D 180-20689-1

Contract NAS9-I S 196 DRL NumberT-1346 DRD Number MA-664T Line Item 3

Solar PoYMr Satellite SYSTEM DEFINITION STUDY

Part I Volume I

Executive Summary

Jur.e 28, 1977

Submitted to The National Aeronautics and Space Administration

Lyndon B. Johnson Space Center In Partial Fulfillment of the Requirements

of Contract NAS 9-15196

Boeing Aerospace Company Missiles And Space Group Space Division P.O. Box 3999 Seattle, Washington

Approved By:

G.R. Woodcock Study Manager

FOREWORD

The SPS syskms definition study was initiatt.-d in De..-ember 1976. Part I was '-"001plded on May I.

1977. Part I includc:d a principal analysis effort to "'valuate SPS energy "'-onv,:rsion options and spa"'-e

l"OllSln.action locations. A transportalion add-on task provided for further anc.'vsis of transportation

options. operaf ions. and coses.

The study was mana,ed by the Lyndon B. Johnson Space Center (JSC) of the National Aeronautk'S

and Space Administr.ation (NASA). The Contr.acting Offit.-er·s Representative KOR) was Oarice

Covington of JSC. JSC study management l'-'am members included:

Lou Livingston System Engineering Dick Kennedy Power Distribution

and Analysis Bob Ried Strudure and Thennal

Lyle Jenkins Space Construclion Analysis

Jim Jones Design F r"-d Stebbins Structural Analysis

Sam Nassiff Construction Base Bob Bond Man-Machine lnterfa"'-e

Buddy Heineman Mass Properties Bob GundefS'-'n Man-Machine Interface

Dickey Arndt Mkrowave System Analysis Hu Davis T rJnsportation Systems

R.H. Diet1 Microwave Tr.ansmitter Harold Benson Cost Analysis

and Rectenna Stu Nachtwey Mkrowave Biologkal

Lou Lc.-opold Microwave GcnerJtors Effeds

Jack Seyl Pha~ Control Andrei Konradi Space Radiation

Bill Dusenbury f nergy Conve~ion Environment

Jim Cioni PhotO\·oltaic Systems Al\·a Hardy Radiation Shidding

Bill Simon Thermal Cyde Systems Don Kessler Collision Probability

The Boeing study manag"'r was Gordon Woodcock. Boeing technical leaders were-:

Vince Caluori Photovoltaic SPS's

Dan Gregory Thermal Engine SPS's

Eldon Davis Con~tn11.:tion and Orbit-to-

Orbit Transportation

Hal DiRamio Earth-to-Orbit

Transportation

Or. JO\: Gauger co~t

Boll Conrad Mas~ Propertii:s

Rod Darrow Operation:!>

Bill Emsley Flight Control

ii

Jack Gewin Powi:r Distribution

llon Grim [k..:tric Propulsion

Henry Hillbrath Propulsion

Dr. Ted Kramer Thennal Analysis and

Optics

Keith Miller Human Factors and

Construction Operations

Jai:k Olson Confi!?uration Design

Dr. lknry Oman Photovoltaics

John Perry Structures

ORIGINAL PAGE Ill O! POOR QUALITY

The Part I Report includes a total of five volumes:

Vol. I

Vol. II

Vol.Ill

Vol. IV

Vol. V

DI 80-:!0689-1 D 130-:!0689-:! Dl80-:!0689-3

D 180-:?0689-4

DI 80-:?0689-S

Ext.-cu1ivc Summary

System Requirements ind Energy Conversion Options

Constrodion. Tm1Sportation. and Cost Analyses

SPS Tr.msportation Syskm Requirements

SPS Trctnsportation: Repn.-scntative System Descriptions

Requ"-sts for informat•on should be d~ded to Gordon R. Woodcock of the Boeing Aerospace

Company in ~attle or Clarke Covington of the future Progr.uns Division of the Johnson Spat.-c

C enter in Houston.

