approach in order to eliminate an unnecessaryof interference, the nature of which

is not completely

sourceunderstood.

The principal disturbance acting on the SPS is gravity gradient torque.
Gravity gradient torque exists whenever the spacecraft principal axes are notorthogonal with the orbit plane and local vertical and the principal moments ofinertia are not equal.
An asymmetrical

SPS, with the center of pressure offset from the masscenter, will also experience torque from solar radiation pressure.
Gravity gradient affects the SPS in two ways.

The first occurs when

the sun is not in the orbit plane (that is, at all times except at the equinoxes)
if the SECS is pointed at the sun (figure A-2).

This torque cycles with a periodof one year, making momentum storage unattractive.

Since the maximum sun angle

is 23.4 °, however, the SECS can be held in a POP attitude at all times.

Thiseliminates the torque and causes a maximum power loss of only about eight percent.

A-4

Becauseof the large saving in attitude control propellant, all studies haveadopted this attitude.
The antenna joint can also be simplified in the POP
attitude, requiring two axes of rotation (one continuous) instead of three inthe solar orientation.
The second gravity gradient effect arises from the continuous attitudechangeof the SECSrelative to the earth (figure A-3). This torque is cyclicwith a 12-hour period. Becausecontinous rotation is required, a fixed attitude
cannot be used as in the previous case. Consequently, the torque has been minimized in most studies to date by reducing the difference in momentsof inertia
about the two axes in the orbit plane. This can be done by increasing thelength of the SECS(perpendicular to the orbit plane) and reducing the width.
The momentof inertia

difference can also be reduced or eliminated bydeparting from a flat solar energy collector.
Several approacheshave been
proposedfor both photovoltaic and thermal conversion, although the geometricalconstraints of thermal systemsmakethemless amenableto such treatment. Power
distribution paths are generally longer in inertially"balanced" configurations,
tending to reduce the total-mass-to-orbit advantage. Most, but not all, inertially balanced configurations appear to be moredifficult
to construct. Since
speed and ease of construction are significant factors in the practicality of
the SPSconcept, most inertially balanced configurations have not survived
evaluation.
Since the gravity gradient torque is cyclic, momentum
storage devices
such as momentum
wheels appear attractive.
However, preliminary studies haveindicated a cost disadvantage comparedto a reaction control system unless thehigh specific impulse projected cannot be achieved.
Satellite Sizing - For a minimum cost of electricity

per kilowatt-hour,

the output per antenna should be as large as possible, although studies indicatethat the cost per kilowatt-hour increases only slightly for outputs substantiallybelow maximum. Maximum output is constrained by two factors, the power density
at the transmitting antenna
osphere(heat rejection) and the power density at the ion-

(ionosphere disruption).

The best current definitions of these limits

A-5

Sun

\
J

,,,,....re,Torque.

Figure A-2.

Long-Period Gravity Gradient Torque

/'

/

'\

/

\

e

\

0

/

I

@

I
!
p

\

/

i

\

/

/

,,

,,,

f

_

Figure A-3.

/

J

Short-Period

A-6

"1

Gravity Gradient TorqueSunare 21 kW/m
2 and 23 mW/cm
2, respectively, although future tests could indicate
revisions. Since the variation of the two factors with antenna size is different, the limits can be used to calculate not only the maximumoutput, but alsoantenna and rectenna diameters. Using the microwavesystem parameters definedelsewhere in this report, maximum
output on the ground is approximately 5 GW
with an antenna diameter of l kmand a rectenna diameter of lO km (figure A-4).
Constructionof this appendix.

- Construction of the satellite is discussed in Section F

However, satellite configuration strongly influences ease of

construction, which is probably necessary if the SPS is to be economically
petitive.

In general, a configurationthat permits a high degree of automationin its construction, such as the reference photovoltaic
more easily.

system, can be built

Conversely, a typical thermal cycle configuration

large number of different operations.
dozen times.

com-

requires a

Many of these are performed only a few

A large number of fluid connections must be made.

Even the highly

repetitive tasks, such as reflector facet installation, are largely discrete
rather than continuous.

All of these are comparatively

difficult to automate,

and as a result the thermal SPS is relatively difficult to construct.
Power Output Variations - Ideally, SPS power output at the rectenna
would remain uniform at all times.

Actually, however, the power output will

depend on the intensity of illumination of the solar collector/converter,
efficiency,

and the efficiency of the microwave transmission collection

conversion system.

its
and

Figure A-5 illustrates the net effect of these variations

for a solar photovoltaic

solar collection/conversion

system.

In this case, the180-day cyclic variation is caused by variation of the sun's declination
with respect to the orbital plane.

Superimposed on this variation

cyclic fluctuation resulting from orbit eccentricity.
a variationin satellite-to-rectenna(angle)is a daily

Orbit eccentricity

causesdistance, affecting transmission/collection

efficiency.
The power output of a solar array depends on the intensity of illumination at the cells and the temperature of the cells, the power output of cells
diminishing as the cells become hotter.

In geosynchronous orbit the temperatureof the solar cell is related to the intensity of sunlight for any given panel
configuration.

A-7

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A-9

i.i

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ORIENTATION

2;oVariations

3_o

|

3_o

The sun is brightest at perihelion, which occurs around winter solsticewhen the orientation of the array is such that the sun's rays arrive at 23.5 °off of normal incidence.

The worst-caseillumination

is at summer solsticewhere the 23.5° misorientationis accompanied by aphelion where the intensityof sunlight is 0.967 ° average.

However, the solar array temperatureis alsodown, being 36.5°C rather than 46.0°C as at the spring and autumn equinoxes.
The net effect of these variations, as shown in figure A-5, is a lower power
output during the summer months than during the winter months.
Eclipses by the earth, illustrated in figure A-5 by the close-spaced
vertical lines, will cause total shutdowns daily around local midnight
about six weeks in the spring and fall.

for

The maximum duration is about 75

minutes.

A-IO

B.

Solar Energy Collection, Conversion, and Power Distribution
I.

Energy Collection and Conversion
The energy collection/conversion

systemfor the SPS has the func-tion of collecting the dispersed solar energy in sufficient quantities for conversion to electrical energy.

This electrical energy is then delivered to thepower transmission system for beaming to ground-based receiving stations as
microwave

radiation.
Several major trade studies have been performed on a variety ofenergy collection/conversion
concept.

systems to determine the most favorable to the SPS

As part of these studies, photovoltaic systems, thermal conversion

systems, space nuclear power plants and orbiting solar reflector systems were
included.

Both NASA JSC and NASA MSFC performed in-house studies (ref. 2,3)

on various systems during 1974-1976.

During the same period and for the next

few years, each center contracted with aerospace companies to perform independent
system definition studies.

This section summarizes certain of these studies

concerning energy collection/conversion

systems that have led to the reference

system.
The MSFC in-house studies were conducted primarily
quantitative

understanding of the advantages and disadvantages

rather than to eliminate options.
mass and costs.
ductionto obtain aof various schemes

The tradeoffs were performed to minimize

The MSFC-Boeing study (ref. 5) evaluated and identified pro-rates, launch frequencies, facilities, etc., so that electric power

cost could be estimated.

Satellite size, mass and life cycle cost were estab-lished within the limitations of the contract.

The JSC in-house study emphasized

analysis of the photovoltaic concept and investigated sensitivity of systems to
mass, performance, and transportation

cost.

Also, a thorough review of pastsystem studies involving several thermal energy conversion
plished.

concepts was accom-

Most recently, the JSC-Boeing and the MSFC-Rockwellstudies both per-formed a comparative analysis of a variety of systems with the objective ofidentifying feasible systems.
as follows.

The systems investigatedA-ll

are discussed briefly

Solar Photovoltaics - Since the SPS concept was first postulated,
the passive system advantages of photovoltaics has provided a standard of comparison for other solar collection/conversion

systems.

The MSFC in-house study(ref. 3) looked at numerous photovoltaic systems to select a solar cell technology suitable for the SPS.

The design utilized silicon solar cells although

their studies indicated some other compositions
study modeled a silicon photovoltaic

definitions were extremely preliminary,

SPS.

The MSFC-Boeing

and gallium arsenide photovoltaic

cept to determine total power generation costs.

studies.

showed promise.

(ref. 5)

SPS con-Although most of the component

cost estimates allowed reasonable trade

The JSC in-house study (ref. 2) emphasis was on the silicon photovoltaic

The JSC-Boeing study (ref. 7,8,9) included evaluations of several photo-voltaic options including single crystal silicon, single crystal gallium arsenide
and thinfilm options and some other less developed thin film approaches suchas copper indium selenide.

The MSFC-Rockwell study (ref. ll) included silicon

and gallium-aluminum-arsenide
ratios.

solar cell evaluations and several concentration

The data base at the beginning of the study was evaluated to determine

which SPS approaches should be seriously considered as candidates

for further

analysis.
In evaluating the various photovoltaics options, a number of factors
have been considered including performance(efficiency), mass, materials avail-ability, susceptibility to radiation damage (performance degradation),ment status and cost.

In addition to the system definition

have been made (ref. 18) to assess materials availability,

develop-efforts, surveys
manufacturing

pro-cesses requirements, and energy payback of several candidate solar cell designs.
This work included an assessment of SPS solar cell requirements with respect
to DOE's U.S. Photovoltaic Conversion

Program.

In comparing the various photovoltaic
silicon cell and the gallium-aluminum-arsenide
promisingfor SPS application.

options, the single crystal

cells have emerged as the most

Other solar cells that have been considered

include amorphous silicon, polycrystalline

silicon, cadmium sulfide, copper indium

selenide and polycrystalline gallium arsenide.

These cell types generally havethe potential advantage of lower costs; however, at present, the performance(efficiency) is low and mass production methods

A-12

have not been devised.

Single crystal silicon solar cells are the only solar cell typethat has been utilized for spacecraft solar powersystems. Researchand development has producedcontinuous improvementsin unit mass, efficiency, structuralquality and reliability;
however, great improvementswould still be required
for effective SPSapplication.
Silicon cells have the advantageof more established manufacturingbase, lower potential cost per cell, and abundantresource materials. Disadvantages, relative to gallium arsenide, are higher massper unit solar cell area,
greater performancedegradation from thermal (temperature) and radiation effects
and slightly lower efficiency.
The thermal degradation effect can be minimized
by designing for low cell operating temperature; i.e., without solar concentration, or by designing so that the concentrator acts as a passive radiator forcell cooling. Performanceloss by radiation degradation is projected to be
recovered by laser annealing the cells in place.
Gallium arsenide cells havebeen under investigation for a numberof years but significant improvementshavebeenmadesince 1972. The developmentof the gallium-aluminum-arsenide "window"which is epitaxially grown on thebasic gallium arsenide cell has led to the improvementin cell efficiency. Since
most solar radiation is absorbedwithin l _mof the GaAscell surface, it ispossible to construct a very thin cell (~5 _m) with good efficiency.
Consequently,
the quantity of gallium neededto makethe cells is significantly reduced. The
advantagesof gallium arsenide cells are low masspotential, resistance to degradation by thermal and radiation effects, and good efficiency.
Useof solar concentration provides self-annealing of the cells at moderate temperatures. Disadvantagesare relatively high cost and less technology base than silicon.
Gallium availability
is also a consideration.
Table l provides an examplecomparisonof gallium arsenide andsilicon cells for a specific SPSconfiguration. Note that with solar concentration (CR=2), the gallium system has a cost advantageover the silicon system,
but with CR=I, the silicon systemwould be either slightly less or competitive
with gallium. Becauseof this close competition, silicon and gallium arsenide
are both viable candidates for SPSapplication. Therefore, they are presented
as options in the description of a reference system.
A-13

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Solartypicalsolar

Braytonreflector

and isusually

an inert

itisheatedthrough
duces

Braytoncycle

focusedsystem.

into

turbinea turbinewhichelectric

passesthroughconditions.

throughpassingtheexpands

The generator

pro-turbine,theheat

absorber.

The working

fluid

receives

final

exchanger

where

waste

is

via

rejectiontothecompressor

where

the

space
its

a radiatorpressuresystem.

iscompressor

uses about

The conversion

efficiency

35 percent
than

at

turbineinlet

40 percent

The higher

raisedwithtransferred

The cooledtotheturbineofthetwo-thirdsof a Brayton

temperaturestemperatures

whereas

heat

inlet

1700°Frequireuse ofmore advanced

There

are

variations

fluid)

radiatorin

may be used at

in
fluid

passesoutputsystemtemperaturealloyscooling

throughpressure

turbine

inlet

metal

preheats

to a coolant

turbinerefractorygas isthengas thencycle

in thefluid,

gas thenresidual

cycle

a

where

The hot

theofabsorber

where

the

heat

workingexchangerforcomponents

or argon,Afterpower.

heat

range.

The cycle

and generator.

intogreatercollected

a compressor

passingtoby a concentratingisdrivesthe

20 toenergySolartemperature

in a recuperator

Typically,

diagraminletcooleda cooler

A-6 shows a schematica cavityabsorber.

furthergas- Figuregas such as heliumtousefulCyclelevel.

work.

rangesfromto2200°Frange

the2500 totechnology

3000°F
ceramic

the

lower

temperature

Braytoncycle

includinglevel.

gas (working
multipleseveralshaft

systems,

hightemperature

end ofofcycle.

thesystems,

thermionicratio

The MSFC-Boeing

study

basic

alternative

working

fluids,singleversuscycle

concepts

usingthermionics

concept

attheor a Rankine

cycle

at

rejectionend

GW satellite

with

of

generatingtheand dualcycle

The MSFC in-housea concentration

of

study

2000:1,(ref.

a heliumloop

in combination

(ref.

5)

the

heat

3) configured

a I0

working

and a highwithinvestigatedfluid,a Braytoncycle

temperature

conversion

a thermionic/Brayton

system.

combined

cycle

includedexami-

systemand a closed

cycle

Braytonsystem.

The JSC in-housestudy

nationofcycle

Braytonconcept

and investigatedseveralsubsystem

study

(ref.

closedsystemtheclosedalternatives.

Followingprobedinto

deeperthemany ofearlierthestudies,apparent

problemA-15

the

JSC-Boeing

areas

of

the

the

7)

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A-J6

Brayton cycle design.
modifications.

Reflector degradation was reevaluated resulting in size

The configuration was altered to conform to new construction

concepts and geometric improvements.

These refinements and more detailed in-vestigation of the construction techniques and materials requirements
in determining

better cost and mass estimates in the study.

the determination

resulted

This resulted inthat the system not only was significantly heavier than photo-voltaic and Rankine cycle systems (due primarily to large radiator requirements)
but was also more complex to construct.

The MSFC-Rockwell concept definition

study comparatively analyzed several solar thermal concepts including the closed
Brayton cycle and found that, although the technology for the Brayton system wasmuch further advanced than the others, the weight penalties made it less competitive than less developed, but higher performance cascaded Rankine systems.
Solar Rankine Cycle - Like the Brayton cycle, the Rankine cycle
utilizes a solar concentrator

to collect and focus energy into cavity absorbers(figure A-7); however, instead of an inert gas, a liquid working

fluid such as

water or potassium is passed into the absorber where it is vaporized(boiled).

The hot vapor is then expanded through a turbine which drives a liquid pump(not a compressor as in the Brayton cycle) and a generator which produces electricity.

After passing through the turbine the vapor is cooled (condensed to aliquid) either directly in a space radiator or indirectly by an intermediate
coolant loop.

The liquid then passes through the pump which raises the liquidpressure to boiler inlet conditions.

As with the Brayton system, there areseveral variations of the basic cycle.

Typical Rankine cycle conversion

effi-

ciencies are 15 to 40 percent depending upon cycle arrangement.
The JSC in-house study (ref. 2) reported on the potential of thesolar Rankine cycle to SPS applications.

Although the technology was not as

advanced as some other thermal systems, the Rankine cycle utilizes higher heatrejection temperatures which results in lower radiator mass.

Several working

fluids were investigated in the JSC study, but the overall analysis was only apreliminary

investigation and no cost and mass estimates were reported.

the JSC-Boeing

Later,

study (ref. 7) developed sufficient detail for comparative analysis

A-17

u
°_

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A-18

with

other

thermalmass and cost

systems

potential

enginesinvestigated.

included

the

to

cost

was a cascaded
arrangement

thelowestalternative

to

the

cycle

and otherthe

Thermionics

bottoming

with

for

least

- Thermionicconverters

theconverters

reportedinrequiringsuch as thethermionic

MSFC-Boeing

study(ref.5).

and liquid

cooledradiatorsystemsusing

Subsequentstudiesconfinedtheirinvestigations

conversion

system.

The generalmore recent

evaluations

would

be at

least

sion

efficiencies

50 percent

3000°Fconsidered

in

Other
reasons

attaining

the

solar

cells.

produce

directwere analyzed

of

these

study

also

cycle

and

systemswerelookedconversion

thermionics

atdirecttechnique.
tothein thesestudies,that

thermionic

the

in

singularas wellassystemthermalcycle

systems.

20 percent)withpeak temperaturesystemhas,been includedinConver-therefore,

not

been

detailed

evaluation

studies.

have notbelow:and heat- low conversion
rejection

Magnetoplasmadynamics

necessary(aboutdefinition

Thermoelectrics

consumption,(2)

systemisother

The thermionicknown options

stated(I)low

and above,ofreached

than

to

Both

thermionicand Rockwell,heavierare relatively

the

or

thermionic/Brayton

The latterconclusion

by Boeing

enginesconversion.

theresourcesa potential

some studies,concept

theprovideas thecost.

energya nuclear

for

solar

thermalsystems

an open absorber.

cycle

temperatureenergystudy

Rankineuse highotherin theanddevicesIn

definition

efficiency

by thermal

withfurthertheinvestigations

overallprovidedcombination

of

cycle

conversion

energy.

II)selected

confirmed

thermaldesignenergyelectrical(emitter)(ref.

sufficient

improved

has a lower

alternative

The configuration

The reportpotential

systemMSFC-Rockwell

have significantly

theRankineand performed

fluid/steam

weight.

offering

passivecurrent

tothat

Braytonconcept

working

provedspecific

systemtheand mass estimates.

Solar

These

cycle

cesiumThis

thermalthan

Simultaneously,

Rankinedetermineand concluded

working

fluid

efficiency,

materials

considerations.
- rejected

temperature

A-19

on grounds
by solarofheating.

problemsin(3)Direct Thermal Conversion

data available for this recently-proposed(4)

Thermophotovoltaics

by Electrostatics- Insufficient

thermal engine,- rejected on consideration of overall

efficiency and problems of waste heat rejection.
In conclusion, the photovoltaic

options utilizing either single

crystal silicon solar cells or gallium-aluminum-arsenide
with the overall lowest cost and mass potential.
(no concentration)

have evolved as options

A concentration ratio of one

is preferred for silicon systems, whereas

designs may be more advantageous with a concentration

gallium arsenide

ratio of two.

Based onthe analysis conducted to date, the results indicate these options would beless complex and would have lower mass than the thermal cycle options.

The

major disadvantages of the thermal cycle systems are fluid system containment,
wear in rotating equipment, construction

complexity, and relatively

high mass.

It should be noted, however, that the cost advantage of photovoltaics
is very sensitive to array blanket costs.

To illustrate

this point, figure A-8

shows a cost trade-off between silicon photovoltaics and the potassium-Rankine
cycle.

At the projected solar blanket cost, the silicon system has a 5 to lO

percent cost advantage.
With respect to the thermal cycle options, the recent system definition studies have indicated a preference
cycle.

for the Rankine cycle over the Brayton

The Rankine system would use either a potassium working fluid or acesium-steam working fluid combined cycle.

This preference

is not strong because

it is based #n small mass advantage relative to Brayton systems.

Therefore,

these systems are regarded as competitive at this time; however, as previously
stated, the photovoltaic2.

systems are preferred over the thermal systems.

Power Distribution
The purpose of the power distribution

late and control power from the power generation

systems is to collect, regu-

system (solar array sections)and transmit this power via power busses and rotary joint/slipring
microwave generators on the transmitting

antenna.

This system would also include

energy storage as necessary to meet power requirements

A-20

system toduring the eclipse periods.

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A-21

A number of trade studies on this system have been conducted aslisted below:(1)

AC vs. DC Transmission

based on power conductor and conditioning(2)- DC transmission

at 40 kV is preferred

equipment mass considerations.

Central vs. Decentralized

Regulation and Control - centralized

system preferred for providing regulated power to the transmiting antenna; decentralized system for solar array power control.
(3)preferred

Fiat (sheet) Conductor vs. Round Cable - flat (sheet) conductor

based on mass considerations.
(4)

Conductor Material (Copper, Aluminum or Silver) - aluminum

preferred based on thermal control and mass considerations.
(5)

Use of Structure for Current Return Path vs. DedicatedConductor - dedicated conductor preferred as a less complex approach.
Even though the above preferences are indicated at this time, continued analysis and optimization

may result in changes.

A-22

C.

Power Transmission,

Collection, and Conversion

The SPS concept for the transfer of energy from geosynchronousto earth is accomplishedby means of microwave power transmission.

orbit

This isbasically a 3-step process which has as one of its major requirements

highefficiency.
The process consists of:I.

Conversion from DC power to microwave

power

2.

Focusing and transmission 6f the microwave power to Earth3.

Collection of the microwave power on Earth and conversion

toDC power.
Some of the more significant design and operational
which have been made are summarized

trade studies

below.

Microwave System Frequency - The primary frequency in the SPS microwave system which has been the subject of the majority of the trades thus far
is the power beam frequency.

(Trade-offs to determine the optimum set of up-

link pilot frequencies for the phase control system are still being conducted.)
The requirement for high efficiency over the microwave

propagation path dic-tates the band of microwave frequencies which are acceptable; molecularabsorption and rain attenuation through the Earth's atmosphere
limit on frequency selection of not much above 3 GHz.

sets an upper

Alternate frequencies

have been considered in most of the microwave system studies to date, the mostnotable being 5.8 GHz.

Trades have been made from many points of view such as

SPS system sizing of transmit antenna and rectenna, etc., (ref. 2,10,23);
propagation effects through the ionosphere
jections(ref. 2,10,23), etc.