iii

D 180-20689-1

SOLAR POWER SATE LUTE SYSTEM DEFINITION STIJDY PART I TECHNICAL REPORT

1.0 EXECUTIVE SUMMARY

I .I THE SPS CONCEPT

Soliir power sateUites represent a proposed means of tapping baseload electric utility power from

the sun on a large scale. Th~ advantages of the space environment for generation of electric power

can thereby be ••brouaht down to Earth ... These advantages include essentially l"Ontinuous sunlight.

the ability to l"Onstruct very large solar l"Ollectors with minimum resource investment. and the abil­

ity to always aim the collectors at the sun. Studies presently in progress sponsored by ERDA and

NASA. of which this SPS Systems Definition Study is one element. are defining the systems. devd­

opment approal-hes, and risks and costs for this ,·enture and evaluating the probable benefits of the

system against these risks and \."OSts.

SYSTEM CONCEP·;

An SPS system for utility electric power would include a number of satellites in geosynchronous

orbit. ead1 with one or two associated power receiving stations on the ground. Receiving stations

can be located near load centers (weather is not a significant factor>: each will provide IOOO mega·

watts or more of basdoad electrical output. A satellite system is pictorialized in Figure 1-1. Power

is tr.msferred from the satellites to the ground stations by high-precision electromagnetic beams.

The tran-.missions would presumably use the industrial microwave band at :!.45 GHz: an alternative

industrial allo1.:ation available at 5.8 GHz l·ould be used but has received comparatively litrle

attention.

A complete SPS system is depicted in Figure 1-2. In addition to the satellites and their ground sys­

tems it will include:

• A space transportation system capable of delivery of the SPS's to geosynchronous orbit and

capahlc of supporting all required space operations needed to establish and maintain the SPS

system.

• One or more constrm:tion bases. located either in geosynduonous orbit or low Earth orbit.

capable of constructing the satellites. Satellite hardware delivered to the construction hases

will be prefabricated to the extent practicable.

• Mainknance and ... cn·ice bascll capable of supporting the maintenan1.:e operationi. required to

kl.'ql th1.· SPS's operating.

ORIGINAL PAGE JS OF POOR QUAI.ll'»

Power Satellite

D 180..10689-1

El..ECTIUCAL OUTPUT: 10.000,000 kW {2.5 TIMES GRAND COULEE DAMJ

figure t-t. The SPS Concept

PH1'.J'tOVOt TAIC SP$

Figure l ·2. Major Program Elements

TH£flMAL ENGH\1£ 1WS

ORiGL~AL lS Of<' Pl·~ Ht LJt /\ l rrv

D 180-20689-1

• One or more Earth-based space transportation ports (launch sites) capable of supporting space

transportation operations.

• One or more spa'--e-based transportation operations support bases, capable of supporting spal·e

transportation operations. This function could conceivably be combined with that of either a

'-"Onstruction base or a maintenance base.

• Earth-based manufacturing facihties capable of producing the hardware and consumables nec­

essary to transport. l"Onstruct and maintain the SPS system.

• An Earth-based logisti<:~ system capable of delivering the hardware ancJ consumabks to the

space transportation parts.

1.2 STUDY OBJECTIVES

The objectives of Part I of the study. as specified by the NASA statement of work. were as follows:

(I) Issues-To derive spedfic. comprehensive supporting data necessary for NASA evaluation of

the following two major SPS system issues:

a. What is the overall most effective means of accomplishing solar energy-to-electrical energy

conversion on an SPS in geosynchronous orbit?

b. At what location (or locations) in space l·ould the various phases of SPS construction and

assembly be done?

( 21 Transportation-To increase the scope and depth of understanding of the space transportation

systems necessary to support an SPS program.

a. Provide a set of transportation system requirements and reference transportation system

elements descriptions appropriate to the conduct of an SPS program as represented by

JSC Scenario "B".

o. Identify and define analyses and tests nel·essary ro advance the confidl'nce levd in pro­

jected SPS transportation systems performance and cost sufficient to recommend initia­

tion of an SPS technology advancement program.

3

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1.3 STIJDY APPROACH AND PLAN

The SPS System Definitions study is being conducted in tw'> parts as indkated in Figure 1-3. It

began with two reference "Ystems: the photovoltak syskm from what is popularly known as the

JSC tireen book CJSC 11568,: and the Brayton thermal system from prior study work by Boeing.

We have also in this phase considered the other options shown: high tedrnology gallium arsenide.

thin films. and Rankine and thermionic thermal options. Part I provided evaluation data that would

allow selection of one or two energy conversion options and evaluate LEO and GEO space construc­

tion locations. In Part II we will analy 1.c the mh:rowavc power transmission system and develop an

integrated system definition. The objective of Part II is to reduce the ~ystem mass and 1.·ost uncer­

tainties as much as possible.

The program under study is basically an operational or commcrdal SPS program. Ground rules arc

summarized in Table 1-1. It starts with the first full capabilit/ I 0.000 mcgawan satdlite and g~s

through a program of many satellites. In a few instances where the cost of money was important wt.•

have used a 7!1~~ discount rate as recommended by Econ in earlkr studies. We are assuming a 30

year system life for purposes of financial horizon analysis. even though there is no particular lifo

limit on the SPS's. There are a few cases where creep rupture analysis was required. A safety factor

of I Yi on 30 year creep-rupture-Hf e material thickness was used. The reft>rcm."e data for transporta­

tion came from the Heavy-Lift Launch Vehicle study and Future Space Transportation Systems

Analysis study data base (contrads NAS 8-3~169 and NAS 9-143~3. res~ctivcly). That data base

was updated as part of this study.

Table 1-1. SPS Program Ground Rules

• IOC date: end of 1995

• 11 ~ SPS program

• I 0-gigawatt SPS size

• 7.5' ~ discount rak ( 1977 constant dollars)

• 30-year satellite life

• Transportation system reference data to be tahn from HLLV

and FSTSA data base and JSC-11568

• Operational system design and performan~:e projcdions bast.'d

on 1987 te1.·hnology baseline

The program s1.·enario U'il'd for SPS in-.1.1lldtion rak was JS( s(enario B. shown in Figure 1-4. Th1.•

fir-.t .. akllite was 1.·ompkted in 1995 and produl."lion rate increases gradually to 7 pcr ycar at tht.' l'lld

of the program. For purro~1.·s of predktmg tran~portation l'O'>t. wt: have takt.•n J ~nap,hot of the

mid-point of the program at the rail' of four SPS\ per year.

4

Point of Departure

Reference Systems :

Silicon photov91taic from JSC study

Brayton thermal from Boeing study

Options

Gallium arsenide and thin-film photovoltaics

Rankine and thermionic thermal engines

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Part I December 1976 to May 1977

&esig : y reference

designs to next level of detail.

Analyze: Construction and transportation systems and operations. CharactPrize: Options and delta off reference designs Compare: Performance, practicality. operations, environt:lental factors, cost, and technical risk.

Select: Q.;0r two energy conversion options and low versus geosynchronous orbit construction.

Part II May 1977to December 1977

Analyze: Microwave power transmission syst'!m.

Design: Integrated SPS systems-energy conversion, power transmission, construction, transportation, and overall operations.

Develop: . • Integrated system

conceptual definition. • System mass estimates • System cost estimates • System de.velopment plan5 • Technology advancement

requirements

Figure 1-3. Solar Power SateOite Systems Study Overview

NUMBER OF SYSTEMS COMPLETED

CUMULATIVE TOTAL

NUMBER OF SYSTEMS COMPLETED

CUMULATIVE TOTAL

• Corresponds to JSC scenario B

0 .. 2

1

1

4

3~

0 1 1

0 2 3

4 5 5

34 39 44

1

4

5

g 0 N

IO .. 0 N

1

5

6

49 55

1 2 2

6 8 10

6 6 6

61 67 73

Figure 1-4. Baseline Operational Program Scenario

5

8 N

2 2 3 3 3

12 14 17 20 23

6 6 6 7 7

79 85 91 98 105

0 -~ 3

26

7

112

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During Part I no power transmission system analysis was done. The data shown in Table I-~ came

from the JSC .. green book." This information was used for purposes of sizing satellites to dl·liwry

I0,000 megawatts of ground output ro rhe receiver antenna system.