(ref. 26); hardware technology

pro-

Although higher frequencies offer some advantages(smaller rectenna), the frequency of 2.45 GHz offers many more.

This frequency

is in the middle of the ISM band (industrial, scientific, and medical) of lO0
MHz, there is lower attenuation when propagating through the atmosphere, andthe projections for microwave equipment technologies are more promising.

A-23

Microwave System Efficiencies- One of the earliest demonstrations

that meaningful amounts of power could be transmitted via microwaves
at Raytheon in May of 1963 (ref. 20).

The collection

cent; however, the RF-DC conversion efficiency
Overall efficiency was only 16 percent.

was doneefficiency was 87 per-was only about 50 percent.

As the experiments

continued, more

attention was devoted to increasing

the efficiency

The concept of a "rectenna" evolved(receiving antenna and rectifier) which

exhibited low directionality
conversion efficiencies.

of collection and conversion.

for the collecting antenna, and high RF-to-DC

Early application of the rectenna concept to the

SPS resulted in the addition of filters to attenuate the radiation of harmonics
and to store energy for rectification.

The rectifier was changed from a full-wave bridge using point-contact diodes to a single GaAs Schottky-barrierstate diode in a half-wave rectifier configuration.
80 percent were obtained

Conversion efficienciesof(ref. 21).

In 1975, the rectenna portion of the microwave
the JPL Goldstone facility (ref. 22).

system was tested at

Microwave power was transmitted from an

85-foot antenna to an array of 270 rectenna dipole-diodetance of 1.6 km.

solid-

elements over a dis-

Of the microwave power impinging upon the rectenna, over82 percent was converted to DC power.

A total of 30 kW was collected anddissipated into a lamp and resistive load.
investigated mechanical and electronic
which has resulted in demonstrated

Since that time Raytheon has

improvements

conversion

certain power density levels (ref. 24).

to the rectenna element

efficiencies of 85 percent at

Current projections,nology advancements, are for conversion efficiencies

assuming tech-of approximately

90 per-cent in the time-frame needed for SPS.
Coupled with this experimental

concentration

on improving

thecollection and conversion efficiency at the receiving end, a concept evaluation and technology assessment was being conducted on the transmit antenna.
Major contributors toward transmit efficiency

in the microwave

I.

Microwave power amplifiers

2.

Beam focusing and pointing (phase control)

A-24(DC to RF converters)

system are:3.

12R losses in antenna

4.

Mechanical tolerances(pointing, subarray/power module mis-alignment, waveguide/slot

tolerances.

These are, of course, dependent on many factors, such as frequency,
subarray sizes, phase control concepts, etc.

As the microwave

system configura-tion developed and converged into the present concept, system efficiencies
came better defined.

Some of the differences in estimated efficiencies

evolved are shown in Figure A-9.

The end-to-end efficiency

microwave reference system is 61 percent.

be-as they

for the current

As an aid to understandingtheterminology used in the microwave system, please refer to Figure A-lO.
Microwave System Sizin 9 - Sizing and power transfer in the microwave
system is dependent on three factors:
I.

System end-to-end efficiencies

2.

DC power output from rectenna

3.

Transmit antenn_ size.

For a minimum cost of power at the grid, the output of the transmit
antenna should be as large as possible.

However, this output is constrainedby thermal limits on the antenna and power density limits in the ionosphere.
Thermal dissipationof approximately

limits set a heat density value on the transmit antenna

21 kW/m2.

For the operating frequency of 2.45 GHz, studies

have indicated that the power density of the SPS beam should not exceed
23 mW/cm 2.

If the power density exceeds this value, it is postulated

that

nonlinear interactions may occur between the power beam and the ionosphere.
With these two limits, and a desired value of power output at the rectenna,transmit antenna size and rectenna size can be traded.

The Raytheon study

in 1975 (ref. 25) initially studied overall microwave system sizes.

Some ofthe parameters established were:I.

5 GW DC output power at rectenna

2.

Rectenna size--approximately

3.

Transmit antenna size, l km diameter

4.

Aperture illuminationlO km diameter (at the equator)- 5 step truncated Gaussian with 5 to lO dB

taper.

A-25

Current
JSC

I DC-RF

Conversion

Antenna,

12R

Transmit

AntennaMech-anicalTolerances

ReferenceStudy

I Study
BAC

.87

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).89

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0.99

).97

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J

Power

Interface*Combined

Efficiencies

Overall.655

Efficiency

FigureA-9.

Microwave

SystemA-26.602

Estimated.642

Efficiencies.611

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A-27

The JSC study in August 1976 (ref. 2) expanded on the previous
studies, and defined a lO-step truncated Gaussian, lO dB taper which approximated a continuous taper and provided for side-lobe suppression.
developments and trades regarding configurations

Otherand subsystems are discussed

in the following sections.
Transmit Antenna Configuration- The early in-depth study of the

SPS microwave system conducted by Raytheon in 1975 (ref. 23) took the systemsizing as discussed above and developed a configuration

for the transmit

antenna with the following characteristics:
The transmit array consisted of 18M X 18M subarrays, with a structure
fabricated out of graphite polyimide,material coefficients of expansion,
the Shuttle cargo bay.

based on a l percent beam power loss andand sized to be effectively

Because of questions of outgassing

packed intoof the graphite

polyimide, an aluminum configuration was suggested with independent 5M segments
within the 18M X 18M subarray.

Motor-driven

array to adjust for subarray deflections

screwjacks were used on each sub-and tilt during operation.

The JSC study (ref. 2) investigated 4M X 4M and lOM X lOM subarraysin order to widen the beamwidth and eliminate the requirement

for screwjacks.

It was determined that the effects of coefficients of expansion on surface
tolerances would allow a lOM X lOM subarray, and the larger size (lOM X lOM)would reduce the complexity of the phase control problem.
Raytheon also performed a trade-off study of transmit antenna types(including space-fed and cylindrical arrays) and on subarray types, including
helical radiators, parabolic dishes, pyramidalhorns, and slotted waveguides.

The latter was chosen because of its potential

for high efficiency

efficient means of distributing

the RF power from the power amplifiers.

as a result of the Raytheon and JSC studies
evolvedand the(ref. 23,2) the configuration

into:
I.

lOM X lOM subarrays (7850 total)

2.

Subarray structure - graphite composite

A-28

Thus,

3. Slotted waveguide
4. Waveguidematerial - aluminum
5. Phaseconjugation - to subarray level
6. Nomechanical alignment of subarrays.
The subsequentBoeing studies (ref. 5,7,8,9) as part of the overall,
indepth trade-offs on the SPS,arrived at basically the sameconfiguration
except for the waveguidematerial. Doubtwas expressed that mechanical tolerances such as slot sizes, subarray tilt, waveguidedimensions, etc., could
be held tight enoughusing aluminumwaveguide.
As a result of that factor and the weight trade-off, graphite
composites (for mechanical stability)
together with thin-skin depth aluminumon the waveguidewalls (for RFconductivity) was investigated and recommended.
Kovar waveguideswere also investigated as an alternative.
The MSFCand Rockwell studies (ref. 3,4,11) resulted in basicallythe samemicrowavesystem configuration, with somedifferences in antenna
structure and configuration.
The basic building block was a 30MX 30M
mechanical module, with approximately lOMX lOMsubarrays as an integral
part. The radiating portion of the subarrays was defined as a resonant cavity
radiator to reduce wall weight of the waveguides.
The reference configuration of the transmit antenna is nowas listed
in the fourth paragraph of this section, with the following alternate
possibilities.
Phaseconjugation to the power module level may be desirable
for more precise control and better focusing. This resulted from the LinCom
effort (ref. 27) on the phase control systemand will be discussed more laterin this section. It maybe desirable to go to resonant cavity radiators _f
efficiency and resonant wavestability showan effective trade-off with weight.
Figure A-ll is a pictorial summaryshowingthe transmitter design concept as
evolved for the reference configuration.

A-29

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A-30

Microwavein1975 (ref.

amplitron23) on the(crossed

The amplitron
klystronhighPower Amplifiers

field

microwave

systemamplifier)and thewas characterized

was characterized

efficiency.

parameters,
concluded
could

that

the

amplitron

forinitialtime

advantage,

and so that

into

the

systemoverallstudy

The JSC studies
amplitronand thewere

investigated.

also

and thebeing

klystrongiven

The major

tothesimplifying

themaintainability
distributionand (d)lower

investigate

klystronthatisseries,microwave

possible

operationtheklystrontube

The study

or 50 KW klystrons

thein parallel

1976 (ref.

3) selected

efficiency

and life-

and mechanicalin Augustassembly

1976,

wave tubes

best

integration

both

the

amplifiers

that

time

the

potentialforuse on theimplementation

at

(a)

that

much higher

and installation

bothhigheroperatingsystem(c)much lower

phase

amplitron

SPS, with

the

edge

considerations.
power

operations,(b)weights;

investigatedand solid-state

requirements;

per

tube,

and ameliorating

voltage,thus

reducing

control

drive

power;

as theamplifiercharacteristics.
studies(ref.

5,7,8,9)and performed

A comparison

shown in Table

into

cost,

chose

an in-depth,

integration

can be integrated

advantages.

in

were:

manufacturability,

designs

system.

klystronRF noisecertaintheoffurther,a klystronintoof system

configurations,scenarios,

versusbecauseThe Boeingtoon amplitron

while

be analyzed.

theinitialbeam amplifier).

and moderately

was concluded

the

the

levels,noise

klystronadvantages

thus

power

exhibitedoflowbecause ofTravelingIt(linearin November

concluding

klystron.

klystronconcluded

electrical

could

on investigation

gain,

performed

The MSFC study,

by Raytheonand low

were

5 KW amplitrons

performed

efficiency

gain,

as on integration

eitherbe used.

by highstudy

focusedby high

Trade-offs

as well- The in-depththeweight,
into
of

the

klystroncomprehensive

microwave

system,study

maintenanceetc.

The study

demonstrated

an overall

SPS and it

does havefeatures

II.

A-31

of

the

amplitronofand klystron

POWER

5 KW with

50-250106 TUBES

85-90%EFFICIENCYCATHODEK LY ST RONAMPLITRONITEM(AVAIL_

COLD PURE METAL
LIFE DATA - 10,000

GAiN

KW WITHHR)<105

TU_3ES

80-85%
THERMIONICOXIDE/MATRIX(AVAILLIVE DATA _ 50,000 HRS)
40 db

7 db

VOLTAGE

20 Kv

40-65

SPUR IOUS SIGNAL
AM(TYP.)-100 db/KHz 10 MHz
FROM CARRIER-125 db/KHz 5 KHz
AWAY FROM CARRIER

TUBE MTBF

COMPARABLE

THERMAL

DISSIPATIONCONCENTRATEDINTERACTIONCOMPARABLEPERHAPS SOMEWHATREGION;PHASE SHIFT
TUBE-TUBE
TRACKING

HIGH POWER PHASE SH1FTERS
2 ° PER%CHANGE
IN BEAM CURRENT+15 ° PHASE TRACKINGSPECI F IC COST_$20/Kw

SPECIFIC

0.4 kg]Kw

ARRAY

WEIGHT

INTERFACE
TURN-ON

X-RAY LEVEL(RANGE FOR 5 mR/YEAR)*SIGNIFICANT

Kv

LESS

DISTRI[3UTEO,COLLECTORCAN RUN
500-700°C,
REQUIREHEAT PIPESLOW POWER PHASE SHIFTERS
10°-30 ° PER % VOLTAGE
CHANGE±20 ° PHASE TRACKING$40 TO-S 2-07K w
FOR ABOVE POWER RAfJGE
U ._-TO- _..", -iC: i'i_-w
FOR ABOVE POWEI;{ RANGESERIES OPERATIONNO FEED WAVEGUIDES

POWER ADJUSTS TO VOLTAGE
CHANGESCORPORATEFEEDDIRECT

MAY

FROM

BUS BAR ?

SAFE @ 1 METERREQUIRELOGIC

SEQUENCE

SAFE @ 1-1.8 km FOR 70 kw DESIGNADVANTAGE

Table II.

Some Features of Amplitron and Klystron Design for SPS

A-32

Boeing also surveyed current developmentsin solid-state amplifiers,
and concluded they should definitely be considered, but were not emphasizedin the studies. Devices surveyed included bipolar silicon, electron beam
devices, GaAsIMPATTS,and silicon FET's. In approximately the sametime
frame, JSCalso independently performeda survey of solid-state devices.
Although no particular device surfaced as being an outstanding candidate atthat time, subsequentresearch indicated that GaAsFET's showedsignificant
potential.
Rockwell studies (ref. ll) investigated amplitrons, klystrons, magnetrons, and solid-state power amplifiers.
For purposes of performing tradestudies on SPSintegration issues, the klystron was also chosen. Similarities
were evident betweenthe Boeing study results and the Rockwell study results.
Basic requirements for the klystron were similar, with somedifferences evident
in implementation into the transmit antenna.
The Boeing study developeda concept for a 70 KWklystron, integrated
into a slotted waveguidepowermodule, with 4 to 36 power modulesper lOMX l OM
subarray. Cooling is accomplishedwith 300°C body and 500°Ccollector heat
pipes with passive radiators on the backside of each subarray.
The Rockwell study developeda concept for a 50 KWklystron,
integrated into a resonant cavity radiator with 6 to 50 klystrons per lOMX lOM
subarray. Cooling is accomplishedwith body heat pipes dissipating heat on
the front side of each subarray. Blockageof RF radiating slots is estimatedto be approximately 3 percent; however, this was required becauseof thethermal environment on the backside of the transmit antenna.
Continuing investigations into solid-state power amplifiers provide
an increasingly attractive alternative to the power tube amplifiers, especiallyin the area of low noise and increased reliability.
A concentrated technology
advancementprogram is neededfor solid-state, coupled with a comprehensive
systems integration study.

A-33

Phase Control
highaccuracy

beam focusing

time-varying

ionosphere,

subarrays,
power

and phase

amplifiers,

The subject

phase

of

power

beam, orperformanceeach

subarray.

the

mit

where

precisearray,
signalproblemproblemon a system

JPL,antenna

distribution

in thesystem.

Rockwell,

Boeing,

LinCom,both

command andtheon (a)

a matrix

covering

thesensorsofa distinctivephasemodulationeachsystem,reliesofandcomponents

examination

to

array

The command approacheachtransmitsub-on theoutputapproachuses a reference

pilot

beam

subarrayor poweron thetrans-

phase measurements

moduleare made and conjugation

onlybasis.

conjugation

1977.

activity

such

large

were

comparedand open loop

for

of

the

arrays.

commanded phaseadjustments.
mutually(a)icallycompensateddistribution

intoreference

distribution

systembutionpath

variations

lengthforcoupledtheconcepts

thesystemphase

distribu-

previouslydis-on an overallbeam steering

control,

study
to

of

(b)

groundThe techniques

has resulted
threesystem.

phase

Basic

phase

oscillators

systemincluded(a)

network

monitor-consideredforand (b)electron-system.

LinCom activity

the system

consistedand evaluated

shift

to

the

a contract

accomplishingconsideredphase

studiedIn an effort

JSC awarded

Initialforincludedmicrowave

basis

techniques

The techniques

mentioned

concepts.

potential

distribution

Thisvariousin Aprilsuperficially,

and uplinkeach organization

end-to-end

phase

the

electronic

providea nonhomogeneous,

wiggle"level,

and beam steering

partitionsoftransmit

1975 investigateda "bitdegrees,to evaluate

engineering

ing

(b)

and documented

LinCom Corporation

phase

in

SPS must

occurs.

controlcussed,referencephase

made overrectenna

To varyingtionin theThe retrodirective

fromanalysistheby Raytheon,by commandingtransmitted

pilot

of

approaches.

measurements

thedistortionand otherstudies(retrodirective)

receivedarrayvariations

in

for

presence

has been investigatedestimation

systemthethermalRaytheoncontroland pointing

conjugators,

MSFC and JSC.
adaptive- The phaselevelsina phaseand whichconceptwhich

has been incorporated

into

The first

level

which

electronicallyto maintaina constantor tree

A-34

controlconsistsofa referencecompensates
phasereferencefor

distriateachradiating element.

The second level consists of a phase conjugation

which receives a uniquely-designed

systempilot signal allowing reconstructionof aphantom carrier at the same frequency as the power beam, thus avoiding problems of squint due to frequency offsets between the pilot and power beam
signals.

The pilot signal design also provides for anti-jamming(security)of the pilot receivers, provides additional processing for isolation of thepilot receiver from the power signal, and allows for multiaccess

operation(i.e., more than one SPS can be intercepted by the same pilot signal without interference).

The third level of control is associated with maintaining

an equal and constant phase shift through the microwave power amplifier
devices.

In addition, this concept of phase locking around the power

amplifiers will provide added suppression of transmitted noise and reduce
the SPS RFI (radio frequency

interference) potential.

As part of the preliminary analysis and system definition
an SPS microwave system computer model has been under development

activity,

which will

provide a flexible system engineering tool for evaluation of SPS microwave
system performance as parameters which uniquely affect the phase control system
are varied.
Efforts thus far have concentrated on modeling and performancetrades associated with the reference phase distributiontree.

Tentative

results show that beam pointing is most sensitive to errors in the first few
branches of the distribution

tree, while errors in the later branches reducemain beam gain and spread the radiation pattern.

These results will be used

to continue definition of the reference phase control system.
Rectenna Configuration - As mentioned
system efficiencies,in Section 3 on microwave

some of the earliest efforts on the SPS were directed

toward studies and experiments to improve the power collection
efficiencies.

and conversion

Raytheon pioneered these early efforts and essentially verified

the rectenna concept(ref. 20,23,24).

The initial studies reviewed several

options for the antenna design at the receiving site, including contiguous
horns, contiguous reflectors,

phased array of small aperture elements with

A-35

common microwave load, and an array of small aperture elements with independent
microwave

loads (rectenna concept).

A comparisonof these approaches

is shown

in Figure A-12.
It was quickly decided that an array of solid-state diode rectifier
elements, each combined with an individual dipole antenna and suitable harmonic
filter, was the only option combining both high efficiency
combination, receiving antenna

and low cost.

The

and rectifier, came to be known as the rectenna.

A simplified schematic of a rectenna element
RF-to-DC conversion efficiency

is shown in Figure A-13.

has steadily increased fromapproximately 50 percent in May 1963 (ref. 20) to 82 percent with the JPL Goldstone tests (ref. 22) to 85 percent as a result of Raytheon's
into electronic and mechanical improvements

investigationsin 1977 (ref. 24).

Concepts other than the Raytheon developed concept have been
investigated.

Renssalaer Polytechnic

Institute (RPI) has been investigatingother types of receiving elements which have the potential of reducing mass
manufacturing costs, as well as reducing the number of rectifying
and in some cases, the number of receiving antenna elements.
a continuing study of the reference system configuration,poles, printed circuit yagis and conventional
investigated.

diodes,In addition to

printed circuit di-construction yagis

have been

The most promising to date is the printed circuityagi

whichoffers both the capability of increased gain and thus fewer elements, aswell as a potential for reducing mass manufacturingcosts.

In addition tothis effort RPI has performed an analysis of the sensitivities of series
parallel combining the DC outputs from the rectenna elements.
Alternate rectenna concepts were investigatedBoeing.

by both Rockwell and

Rockwell's approach concentrated on stripline and bowtie dipole

rectifiers.

Analysis indicated that cost and high RF attenuation

traded against reduced number of elements.

should be

The Boeing approach concentratedon (a) increasing the dipole spacing to reduce the total number required;reducing structural cost by trading RF losses versus flat rectenna layout;and (c) increasing the antenna gain (reducing the number of diodes) with a

A-36

(b)

L

*--:3_

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r,-

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A-37

4_
c"
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IlJ

c"

O
IlJ

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A-38

novel "hogline" antenna concept.

The "hogline" combines the characteristicsof the hoghorn (horn parabola) and the line-fed cylindrical

horn.

See

Figure A-14 for pictorial concept.
It is recognized by all that additional studies are required toobtain the desirable reduction in the number of rectenna elements.
some combination

Possiblyof the approaches outlined above, based on incoming RF power

density levels which are higher in the center and lower at the edge will beshown to be more effective.
is the manufacturing

The other area which requires additional effort

methods area, which has the potential of greatly reducing

the cost of each rectenna.
Perhaps the greatest cost reduction will result from developinga novel, low-cost structure and construction technique.
the rectenna is 78.54 km2 at the equator.

The basic area ofThis area in itself is sufficient

to require a thorough study program to develop methods of lowering the structuraland construction

costs.

In addition, at higher latitudes,

the rectenna groundarea will increase as much as 35 km2, or a factor of 50 percent.

(Even though

the rectenna is still only intercepting the circular power beam, since the beam
arrives at the Earth at greater angles at the higher latitudes, the groundfootprint

is changed from a circle to an ellipse.)
The rectenna configuration,

investigations

and experiments

although in many respects has had morethan other areas of the SPS, offers the potential

for greater change in both electronic and mechanicalgreater cost reductions.

A-39

configurations,and thus

Figure A-14.

Hogline Rectenna Concept

A-40

D. Structures and Materials
I.

Satellite
GeneralCharacteristics

theminimalstructural

Althoughthishas been recognizedobviousin conceptualenergyis

as a prominent
generally

percent

ofment oforbiting

teristic

is

the

tural

Studies

less
total

than

ten

systemmass.

systems,

the

of

the

systemtodateindicate

that

the

and verylikelyon thepercent

onset

The reason

(ref.

I),

where

the

is

the

for

satellite.

it

is

not

structure

may appear

entireextremely

alwaysstructural

order

of

low

load

five
environ-terrestrial

engineeringexperiencewhere

structures

mass and energy

investment

requirements.

In light

ratherprime

struc-

is

alien

of

to

the

one ofproviding

structures

which

meet overall

philosophy

thatadequate,

the

large

SPS concept

This

Feasibility

of

for

a verythefromparticularly

pointing.

seeminglyoforbit.

functionisadvantagein geosynchronous

somewhatcan dominate

requirements

portrayals

feature.

mass is- A major

assessments

achievable,

MPTS (ref.