Table 1-2. Reference MPTS/MCRS Nominal Efficiencies*

Antenna power distribution .98

DC-RF conversion .87

Phase control ** Waveguides cJ.:!R) .98

Mel·hankal alignment .98

Atmosphere .98

Energy collel·tion .88

RF-DC l'Onversion .90

Power interface .99

Overall .6.:!9

* From JSC 11568. fig. IV-A-I-I. .. SPS Effo:iendes"

** Included in energy coJlection

The Part I analysis effort was grouped into four higher-level task types as shown in Figure 1-5.

These analyses were based on the point-of-departure configuratiom. Results of the analyses were

integrated as indh:ated in the figure. Additional tradeoffs and analyses dfort after the midterm led

to the evaluation data summarized in the summary sc:ction of this report.

The silii:on single crystal photovoltaic SPS system and the Brayton thermal engine system referenn~

designs wert" used to 1.:arry out analyses such as performance. structures. and powt•r distribution.

The constn11..:tion and transportation analysc:s consider~d both geosynchronous orbit and low earth

orbit as construction locations. The analyses resulted in the configuration ~volutions shown in Fig­

un.• l-6. These analyses led to concepts for fai.:ilitization of construction. influencing the satellite

design, e~pecially in the thermal engines where a configuration change from compound curvature

concentrators occurred. In both cases modular SPS designs at approximately 1.000 megawatts of

onhoard husbar power per module were developed.

The thermioni.: syskm was re.:ommended for 1.li~rnntinuanl'l' about halfway through the Part I

efforr.

1.4 DOCUMENT STRUCTURE

Tim, l'Xt'l·utiVl' 'unHnJry dc~nihl'~ thl· re~ults of the Part I '-·valuations of enl·rg~ l'onwrsio:i and l'On­

strw.:tion l<Kation options. Voh1ml'S If through V prm idl' dda1lcd rl'porting of the Part I efforr.

6

ORIENTATION

CONFIGURATION ANALYSES

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PART 1 MIDTERM

PART1 FINAL

EVALUATION DATA PRODUCTS

ISSUEN0.1 ENERGY CONVERSION METHOD

•PERFORM· ANCE

•PERFORM· AND DEGRADA· TION

ISSUE NO 2 CONSTRUC· TION LOCATION

• O\I qALL DESIGN REOUIRE· MEI-HS

IOPTION&ORIENTEDI ITRAOESORIENTEDI •SIZE •MASS

• PERf-ORM· ANCE• DCGRADA· TION

SILICON PHOTOVOLTAIC GEO CONSTRUCTION

BRAYTON THERMAL ENGINE-LEO CONSTRUCTION

4MODULES

ORIGINAL PAGD:. IS OF POOR QUALtrf

•TRANSPOR· TATION

•CONSTRUC· TION

•MAINTAIN· ABILITY

• ENVIRONMEl'fT •MATERIALS

•SYSTEM COMPLEXITY

• TECHNOt.OGY ADVANCE· MENT REOUIRE· MENTS

eCOST DIFFEREN­TIAL FACTORS

e TR,NSPORTA TIOl'f ROTS& COMPLEX ITV

• CONSTRUCTION • LAUNCH SITE

DIFFFHENTIALS e SYSTEM

STARTUP •OPERATIONS •COLLISION

•COST DIFFERENTIAL FACTORS

Figure 1-S. Synopsis of Part 1 Study Logic

STRUCTURES, CELL CHARACTERISTICS, PERFORMANCE, Iii POWER DISTRIBUTION

GEO COl'l:STRUCTION 6 TRANSPORTATION

LEO CONSTRUCTION 6 lRANSPORTATION

PERFORMANCE, SUB~YSTEMS, STRUCTURES, f'OWER DISTRIBUTION

THERMIONICS DROPPED­EXCESSIVEL V MASSIVE

SILICON

a •CR•1 • • • CR•2

~~· • MAINTENANCE ~ IGEOt OPTIONS

/7) • GALLIUM ARSENIDE

~ ILEOt ntlN FILM

18MODULES

BRAYTON

• METALLIC TURBINE

e SIC TURB1NE

RANKINE

e POTASSIUM • STEAM

Figure 1-6. Part 1 Analyses and Configuration Evolution 7

ORIGINAL PAGB JS or POOR QUALl'IT .