MPTS and array

28).

which

benignload

adequate

stiffnessand systemsystemapproaches

even for

the

might

best

attitudethedynamicofan underlyingThisapproachcontroltolerances

joint

betweenstructural

isolationandon passive

effective.

dimensionaliscontrol

have focused

be coststringentachievetheas a result

will

exceptionforstudiesrequirements

simpleA possible

environment,charac-throughthea "smart"

structure.
The basic
imum gauge material
ratherlarge

buckling.

operatingofa representative

under

low stress,

dimensions

and sized

The structuraldesignofconstruction,areevolvingtherelatively

ideas

generated

as a result

withsquaretherequirements- Earthgravitational
distance

from

orbit

of

the

is

fundamentally

centerinto

shouldisa truss

an adequate

environment.

ofattraction
theofand configurationand thenovel

SPS structure
tieredbasis

operationwithassociatedon thesystemLoads

offeaturesone ofmin-elementofmarginreflect

theStructuralan improvingforelastic

requirements

designapproaches

understanding

of

SPS.

(whichof thea balanceisearth)

A-41

inversely

of

the

body

forcesproportionaltoand centrifugal

theacceleration

(which is proportional to the distance from the center of the earth). The
finite size of the SPS, therefore, gives rise to a distribution of body forces
which, dependingon the geometryand orientation, characteristically
represents
the largest operational applied forces and moments. For a rectangular, 5 GW
ground output configuration in the worst orientation, control forces on theorder of 300 N at the corners of the array would enable attitude control. The
solar radiation pressure of about 5 x lO-6N/m2 acts on illuminated surfaces asa function of the surface solar reflectance and orientation.
The greatest
influence of this force is a potential daily and six-month periodic perturbationof the orbit.
There is also an antenna recoil from the microwavepower transmission of about 25 N. Solar and lunar gravity and earth eccentricities giverise to potential orbit perturbation but insignificant structural loads. Thereis no atmospheric drag at geosynchronousorbit. However, in low earth orbit(_500 Km)this pressure (_lO-4N/m2) can give rise to a force which is significant
to orbital

decay.

Operational system induced loads on the SPSstructure must be considered, although they are quite dependenton configuration and system design.
Current interaction forces are generally small, although their greatest influenceis at the joint where the largest currents and least separation distance between
conductors can occur. Dependingon the configuration and systemoperation, centrifugal acceleration about the center of masscan also contribute to structuralloading. The applied forces and momentsfor attitude control and MPTSpointing
are significant inputs to the structural loading as discussed above.
Structural loads associated with maintenance, construction, transportation, handling and all relative aspects of the SPSactivities must be considered in the structural design. Any governing loads other than operationalmust, however, be weighed against the impact to the systemand, ultimately, tothe cost of delivering electricity.
Environment
structures(e.g.,

etc.)

not encountered

are

wind

- The normal
loading,inenvironmentaloxidation
earth

orbit.

A-42

and moisture
However,

concerns
effects,
there

are

of

terrestrial
soil

mechanics,

environmental

factors in earth orbit which must be considered: heat transfer, vacuum,particulate and ultraviolet radiation and tenuous plasma interactions.
Heat transfer
is essentially limited to thermal radiation, since conduction and convection
effects are generally non-existent.
In normal operation, the MPTSis most affected by the thermal environment due to the waste heat generated and due to the daily cycle of orientation relative to the incident solar flux. Since the waste heat must ultimately
be radiated awayto space, the characteristic operating temperature levels of atapered microwaveemission profile can range from 550°Kat the center to_300°K
at the edge. The centerline temperature imposesa limit to the local power
emission and, therefore, the extent of emission tapering. Temperature levels
can limit systems designs, material selections and lifetime characteristics.
An important aspect to structural design is the distribution of temperatures andthe time variations brought about by changesin orientation relative to the sunor shadowingeffects (local or system-widevia occultation).
Structural temperature levels in space are greatly affected by surface properties, overall geometric configurations and orientation relative to the sun. Temperaturedifferences can give rise to significant local structural distortions, degraded structural
performance and overall configuration distortion.
Theseeffects are influenced
by structural material, structural design and overall configuration.
They canbe particularly significant to the flatness of the MPTStransmission surface
and the dynamic behavior of the entire system.
To illustrate the magnitudeof thermal environments, temperature
differences across simple structural memberscan be on the order of 50°K (sunside to space-side), temperature differences betweenstructural elements caneasily be greater than lO0°K (due to orientation relative to the sun), and thetemperature changesdue to an occultation are nominally 200°Kand can be 400°K.
To accommodate
this thermal environment, the structural material must be insensitive to temperature gradients and transients (low coefficient of thermal
expansion, _2 x lO-7/°K), or the structure must be active, or the structuraldesign and configuration must be insensitive to thermal effects (environment

A-43

and/or distortion).

The latter is difficult to achieve without compromisingother structural requirements.
The significance of the other environmental
culate and ultraviolet radiation and plasma interaction,effects, vacuum, partiis difficult

due to our limited experience with exposure to this environment.

to assess

The vacuum

environment mainly affects the loss of volatile ingredients and surface deposition of effluents.

Prime concern with particulate

and ultraviolet radiation andplasma interaction is the stability of surface properties
tance and infrared emissivity.

such as solar absorp-

It is possible that structural properties

be affected particularly for minimum gauge materials.
have maintained operational performance

may

There are spacecraft whichin this environment

and there are spacecraft which have suffered degradations

for a number of yearsof performance which

can be correlated to these environmental effects.
System Dynamics - The low structural mass fraction of SPS concepts
to date, the stringent pointing and flatness requirements
disturbancesof the MPTS, cyclic(such as gravity gradient and the configurationseasonal occultations

dictate consideration

associated configuration requirements
array, the most straightforward

kinematics) and theof the SPS system dynamics and thesuch as structural

stiffness.

For theapproach to achieving dynamic stability with apassive structure is to have the natural frequency of the solar blankets be greaterthan that of the overall array which must in turn be greater than the control
frequency which is in turn greater than the disturbancedient ~2x IO-5HZ).

frequency (gravity gra-

The array natural frequency for a given size array and agiven structural material is established by the geometry of the structure (e.g.,
depth), while the solar blanket natural frequency
level (_4N/m).

is controlled

This requirement of passive dynamic stability

design structural compression loads (nominally~lO 3 N).

material has a high coefficient of thermal expansion

by the tension
is the source ofIf the array structure(such as aluminum),thedynamic response of the SPS system to occultation would set the stiffness requirements and many of the design loads.

A-44

The stiffness
ing requirements(~arcthat of the array.

requirements for the MPTS are dictated by its point-

minute) and a separation of its natural frequency from

The reference system meets both of these requirements witha frequency margin of more than an order of magnitude.
Materials- Early SPS concepts employed aluminum as an efficient

structural material with a wealth of aerospace experience.
tural and thermal/structural/dynamic
desirability

As the thermal/struc-interactions became apparent, however, theof a structural material which was insensitiveto

the thermal en-

vironment also became apparent.

Since this insensitivity can be readily obtainedby the use of graphite composites(more than two orders of magnitude lower co-efficient of thermal expansion than aluminum), this material has been consideredas the prime candidate for an SPS structural material.
materials

have a higher Youngs modulus-to-density

The graphite composite

ratio than aluminum; however,its raw material costs are an order of magnitude higher than aluminum.
The trend of graphite composite material costs is downward due to an expanding
market; however, for the raw material cost to approach that of aluminum would
require a major market acceptance such as the replacement of steel in the automobile.

Energy investment requirements for the production of graphite composites

are on the order of one-half to one-quarter that of an equal mass of aluminum.
Raw material costs are only one facet of the structural systemcosts.

The ease by which a material can satisfy overall system requirements

or conversely

restrict system performance could have a much greater influenceon the final cost of electrical energy.
material

is a thermophastic

should be emphasized,

For example, the Reference System

resin which provides ease of forming.

however, that this is a preliminary

on the current level of understandingItselection basedof the structural material

requirements.
Development Features - Our lack of experience

with a space structuresuch as required for an SPS was pointed out in the beginning of this section.
There are other unique features which should also be mentioned.

A-45

First, there is no known way of testing a full-scale SPS structure,
as it is currently envisioned, on the ground.

The structural

one of stiffness which requires a zero gravity environment
mental verification of structural performance.

requirementisto achieve experi-

Even component testing would bequite limited since some of the structural concepts would not support their own
terrestrial weight.
ance.

This implies a design predicted on analysis of the perform-

A logical way to achieve confidence

in this analysis capability

scale model similitude testing and component testing.

is throughIf the analysis predicts

the data on a scaled system test, confidence would be obtained in the predictionof full-scale system performance.
Second, the design and performance of the SPS structural

system isclosely coupled to the design and performance of the attitude control and pointing systems.

Major load path characteristics

and dynamic

inputs to the structure

depend on the control system design and operation.
Third, the thermal and structural analyses of the SPS can only be
decoupled if either a low coefficient of thermal expansion

structural material

is used or an active structure is employed.
Fourth, the constructionof the SPS structure and the assembly of

all the systems is a major considerationin the structuraldesign and in thedevelopment of the overall system configuration.
Structural Configurations- An initial considerationof a structuralconfiguration for the photovoltaic SPS array used the relatively massive power
distribution

system as a prime structure "mast", with a solar cell support truss

made up of minimum gauge aluminum "venetian blinds"(ref. lla). This approach afforded

adequate structure for the minimal loads and afforded an elastic buckling mechanism to accommodate local off-design

loads.

An initial MPTS structural

guration (ref. 23) employed a two-tier truss structure.
afforded overall stiffness, while the secondary
radiatingsubarray elements.

flatness was recognized

The prime structure

structure accommodated

The significance of thermal distortion(ref. 23) along with the attractive

low coefficient of thermal expansion material

to achievablefeatures of astructuralunits (ref. llb).

A-46

the

such as a graphite composite.

Studies of thermal engine concepts led to different
large concentrated masses in modular

confi-

concepts due to theLater studies of photovoltaic systems (ref. 2) achieved symmetry
via a dual system of MPTS.

This eliminated the need for any interference between

the microwave power beam and the structure.

Structuralconfigurations

weresized for dynamic stability and incorporated tiered triangulartrusses withcylindrical or geodetic cylindrical elements.

of construction

The significanceon the system and structural configurations was recognizeddirected from structural efficiency

and attention was

to ease of construction.

the SPS (ref. 5) suggested the incorporation of the modular

Further studies ofunit concept forphotovoltaic, as well as the thermal system.
More detailed considerations of operationala photovoltaic

loads and dynamics of

SPS with aluminum triangular trusses made from beam builders aredeveloped in ref. 6.

The study of space shuttle orbiter delivered erectablestructures resulted in the use of a planar truss based on a tetrahedron.

Thisstructure

(ref.

lO).

(called a tetratruss) was employed in a two-tiered

A concept for an efficient space-fabricated,

MPTS antenna

graphite-composite

cylindrical element has also been developed (ref. lO).

geodetic

A first order analysis

for a 1/15 scale similitude space test of the SPS structure and its major systems
under appropriately

scaled loadings and environmental

factors has also been

performed.
Further studies (ref. 8) of the SPS structure addressed the potentialof an erectablestructure via an efficiently packaged graphite-composite,cup" element for establishing

firm cost estimates of the structural

Recent innovative structural concepts for the MPTS and photovoltaic
been developed(ref. ll).

Ref. II provides a good discussion

system.
array haveof structuralelement and configuration trades and addresses control/structure2.

"dixietrade-offs.

Rectenna Structure and Materials
The rectenna is the ground unit of the SPS which receives the beamed

microwave

power and converts it to grid compatible electric power.

The receiv-

ing area is elliptical due to the location of the receiving station relative tothe equatorial orbit plane of the MPTS antenna.

As now envisioned conceptually

the rectenna is composed of rows of panels oriented normal to the incoming beam.

A-47

Theweight of the rectenna structure is a significant part of thetotal design load and, thus, enters into the material selection process. Environmental loads include wind, rain, snowand ice, etc. Studies (ref. 35) haveindicated that steel is the most cost effective material for the rectenna structure. Other materials studied were aluminumand wood. The disadvantages ofaluminumare relatively high cost (comparedto steel) and electrical energy
demandfor its production. Woodis projected to have a higher cost and its
lifetime is questionable.

A-48

E.

Space Transportation
Systems

Considerations

The transportation
Systems

selectionin terms

ofisdollarsby theper mass toorbit.

ofbetweeneconomic

considerations

the

earth

that

payloads

entry

and landing,

A single

and theloads

and highsuitablerequiretive

separate,functionallyatapproximately

their

a compromise

theiroperational

supportregimesThe Heavy

Lift

LaunchVehiclematerialand personnel

traffic

Transfer

Vehiclethe(COTV) and therequirements

between

LEO and GEO.

studies

have

most

cost

effectively

spectrumofpotential

Severalthatof

responda widebase ofStudy
ofdesignfeatures(NAS 9-14710),

payloadand annual

contracted

tomass toandbetween

earth

earth

the

appropriate

flightdesign

that

not

of

four

basic

which

respond

to

needs

of

with

be costthe

Personnel

LaunchtheTransfervehiclerequirements

(PLV)

Cargo

Orbital

Vehicledesigncharacteristics

These

studies

provided

A-49

The Heavy Lift1975,

on vehicle

handle(POTV) satisfy

rate

have

of

cargo.

Vehicleand launch

LEO requirements

site.

types

payloadJulylocationlaunchvehiclethepresent
competi-human and materialand LEO while

in

non-

propulsion.

The LEO staging

inclinationofforbe difficult

would

vehicles.
thedesigned

low thrust)would

SPS study.

Boeingand low

atmosphericregimes

the

the

the

(possibly

Orbitalmissions.

to

Performance(OTV's)earth

investigated

supporting

space.

vehiclesand thePersonnel

andinlaunch,consists

betweenpersonnel

orbit

differing(HLLV)

possible

to

for

carry

cost

LEO and geosynchronous

facilities)

and thelowestbetweenand atsystemthe

SPS program.

re-optimized

500 km altitude

associatedbothoverallflight,impulsefor

The SPS transportation
(plus

with

transfer

specific

and wouldisdesignedtheistransportationtransportation

technology
withlocationsassociatedby orbital

designsystemby vehicles

whereasloads
vehicletransportation

dictaterates,

inneed to achieverequired(LEO) be accomplished(GEO) be accomplisheda key element

and theorbit

atmosphericisdominated

The mission
material

systemrequirements
a dataLaunchinvestigatedconfigurations

theVehicleseffectforthemost cost effective performance. The Future SpaceTransportation SystemsAnalysis
Study (NAS9-14323) contracted to Boeing in September1974, investigated thetransportation systemconcepts best meeting the requirements of several future
program scenarios. These studies, plus several in-house heavy lift launch vehicle studies, provided an initial appraisal of candidate vehicle concepts for
SPSrequirements.
The Heavy Lift
The function
site

to

LaunchVehicle(HLLV)ofthe

HLLV istotransport

LEO.

There

were

five

basic

vehicleI.

Single

Stage

to

Orbit

(SSTO)

2.

Two Stage

to Orbit3.

Modified

SSTO - Winged,

4.

Two Stage

to Orbit

5.

SingleWhere:Stage

to(SSTO)

takeoff,

Horizontal

takeoff,

Study.

Its

shutdown after

until

a 92.6theLaunchcharacteristics
127.4

are

secondsofAfterdischarging

and reenters.

Final

deceleration

ciallyconstructed

The vehicle

then

to

near

5 km diameter
undergoes

and theand theachieved.

A small

The LOX/RP-I

separate

payload,forlandingiszerovelocity.

The vehicle

fresh

water

pond adjacentA-50

and isthevehicleaccomplished

ready

for

Definitionengines

continue

stage

the

refurbishment

extensively

SPS Systems

LOX/LH 2 engines

orbit.

vehicleStudy

was investigatedflightpayload's

theconcept

A-15.

is

bring

landing

shown in figure

x 500 km orbit

which

landing

Horizontal

VehicleHTOHL

landing

VTOVL - This

Lift

considered:
VTOVL

Winged,VTOHL = Verticalthe HeavylaunchVTOHL

Vertical

in

earth

VTOHL

takeoff,

Horizontal

the- VTOVL- Winged,

SSTO - Ballistic,
by Boeing- Ballistic,

VTOVL = Vertical

HTOHL :

fromconfigurations- Ballistic

Orbit

cargo

then
is

are

alone

circularizes

deorbitedby LOX/RP-I

engines

retrosinto

a spe-

tolaunchsite.

thereuse.

0
0

O,

0

.J

0

>

o

o(

i-m

d

°r-

>.
r,-

°p,-

Z

¢o

UJ

-I
e_

_4

i

°r-

y

=;

XZ
-Jl-I

0

0000

\
\

/
_W
_

_,__

_

o_
O

I

0
>>

A-51

_

Two Staqe
siderable(EDIN)

detailto

Ballistic,

by Boeing

evaluatethe

VTOVL - Thisand JSC.

following

concept

JSC utilized

four

was evaluatedcomputer

aided

in

con-

designtechniques

configurations:

Propellant
Ist

Stage

2nd Stage

Payload

LOX/LH 2

450 MT

02/RP-I
02/Propane"

450

02/RP-I"

90002/Propane"

900

The characteristics
forcomparison.

forsizingminimumof
Weightlift-off

weight.

was designed

fordifferential

throttlingon watervehiclesestimating

and performance

Boeing

JSC and utilizes

retro

stage

Single

cent

the

site.

Its

Two Stage

Winged,atinvestigated.

a SaturnforSaturnwere

a waterinexternal

A-17.

The vehicle

It
stage

VTOHL - This

characteristics

basis

is

steeredislandedpond adjacentvertically

tank
but

used

on thefigure

The first

hydrogenVoflanding.

and operation.

Orbit,

launchedwithto thebydownrange

launchpad.

concept

was investigated(hence,modified

lands

SSTO)

horizontallyand configuration

are

adja-

illustrated

A-18.

JSC utilizing

propellants,

is

to

A-16,

optimized

illustrated

no gimbals.

Stage

The vehicle

figuretoon a 5 km diameteran expendable

launchdeveloped
were

in construction

winged.

instagesa concept

and istorelationships

Bothand usessecondshown in figure

pointssimplicity

and theareand staging

developed

Modified
at

these

EDIN.
propellant

VTOHL - This

Seventeen

variations

feedarrangements,

These variations

included:

was analyzed

representing

different

payloads

Payload

500 toPropellant

LH2,

A-52

configuration

I000

C3H8,

and structural

K Ibs
CH4 and RP-I

extensively

engines,material

wereII0 MT oayload
•

^

• I
i

/--33' O
,,"3

-

i

-_ll,,---i i -L)!,.
Booster

2nd Stage./-600

--6,

°

312'

l
1--80' O,\

A_'-- 80'

O

/

%1,

I
-Uv

t

i I

!

'I

LO X

RP-1

i

,L

,

/
I

I

02ARP-I

02/Propane

Booster

9O7

45--4

90--7

Stage I inert, tons
Stage 1 propellant, tons

500
4441

889
8236

485
4410

865
8177

Stage 2 inert, tons
Stage 2 propellant, tons

233
1937

40O
3599

245
2065

421
3832

Gross lift-off weight, tons

7565

14031

7659

14203

Payload,tons,

90 x 500 km

Number of engines, stage 1
Number of engines, stage 2
Staging altitude, km
Staging velocity (REL),

km/secBooster maximum down range

Figure A-16.

45__4

Booster

12
6

24
12

12
6

24
12

43.4
1.84

43.5
1.91

41.3
1.70

40.6
1.78

381

396

346

357

Two-Stage Ballistic Launch Vehicle

A-53

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Stage oxidizer, tons
External tank inert, tons
External tank fuel, tons
Gross lift-off weight, tons
Number of engines (uprated SSME's)
Tank staging altitude, km
Tank staging velocity, km/sec
Thrust/weight,ratio

Figure A-18.

140

Modified Single-Stage-to-Orbit

A-55(shroud)- 45 (inOrbiter)
61
166
2352
69
392
3143
15
111
7.82
1.25

Launch Vehicle

Stages

Tandem and parallel

Materials

Aluminumand Composite

I0

3 and 30 Ibs/ft

PayloadstagedCrossfeedWith

and without

Stage

Flybackis

illustrated

inSectionfortheprovideflyback

capability

adjacent

to thelaunchsite.

MSFC and Rockwellarecurrently

and preliminary

sizing

of

burn

with

The vehicle

climbstodives

to achieve

optimumthement ofauxiliary

III

first

of

a two stage

landing

this

stage

engagedHTOHL - This

takes

off

conceptinreport.

Jet

and bothstagesthewingedland

configuration
vehicle,was investigatedhorizontally,transonic

flightand accelerates

in

with

operateengines

550 km orbit.

propulsion
Each ofsystemthese

fivecruisetoparallel

the

using

for

airbreathinga 91X

and downrange

altitude

At Mach 3, ramjets
wherecombustion,

featuringcrossfeed.

SST0 - Winged,
Rockwell.

crossfeed

Returnenginesparallel3

Gas generator,
Tripropellant

concept

development

Ibs/ftEnginesFirst
The current

densityturbofanequatorial

theby MSFC andjetplane.

a supersonicturbofansup toand theSSME's complete

The rocket

enginesarethenPayloadorbit.

optionsthewas evaluatedfeasibility.

A-56

It

into

shutdown,

circularizes

engines,forclimb.

130,000

shutdown

technical

then

the

feet

achieve-and anand economic

Important distinguishing

characteristics

of each concept aresummarized below:

VehiclePropellant

SSTO-Ballistic

PayloadMT

LOX/RP-I, LH 2

Two Stage Ballistic

230

Recovery Area
Landing pond at
Launch- Stage l

LOX/RP-I- Stage 2

LOX/LH2

390

Launch site - Water

SSTO - Winged Vertical Takeoff

LOX/LH2455

Launch site

Two Stage Winged - Stage l

LOX/CH4- Stage 2

LOX/LH2424

Launch site

Air/LH 2

91

Launch site

SSTO - Winged - Horiz. Takeoff

Downrange - Water

Launch site

LOX/LH2

The following considerations
winged, vertical takeoff, horizontal

led to the selection of the two-stage

landing, configuration

as the reference

system:I.

In both ballistic and winged concepts a two stage configuration

has fewer technology development requirements and is less sensitive to designand operations constraints which result in lower operational costs.
2.

Winged vehicles are preferred to ballisticof the operational

vehicles because

simplicity and reduced recovery time of horizontal land

landings adjacent to the launch site compared to vertical water landings andthe recovery of first stages downrange in the sea.
3.

Hydrocarbon fuel is preferred to hydrogen for first stages

because of its lesser volume and cost.
of methaneby coal gasification

period involved.

The relatively

simple production

guarantees its availability

in the time

For these reasons methane was selected for the first stage

A-57

while hydrogen is to be used for the second stage because of its superior• performance.
Personnel Launch Vehicle (PLY) - The PLV transports crews andcrew supplies to LEO.

The vehicles consideredderivatives of the STS.

for this function were all

MSFC and Boeing have investigated

for replacing the STS solid rocket boosters

various concepts($RB's) and have, in particular,

studied a twin LOX/LH 2 strap-on booster configuration.

JSC also investigatedseveral alternatives including series and parallel burns, winged andballistic configurations and the use of LH2, RP-I, and CH4 (methane) as fuels.
Rockwell has proposed a concept for modifying the Orbiter tocarry 68 passengers and other concepts have ranged between 50 and lO0.
Improved booster performance would permit even higher passenger payloads.
The booster configuration

actually to be utilized for the SPS

program will depend on whether it is designed strictly for the SPS traffic
requirements or if a suitably modified STS will be available.
system is based on a design for SPS requirements
SPS HLLV technology.

The reference

and compatibility

with theIt utilizes a methane fueled winged booster, seriesburn and a resized, smaller external tank.

(See figure 19.)

Cargo Orbit Transfer Vehicle (COTV)- The COTV transports cargo from
LEO to GEO.

Several approaches

to meeting

this requirement were considered:

A conventional chemically fueled high thrust COTV with a short round trip
time and a high degree of reusability,

a nuclear COTV featuring

a propulsion

system capable of high thrust and high specific impulse, and an electrical
COTV requiring an extended transit time, but deriving the electricalto run its ion engines from a solar cell array.

The electricalpower

approach was

subdivided into two options, the self powered option in which part of the
SPS solar array is used to power the COTV and the independent option in which
the array remains dedicatedto the COTV function.

The chemical option is a space-based

common stage system withboth stages having the same LOX/LH 2 capacity and utilizing conventional

A-58

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types of rocket engines.

The first stage supplies approximately

2/3 of thedelta-V requirements, after which it is separated and returns to the LEO
staging depot.

The second stage completes the boost from LEO to GEO andalso provides the thrust for returning the stage to the LEO staging depot
for reuse.

The vehicle gross weight at start burn of 1,290,000 kg of which

400,000 kg is payload and 830,000 kg is propellant
than two out of every three HLLV flights would
the COTV.

(Figure A-20).

More

be required to carry fuel for

The chief advantage of the chemical COTV is the short trip timesof 30 hours for the round trip from LEO to GEO and back to LEO.
The nuclear COTV concept analyzed combined the desirableof the chemical COTV and the electricalimpulse, respectively.

features

COTV - high thrust and high specific

The stage, shown on Figure A-21, has a nuclear gas

core, light bulb-shaped engine with a theoretical specific impulse of2250 seconds and a thrust level of 890,000 newtons.
breakdown is given in Table A-3.

The component mass

Although such a system could meet the

short trip time requirement for personnel transfer and the high performancerequirement for cargo transfer, the development
nuclear materials in LEO eliminatedrisks and the presence of

this system from further consideration.

The electric approach utilizes low thrust engines with high Isp
and round trip time measured in months rather than hours.
conducted to determine the optimum thruster, propellant
Thruster types considered were the nuclear-thermal,
jet, MPD arcjet and ion bombardment.
cesium and xenon.

readiness.

type, and trip time.

resistojet, thermal arc-

Fuels considered were mercury, argon,

Of these, the ion bombardment

satisfactory performance

Studies were

thruster had the mostin terms of thrust and specific impulse technology

Argon, stored as a cryogenic,

because of its ready availability

appears to be the best propellant

as a by-productconsequent low cost (about $0.50 per kg).

of LOX production

Furthermore,experienceand its
with thedevelopment of 8 and 30 cm diameter mercury ion thrusters is sufficienttoanalytically predict the performance of argon ion thrusters as large as 120
cm.

A-60

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A-62

Table A-3.

Nuclear COTV

Mass

Breakdown

Stage Element

Weight I(9Structures and Mechanics

18,780Main Propulsion

56,850

Auxiliary

600

Propulsion

Avionics

260

Electric Power

480

Thermal Control

1,220

Growth (15%)

II,730Dry Weight

89,920

Other Propellants and Fluids

2,000

Total Inert Weight

91,920Mainstage Propellants
LOX
LH2

124,280

Stage Weight

206,204

A-63

Oncethe thruster type and propellant were selected, investigations
were conducted to determine optimumtrip time, and the desirability ofutilizing the SPSarray itself to provide power for the ion engines. In theself poweredapproach, studies were completed of assembling the complete SPS
in LEOand self-powering to GEOas well as constructing the SPSin eight
modules in LEOand self-powering each to GEO. For the secondoption, thruster
and power processing systemsare located at four corners of the satellite
module and connected to a gimbal systemto enable required pointing (FigureA-22). A joint cost optimization of Isp and trip time resulted in a selectionof a 180 day transfer at 7000 secondselectrical Isp. The effective Isp,
after accounting for losses for attitude control thrusting and the use ofchemical propulsion during transits of the Earth's shadowis about 3000 seconds.
This high specific impulse option therefore requires about 0.25 kg of propellant per kg of payload to GEOcomparedto about 2.1 kg of propellant per
kg of payload to GEOfor the chemical COTVoption.
Approximately one-quarter of the SPSsolar blankets on each module
are deployed for the transfer from LEOto GEO;the remainder are deployedfrom their shipping boxesat GEO. The deployed arrays will be degradedby
Van Allen belt radiation absorbedduring the transfer but will be annealed
at GEOto regain most of the lost efficiency during the final checkout andpreparation process. The ion thrusters and propellant tanks would remain in
GEOas an integral part of the SPS,thus incurring no propellant penalty for
returning them to LEO.
A secondelectric solar array option, called the independent
electric OTV,involves the construction of a fleet of reusable electric
poweredround-trip vehicles and their associated solar array in LEO. Thesevehicles are used to transport SPSmaterial which is fabricated in GEO.
Potentially, this option offers the performance advantagesof the LEOconstruction/self-power concept and the monolithic satellite construction
operations associated with the GEOconstruction/chemical COTV.

A-64

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Two configurationsof the independent electric COTV have beensynthesized and are currently under investigation.
are large, lightweight structures constructed
an all-GEO SPS construction

mode.

Both COTV configurations

in LEO and are associated with

The first concept utilizes a self-annealing

gallium aluminum arsenide array with a concentration

ratio of 2 and carries apayload of 4000 MT for a LEO-GEO trip time of 133 days and a total round trip
time of less than 180 days.

Ion bombardment thrusters of lO0 cm diameter areused with an Isp of 13,000 seconds and argon as the working fluid.
cept is shown on Figure A-23.

The con-The primary thruster array of 259 thrusters issuspended by cables and located at the vehicle c.g.

Additional attitude

thruster control packages are located at the structural extremities.

The

component mass breakdown is given in Table A-4.
The second concept utilizes a silicon photovoltaic
a planar configuration with no concentrationfrom LEO-GEO-LEO is approximately
year for each COTV.

reflectors.

Round trip time

160 days which also allows two trips per

Ion bombardment

thrusters of 120 cm diameter are usedwith an Isp of 7,000 seconds and argon as the working
shown on Figure A-24.

solar array in

fluid.

The concept is

Thruster modules of 296 electric thrusters each and an

appropriate number of chemical thrusters are located at the four corners of
the COTV.

The component mass breakdown is given in Table A-5.

The COTV start-

burn mass is seen as 2085 MT for the second concept as compared to 1442 MT for
the first concept of the independent electric COTV.
of both configurations

is quite sensitive

base such as maintenance/refurbishment

The cost effectivenessto items which have very little datarequirements, design life of various

components and unit cost.
Both electric approaches

have their advantages--therequiring less HLLV flights and the independent minimizing
solar cells while traversing the Van Allen belts.

self-powered

the damage to SPS

The cost comparison

two approaches is difficult to assess due to the lack of data base.
relative advantages and disadvantagesof conducting

construction

in LEO vs GEO have been explored, and GEO construction

A-66of theThe

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has been selected for

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A-67

Table A-4

GaAIAs Independent Electric COTV Mass Breakdown

Vehicle

(Dry)

,.

(MT)

Power Generation/Distribution

249

Thrusters

26

Propellant Tanks and Lines

39

Structure/Thermal

229

Control

Rotary Joint

7

Attitude Control/IMS

22

Primary Power Unit
Total (Dry)572

25% Growth Margin

143

Payload3469

Propellant Up

185

Propellant Down

27

Total in LEO
Table A-5

4396

Si Independent Electric COTV Mass Breakdown

Vehicle

(Dry)

(MT)

Power Generation/Distribution

570

Thrusters

70

Propellant Tanks and Lines

60

Structure/Thermal

8O

Control

Rotary Joint
Attitude Control/IMS
Primary Power Unit

185

Total (Dry)965

25% Growth Margin

241

Payload4000

Propellant Up

835

Propellant Down

150

Total in LEO

6191

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A-69

the reference orbit.

In the comparisonof chemical versus electric COTV

options, the chemical option has a shorter trip time and requires significantly
less technology development

for SPS application;

mance, significantly more propellant,
support the chemical transfer.

but due to the lower perfor-

i.e., HLLV flights, is required toConsequently,

the best approach at this time

appears to be an electric propulsion COTV augmented by a chemical propulsion
system to overcome gravity gradient torque at low altitude (<2500 km) and toprovide attitude control during occultation.
Personnel Orbital Transfer Vehicle(POTV) - The functions of the

POTV are to deliver personnel and supplies from LEO to GEO and to return
personnel from GEO to LEO at 90-day intervals.
function, the electricallytrip time.

For this orbital transfer

propelled option is not viable due to its extended

Therefore, a LOX/LH 2 fueled vehicle which can make the trip in

the order of one day was chosen for the crew rotation function.
For the option of electric propulsion
dedicated

LOX/LH20TV

with two common stages is chosen as the most costeffective POTV configuration.
is shown on Figure A-25.

The propulsion

vehicle with associated modules

The vehicle transports

supplies for 480 man-months
for rotation.

for cargo transfer, a

160 personnel and crewfrom LEO to GEO and returns 160 personnel to LEO

The propulsion vehicle has a start burn weight of 890 tons anda payload up of 151 tons and a payload down of 55 tons.

The payload up

consists of 160 personnel in a passenger module, 480 man-months

consumables

in a resupply module, and a flight control module piloted by a crew of two.
The payload down is identical except the resupply module returns empty to LEO.
For the option of chemical propulsion

for cargo transfer, only one

stage of the two stage LOX/LH 2 COTV (Figure A-19) would be required due to thelower mass of the crew rotation module and flight control module as comparedto the cargo payload delivered
this approach minimizes

to GEO by the common stage COTV.

the DDT&E expense, as well as maximizingtion of the vehicles.

A-70

The use of
the utiliza-

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A-71

Ground Support Facilities- All launches, both the HLLV and PLV

were assumed to take place from Cape Kennedy.

Since the PLV is a Shuttle-

derived vehicle, a large portion of the facilities
shuttle can be used.
extensive.

The launch facilities

and equipment

built forrequired for the HLLV are more

Several launch pads will be required to support an SPS implemen-tation rate of lO GW per year.
for the construction

A preliminary cost estimate of $3.4 billionof the required facilities

and the fabrication of

specialized ground support equipment has been made.
Consideration 9_ 28.5 Degree Versus 55 Degree Inclination - A
comparison was made by the Marshall Space Flight Center which consideredpotential cost impacts of launching SPS construction

materials

theinto a 55 degree

inclination as opposed to the baseline inclination of 28.5 degrees for low
Earth orbit (LEO) assembly.

Potential benefits of the higher inclination

were elimination of chemical propulsion

subsystem requirements

shadowing.

of the already on-board SPS ion pro-Such chemical augmentationby earth

pulsion subsystem would not be necessary at 55 degrees.

Eliminating thechemical system was estimateddollars in development

to save about lO0 million

cost and about 500 million per satellite
propulsiun hardware.

in launch costs associated with the

These savings were reduced however, by the lower launchvehicle performance into 55 degrees.

A number of other factors would have to

be considered in more detail before a conclusive
regarding the most desirable inclination
SPS.

decision could be reached

for a low earth orbit assembly of theThis has not been pursued due to the subsequentnous orbit construction|

selection of geosynchro-

however, the trade data does tend to suggest thatfurther analysis is warranted should the LEO assembly option be reconsidered.

A-72

F.

Construction
I.

Satellite Construction
The early SPS configuration studies indicated that construction

inspace was desirable because fabrication of a large area of lightweight

structurein orbit takes advantage of the absence of severe loading conditions.

Favorable

packaging density requirementswith space fabrication.

for heavy lift launch vehicles can be accomplishedThis allows a reduction in the numbpr of flights andgreater load capability of satellite components when the components are mostefficiently

packaged.

As concepts for construction

have been developed,

concepts have been reflected in various satellite configurations
tability of a configuration
ration.

and construc-is a major parameter in selecting a design configu-

The difficulty of constructing a configuration

the concentrationtheratio, where concentration

become more difficult to construct.
GEO or LEO, the transportationis generally related toratios greater than two begin toWith regard to construction

and operation requirements

location inare the principal

drivers in the selection.
A variety of construction and support equipment will be necessary
to complete a satellite.

Because of the very large scale of the operation,ahigh degree of automation will be employed to keep the number of personnel to areasonable level and the total construction time to a minimum.
will principally perform monitoring and repair functions.
been defined, the need for a maintenance capability
operational

These personnel

As SPS systems have

has become an importantconsideration.
Construction

Location - Low Earth Orbit(LEO) vs. Geosynchronous

Earth Orbit (GEO) - The issue of where to construct the SPS received considerable
study effort.

Conclusions varied due to the sensitivity

to assumptions andperformance parameters.
Constructionof the satellite in GEO offers many desirable features.

Gravity gradient loads are two orders of magnitude
drag loads are not significant,lower than in LEO, aerodynamic

thermal effects from passing through the earth's

A-73

shadoware negligible, collision hazard from other satellites is low, and theconstruction sequenceis simpler. Personnel logistics requirements, on the other
hand, are greater than in LEO,but the percentage cost impact of personnel is
not significant.
The most effective modeof construction in LEOis to build the satellite in modulessized to be compatible with the thruster requirements for thecontrol of the SPSin GEOoperation. The modulesare berthed together in GEO.
Building the satellite as a complete unit in LEOfor transport to GEOis notpractical becauseof control requirements and loads to the structure due togravity gradients.
LEOconstruction offers a potential cost saving by using a self-powered
modewhere the output from the partially deployed SPSsolar cells is used to
power a LEO-to-GEO
propulsion system. The degree of degradation of the deployedsolar cells by VanAllen belt radiation is an important parameter in the LEO-GEO
trade. For self-powered transfer, the satellite solar array must be oversized
to maintain the specified output or the cells subjected to an annealing process
to restore efficiency.
The use of an electric OTVconcept for GEOconstruction
mayreduce the cost differential
betweenLEOand GEOsites; however, radiation
effects also affect the efficiency of the electric OTV.
Studies to date have indicated that either LEOor GEOconstruction
appears feasible.

The GEOconstruction location is used as the reference.

Configuration
studies,totheconstructsatellite
theconfigurations

configuration.

used as a parameter
of

space

for

structureinconstruction

originatedtoimproved

improvethetoability"trusswereofease ofpower

percentage

be small,to

In thegeneratedvariousand theby two early
and the-Relativecomparison

was determined

may be illustrated
configuration"

Constructabilitytherapidly

configurations,configuration".

A-74

initialphasesand thenmethods

construction
systems.
ofsatellite

and simply
whichconstruct
the

SPS

were

devisedunderstanding

weight
were

are

the

was sometimes
As theconfigurationsofrequired

modifiedthem.

"column-cable

or
This

The "column-cable" concept (figure A-26) emphasizeda very efficient use of compressionand tension membersto minimize structural weight. In
the "column-cable configuration", orthogonal central columns were stabilized bycables with the solar cell blankets stretched on the cables also. Construction
facilities
necessarily were dispersed becausethe configuration had to growsymmetrically to be structurally stable. Equipmentto install solar cell blankets
and power distribution had to be supported on the satellite structure and cables.
In order to provide a continuous automatic assemblyprocess, along narrow geometrywas conceived called the "truss configuration" (figure A-27).
Structural efficiency was sacrificed to improve construction. Structure wasautomatically fabricated and the concentrator membranes,solar cell blankets
and power distribution were installed from a facility which was the full widthof the SPS. In later designs, this approachis exemplified by the Rockwell
International extrusion concept (figure A-28). This concept results in a facility
with a size and shape similar to the cross section of the SPS. The "extrusion"concept features uniform generation of longitudinal structure with solar cellblankets and powerdistribution (figure A-29) addedas the structure leaves thefacility.
An alternative concept is the Boeing configuration (figure A-30)
which builds a structural bay then steps one bay to build the next bay whilesolar cell blankets and power distribution are installed in the just completed bay.
Thermal engine satellite configurations are relatively more difficult to construct than photovoltaic systembecausethe geometry is irregularand they require large active systemswith fluids systems associated with theheat absorber cavity and thermal radiators (figures A-31 and A-32). The concentrating reflectors must approximate a parabolic segment. Building the concentrator support in a cylindrical shape (figure A-33) was an improvement, butconstructability
has remainedone of several disadvantages of the thermal conversion approach relative to photovoltaic configurations.
LaunchVehiclePacking

struction

logistic

requirements

in

HLLV isan importanteconomicfactor.

Density- The capability

a manner thatminimizes

The HLLV payload

A-75

to

handlethetheimpacton thevolumefor

efficientcon-Antenna construction Facility

Inte_ation

Cons truction

management

Facility

C ol lector:onstruction
facility

Construction
Facility

Figure A.26.

Construction

Base Concept for Column/Cable
A-76

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operation

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be as small

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toThe automatic

fabricationofpackaging.

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bussescan be densely

also

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hardware.

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A-35

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A-6 provides

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for

the

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a chronologyofEquipment- For thecategories

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were established.
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stations,

compo-

personnel

facilityfluctuation

fabrication

installation.

on theand costing,isofmonitoring,

concepts

equipment

deployment,ormaintaining

construction

degreeand Construction

and construction

construction

positioning

struction

FigureSPS studies.

construction

A primary
of

while

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have been made.

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a relatively

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has been little

estimates

equipment

avoiding

wave-

packing

antenna

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manufacture

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construction

requirementsofon individual

operationscrew

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anticipatedConstruction

components.

by mixing

maintenancethatvarioushave shown thatofare derived.

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launches.

SPS isofarestudiesdensity.

monitoringand power distribution

to avoida functionofmembranestodensity

HLLV'sbuilding

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number ofamountresultsnominalPersonnel

and the

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Space Construction

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payloadto reduce

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vehicleA-34

high-density
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densities

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in

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constructionisprovidinga means

equipmentso it

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function.

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A-84

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tracks to position equipment and provide a logistics capability
supplies and operating personnel.

to transport

The base also acts as a receiving depot for

0TV traffic and it provides personnel housing.
Also considered as construction
remote work stations as illustrated

support equipment(CSE), are mannedin concept in figure A-36.

This concept

provides close contact with work to be performed while maintaining
a controlled environment.

A variety of remote manipulators

will be needed

ranging from short precise ones to space cranes to accomplish
meters from facility supports.

the crew intasks at over lO0

Some CSE tasks are the assembly of automaticallyfabricated beams into the deep trusses needed for the primary structure,servic-ing deployment machines, and installing equipment modules.
The task of constructionof very large lightweightsulted in the concept of a "beam builder".
takes preprocessed flat strip material,
for the cap member of a triangularThis is construction

structure reequipment whichforms the strip into a structural

shapetruss, and attaches cross members to thecap to complete the structural truss (figure A-37).

An alternative approacha automatic beam assembler which builds the lightweightistruss structure from

prefabricated structural elements.
The solar cell blankets are to be stretched in membrane fashion
between bays of the structural truss.

The blankets can be densely packaged ina container which is loaded into construction

equipment which performs the de-ployment.
If concentrating reflectors

are used in the configuration,

ment directly from the shipping container by the construction
conceptual approach.
required flatness.

equipment

deployis a

The reflector will require tension support to achieve theThis requirement complicatesthe installation by the need tojoin the packaged strips at the edges and to apply a lateral stretchingforce

within each structural bay.
The power busses are thin flat strips several meters wide to permit

A-88

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effective radiation of waste heat.

The strips can be coiled for transport anddeployed and installed by construction

equipment with little direct labor. Con-nection of power busses to solar cell blankets is less amenable to automationand may involve more CSE.
In summary, construction of the SPS represents a significant

tech-nology challenge because the size and operational location have no valid analogies.
The number of parameters and options in developing concepts is almost unlimited.
Analysis of construction

has been concentrated

in defining feasible approaches.

A technology base must be developed to establish the credibility of construction
techniques and of productivity and cost estimates.
2.

Rectenna Construction
The initial concept of an SPS rectenna advanced by Raytheon utilized

a lO km diameter rectenna with its groundplane perpendicularbeam.