D 180-20689-1

l.S ENERGY CONVERSION EVALUATION

The statement of work spedfied the evaluation factors listed in Table 1-3 for energy l·onwrsion.

The cost differential factors are presented at the end of this summary because they combine re!-iults

of b1..1lh evaluation efforts.

Table t-3. Energy Conversion Evaluation Factors

Evaluation Factor

a) SPS Performance

h) Performance Degradation

c) SPS Size

di SPS Mass

e) System Complexity

fl Maintainability

g) C'onstmction Requirements

h) Transportation Requirements

i l Technology Advancement Requirements

j I System Cost Differential Factors

k) Environmental Effects Differential Fal·tors

ll Materials Differential Factors

SPS Performance-The first factor is performance c efficiency). When this study began it wa'.'>

bdieved that thae was quite a diffrrence hetwt•en thennal t:"ngine and photovoltaic performanl"t.'.

There is mm:h less than we had thought. Effidency generally follows the technology advances and

advanced ~ystcms t1..nd to he more efficient. except for the thermionics option. Efficiency is not an

important discriminator unles., very low. Effickncy re!-iults are summarited in Figure 1-7.

Perfonnance Degradation-Degradation effects were found in ali of the systems. The silicon photo­

voltaic degrades more than thermal engine systems. with gallium arsenide in between.

On the left hand -.ide of Figure 1-8 " the magnitude of the degradation effect and on the right lwnd

it 1s normalized to indicalt' what pen:entage of SPS mass h affoded. for example. thin film rdlel··

tors are the Jegrad<ition mode for them1<1I engines hut rt.>pre ... ent only a small fradion of satd ... e

nwso;. In .tll l·a-.l''> maintaining th~ ~atdlite output .,cems lo he promi-.ing. We comp1..·nsatcd for tkgra·

dation in our rc1.:ommendcd SPS\. Pcrformanct• degradation is rdleded in size. mass. and l"<ht and

therefor~ l·arries litth.· weight a~ :111 independent evaluation fal..'tor.

SPS Size-Figure l-9 ;..; ;1 -.i1c ..:omr:~o1.,on of the prin..:ipal sysktn.... The small~st sy.,tem io; the

gallium ar.,t.'nidt.' amwalahk follow~d by thl.! Brayton. Tht• -;ili..:on ... y~tcms show that t·on..:entration

8

ID .

1&

10

&

0

1

.15

.25

0 0

--

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EFFICIENCY FOLl.OWS TECHNOLOGY ADVANCE EXCEPT FOR THERMIONICS

-

..---

--

~

-

-

-

SILICON REF CR2

SILICON ARRAY ADDITION cu

SILICON ANNEAL· ING

GALLIUM ARSENIDE ANNEAL· ING

THIN FILM

BRAYTON BRAYTON RANKINE ADVANCED THERM· 1422K 1660K 16SOK BRAYTON IONIC

CR1 .EOL CR1

Figure I· 7. Energy Convenion Comparison SPS Performance

RECOMMENDED SPS OPl IONS COMPENSATE DEGRADATION BY ANNEALING, MAINTCNAloCE. OR INITIAL OVlRSIZE

DEGRADATION EFFECTS PREDICTED MAGNITUDE

10 20 TIMt. ON ORBIT, YEARS

GALLIUM ARSENIDE CELLS

SILICON CELLS

30

.15

0

NORMALIZ.OD DEGRADATION

~ ~ -THERMAL

I '-----------J-GA A~ . SILICON

0 10 20 Tlft4E ON ORlllT, ~EARS

Figure 1·8. Ene19y Convenion Comparilon Performance Depadation

9

ORIGINAL PAGE IS OF POOR QUALITY

D 180-20689-1

ralio I kads to a sipificanl redudion in size. The tJ1in fdm size esrimate is a guess bc~au~ of the

unc.-ertainty in the dala base. Size does nor sttm &o lk:- a slron1 diSl."rimir.alor: m~h more imporlanl

is mass.