This was accomplishedby segmenting the groundplaneto the microwave

into a series oftilted (relative to the ground) panels.
An in-house study (ref. 2) was conducted at JSC to further define
the Raytheon concept so that materials requirements and costs could be analyzed.
Primary ground rules for the study were:(1) groundplane must be perpendicular

to the microwave beam, (2) maintenance of the entire area is required, (3) land
must be available for other purposes, (4) terrain is flat, and (5) location isnear the Houston area.

The preliminary design selected was a very simple con-ventional structural concept and no attempt was made to optimize the structure
as to concept, weight, constructability,

or cost.

structure was basically structural aluminum shapes.

Material selected for the
The unit cost for a 5 GW

rectenna system was $O.60/ft 2 based on structural materials,

construction

labor,and site preparation 1975 cost indexes.
The quantities of aluminum required to construct the number ofrectennas projected for the year 2000 would result in a 7 percent increase in
the projected annual United States aluminum demand for the year 2000.

Also, theenergy required to produce the aluminum would require a longer payback time thansteel.

Because of this high aluminum usage, a small study contract was initiated

A-91

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structure

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and concrete.

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mentioned

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to

reducedcandidate

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determine
for

matic

fabrication(or,determiningyear

systemin thePart

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for

2.

The
or

and construction
study

utilized

total

cost.

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be a galvanizedmaterials

2000 aluminum

the rectenna

However,the

SPS program.
if

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reducing

has been investigatedconcept

2 for$O.06/ft

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add anotherstructure

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and materials

and theremaining

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analyzed

expensive

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least

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20 percent

above

the

ventional

over

(ref.

A-38

preparation

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steel
Site

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of

the

concludedlabor.

Inc.,

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Becauseofcan be reduced.

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is

for

the

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for

the

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III

Systems

requiresstructure

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further

A-93

is

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on the

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for

the

structure

is

this

cost,

this

area

needs

The areabelievedtobebecauseno automatic

fabri-An artist's

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A-39.

This

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9).

This

Study

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A-94

G.

Natural Resources
In a program as large as the solar power satellite

advanced technology, the selection of materials
ability of a particular material.

program which utilizes

is often influenced by avail-The material availability

may be limited bytotal natural resource or limited by the amount of yearly production.
program materials

list was published by NASA in 1976 (ref. 2).

shown in column l table A-7.

The first

This list is

However, before that time several modifications

had been made to informal working materials lists based on design changes required
for supply reasons.
I.

The following is a list of significant changes

prior to 1976.

Original concepts for photovoltaic cells utilized gold for elec-

trical contact; however, size and number of cells required for each satellite
precluded the use of gold.

Designs are now using silver, copper, and aluminumin various amounts.
2.

The rotary electrical joint as originallyamounts of silver.

designed contained large

The joint was redesigned to reduce the requirement

for this

precious metal to less than 40 metric tons per rotary joint.
3.

Both klystrons and amplitrons have been consideredwave generators in the satellite.
disadvantage

Each has advantages

for use as micro-

and disadvantages.

One

to the use of amplitrons as presently designed is the 7 to lO metrictons of platinum required (15 to 20 percent of U.S. production)
This demand for platinum would restrict the use of amplitrons
solar power satellite program.

in the cathode.

for use in the

Alternative designs to replace platinum withthoriated tungsten are being pursued.

Subsequent

to the mid-1976 materials

lists, two other significant adjustments to design have occurred because ofmaterial demand.
I.

These are as follows:

A redesign of the rectenna structural members eliminated over 75

percent of the aluminum from the rectenna design.
steel--a more abundant natural resource.

The aluminum was replaced by

Further detailed design studies arebeing pursued to obtain the best minimum resource design for the rectenna.

A-95

2. Gallium arsenide photovoltaic cells have certain advantagesoverthe more common
silicon cells, especially in orbital applications.
Early inthe program, gallium arsenide cells were not strongly considered becauseof
limited supply of gallium. This limit arises from the fact that though widelydistributed, gallium is generally found in very low concentrations averaging lessthan 15 parts per million.
Gallium is found in higher concentrations in bauxite
and can be recovered in association with aluminumproduction. At this time,
gallium recovery from bauxite is the only economically available process. Gallium
arsenide cells are being considered becauseof technical advantagesover siliconcells. Howeverto avoid exceeding the projected availabilityof gallium, thegallium aluminumarsenide solar cell design was driven toward very thin cells(5 micrometers) and synthetic sapphire replaced additional galliumarsenide as
the substrate material. In addition, a further demandreduction is achieved bydesigning a solar concentration of two into the collection system.
Table A-7 showschangesin the estimated materials requirements forthe silicon systemsbetweenmid-1976 and January 1978 for both the satellite andrectenna in columns 2 and 3. In columns4 and 5, the estimated materials requirements for the gallium arsenide systemas projected in mid-1977and January 1978
are shown.

A-96

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A-97

REFERENCES

The following is a partial listing of the reports of studies dealing withsystems

definition of the Solar Power Satellite and related subjects.

are grouped by general subject.

Reports

Listings under each subject heading are chrono-

logical.
Most of the reports listed are available through the National Technical
Information Service.. NTIS numbers are listed in parentheses after each reference if applicable.

Where a series of numbers is given, the report consistsof more than one volume.
These restrictions

Systemdistribution.

are listed in this document as follows:N_

NASA onlyG-

Gover; nent

agencies

only

GC-

Government

agenciesand contractors"PowerFrom theonly

Definition

I.

Glaser,November

22,

2.

"Initial

Solar

An "X" prefix indicates restricted

P.E.,

Sun:

Its

Future,"Science,

Vol.

162,

and Economic

Evaluation

of

Space

1968.
Technical,

Power Concepts,"

Environmental

gohnsonSpace Center,

JSC-I1568,

August1976 (N77-16442

Engineeringand Economic

Analysis,"

TMX-73344,

November

1976 (N77-15486)oPower Systems(SPS)

and 16443).
3.
Space

"Satellite
Flight4.

Center,"Satellite

Corporation,
5.

SD76-SA-0239,"Systems

Aerospace
I0101

Definition

Corporation,and 10102
6.

Power

Contractof

Feasibility

Study,"

NAS8-32161,

DecemberSpace-Based

D180-20309,

MarshallRockwellInternational

1976.

Power Conversion

Systems,"

Contract

NAS8-31628,

February

Power

Conversion

and Delivery

Systems,"

Corporation,

Contract

NAS8-31308,

SecondBoeing1977 (X77-(GC)).

"Space-Basedand Grumman Aerospace
June

System

Solar

1976.

A-98

Econ,

Interim

Inc.

Report,7. "Solar PowerSatellite SystemDefinition Study - Part I," Boeing
Aerospace Corporation, D180-20689,Contract NAS9-15196,June 1977 (N7813099through 13103).
8. "Solar PowerSatellite SystemDefinition Study - Part II," Boeing
Aerospace Corporation, D180-22876,Contract NAS9-15196,December1977 (N7820156 through 20163).
9. "Solar PowerSatellite SystemDefinition Study - Part III," Boeing
Aerospace Corporation, D180-24071,Contract NAS9-15196,March 1978 (N78-20164).
I0. "Solar PowerSatellite ConceptEvaluation," JohnsonSpaceCenter,
JSC-12973,July 1977 (N78-12116).
]l.
"Satellite PowerSystems(SPS)Concept Definition Study," Rockwell
International Corporation, SD78-AP-O023,
Contract NAS8-32475,April 1978.
ll.a.
"Feasibility Study of a Satellite Solar PowerStation," NAS3-16809,
Arthur D. Little, Inc., NASACR-2357,February 1974, NTISN74-17784.
ll.b.
"Derivation of a Total Satellite EnergySystem," AIAA Paper 75-640,
AIAA/AASSolar Energy for Earth Conference, Los Angeles, Ca.,April 24, 1975.
Other Concepts
12.

Ehricke, Krafft A., "The Power Relay Satellite:

Transmission

A Means of Global Energy

Through Space," Rockwell International Corporation,

Report E74-

3-I, March 1974.
13.

"Applicationof Stationkept

Array Concepts

to Satellite

StationDesign," Aerospace Corporation, ATR-76(7575)-l,November

1976 (X77-I0236 through I0238 (GC)).

14.

15.

Contract NAS8-31842,

Billman, K. W., W. P. Gilbreath and S. W. Bowen, "Introductoryment of Orbiting Reflectors for Terrestrial
Center,Solar PowerPower Generation,"

Assess-Ames Research

TMX-73230, April 1977 (x77-73872 (N)).
"Orbital Motion of the Solar Power Satellite," Otis F. Graf, Jr.,

Analyticaland Computational

Mathematics,

May 1977 (N78-15148).
A-99

Inc., ACM-TR-IO5,

Contract NAS9-15171,

PowerConversion

16.

"To Design

University

of

72009

Variable

Georgia,

Contract"Design

UniversityofForwardof Georgia,

18.

"Evaluation

cations,"

Arthur19.

of

D. Little,"Development

of

W. C.,

Cells

in

Final

"Free-Space

Report,"September
22.

theOctoberforSpace Applications,"

1975 (X76-72006(N76-I

Potential

forContract
EnergytoSpacethrough

Space Applications,"

Power Appli-

June

1977 (N77-30612).

Annealable

Solar

Cell

NAS9-15196,

July

EnergyConference

Record,Vol.

Power Transmission

Raytheon

Satellite

NAS9-15294,

ofMicrowave

W. C. Brown,

Concentrators

Transportation

ReportSystem,"

Spire

1978.

by Microwave

Beams,"

12,

2,

1964,

pp 8-17.

Combined

Phase

Ill

Contract

NAS8-25374,

Study,

No. PT4601,

Part

Brown,

and1975 (N77-16619).
"Reception-Conversion

System,"for

Subcontract

1964 IEEE Intersociety

21.

Raytheon

ReportSubsystemNo.

(RXCV)

ER75-4386,

forMicrowave

JPL Contract

Power Transmission

NO. 9533968,

September

19755598).

23.

"Microwave

Contract
24.

Power

Transmission

NAS3-17835,

December"Electronic

Free-Space
PT-4964,"Microwave
J.,

Research

1975

NAS3-19722,
SystemAugustStudies

performed

University,(N76-15594
ofSystem,"RaytheonthroughtheCompany,

ER75-

15597).

Receiving

W. C. Brown,

TerminalofaRaytheonCompany,

Performance,"

Gutman,1977 (N77-31613).

Affectingat JSC,

Contract

Studies,"

Improvement

Power Transmission"lonosphere/Microwave
Rice

Systemand MechanicalMicrowave
Contract25.

Gordon,Concentrators

NAS8.32149.

a Directed

Power Transmission"Experiments

Solar

Inc.,

Microwave
20.

LightedSolar

FR-20066,

26.

NAS8-31565,

Contract

Corporation,

Ronald

Solar

(N)).

17.

4368,

Orientation

AugustSPS Rectenna
1977.

Beam Interaction

Study,"

L.

NAS9-15212,

September

1977

A-IO0

Mo Duncan and W. E.
(N77-33389).

27. "A Solar PowerSatellite Transmission SystemIncorporating Automatic
BeamForming, Steering and PhaseControl," LinCom, Inc., TR-7806-0977,Contract
NAS9-15237,June 1978.
28.

"Achievable Flatness in a Large Micro;vavePowerAntenna," General

Dynamics,Convair Div., Contract NAS9-15423,September1978.
29.

Perkins, F, W. and R. G. Roble', "Ionospheric Heating By Radiowaves;

Predictions for Arecibo and the Satellite
Research (to be published).

PowerStation,"J, Geophysical

30. Holway, L. H., A. H. Katz, G. Melty, "Ionospheric Effects of a HighPowerSpace-BorneMicrowaveBeam,"R.aytheonTechnical Memorandum
T-I028,Waltham, M.A., Nov. 1977.
31. Arndt, G. D. and L. Leopold, "Environmental Considerations for theMicrowave Beamfrom a Solar PowerSatellite,"
13th Intersociety Energy
Conversion Engineering Conference, SanDiego, CA, August 1978.
Transportation
32.

"Future Space Transportation

space Corporation,
N77-31235;
33.

D180-20242,

Systems Analysis

Contract NAS9-14323,

Study," Boeing Aero-

December 1976 (Vol. l:

Vol 2:X77-78969(N); Vol 3:X77-79819(G)).

"Orbital Propellant

Handling and Storage

Systems for Large Space

Programs," General

Dynamics ConvairDiv., CASD-ASP-78-OOl,

Contract

NAS9-15305,

April 1978.
Orbital34.

Construction"Orbital Assembly and Maintenance

MCR-75-319,
35.

Contract NAS9-14319,"Orbital Construction

NSS-OC-RP-O08,
36.

MCR-77-234,

August 1975 (N75-32144).

Demonstration Study," Grun_an Aerospace

Contract NASg-14916,"Orbital Construction
Contract

Study," Martin Marietta Corporation,

Corp.,

December 1976 (N77-23136).

Support Equipnent,"

Martin Marietta Corporation,

NASg-15120, June 1977 (N77-27157).

A-IOI

37.

"SpaceConstruction

GeneralDynamics

Convair(N78-2511

through25113).

37.a.

"LargeSpace

Erectable

Company, April

1977.

Aerospace
Rectenna

Div.,

Fabrication

CASD-ASP-77-OI7,

Structures

ExperimentContract

Study,"

Definition

NAS9-15310,

NAS9-14914,

Study,"
May 1978BoeingConstruction

38.

"Feasibility

Arrangement

of

Inc.,

Automated

Contract

the

Study
Groundfor

Various

Rectenna

NAS9-15280,

forApproaches

to

the

Satellite,"

ProposedtheStructural

DesignBovay

andEngineers,

May 1977.

Technology Advancement
39.

"Preliminary Assessmentof Technology

Solar Power," Johnson Space Center, JSC-12702,
1
Economic and Political40.

"An Initial Comparative Assessment

Power System," R. Caputo, Jet Propulsion

Advancement
March

Requirements

for Space

1977.

of Orbital

and Terrestrial

Lab., JPL-DOC-900-780,

March

Central
1977(_77-22612).
41..
Utility
July"Impacts

and Benefits

Industry,"

Arthur1977(N78-24255).

42.

"Political

SatelliteofD. Little,

and Legal

Power System,"a Satellite
Inc.,

Implications

Econ,

Inc.,

Power
C-80020,of

SystemContract

Developing

ECON-77-195-1,

on theElectric

954639(JPL),and Operating

Contract

954625a(JPL),

August 1977 (N7S,25003).
43.

"A Study of Some Economic Factors Relating

Implementation of Satellite

to Development

Power Systems, NASA-CR-150602,

32698 (N78-I 9654).

A-102

Contract

and
NAS8-

Biologicaland Environmental

44.

"Research

Plan

Satellite

Transmission

NAS2-9655,

March

45.

"Compilation

west

Laboratory,
46.

ment,"

Study

System,"ofBiological

Bernard

Effects

of

the

Solar

D. Newsom and Associates,

Power

Contract

1978.
and Assessment

AO-O2-OI/EA81028,"Satellite
Floyd

for

Power

R. Livingston,

SystemJetof Microwave

Bioeffects,"

Contract

EY-76-C-06-1830

Environmental

Impacts

Propulsion

Laboratory,

Pacific

North-(DOE),

May 1978.

Preliminary

Assess-900-822,

Rev.

A.

May 1978.
Work

Breakdown
47.

Space

Structure"Satellite

FlightCenter,Power System
TM-78155,

Work Breakdown

January

A-Z03

1978.

Structure

Dictionary,"

MarshallAPPENDIX B

CONTRACTEDSYSTEM DEFINITIONSTUDIESI.

ROCKWELL INTERNATIONAL

2.

BOEING AEROSPACE COMPANY

,S-I

I.

ROCKWELLINTERNATIONAL

The following is a description of the SPSconcept produced by Rockwell
International under contract to the NASAMarshall SpaceFlight Center. A
documentationroadmapis provided at the conclusion of this section to aid thereader in locating more detailed information from study reports.
A.

Guidelines

and Assumptions

The guidelines
as fortheB.

and assumptions

Referencetransportation

definition

and their

converted

concentrateenergytoelectricaltheenergyontoessentially

the

same

and spacesystems,

and

Solar

energy

using
solar

illustrated

in Figure

R-I.

solar

havingreflectors

arrayscells.

whichis

The solar

conducted

thisconcept

includeuse ofthreetroughconfiguration

klystron

dc-RF

of

designstructural

cells,

structure,material,

kV power distribution,

phase

control.

5 GWoftothe

km in thetoutility

acceptpower

interface.

to

solar
micro-RF microwave.
(CR=2),

designs,

from

the

locatedmicrowave

and subarraythatconvertconcentrators

mounteddesigned

34°N latitude

dc power

center

is

power

the

a single,

Each rectenna

cells

to a centrallytransforms

frame

km2.

is

antennaantenna78.2

groundThe microwave

compression

of

the

energy

solar

by I0

satellite,

concept

energyaluminum-arsenide

aluminum

includessystemdc electrical

Key features

at

were

relationship.

towave antenna.

providestudy

System.

The satellite

45.5

this

System Overview
The system

is

for

gallium-tensionweb/

converters,

antenna,

GEO construction,a single

satellite

A typical

rectennasite

andlocatedcoversan ellipticalarea

13 km in thenorth-south

direction

east-west

direction.

This

area

providesan active

intercept

comprised

of

strip

line

was basedprimarily

A phasedarray

was selected.

This

selectionand decreased

diode

count

employspreformedhatbraces.

The l-beams

obtained

using

this

on theapproach.

sections,

standard

1-beams,

and braces

supporttheB-2

patterns

bow-tieincreasedinch

structure

diameteron concrete

dipoles

efficiency

The supportand 3.5

structureofpiers.

area

tube

CD.
(]J
{.J
cO

cCn
.t-u_
(D
r-_

(D
0
C)(.}

0
0
t-__

S-

0r-

B-3

The transportation

systemselected

for

the

point

(I)

a horizontal

takeoff

HLLV (one or

two stage)(2)

a dedicatedelectricion

thruster

OTV forcargo

personnel

and priority

cargo

transport;

fortransfer

LO2/LH 20TV

for

LO2/LH 2 vehicle

orbit(shortand cargo.

These elementsareFigureAn alternative

2 stageR-2.

has also

been definedOTV utilizes

theconcentration

ratio

satellite

The basicofsatelliteisconstructedstructionmaterial

The OTV's

are constructedThreeorbital

(3)

low

earth

orbit

base.

A crewsizeofFigurehas been sizedan overall
tointercept

79.0

SPS.

The electriciondesignof

GaALAs solar

cells

at

a

that

the

entirenecessarycon-

assumes

GEO usinghave been identified

tosupportbase;

and maintenance

low

(2)

operations

An additionalinthrusterthe

OTV's.

GEO construction:
base;

and

for

accomplishingof

20 has been estimated

crew

the

construcfor

and operations.
theprovidesystemand subsystems

end-to-end

5 GW ofelectrical

to

the

utilitynecessary

to

size

percent,

it

GW of

solar

energy.

The quoted

worst-casereflectivity

solar

efficiency(15.4%).

Solar

Cells

FigureR-4 shows thein Table

for

power
is

given

chain

6.29

the

is

efficiency

ofconcentrator

The pointforto(30-year)istics.

LEO operation

orbit

includingC.

toconstructionfromefficiency,

cell

earth

640 has been establishedofefficiency

takeoffearth

R-3 shows thetoscenariothe

90 days.

A description

personnel

in LEO.

construction

maintenanceAV)of

and thattransported

satellite

satellite

and smalla single-stage

orbit

bases

in a scheduledand (4)a two stageorbit

in geosynchronous

(I)

tion

satellite(3)

2.

approach

isfor

LEO operation;

earth

option
array

to

earth

in a low

verticalsolar

includes:transport;

distanceillustrated

as a viablefordesignefficiency

seasonal

variation(91%),(91.5%),and theend-of-life

the

is

R-I.

systemwhich

busbar.

With

the

solar

arraystheminimumtheend-of-life(30-year)and Blankets

designsolar

utilizesarray

blanketa GaAIAs solar

B-4

descriptioncelland arrayefficiency

ofcharacter20-percent

\
\

B-5

TableR-I.

Photovoltaic

Point

DesignCharacteristics

Overall Description
5-GW

power to utility

Geosynchronousinterfaceconstruction

location

Single microwave antenna
Geosynchronous

equatorial

operational

orbit

Subsystems
Power conversion•

GaAIAs solar cells•

Concentration

Attituderatio = 2

control/stationkeeping• Y-

I_P,X-]OP• Argonion

thrusters

Power distribution•

45.6 KV DC•

Structure/wiring

Microwave

not integrated

antenna•

Gaussian beam•

2.45 - GHz frequency••

Subarray

RCR waveguide•phase control

Tensionframe

panels- web, compression

structure

Structure•

Aluminum•

Beam machine

Information•(graphite/thermal

plasticalternate

construction

management

Distributed

8-6

as needed)_ 3 GW.

lO,3 GW

66,1 GW ..

9.68 GW

79. 4 GW I
ARRAY(SOLAA]_t
I

7.36 GW

7u.

6.0GW

1
o,,o.,o:1
I o,oo.o
I ,L_, I I ,3.3,

I
I

FigureR-3.

OVEP.ALL
EFTICl ENCY- 6. 29 li

Photovoltaic

End-of-Life

t

Point

Efficiency

DesignChain• G,,,A,.

NG• ARE*

3O6,M2

SOLA_

•

__,_

IC_Ra;SOLAR

g 70Gw
C_OUTPUTJI_ACTERtSTICS
VOLTiGE
4_KV

e12

6

.i,i

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GII_O
CON tACT$;;;_,0_0,_.
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8

_

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FEP

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KAPTONO

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R-h.

Photovoltaic

B-7

EnergyXAP'rONCl,lg&.........o..,.,_.
G.A,..O_,_$
tvl_E

WIE_GHT•CO-CE.,_"O_..,O ,_,O,,0,

FigureII CELL

CONCENT_IATOR
--_-_
•

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• ,_{:

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IAmRAvARRAY
I ARRAYQIILJNK_T

"/

TEUP

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40
_

311

COATING

Conversion_oL.o..
C_

0 II_'_ 16GrCM2

System

" '

?7,E

O L I

AMO, 28°C, and the sizing of the array is based on 125°C operating temperature(18.2% cell efficiency).