SPS Mass-Mass c."Omp .. ;ison fort~ systems is shown ;n a igure 1-IO. Note rhal annealin(! is a mu~h

~ner syslem from the mas." s1andpoinl lhan arr.ay addilion. The annealing melhoJ employ"

eleclron-beam or laser healing. A t'1ennal pulse di~-tty inlo lhe solar ~di r.ai~ the tempcr.ature

momen!aril)· .and anneals out rhe deg.Old;ation. Heat j,!cner.a1ed in rhe \."'ell diff US&."S only sli"1tly in10

the subslrale. Rough es1im1t1es indicate that 'lie need aboo• a half dozen annealing ma~hines. 1wo

meters squaR by thr« ureters long. This nun:~r of u1ad1ines oper.uing c."Ontinually durin~ satdlite

operation will keep lhe performanc.-e iAp. These madlines \."OUld be oper.ated remotely by oper.tlors

via RF links. Annealing is like painting the Golden Gale Bridge: as~ as it is finished. ii ·n~•SI

Sl3rt dpin.

The lighlest system of all is plli•1m arsenide. We have found lhat there is a c."Onsidcr.tble va.riation in

Brayton system mass as a funclion of ledmology. The steam Rankine sys1em was ex"·essively mas­

sive and c."Ould nol be ploned on the dlan. Thennionics • .:onversion was also quile massive.

SySlem Compleury-There are two ways to measure "-omple"ily: one way is 10 eslimale lhc: num­

ber of uniqu~ parls or subas..;emblies. The thermal engine system h..s abou1 five ti:nes as many

u,ique paras as lhe (lho1ovol1ak. Total parts is lh(' olher measure. If one 1.."0unts indi,idual solar

cells. photornltaks have about 1.000 limes as many 1otal pans. lnlegration \."Ompkxit)· of 1he sys­

!c:m is delc:nnined primarily by the number of unique parls. Syslem romplexitites are \."Omp:1red in

Figure 1-11.

Main1ainabili1y Faclors-For both types of ene~ l·onversion we found mainlenance problems. and

in borh ba~s \. ~ found )()h.aions. The results after applying the solurions are summarized in Table

1-1. Roughly 5 tt• I 0 manhours per hour for annealing are needed wilh lhe pholo\'Ollaic syslem and

~ligiltl} r111m· than 10 mh;h with lhe thermal syslem for mel·hanical repair and repla•:e. h i:> l·on­

l·ei,ahle 1hat rho~ manhours might be spenl on lhe ground if '.\'e can dewlop suitable automakJ

'~\fems lhJI 1.-an he man-Jirech:u from a remole dislanl·'-'· It does not nec.·essarily mean thal lhe

Sfi~ "s haw to ~ manned. It j, likdy that these maintenance requirements will be overshadowed hy

th;·I for lh.: miaowa\e transmirter.

10

118

125

I I •

•

I-

-,_

-

GAU.W ....... ·~ C.""'

0

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llml 111ml

Fagure 1-9. Eaeqy Coavenion Comparisoa SPS Size

-,..._ - -,..._

- -..-

-

-

llLICON SILllCOll llUCION GM.Lii# ANNEAL· MllNIOE

mm ftUI

llllAY"ON lllAY10N AOVMCID flOTA'S- TttlflMIONIC lllf Allflill\' tGl!X t&A IRA\'TON SIUM lllCllADINGI AOOITICllll .... ANllUlllWG llWdUNI Cll2 all an Cllt