The total output of the array is I0.3 GW with a voltage

output of 45.5 kv for each array panel.

The solar blanket weight is 7.65 x lO6 kg,

and the total array weight (including the concentrator)

is 8.83 x lO6 kg.

Thisweight is based on a specific weight for the blanket of 0.25 kg/m 2 and 61.2 x
lO6 m 2 cell area.

A cross-section of the solar cell is also shown (Figure R-4).

The 20-_m synthetic sapphire (Al203) substrate, used in an inverse orientation,
also acts as the cell cover.
The solar blanket layout is shown in Figure R-5.

The solar panel in the

top trough (effective cell area) measures 600 x 750 m x 2 for 900,000 m 2.

Twelvepanels are required for the top trough.

The panels for the two lower troughs2
are slightly smaller in width and measure 550 x 750 m x 2 for 825,000 m(effective cell area).

Twenty-four panels are required for the bottom troughs.

The total deployed solar area for the SPS is 30.6 x lO6 m2 which is comprisedof I0.8 x lO6 m 2 in the top troughs and 19.8 x lO6 m 2 in the bottom troughs.
The basic building block is a l m 2 module configurationand the cells

are connected together in a series parallel arrangement.

The voltage output ofeach l m 2 module is 30.3 V with a current of ll.ll amps.

The module output iscalculated to be 336.6 W/m2 at the end of life.
D.

Solar Array and Structure
Reflectors - Thin reflector membranes

are used on the SPS to reflect

the sun onto the solar cell surfaces and obtain a nominal concentration ratioof 2.

The reflector is made of 12.5 _m (0.5 mil) aluminizedtivity of the reflector was taken at 0.9 BOL and 0.72 EOL.
has a mass of 0.018 kg/m 2.

The reflective membranes

using attachments and tensioning devices.

Tensioning

Kapton.

Reflec-The reflector membrane

are mounted on the structure
based on structural limitof the existing beam design (with safety factor of 1.5) indicates that tensioning of up to 75 psi can be used.
The selected design concept

is the 60 ° Vee trough configuration.

point design solar array sizing allows for 20% reflectivity
30 years.

Figure R-6 presents a cross-sectionalB-8

degradation

view of the satellite.

The
over

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The structure

structure.

strutsofthecables.

stresses

and/orcomposites.

girders

subjected

chronousorbitmaterial

thickness
The girder

operational

of

0.254

mm (0.010

is

50 m on a side,

The threeiscap sectionsatand X-tension

bracesaluminum,

approximately

and weldedby theinFigurethrough

joint

of

the

Power obtainedrings

by theklystrons.

a switchgear.

Subsequently,

toon theFigureR-9.

Excessivetheregimebe adequateofconsidered

bay is

50 m inand thewhen theinsafetygeosyn-forminimuma basic
gauge.

length,stabilized

transverse

struts,by a beam machine.

transverse

struts

The cap

sections,

sheetsof

0.254

cutouts,

which

every

mm (I0mil)

struts,2000 seriesroll-formed,

beam element

The

2 m and modified

transverse

isofflanged,2 m on a side.

power

at

subarray

the

dc/dc

is

operated

member to

subsystemtransferredtoiscircuit-breaker.
member of

On therotating

member,

power

which

outputthe

five

primary

conducted

toa summingisvoltagebusPower isrotating

Each voltage

presented

a summingtheconverters

each

distribution

the

then

rotaryisconducted
voltagesbus throughisconducted

from

the

summing buses

diagramof

power

distribution

sub-

is

in

klystrons.

A more detailed
systemtoison orbitoveralland brushes.

required

135,864

whichform a basic

nonrotating

switchthemarginelements

three

the

throughtoaluminum.

selectionand torquesand each

from

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structural

material

in.),(SG) and manually

slip

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forcesa switchtransferredby theisand Distribution

diagramR-8.

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tomembers andhave been verified

withsystem.

stabiliza-

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88 percent

beam machine

two-tier

designShearfabricatedare made fromPower Collection
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SPS point

by a beam machine.

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ties.

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and thedeflections

isE.

tri-beam

Current

vehiclebasic

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are fabricatedtri-beamX-tensionareR-7 shows the configuration

is made up oftransversetion- Figure

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or nonrotating

portion

B-II

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on theThe latterfunctionofkV blanket

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voltageschematic

being

accomplishedtoassureRotary- The rotarythrough

structure

is

set

inner

of

tri-beamby selective

voltagefabricatedinFigure.

rings

the

outputs

and brushes

from

the

switch

gear.

last

segments

shortingofthecompatible

withsummingbus

AttitudeFigureand use of

R-II

the

attitudecontroltotal

RCS propellant

highspecific

argon

ion

X-lOP

are modified

50-m

by beam machines.

isutilized

to

fixed
is

in Table

Rotarytransfermember to

located.

joint

dimensions

energythroughthesatellite

rotating

The rotary

joint

assembly

R-2.

(ACS)
high-performanceand inertia

electricbalancingrequirements(see table)

are

low,

due primarily

impulse(13,000isbelievedtobe feasible

bombardment

thrusters.
referenceand sun sensors

locationson theand todesensitize

thefeaturestate

command policy

tosystem

as electrostatic

Five

attitudereference

spacecraft

to

sense

these

disturbances.

estimatorsfordeterminingor estimators,
preciseThe ACS hardware

controlmass isrequirement.

B-15

very

RCS quads.

the
the

charge-coupled
laser

gyros

and

units

and dynamic

body bending

The control

algorithms

principal

aixs

and a quasi-linear
and minimize
small

to
with

featuresorThe

determination

thermaltoobservers
providedetermination
as wellsystemstatisticalbody-bendings) whicheightthrusters,

to minimize

employsvariouspropellant

The rings

This

microprocessors.

excitation.

rings.

approach(CCD) star

dedicatedrotatingorientation

propellants.

The ACS attitudedevicea doubleshows an ACS employing

Y-POP,

carryofantenna

Systemcenterconsists

microwave

Controlhexagonal

The joint

joint

given

the

R-IO.

satellitearetoin Figurethecharacteristics

G.

attachedon orbit- The rotary

member upon which
designand outer

girders

noted

joint

illustrated

stationary

Electricalwill

and regulating

Joint

Mechanicalslip

switching

characteristics.

F.

are

are therelativetoorientation,

RCS thrust

structural
the

are

bending30-yearat

cO°re°r-

e-

0

c_
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A:"
v

0..

B-16

TableTOTAL

R-2.

RotaryJoint

DesignASS£I4BLY

OPERATING VOLTAGE (kV)
AMPS PER RING ASSEMBLY
TOTAL MASS (kg)
SLIP

RINGS

42.9
I13,0OO
173,4OO(4)CORE
CLADDING
CORE SIZE (cm 2)
DIAMETER(km)
LENGTH (km)SHOE

Characteristics

BRUSH(I6/SLIP

ALUMINUM
COIN-SILVER
41.3 (CROSS
1.13
3.55
RING

SECTION)ASSEMBLY)

MATERIALSHOE S[ZE
CURRENT(A/cm 2)
CONTACT
AREA (¢m 2)
QUANTITY

75% Mo Sz, 25 Mo + Ta
I I.7 ¢m x 12.7 cm x 3 m
7.75
8.68
64 BRUSHESPER SHOE ASSYGROUNDINGSINGLEPOINT- COPPER

BUS

B-17

(THERMAL

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B-18

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Antennasatellite

antennacomprisethestructure

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made fromcomposite

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Three

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3.40

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R-14.

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described

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major

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rectennacharacteristics

FigureR-21 presents

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rectenna

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19.79

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R-3.

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designed

of

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transmittedisA summary of

the

the
The

to dc power and subsequently

and level

R-20.

array

total

the

of

R-18.

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Figurethebased element

energyThe phased arrayofresultingFigurelevel

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- The power

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50 at

the

subarrays

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required

type

from

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array

each

variesgaussian

primaryissubarray

modifiedcatenary

types

lO.3-m

thus

the

nine

x 30.6

moduleby definition,
the

forms

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50 kW klystrons

efficiency

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characteristics

with

11.64

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B-28

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The limits

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Table

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23 mW/cm2 atthecenter,x

TableType

R_3.

MechanicalNo. ofMech'l.

ModuleMass

and

Km/M 2per density

Mod.

Power

Characteristics

Wt (Kg)per Mech'l.

I

17

21.05

756

2

32

18.84

725.7

3

44

16.64

674.5

4

44

15.18

651.6

5

62

12.83

610.3

6

68

i0.52

671.9

7

82

8.4?.

626.1

8

I10

8.31

601.1

9

175

4.21

476.2

I0

142

2.52

406.6

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TableR-4.

AREA

Sumn_r_AVAIL.

of

Recten_na

Definition(TOTAL

USED
PANEL

AREA

INCIDENT(ILLUMINATED)
POWER

NO.

PANELS

NO.

40

102.1

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99

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179.8

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5.53

(TOTAL)USED

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Incident

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Power

Densityreducing to 1 mW/cm
at the perimeter of the rectenna is compatible with present
environmental limit as stated in the programguidelines.
I.

Mass Statement
A summary of

The two major

thesatellite

segments,

the

equal

in mass.

The major

source,which

includesblanketispredominant

theby themicrowave

Total

satellite

kilograms.
very

small

J.

collector

contributor

the

power

and theantennatocollector

array

the

blanketmass.

Antennaand theR-5.

section,arenearlymass is

the

power

mass propertiesare driven

mass,

includinga 25-percent

mass growth

Propellant

resupplyforattitudecontrolmass compared

to

the

annualTable

The solar

sectionincludesinconcentrators.

which

the

RF radiatorssatellite

and thefactor,is

klystrons.

38.09-million

and stationkeeping

is

a

mass.

Space Transportation

be capableofsystem

(for

short

transporting

operational

regimes

includedistanceand durationflight).

bothand cargo.

HLLV - SPS transportation
factors,ofpresented

segment

GEO, and on-orbit

must

is

array

solar

The transportation
tomass properties

elementofrisk

two HLLV concepts;

configurationcrew

requirements

and operations

a winged

earth

to

LEO, LEO

These systems

toLEO, environmental

complexity

and cost

verticaltakeoff-horizontal

technology

option;two-stage(VTO) and a more advanced

from

earth

led

to

the

selection

landingan HTO-SSTO configura-tion.
Winged VTO-HLLV - A potentialstagesisdepictedwithpropellant

incrossfeeda 225 x 106 kg payload
stageshave fly-back

The boostFigurephase uses

HLLV candidate

R-22.

from

The vehicle

thein an orbit
capability;
LOX/RP in

with

equal

is

a parallelburn

2nd stageand iscapable
of

Ist

to

of

500 km at

an inclination

Ist

stage

only

stage

engines.

the
both

B.-34

size

employsair(volume)configurationofplacing28.5 ° .

breathingThe 2nd stage

employsBothengines.

Table R-5. Photovoltaic Point Design Mass Statement
5 GW
I

WE IGHT,sUBSYSTEMCOLLECTOR(_MI LL ION

i

KG)

ARRAY

iiii

STRUCTURESANDMECHANISHS

POWER

SOURCEPOWER

DISTRIBUTION

ATTITUDE& CONTROL

0.116

MANAG-:HEi'IT

TOTAL

ARRAY

0,050

& CONTROL(14.17)(DRY)

SECT I0;4

STRUCTURE

AND

THERMALCONTROLMICROWAVE
POWER

DISTRIBUTION

TOTAL
SPS

1.685
2.457
7.012
4.516
0.630

MECHANISMS

POWER

INFORMATION

TOTAL

8.831
1.347

CONTROL

INFORMATIONANTENNA

3.825

DRY

& CONTROLMANAGEME_IT
ANTENNA& CONTROLSECT ION(16.3)

(DRY)

30.47

WEIGHT

7.62(257,)GROWTH

i

TOTAL

SPS

DRY WEIGHTPROPELLAiIT
I i

I

PER

WITH

YEAR

I

GROWTH

(W/O
in

B-35

38.09

GRO_wTH)
I

0.093
I

i

!

Mass Properties

Concept

Features•

LOX/RP

1st stage

12.00

•
•

LOX/LH 2 (dual mode) 2ndstage
Propellant crossfeed - Parallel burn

10.2

•

Region of minimum

2 50

•

Staging velocity

2377

0.966

2.13

•

Staging altitude

61 Km (200,000

0.227

0.500

106 kg106 LB

Glow

6.804

15.00

Blow

5.443

WP I
U low

4.627
1.134

WP 2
PayloadI
I

J\

FigureR-22.

Preliminary

B-36

/

VTO HLLV Concept

glow
rn/sec (7800 ft/sec)ft)s,multi-mode

engines which operate on LOX/LH2 during 2nd stage burn only.

configuration
payload.
minimumvery nearly approximates a minimum GLOW vehicle

The

for the prescribed

Staging conditions are such that a "heat-sink" booster requiring
thermal protection may be utilized.
Horizontal Takeoff SSTO HLLV - The horizontal takeoff, single-stage-to-orbit concept represents a more advanced technology option for SPS.

This concept

was considered because of the operational problems related to multiple daily
launches of very large vertical takeoff concepts and the overall operational
flexibility.

Without a commitment

toward accelerated advanced technology

pro-grams, it is not apparent that this configuration can meet current SPS technology
readiness requirements.
The winged vehicle, Figure R-23, is a delta flying wing consistinga multi-cell pressure vessel.

of

The wing contour is a supercritical air-foil withleading edge modified to improve supersonic and hypersonic performance.
The cargo bay floor is designed similar to the C5-A military transport
to permit the use of Airlog cargo loading and retention systems.

Cargo isdeployed in orbit by swinging the forebody to 90 or more degrees about a verticalaxis and transferring cargo from the bay on telescoping rails.
Ten high-bypass,supersonic-turbofan/air-turbo

exchanger/ramjetwith a combined thrust of 1.4 x lO6 Ib are mounted under the wing.

engines

The inlets

are projected by retractable ramps that close the inlets and fair the bottom
surface into a continuous surface suitable for reentry.
Three uprated SSME-type rocket engines(total thrust = 3.2 x lO6 Ib)

provide the required thrust above the sensible atmosphere.

The vehicle is capableof placing a 91,000 kg payload in a 550 km equatorial orbit.
Cargo Orbital Transfer Vehicle - The payload required for construction
is transported to GEO using a dedicated electric OTV.
in Figure R-24.

This concept is illustrated

The OTV is sized to carry 4 x lO6 kg (8.8 x lO6 Ib) of payload

for a LEO-GEO trip time of 133 days.

Approximately ten OTV flights are required

to transport the mass required for the constructionof each SPS.

GaAIAs solar

cells also are used in this concept to provide power for propulsion.

8-37

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theduringsolar

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I00

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fixed

power

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To avoidsignificant

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power

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passenger

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was developed.

theby voltage

every

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near

can be regulated

rotationforflights.

50 kW per

IPOTV) - The construction

studies

provided

only

S_stems

Rockwellis

BecauseTransport

Based on previous

isconvertersarrayspropellant

moduletaken
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voltagesmallofand resupply

area

array

impulsea crewarearelow-voltage

theOTV solarlines.

sources

13,000thedistribution

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the

of

only

beam voltage

mass.

voltage
unit

the

power

power

of 48.

curve

diameterprocessfromspecific

SPS required

multiples

power

the

obtainedis278 W toBecause ofisdischarge,

reducesfurnishes

The otherthruster).

plasmalocation

An auxiliary

per

from

This

thermallysystemloss

2000 V; thisby theSPS propulsion

The beam powerand discharge

limiters.

for

(278

the

The accelerator

generated,array,

ofarrays.

collection

thrusters.

thrusterpropellant.

The power
low-voltage

bombardment

48 persons,kg.

passenger/
module,andspares.

layoutoftheCRI_ isto monitor

and control

and layout

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shown in

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commercial airline practice.

A nominal packing density of 160 kg/m 3 (I0 Ib/ft 3)

is assumed for resupply consumables.
each mission.
storage module.

The resupply modules are to be exchanged

While at GEO, the resupply module may be used as the consumables
Thus, multiple access aisles are included in the sizing.

Agross packing density of 93 kg/m 3 (6 Ib/ft 3) allows for a large growth factor.
The POTV is a common-stage

chemical(two stages having the same propellant

capacity) vehicle, utilizing LOX/LH 2 propellant.

(See Figure R-26.)

The pro-pellant tankage for both stages is the same, but the propulsion system for the
first stage has twice as many engines as the second.

This allows the initialthrust/weight for both stages to be kept at about 0.15 g.
specific impulse of approximatelyK.

470 s at a mixture

The engines

have aratio of 6:1.

Natural Resources
Table R-5 identifies the type and amount of materials required in theconstruction of one satellite power system using GaAs solar cells at a concentration ratio of 2.
L.

NOTE:

Totals do not reflect growth allowance.

Operationsa.

Construction
This section describes

the space-related aspects of SPS operations.

The overall operations scenario is first described.
tion of the construction

approach,This is followed by a descrip-

including the construction

base and theapproach to handling cargo.
Overall Operations Scenario - Figure R-27 summarizes, by location(i.e., earth, LEO and GEO), the space-related

operations

the construction and operation of the SPS satellitesof the necessary ground and space bases.

required to support

subsequentto establishment

The right hand side of the figure

identifies the four types of transportation

vehicles used in the space operations:the HLLV transfers cargo and crew to LEO; the electric cargo orbital transfer
vehicle(EOTV) transfers cargo between LEO and GEO; the personnel orbital transfer

vehicle(POTV), which is a two-stage chemicallypropelled vehicle, transfers crew

and priority cargo between LEO and GEO; and the intra-orbital

B-42

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TableR-6.

Summary - SPS Mass
(Photovoltaic

and Materials

GaAs,

CR=2)

MASS

(LESS

MATERIAL
ALUMINUMSTEEL
TITANIUM
COPPERALNICO-V
SILICONALUMINUMOXIDE(CERAMIC)
KEVLAR/RESlN
GRAPHITE/RESIN
PLASTICSSAPPHIREGALLIUM ALUMINUM
ARSENIDEGALLIUM ARSENIDE
TEFLON(FEP)
KAPTONsiLVERSILVER-PALLADIUM-TITANIUMSILVER MESH (INTERCONNECTOR)
CRYOGENIC ARGON
HEAT TRANSFER FLUID%

12.104

39.72
3.84
O.08
12.01
2.65
0.02
I .43
0.12
8.50
2.06
10.13
0.37
3.71
3.46
8.32
1,89
O. 68

1.171
0.026
3. 662
0,808
0,006
0.437
0.038
2.590
0.630
3.087
0.113
1.133
I .055
2.536
0.576
0.208
O.010
0.042
0.215

TOTAL

30.47
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B-45

(IOTV) transfers cargo and crew between the HLLV and the EOTV/OTV/LEO base,and at GEO transfers cargo between the EOTV and the GEO bases.
All launch operations
addition to HLLV launch facilities
for:take place at a single launch site.

In

per se the launch site includes facilities

receiving, storage and processingof satellite and EOTV construction

materials, spares, and propellants; fabrication and subassembly of selectedsatellite elements; payload packaging;storage, refurbishment

and checkout of

the HLLV, POTV and IOTV fleets; space crew support operations; and space andground crew administration.
The continuing operations
and maintenance;

taking place in LEO are:

IOTV servicing and maintenance;

EOTV servicing

transfer of up and down payloads

between HLLV's and OTV's; integration of POTV stages and assembly of crew/
payload modules to the upper stage; and staging of spent POTV stages and downcrews and integration of same into HLLV's for the LEO-to-earth(Additionally, the EOTV's are constructedtransfer.

in LEO, an operation which takes

place during the early program, with the possibility of additional

units being

constructed late in the program.)
Construction

material for each satellite is transported

175 separate HLLV flights.
EOTV availability

to LEO in

The HLLV flights will be scheduled to coincide within LEO to permit direct cargo transfer.

Five days are requiredfor EOTV loading with their departure for GEO scheduled approximately
five days.

The LEO-GEO-LEO

transport cycle requires 161 days, which includes

allowances for loading, unloading and refurbishment.
the down cargo (packing materials,
containers)

every

Upon EOTV return to LEO

damaged equipment/materials,

is off loaded to the HLLV and the scheduled

consumables

EOTV maintenanceof

replacing thruster grids and replacing the empty argon tanks with full tanks
is accomplished prior to rescheduling.
Satellite construction and maintenancecrews are carried to LEO by

HLLV's, utilizing crew modules which can accommodate
sumables for 90 days.

For GEO-destined

48 people each plus con-crews the module,

is mated with two

chemical propulsion stages in LEO to become a POTV which transfers the crew to

B-46

GEO.

The fueledstagesarecurrent

with

crew

module.

the

cycle

are

transported

crew

and thetotwo spentFeatured

from

the

HLLV todepot

and double

tions

are

small

permanent

of

all

are

this

scenario

EOTV thus

and allconstruction

operations
of

these

the

a largecrew

and light

isintegrated

on theconstruction

satellites.

All

EOTV cargois

SCB, POTV's docksized

a 24-man crew.

the

Operationalsite

rectennaand issite

satellite

designatedcontrolof

The construction
fixture,

the

the

SCB,

permanently

mannedLogisticalsupportthroughits

function

of

both

the

onboardcrew

aretransported

be rotatedBase (SCB)

the

next

the

crews

Construction

the

with

IOTV

is

to

Upon completion

location

directly

thesatellite

transported

at

via

each

and maintenance

GEO longitudinal
satellites

location

supports

by POTV's and will

are

the

and LEO base whichofan integrated(SCB) isSCB.

a small,

site

a

crew.

operations

materials

program.

same launchcargo

LEO staging

onlytransferredbasethroughof

LEO opera-

base

The SCB isand maintenanceis

its

projected.

of

for

on-going

maintenance,

currently

satellite

fleet

the

is

assigned

constructionforsatellite

groundoperational

requirementin each(SCB) produces
the

the

entireGEO Satelliteofprecluding

activities.

Satellite

its

transfer

BecauseGEO duty

HLLV's.

directIncorporated

rectennaatvia

on thedesignatedMaintenanceearth

and work

construction.

their

to

con-

with

of

a 30-mancrews

moduleconcept

controlarea

90-daythetraffic

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their

is

of

satellite

toreturned

those

the

HLLV and arrive

The crew

live

bases

and thethenand crews.

GEO construction
toPOTV's.

cargo

building

LEO by thehave completedofbase withdirectly

Crews whichhandling

primarilyinto

LEO by returning

At GEO a single
employeddeliveredstages
in

the

also

satellite

location.
rate
of

of

base,

one satellite,

Figureequipment,

B-47

90-daythebaseconstructedmaturebase
portionis

moved toisinitiated.

R-28,consists
and theisconstruction

duringtheintervals.

HLLV/EOTV.

- Each satellite

and construction

shown inconstruction

via

The satellite

4 per year

satellite

at

directlyofthebase support

thesatellite
facilities

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B-48

and equipment. The construction fixture is a rugged heavy gage metal structure
upon which all of the elements of the construction baseare mounted. The
fixture constitutes the reference surfaces for the construction operations
and the locating jig for the equipmentwhich either constructs or installs
various elements of the satellite.
The location of the major construction equipment is shownon the
figure. Thirty-three tri-beam fabricators, installed at thirteen different
locations as indicated by 7A and 7B are required. Dispensing devices for thesolar blankets, reflectors, solar array retention cables, and power distribution
conductors 9 and I0 are located at the bottom and sides of each of the threetroughs. The assemblyfacility for construction of the microwaveantenna frame,denoted by 6A, including its tri-beam fabricators 6D is located in the centerof the SCB. Also shownis the RFelements assembly and installation facility,
6B which translates towards the completedantenna frame and provides a platformfrom which the RFelements are attached to the antenna as it translates past
the platform in guideways6C.
Facilities for docking IOTV's and for storing cargo are provided on
the platform 4 located to the left of the antenna fixture in Figure R-28. GEO
construction base support facilities
and their locations also are identified
in Figure R-28. A crew size requirement of 680 has been estimated for accomplishing the construction in the scheduledtime. The crew and their facilities
are divided equally and are located on each side of the hex portion of the
fixture as shown. Oneof these 340 menfacilities
shownin more detail in
Figure R-29, consists of 7 three-module crew habitability complexesplus 2 basemanagement
modules, 2 pressurized storage modulesand solar array powermodules.
The modulesof the crew habitability complexare described in moredetail in the lower right of the figure. Eachcomplex is composedof two of
the crew hab modules, each of which provide staterooms, personal hygiene
facilities
and support subsystemsfor 24 crewmembers;and one crew support
modulewhich provides galley, recreational and medical facilities
and subsystems
for the 48 crewmembersof the two crew hab modules. The base management
modulesB-49

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house the communications

and control systems for the construction

pressurized storage modules include workshops for maintenance

base.

The

of construction

facility elements and satellite hardware as required.
Seven of the modules(enclosed by the dashed lines) are hardened

against solar flare radiation and serve as temporary quarters for the entire
crew when the base is subjected to that environment.
Satellite Construction
construction

Operations - Identification of the major

operations and their time-phased relationship with each other andwith the overall construction

schedule for a single satellite are given in

Figure R-30.
Construction starts at one wing tip and progresses

toward the center

section where the microwave antenna rotary joint is located, and thence
continues outbound building wing No. 2 and terminating at that wing tip.
first eight days are designated for preparation of the construction
Prior to the eighth day sufficient materials have been delivered
satisfy the first several days of construction:The

facility.

by the EOTV toprimary structure material(beam machine cassettes) for I/2 the satellite; solar blanket and reflector rolls,
electricalconductors and switch gear for the first two bays; andcomponents.

antenna

Since the rear side of the facility is always exposed to space withno interference from the main construction activities,
jig for building theantenna

it is implemented as theframe and as the location for assembly, andinstallation of the 30 x 30 m RF mechanical modules.

Fabrication of the microwave

antenna for this Nth satellite was started on the 50th day of constructionofthe previous (N-l) satellite and is continued up through the 48th day of construction of this satellite; at that time it is ready for installation
ring-mounted

into the slip-

trunnions.
Each satellite wing consists of 12 bays 800 m long.

These areconstructed at the rate of one every two days using three 8-hour shifts.
to the start of longeron fabrication,

Prior

the solar array blankets and reflectors

for one bay are placed in position for deployment and attached to the frame ofthe preceding

bay so that they may be unrolled as beam fabrication progresses.

B-51

t

t
B-52

(Figure

R-31

power

distribution

feedersarepositioned

forshift,

areaccomplishedduringmachineoperation

longerons

illustratesthesystemtheunrolling.

as they

the

third

shift.

or 800 meters
2 shifts

toarefasteningoftheshift,

end fittings

are

into

position,

machines

crews

are

are

blankets,

39 men per

shifts.

Duringvarious

at

the

of

The operationisspread

over

etc.,

as thelongeronadvances.

oftwo shifts.

longerons

the

first

added to them and theto thebeams arelongerons.

and securing
moduleswiththeand installation

of

the

previous

tion

of

the

centerisassembly,

frames.

test,

(This

satellite.)takingactivity

The antennaassembly

The synopsis

of

the

shown in Figure

R-32 whichisdescribed.

The antenna

frame

is

constructed(otherwise)

inactivelocatedeach corner

fixturelocatedon thefigure.

Beam machines

antennaframe.

The antennacornerby theconnecting

beams.

The catenary

which

the

RF elements

antennaby two track-mounted

and connected

at

are

elements

by a closed-loop

B-53

antenna

side

of

the

of

the

hex fixture

conveyor.

All

beam

the

antennaoftheduringconstruc-duringconstruc-

SCB, Panelareantenna

and assembly

hexagonal

and suspension
mountedremote-manual

construction

dedicated

operating
cable

continues

on itssubsequently

vehiclesplace,are constructedcablesaretrans-

operations.

was initiatedsection.

Duringofeach beam end.

beam joining

2 m/minute,translated,in placeisantennatransverse

at

rate

Installation

No. 1 construction

the

the

the

wing

proceeding

tostarting

While

the

rolls

beam

2 shifts

during

tion

frame

requiring

shutdown

intotensioned

and mainframe

duringcables,duringmanned manipulator

elements

operations

producebeam translation,requiringon thetheautomaticallyand attachedbeam end fittings,

operations

and fastens

one 8 hour shift.

fabricatedverse

Similarly,

two and one-halffabricatedbeams areinstalled

first

The beam machinesforinstalled

operation.)

deploy.

The transverse
next

are

installs

in approximately
allow

installation

These operations,thesame crew

and cables

blanket(PDS) switches

The longerons
with

solar

A of

the

producetheinitially,followed

web cables

upon

then

on opposite

work

deployed
sides

of

and
the

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Figure

R-31

Solar

BlanketInstallation

Concept(concluded)

0 S/A BLANKET ROLL TRANSPORTER - LOADER@ S/A BLANKET ROLLS® USED ROLL CORES® BLANKET ROLL INSTALLER - REMOVER@ TRANSPORTER TRACKS@ S/A BLANKET DISPENSING SPINDEL QUAD
@ BLANKET STRIP GUIDE ROLLERS@ DEPLOYED SOLAR BLANKET STRIP® LEADIN G TRANSVERSE CATENARY@ BLANKET STRIP - TRANSVERSE CATENARY JOINT LINE@ UPPER VERTEX OF SATELLITE (50M TRIBEAM GIRDER) CROSS BEAM
BEAM ATTACH
POINT@ TRANSVERSE-CATENARY-TO-CROSSCABLE DISPENSER@ LONGITUDINALLONGITUDINALCABLE@CATENARYDISPENSING
SPINDEL@ LONGITUDINALCATENARYROLL@ LONGITUDINAL@ UPPER VERTEX OF SATELLITE (50M TRIBEAM GIRDER) LONGERON IN
BOTTOM

CORNEROFTROUGH

ATTACHMACHINE@ BLANKET-EDGE-TO-CABLE
CATENARY
ATTACHMACHINE@ BLANKET-EDGE-TO-LONGITUDINALDEPLOYEDLONGITUDINALCATENARY@
ATTACHPOINT® CATENARY-TO-LONGERONSWITCHGEAR
MOUNTED
ON
CROSS
BEAM
@
@ RETRACTING PLATFORM FOR SWITCH GEAR AND SECONDARYFEEDERINSTALLATION@ MAIN FEEDER DISPENSER
FIXTURE@ CONSTRUCTION@ TRANSVERSE CATENARY-TO-CROSSATTACH@POSITION;LONGITUDINALATTACH@ ATTACH

26A

27A

EQUIPMENT
ARM

IN

50M

CROSS

BEAM

IN

FABRICATION

B-55

IN

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POSITIONSUPPORT

CATENARY-TO-CROSSGIRDER)

MACHINE

POSITIONNON-ATTACH

TRANSLATINGCROSS BEAM (50M
TRIBEAMFABRICATIONPOSITION®

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TRANSLATINGBEAM

NON-ATTACH

ARM
BEAM

FABRICATIONPOSITIONATTACHPOSITION

FACILITY

IN

BEAM

0

/

B-56

Themovementsof the antenna required to initiate installation of
the RFelements are indicated in Panel B. Uponcompletion of the tension web
installation the antenna structure is released from its hex fixture andtranslated in the -Y direction to the X-Y plane of the vertical section oftranslation track which allows the assemblyto be movedin the ±Z directions.
The RFassemblyand installation facility is then extended in the -Y direction
to be in close proximity to the plane of the tension webto support installation of the RFmodulesonto the web. The antenna is first movedin the +Z
direction until the bottom row of the tension webis accessible to the RF
facility and that row of modules is then installed.
The antenna is subsequently
translated row-by-row past the RFfacility in the -Z direction for RFmodule
installation;
whenall moduleshave beeninstalled the antenna assembly iscomplete.
Completion of the antenna occurs simultaneously with completion of
the rotary joint.
In Panel C the antenna is first translated in the -Z direction
to the end of the verticle section of track, rotated, and then translated in
the +Y direction along the lower track to the position shownwhich locates itover its mounting trunnions for attachment to the rotary joint.
The transfer/installation operations have been designed for minimal transfer distance ofthe completed antenna.
The entire structure and the power conversion system (solar blankets,reflectors and power distribution system), for wing No. 1 is completed on the
34th day. Subsequentto completion of wing No. 1 the construction facility
constructs the longerons and frames in the center section, installs the sliprings, constructs the trunnion supports, installs the trunnions, antenna andthe power wiring in the center. Although 16 days are scheduled for this activity,
the timeline requires only 12 days with two additional days scheduled for transferof the antenna to the trunnion mountsand two days allowed for contingencies.
Immediately upon completion of the center section primary structure thefacilities
for the operation and maintenancebase are installed and the first
operational maintenancecrew arrives to support installation of the antenna
control electronics and satellite checkout, which takes place from day 50
through day 69 as wing No. 2 is also being fabricated.

B-57

By the 51st day all satellite hardware has been delivered. On-site
logistics activities are therefore greatly reduced freeing construction support
personnel for subsystemshookupand checkout during the wing No. 2 construction
period.
Useof the construction facility
is completed on day 78 and it istransferred (flyaway) to the construction site of the next satellite on day 84.
Final satellite checkout and acceptance testing is completed on day 86.
LEO Base - The continuous
supervisionof cargo

and crewmaintenanceofand IOTV's,and up and down trafficfornormaltransfercanbe supported

It

includesEOTV'ssame configurations
trol

cargo

by a crew

one crewand staging

transfers(changeoutofsupportofbetweenthrusterscreens

Itof

HLLV'smonitoring.

and crew.

LEO baseand OTV's,and argon

No depotistheestimatedpropellant

functionthattanks)

isthese

providedactivities

The permanent

LEO baseishabitatmoduleand one crew

supportmoduleas thosemodulesused on the

GEO SCB.

The operations
emergency

providesmultipledocking

ports

for

of

the

electric

OTV'sscheduledtheentireincludescheduled30.

moduleoffunctionsshown inwhich

Figure

R-33.

are

the

of
con-

staging

support.
Construction
LEO.

The fleetbuy for

These vehicles
crew

of

can be constructedat

188 men (36 construction

The EOTV construction
initiationoffacility

satellite

schedule

demand scheduleofexcepttheantenna

Payloadcompositionsof

Figure27 days

workersplus

IIsupporting

be manned earlybutreference- Delivery

FigureThis

R-30

schedule

be deliveredand delivery

only

in

in

to

take

place

in

year

life.

utilizingabased on a I0ofMass Flowcomponent

70 units

rate

will

R-34.

is

the

construction),

Cargo Handlingconstruction

program

is

per

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crewmembers

theprogramper(priortointermittently

thereafter.

requirements

to meet thearedefinedrequirestheby the

that

all

first

72 daysthesequence

must

support

solar

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shift).

mass flow
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individual

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demands.
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requirementstheweightcapabilityofstraints.

Table

R-7 illustrates

satellite.

These

packagesatellitearedesignedand construction

produce

block

for

but

is

225,000needs.

and inon thetransportation

provides

the

thus

packagingin quantities

that

fully

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major

with

the

satellite

construction

areinstalled

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each

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the

basic

building

structurecassettes

quantities

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containenough material

fabrication

to

the

compatibility

primary

sufficientofand specified

earlier.

the

con-

sizes,

described

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was conducted

packaged

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launchpackages

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37.2

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to

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satellite

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and crews.

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of

that

requiredand sequenced

completion

assurearepayloadfigure.

considering

Cargomass flow

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3.2

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one satellite,

mass,

for

have different

make it166 HLLV flights

the mass ofcurrent

thus

allocated

subarrays

characteristics

payload,ofthe 90 days

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representingfor

day ofstructure

and arereplacedbeam machine

to complete

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supportmust

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in the blanket layer and deployed simultaneously.
materials and hardwareare packagedseparately.

Endand side attachment

The reflector panels are 600-mwide and 800-mlong whendeployed.
Whenpackaged, the reflectors havean accordian-fold 25 m wide. The resulting
25 x 600-mstrip is then rolled for packaging as shownin Table R-7.
The 6993 waveguidepanels are the lowest density payload item and,
therefore, becomea major driver in packaging_nd scheduling payloads. Basedon the average panel shipping dimensionsand massgiven on the table, a maximumof 50 panels for a total massof 35,000 kg can be carried in the HLLVcargo bay.
In addition, klystrons (which do not present a packaging problem)
are a major payload item. The microwaveantenna contains a large numberof
subarrays that, in turn, are composedof up to 50 power modules. Eachpower
module has a klystron which is shipped to GEOseparately and inserted before
the subarray is secured to the antenna. Each klystron has an average volumeof 0.092 m3 and weighs 45 kg; 135,864are required for each satellite.

B-63

b.

CommercialOperations

Commercialoperations consists of operating and control of thepower systemas it supplies power to a power grid and maintenanceoperations
required to keepthe systemwithin performancelimits.
Loss of total or partial power from the satellite would occur
several times a year due to satellite routine maintenance, shadowingeffectfrom the earth and from other SPSsystems. Whenthese outages are predictable
and scheduled it should have a minimal effect on the utility
systemsintegrity
and operations since the timing of such outages would be during the low load
periods.
However,assuminga significant penetration of SPSpower systems
in the future, the generation reserve neededto maintain the utility
servicereliability
would be expected to increase to cover the emergencyshutdownsof
chc satellite.
More detailed utility system studies are neededto predict the
impact on reserve levels from emergencySPSpower systemoutaQes.

B-64

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B--65

2.

BOEING AEROSPACE COMPANY
The following

AerospaceA.

isa description

Company under
Guidelines

contract

to

of

the

SPS concepttheNASA Johnson

producedSpaceby theBoeingCenter.

and Assumptions

The guidelines

and assumptions

same as for

the Reference

System.

I0 GW using

two 5 GW ground

for

this

The only

power

outputstudy

were essentially

exception

is

that

microwave

transmission

configuration,

which

the

the

SPS sizeisand reception

systems.
B.

System Overview
FigureSPS without
silicon

B-I

solar

solar

shows thesatellite

concentrators

employing

glass-encapsulated,groundmission

links

for

this

concept

isemploys

klystronmicrowave

each rated

outputis

5 GW.

A summary ofatpresented

I0

in Table

GW throughB-I.

generators,
therectenna

without

in

I00

km2.

The satellite

a crewofa bufferzone,

8 elements

elements

arepoweredby partiallyemployingtransferred

Elements

toof

low earthorbit

each witha 400 metricreferencelaunchthewinged,ton

Configuration

SPS solar

approximatelyby 2-staged,

complex,pending

structuresolar

blanketisdivided

into

into

eight

modulesdiametercrystaltransmitter

antennaorbit

microwave

antenna

distribution

and aconstructed

at

earth

500.

low

The satellite

electric

thrusters

ton

heavy

SPS arelift

Space Centerlaunchlaunched

into

vehicles,was assumed as thestudy.

eachhaving

B-66

chain

area,

metricin figurelocatedefficiency

land

using

B-I,

approximatelyeachtrans-

arrays.

further

256 bays,

is

is

power

The rectenna

approximately

Kennedysupporting

nominalpower

land-landing- As illustrated

planarthe

23 mw/cm 2.

I00,000

payload.

simplegroupedofgeosynchronous

deployed

two microwave

a Gaussian

at

are

singleThe satellite

maximum power densityina photovoltaic

cells.

The nominal

orbit

is

at

the

configuration

102 km2 of
667.5

four

metersby eight

each end ofsolar

square.
bays.

the

is

arrays.

a
The

The bays
A lO00-meter

5300 m X 21280 m solar

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array structure.

The array primary structure is 470 meters in depth.

Eachtransmitting antenna is mounted within a yoke structure which is, in turn,coupled to the solar array structure through a rotary joint/slip

Table B-l

ITEM

Efficiency

Chain

EFFICIENCY

Summer Solstice Factor
Cosine Loss (POP)Solar Cell Efficiency
Radiation Degradation
Temperature Degradation

0.97650.919
0.173 x0.97
O.954 >

Cover UV Degradation
Cell-to-Cell
Mismatch

O.956
0.99 /
0.951
O.998
O.934 "

Panel Lo_t Area
Strin_I_R
Bus I_R

O.151

0.932

/

DC-RF Conversion
Waveguide I_R
Ideal Beam

l.O
O. 97
0.85
O. 98
O. 965 "IInter-Subarray Errors
Intra-Subarray Errors
Atmosphere Absorp.
Intercept Efficiency
Rectenna RF-DC
Grid InterfacingO.
O. 956
981 t
0.98
0.95
0.89
0.97

Rotary Joint
Antenna Power DistrO. 0712108.8

Products/Sums
Sizes (KmZ)

B-68

0.86

ring mechanism.

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B-69

C.

Solar Cells and Blankets
Figure B-2 shows details of the basic solar cell blank structure.

The blanket consists of 50 um thick single crystal silicon solar cells withborosilicate cover glass electrostaticallybonded to the cells front and back.

The cells are designed with both P and N terminals brought to the backs of thecells.

This feature makes it possible to use 12.5 um silver-plated

interconnections which are formed on the substrate glass.
assembled electrically by welding

copperComplete panels are

together the module-to-moduleinterconnections.

Other details and features of the blanket design are shown on figure B-2.
The nominal cell efficiency is 17.3 percent(AMO, 28°C) at beginning of life.

A key feature of the blanket design is the ability to perform in-situ annealingof the solar cells using a laser annealing concept.

Annealing

is required torecover radiation induced degradation of the cells.
D.

Solar Array and Structure
Figure B-3 shows the buildup of a solar array fundamental elements,

which is a blanket panel.
series by 14 in parallel).
assembled

A blanket panel contains 224 solar cells
Its dimensions

are as shown.

(16 inThe blanket panels areinto installable blanket segments by welding the interpanel connectors

and taping the panels together (see figure B-4).

The blanket segments are 14.9 m

wide by 656 m long (about one bay length) and are shipped accordion-folded
suitable box.

in a

The method of supporting the solar blanket within the primarystructural bays is shown in figure B-4.

This method of support will provide

a uniform tension to the end of each solar array segment by the use of constantforce tensioning springs at each blanket support tape.

These springs are also

attached to a catenary cable that is then attached to the primary structure,(upper surface), beams at 15 meter intervals.

The springs are in compression,for better reliability, and exert a uniaxial force of approximately

3.5N toeach blanket support tape.
The buildup of high voltage in the solar array is accomplishedbyconnecting approximately 78,000 sets of solar cells in series (a string).
Solar cell strings approximately 5.1 km long were selected for the reference

B-70

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B-72

photovoltaic
directlycellsystemconfiguration.

the

solar

array

are

identical.

orby thefromstrings

by conductors
are neededforCurrent

generatedby the

solar

cells

themselves.

It

in

the

current

fromthearranged

in

loops

which

start

bringing
being

the

edge ofthearray,

ofsolar

and return

the

start

at

the centerline

back tothecenterline.

cells

outer

edge and thentheprimary

structural

beams,

betweenofaluminum

cable

are

used toelectricallystrings

injumperofthetruss-type

the

using

uniaxial

longcolumnby minimumE.

istensionofbuckling

and local

function

prime

condition,and regulate

generated

for

the

cells

are

noted

that

no conductors

edges

of

array,

fromone centerbus loopbus atthecenter

of

of

the

satelliteisthecarried

the

solar

the

aroundarray.

and go totherefore,

eight

times.

Interbayconnectstrings

in one bay tostringlength.

Figures

B-5 and B-6

mustjumper

cross
mode

the

supportmethodand theinterbay

The solar

array

primary

structure

is

basic

structuralelement.

designresults

beams as the

The trussblankets.

The trusscrippling.

Lightly

strutsloaded

power

of

the

from

the

the

quantity

microwave

power distribution

lead
were

strutssiliconsystemcelland quality

of

the

electricalprovidefortherequired

generators;

energyoccultationor system

maintenancemonitoring

fault

detection,and fault

isolation

array

is

dividedand can be isolated
operations

diagramand/or

of

into
from

other

the

the

is

solar

solar

shows a schematic
solar

solar

a

sized

were

by

sized

Power Distribution

and control

for

All

solar

composite.

solar

electronics.

voltageblanketgraphitetherequiredgauge material.

The prime

duringarray

theA string,bays,

20 m triangular

The structuralmaterialotherbay of therespectively.

designfromnextreference

installation,

power

to

the

show sketches

generatingintervening

strings

appropriate

permits

withoutcell

The strings

This

solar

array

228 power sectors.
main power bus,

servicing.

B-73

collector

shut-down;

distribution

Each power
facilitating

accumulate

panels;disconnects.

power

to

sectorsolar

control,powerenergystorageand provide
FigureB-7

system.

TheisswitchablecellannealingIpIFl143
BEAM

FigureB-5

Reference

Array

BlanketSupport

S_1752f

INTER

BAY JUMPERS

A TYPICALBLANKETJINTER

SEGMENT\.

BAY JU! |PERS /(NO. 12 ALUMINUMFORTOSEGMEN

T

co.. c,o.

BLANKETTENSIONING
DEVICECABLE)JUMPERNTERBAY
CONNECTION

SOLAR

Figure

B-6

B-74

Inter

Bay

Jumpers

STRING

PANEL

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B-75

Solar array power is controlled by vacuumcircuit breakers near thebuses. Voltage is controlled by turning groups of strings on or off, dependingon load requirements. Two sections of the array provide the required voltage
at the sliprings using the sheet conductor voltage drop to achieve the required
voltage at the sliprings.
Powertransfer across the rotary joint is accomplished by a slipring/brush assembly. Mechanical rotation and drive is provided by a mechanicalturntable 350Min diameter. The antenna is suspendedin the yoke by a soft
mechanical joint to isolate the antenna from turntable vibrations.
F.

Rotary Joint
As previously stated, power transfer from the solar array section to

the microwave antennais accomplisheda slip ring/brush assembly.

via a rotary joint (figure B-8) using

Mechanicalmechanical turntable 350 m in diameter.

rotation and drive is provided by a
The antenna is suspended in the yoke

by a soft mechanical joint to isolate the antenna from turntable vibrations.
The antenna is mechanically

aimed by CMG's installed on its structure.

position feedback with a low frequency bandpass allows the mechanicalto drive the yoke to follow the antenna and also provide sufficientAturntable

torquethrough the soft joint to keep the CMG's desaturated.
Figure B-9 illustrates the electrical componentsof the rotary joint.

Coin silver (90% silver and I0% copper) was selected for the slip-ring material
and a silver-molybdenum

disulfide brush with 3% graphite was selected.

With adesign using a brush current density of 20 amps/cm 2 only about 40 kW of power
is dissipated in the rotary joint.

The projected brush/slip-ringwear is verysmall (.0289 to .0617 cm3/year).
G.

Attitude Control System
The attitude control system (ACS) includes all operational

elements

and software required to maintain orbit station keeping and attitude control of
the SPS in the operational orbit or to establish attitude control from an
initially uncontrolled condition.
four installations,

The ACS is an electric propulsion

one at each corner of the SPS solar array system.

B-76

system with
A typical

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B-78

corner installation is illustrated in figure B-IO. The ACSconsists ofthruster, power processor, structure, propellant feed, control systems,
and instrumentation. The thrusters include the primary electric thruster
and auxiliary chemical thrusters required for establishment of attitudecontrol whenelectric power is not generated by the SPS.
The electric thrusters are lO0 cmdiameter ion engines using argon
as primary propellant. Thruster Isp is 20,000 sec. Eachpanel contains 25
thrusters. The net thrust per panel is 150 newtons. The operating lifeof the system is 2 years with a 50%duty cycle and 80 ampbeamcurrent.
The chemical thrusters use LO2/LH
2 as propellants and are used forcontrol during equinoctal occultations.
The engine Isp is 400 sec.
H.

Microwave

Power Transmission System

General - The microwave transmission

system includes the entire

spaceborne array power transmitter which includes the dc distribution
from the rotary joint to the rf transmitters, the rf transmitterssystemthemselves(klystrons), their dc and rf control and monitor circuitry,

and the rf antenna

elements composed of slotted waveguides,

rf feed circuits,

support structure,mechanical pointing control, all the components
control of the phase of the _etrodirective
receiving stations (rectenna).
ponents of the transmitterandantenna subarrays, and the groundFigure B-ll shows sketches of the major com-system.

The design utilizes a retrodirectivearray powered by dc-rf klystron converters.
distributed

required for distribution

phasedDC power from the rotary joint isin a manner to minimize 12R losses to the klystrons, utilizing

unprocessed power with a maximum voltage of 42 kv.
to provide a tapered(lO db quantized Gaussian)85%

The klystrons are combined

illuminationof the array

resulting in low sidelobe levels and high antenna efficiency(over 95%).

The

thermal loading in the center of the array (22 kw/m2rf) permits a design fora l km diameter array which provides roughly 5 GW of dc power on the ground perantenna.

The phased distribution

system is designed to minimize line lengths

and cumulative phase errors in the distributing
3-node reference distribution

transmission

lines by using asystem with line length compensation.

B-79

The pilot

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reference signal from the ground utilizes 2-tone modulation

with a suppressed

carrier near the power beam frequency, to effect conjugation(i.e., electronicfine beam steering) in an efficient manner.

Correction for some systematic

propagation errors is provided through multiple pilot beam transmitting
Antenna Power Distribution - The power distribution

antennas.

system provides

power transmission, conditioning,

control, and storage for all elements of the

microwave transmitting antenna.

Figure B-12 shows a schematic diagram of thissystem.

The antenna is divided into 228 power control sectors, each providing

power to approximately 420 klystrons.

Two of the kylstrons' depressed collec-

tors "A" and "B" which require the majority of supplied power are provided withpower directly from the power generation
losses.

system to avoid the dc/dc conversion

All other kylstron element power requirements

ditioned power from the dc/dc converter.
isolation of equipment

are provided by con-

System disconnects

for repair and maintenance.

are provided for

A nickel hydrogen battery

system provides 12.186 megawatt hours of energy storage for klystron cathodeheater power during solar occultations.

Up to two hours capacity is provided.

DC/DC Converter and DC/RF Generators
B-13) provides power to approximately
klystrons.

42 KV.

(figure

0.5% of the total number of antenna

The klystron with five depressed collectors

efficiency of 85%.
unit.

- Each dc/dc converter

has a calculated

tube

Figure B-14 shows a sketch of the reference 80 kw klystronThe selected design is a continuous wave
It uses a compact, efficient(CW) amplifier operating at(82-85%) solenoid wound-on-bodydesign.

To achieve long life, a cathode loading of 0.15 amps/cm 2 was chosen.
Figure B-15 shows a drawing of the integrated

klystron module.

The

figure shows the klystron mounted on the back of the slotted waveguide antenna
array.

The passive cooling system can be seen.

Not illustrated

here is thephase control system required to insure that the radiation from the modules
will be in phase at the rectenna.

This system will tie the modules within asubarray together with waveguide and all the subarrays together with coaxial
cable or an equivalent transmission
I01,552 of the 70 kw amplifiers

link.

Each transmitting

antenna contains

and operates with a gain of 83-85 db.

B-82

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utilizes

Integrated

Subarray7220-10.4

x 10.4

configuration

array

meter

conforming

maximum RSS error
whichradiation,

composed

of

to

loss

antennathedesired

4 in

the

the

variesfromin

Figure

which

the

I0

db taper.

outeredge ofI0 db taperwill

resultin

B-16 shows a 4-module

radiation

waveguides,

sub-thermalstructureisl-beams.

FigureB-17

arrangementofsubarrayson theThe numberofklystronsperantenna

36 pertheasupportand longitudinal

illustrating

configurationa quantized

The subarraylateralarraytosubarraysubin

center.
Figureradiusthedesireda lO-step- The microwave

overallshapetoand form

open trussofaremetersthewidthmade up of

7-I/2

meter

polysulfone
Secondarycomposite
structureofapproximates

deep,

with

and thestructure

structurerestelements

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Mass Statement
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was derived

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B-2

PHOTOVOLTAIC CONFIGURATION MASS SUMMARY, WEIGHT IN METRIC TONS
I0 GW
Component
1.0

2.0

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1 1

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Power

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7,155

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2,425

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2.5

RF Generation

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5OO

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254
42

Distribution

4,986

and Distribution

19,762SUBTOTAL

81,293

GROWTH

17,590TOTAL ............

B-95

98,883

J.

Space Transportation

system.

This

sectionprovidesa description

Bothlaunchand orbit
launchincluded.

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for

cargo

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facilityrequirementsareand operations/support

are discussed.
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figureB-23.

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439 metricofa payload

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cargo
at

orbit

LEO.

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launched

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simplify

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ofa vacuum thrustweightofanof

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as illustrated

164 meters.

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requirements

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The configuration arrangementand characteristics of the systemelements usedin the transfer of each satellite moduleare shownin figure B-27.
Eachmoduleis 2.7 x 5.4 km. Six of the modulesare transferred without amicrowaveantenna attached. Their massis 8700 metric tons each. The othertwo moduleseach have a fully constructed microwaveantenna attached making
their mass23,700 tons each. Propulsion for the orbit transfer is provided
by a combination of electric poweredion thrusters (argon propellant) andLO2/LH2 thrusters. The ion thrusters provide most of the orbit transfer
energy whereasthe chemical system is used to counteract gravity gradienttorques on the moduleand to maintain desired attitude during the transfer.
Powerfor the ion thrusters is provided by partially deployed SPSsolar array.
The 8,700-ton modules require 13%of the SPSarray deployed to power four
panels of 600 ion thrusters. The heavier modules (with antenna) require 36%of the on-board solar array to power four panels of 1,600 ion thrusters. The
trip time for each module is 180 days.
The SPSsolar array was oversized 5%to compensatefor the radiation
degradation of the silicon solar cells during passagethrough the VanAllenradiation belt. This also compensatesfor the inability to anneal out all of
the damageafter reaching GEO.
L.

Operations
I.

Satellite System Construction

struction, maintenance and transportation

Operations

operational concept for low earth

orbit (LEO) construction of the CR=I photovoltaic
figure B-28.

- The integrated con-satellite is shown inSpace operations crews and all hardware and consumables requiredin space are delivered to LEO by launch vehicles.

The crew launch vehicle wasassumed to be an improved space shuttle with the solid rocket boosters replaced
by a reusable liquid propellant

booster.

wing-wing vehicle capable of delivering
per flight.

The cargo vehicle is a two-stage
approximately

400,000

Kg of payload

Crew flights occur every two weeks while three cargo vehicle

flights are required every two days to each construction
case of constructing one lO GWe satellite per year.

B-104

facility for the

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MAINT CREWSAND COMPONENTS,OPERATIONAL
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FigureB-28

Integratedr.,,s.14ol

Space Operations

[JAYS
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FINAL CHECKOUT&COMMISSIONING
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BASE

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FigureB-29

Photovoltaic

Satellite
B-Z06

LEO Construction

Timeline

The LEOconstruction base is nominally located in a 478 Kmcircularorbit at 31° inclination.
This base housesa crew of 480 with overflow quarters
for transients, e.g., those crew membersawaiting transportation to someotherlocation. The primary purpose of the LEObase is construction of eight SPS
power generation modulesand two antennas. The satellite construction timeline is shownin figure B-29. The basealso serves as a staging depot fororbit transfer vehicles used to carry construction and maintenancecrews, crew
supplies and replacement parts to the GEObase. A construction crew OTVflightto the GEObase normally occurs once every three months. Maintenancecrew andreplacement componentsare also transferred to GEOevery three months.
The satellite modulesare equippedwith electric propulsion systems
and flight control systems for the self-powered trip to GEO.
The GEObase is used for final assembly and maintenanceoperations.
The final assemblyoperations include moduleberthing, antenna placement, anddeploymentof solar array. The maintenanceoperations include refurbishmentof failed SPShardware. The GEObase is also used as a staging area for thesatellite maintenancecrews, mobile habitats, spare parts, and their orbit
transfer vehicles. The GEObase houses60 final assembly crew membersand up
to 240 SPSmaintenancecrew members.
The maintenancecrews are dispatched from the GEObase in an OTVpropelled crew modulealong with an OTV-propelledreplacement parts module
destined for an operational SPSthat is scheduled for regular maintenance.
The maintenancecrew will visit each SPStwo times per year and will spend
four days replacing defective componentsbefore returning to the GEObase orproceeding to the next SPS.
Construction

Base LEO - The LEO construction

satellite consists of two interconnecting
is used to constructfacilities.

base for the photovoltaic
One of the facilities

the module and the other is used to construct the antennas

as shown in figure B-30 and B-31.

B-107

Ilml-ll_l

CONSTRUCTION
FACILITY_..9 kr n

1.0

PRODUCTION

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AND

DIRECTION

FLIGHTkmt..
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FigureB-30

ACS/ORBITKEEPING

THRUSTERMODULELEO Construction

BasePhotovoltaic
PART

III

Satellite

MODULEFRAME
ASSEMBLYBAY

FACILITYSOLAR ARRAYDEPLOYMENTBAY_-----

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B-Z08

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B-31

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The module construction

facility is an open-ended

structure whichallows a 4-bay-wide module to be constructed with only longitudinal
There are two sets of internal working bays.

indexing.

The aft bays are used forstructural assembly using moving beam machines and crane/manipulators
to both the "A" level and "D" level surfaces of the facility

attached(see figure B-31).

Solar array and power distribution components

are installed from equipment

located on the "A" level in the forward bays.

The satellite module issupported and indexed by movable

towers located on the "D" level of the facility.

The antenna facility is located with respect to the module facility
in such a way that the antenna is constructed at a location where the completed
antenna can be mated to the yoke without any vertical movement.

The antennaconstruction

in an open-

facility (also shown in figure B-31) is configured

ended structure that is five antenna bays wide which allows the antenna to beconstructed using both lateral and longitudinal
are used to assemble the primary

indexing.

The two end bays

structure and the inner bays are used to

deploy the secondary structure and subarrays, and to install the power distribution system and maintenancegantries.

Construction

equipment operates

from both the "B" and "C" levels of the antenna facility.
The module construction

sequence for the structure,solar arrayand power buses begins with building the first end frame of the structure.
This completed end frame is indexed forward one structural

bay length.

can then form the remainder of the structure in each of the bays.
row of four bays is then indexed forward to allow construction

Machines

The firstof the second

row of structural bays in parallel with installationof solar arrays in bay l

through 4.

of structure occurs

Solar array installation and construction

simultaneously

across the width of the module although neither operation

depends on the other.

At the completion of the 16 bays (four rows of bays inlength), the power buses and propellant tanks are installed.
the structure and installation

Constructionofof solar arrays of the remaining four bay lengthsof the module are done in a similar manner to that previously

B-109

described.

Thruster modulesfor the self-power system are attached to each of the four
corners of the module. An annealing device gantry is installed on each module.
Construction of the antenna takes place in parallel with module
construction, lhe first antenna is completed during construction of the fourth
satellite module; the second antenna is completedwith the eighth module.
The antenna is indexed laterally through the facility
one bay at a time. Whena full width of bays is constructed the antenna is indexed longitudinally outof the facility so that the next strip of bays can be assembled. Whenthe
antenna is completed, it will be located at the proper position so that it
can be mated to the yoke.
The yoke for the antenna is constructed in the moduleconstruction
facility becauseof its large dimensions. This requires the yoke to be madebetweenthe third and fourth moduleand betweenthe seventh and eight modules.
Following yoke construction, it is movedto the side of the module facility.
At that time, either the fourth or the eighth modulewill be constructed.
During the construction of these modules, the antenna is completed so thatit can then be attached to the yoke. After five bays of either the fourthor eighth modulehave been completed, the antenna/yoke combination can thenbe attached to the module in its required location.
Construction of two morerows of bays pushesthe antenna outside the facility where it then can behinged over the module for its transfer to GEO.
In addition to the construction base facilities,
the LEObase
includes crew modules, work modules, cargo handling/distribution
equipment/
vehicles and base subsystems.
The LEOconstruction base includes five primary crew modules. The
modules have an earth atmosphereenvironment and are sized to accommodate
crew
sizes between50 and lOO. The moduleshave dimensions of 17mdiameter and up
to23m inlength.

B-IIO

Workmodulesare used for operations, maintenance, and training.
They are similar in design to crew modules but have specialized functions which
include clinic, satellite componentsmaintenanceand checkout, and new
personnel training.
Basesubsystemsinclude electrical power and flight control systems.
The total power requirement estimated for all LEObase operations is 3725 kw.
The primary power supply is solar arrays similar to those used in the SPS
with nickel hydrogenbatteries used for occultation periods.
The flight control systemincludes guidance/navigation/attitude typesensors such as IRU, star trackers, horizon sensors and the propulsion system(LO2/LH
2) to perform attitude and orbit maintenancemaneuvers.
Construction

Base GEO - The GEO construction

base is a 2 x 2 bay-wide platform that is attached to and indexed across the solar array side of
the modules, as shown in figure B-32.

This platform has four solar array

deployment machines that are used to deploy the undeployed
are also a variety of crane/manipulators,

solar arrays.

logistics and SPS maintenanceThereequip-

ment aboard.
The first operation to occur once the modules reach GEO is that of
the berthing (or docking) of the modules.

The modules

single edge as indicated in figure B-33.

The major equipment used to performthese berthing operations are shown.

are berthed along a

The concept employs the use of four

docking systems with each involving a crane and three control cables.

Varia-

tions in the applied tension to the cables allows the modules to be pulled in,provide stopping control and provides attitude control system involving
thrusters which are not shown.
During the transfer from LEO to GEO, the antenna is attached below
the module with a single hinge line.
rotatedOnce GEO is reached, the antenna is

into position followed by the final structural

nections, as indicated in figure B-34.

B-Ill

and electrical con-

KLYS_[RON
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GE0 Final Assembly Base/0perations

B-II2

•

MODULEMASS
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MODULE(N)

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FigureB-33

GEO BerthingiConcept..Photovoltaic

Satellite

2.7 KMQ•CONSTRUCT• TRANSPORTGEOIN LEOTO

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FigureB-34

Antenna

Final
B-ll3

Installation-Photovoltaic

Satellite

The GEO base has overallof

operation.

The upper

dockingfacilities

of

the

facility

cranes

used

for

deck

dimensions

supports

thetransportation

supports
in berthing

four

the

modules

GEO base primarysolartransferredtosubsystems

includeand flightcontrols

keepingsystems,which

use LO2/LH 2 propulsion.

Other

major

elementsofmaintenancesystems,

and crewsatellite
similartoareneeded on thenotmicrowave

LEO base construction

antennations.

Both

areas

and forLEO base.

klystronrequiretube

special

base

into

a single

crew

based

at

GEO facility.

theelectricaloperations

power,

fordocking

cranes,which

systems

centeraroundannealing

opera-

array
for

is

accessto maintenance

65 and only

a minimum

operations.

so,

crew

consequently,

used atthe

LEO modulesareas follows:
three

of

functionsmaintenancecontrol,

modulessize

allTransportableenvironmental

crew
life

modules(I)

crew

moduletothepeoplearehousedinsince

a solarflare

gm/cm 2,theshelterwill

add an additional

115,000

Withintheshelterwill

be provisions

up tooperatethecomplete

of
only

and

(2)

add an eighth

shelter.

Assuminga shielding

B-II4

status.

five

designThe major

module,base on standby

structure,in

one offor

also

similar
base.

decks

includesareaccommodations,

incorporation

personnel

modulescrew

are

LEO construction

can be incorporated

support,ofradiationand stationequipment

and provisions

The construction

theKw solarConstruction

and solar

quarters

place

(260

equipment,

replacement

The crewininconstruction

maintenancesystems.

thecontrol

The major

and information

toof

exceptequipment

module.

power

includeDocking

location.

electricalattitudesurfacebase when notlongitudinal

modules.

The GEO base has a construction
construction

theequipment

repair/refurbishmentofattached

consistingandThe lowermachines.

totwo decks

modules

deployment

anotherarray)

theand maintenanceand payloads.

array

are also

systems

1400m x 1600m withcrew

systemstheuse or when the GEO base isofan operationdeck

65 rather

than

deck

serves

Kg to
days

the

crew

modifications

which

requirement

Subsystems

the

basic

of

as

20 to 25

moduleand controls
used within

I00

mass.

to
the

modulesare the sameas for the LEObase modulesdescribed previously.
Figure B-35 showsa summaryof crew requirements for the LEOand
GEOconstruction base operations.
Component Packaqing for Launch - Component packaging
a very significant factor in construction

for launch isas well as space transportation.

Packages must not only meet dimensional and weight constraints of the launchvehicle, but must have appropriate mass density to far cost effective
portation.

Figure B-36 illustrates the dimensions,of various SPS components.

trans-

density, and part countAs indicated, densities vary from a low of 12 kg/m 3

for antenna subarray elements to about 2500 kg/m 3 for power conductor.
obtain desired densities, components must be packaged in appropriate
as indicatedin figure B-37.

Such packaging minimizes

thereby reducing transportation
Crew Considerations
operations.

the number of launches,- Figure B-28 illustrates the integrated

timeline shown in figure B-29.

During construction,aorbit.

The reference crew scheduling concept is summarizedo

90 day staytimeo

6 days on/l day off per weekolO hours work shift per dayo2 shifts per day (2 crews)o0.75 operator productivity

Radiation protection

spacephase requires 580 days as indicatedcrewman will be located in geosynchronous

exposure.

mixes

costs.

The satellite construction

on the construction

To

below.

factoris provided to limit crewmen to 35 REM/year

Shielding of 2 to 3 gm/cm2 in LEO and 20 to 25 gm/cm 2 in GEO will

provide this protection.

Habitat walls will provide the 2 to 3 gm/cm 2 shielding,however, special "storm shelter" facilities must be provided to obtain 20 to 25
gm/cm 2 shielding.

iThis type of facility would be needed during solar flares only.

B-II5

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