Cllt

figure 1-IO. Energy Conversion Comparison SPS Mass

JI

• Ill ::; • ~ Ill

! ! a: ~ Ill

8 z :>

1

1

1

D 180-20689-1

I THERMAL ENGINE kAS ABOUT 6 T•ES AS MANY I _ UNIOUE~ARTSASPHOTOVOLTAIC _

® THERMAL ENGINE

10

TOTAL PARTS

PHOTOVOLTAIC

1

Figure 1-11. Energy Corwenion Compariwn System Complexity

12

D 180.20689-1

Table 1-4. Energy Conversion Comparison Maintainability Factors

Photovoltaic

Problem

Sensitivity of Solar Cell

Strings to Cell Failures

Solution

Parc1lleling and Diode

Shunting

Result

Equivalent maintenance

load roughly 5 to 10 MH/H

for annealing or arrc1y

addition

Thermal Engim.·

Problem

Reliability of NaK System

Solution

Technology Advancement Progrc1m

or redesign to eliminate NaK

~ Equivalent maintenance

lo.id roughly 10 Mff/H

for med1anical repair anJ replace

Construction Requirements-A constructability rating is shown in Figure 1-1 '.! that involves a num­

ber of factors as devdoi<d by the construction analysis task. The LEO/GEO comparison is shown

as w .:II as the thermal engine and photovoltaic comparison. This is a weighted Sl."Ore l'.'Omparison.

The numbers in parenthesis are weighting factors; a long bar is good. A long bar means a smaller

facdity. less complexity. and a smaller crew size. It is a "goodness" rating and not a measure of the

physil:al size or numbers of people. Tht>re is not a dramatic difference between th..- sysrems bur

some preferenl'.'e for the simpler ph'ltovoltaic, as expected because the system is less med1anically

complex.

One of the reasons that the conl'.'entrc1tion ratio = I photovoltail'.' satdlite is preferred is rhat it is

simpler to construct. It avoids having to install rhe large tlat V-ridge reflectors: that is a signitkanr

advantage.

Transportation Requirements-In transpoHarion requircmenrs. there was nor a great deal of difkr­

ence in the mass between the systems but the photovoltaic systems packaged to roughly 20 times

the density of the thermal engine systems. Some diffaully was experienl'.'ed in packaging the latter.

We finally gor down to a Jen~ity compatible with the laund1 vchide capability: further improve­

mentc; appear possible. Most of the demcnts of the photovoltaic system can fold into a dense pack­

al,!c. For the thermal system unless plumbing is produced in spal'.'e or prefilled. ther..- an: limits on

aaainablc packaging density.

Values ~hown m Figure 1-13 are an average but we ad1ieveJ that average in a(tual pa1.·kaging for

ea\:'h type. The size ~hown in the figure represents a volunw large enough to \:'Ontain the entire SPS

as pa\:'kagcJ for laundi.

13

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CJ> HIGH SCORE IS BEST

- 131

128 112

,.. 171

e NUMBERS INC t DENOTE WEIGHTING FACTOR

- NUMBER OF MACHINES 111

-MINOR FACILITY SIZE C1J

- MAatlNE COMPLEXITY C2t

-CREW SIZE 131

(j 180 97

- FACILITY COMPLEXITY 141 .. IC

I ..

- MAJOR FACILITY SIZE 141

lilO

-ASSEMBLY COMPLEXITY 151

LEO GEO

l'HOTOVOt. TAIC SATELLITE CCA•2t

LEO GEO

THERMAL ENGINE SATELLITE

LEO GEO

PHOTOVOLTAIC SATELLITE CCA•1t

figure 1-12. Preliminary Relative ConsbUctability Rating

PHOTOVOLTAIC

52000m3 PERSPS DENSITY• 1300 kg/m2

a1 lblftl

t3Smj ,. -1 .... ~

53m

. .l ----

MOST IMPORTANT DIFFERENCE IS PACKAGING DENSITY

L--98m -

THERMAL ENGINE

CBRAYTONJ

1.1x1o6m3 PERSPS DENSITY • 72 kglm3

•4.S lblftl

Figure 1-13. Energy Conversion Comparison T ransportalion Requirements

14

147m

D 180-20689-1

Technology Advancement Requirements-Technology advancement requirements listed in Table 1-5

are the most important ones. The Brayton lhennal qde requires the leasl tt>chnology advance. The

silicon photovoltaic system is the next least. Continuou:. photovoltaic "-ell and blankd production is

mu'h more impo.tant than obtaining maximum possible efficiency. 14';f solar cells made by con·

tinuous production proc..""esses would make the silicon !.YSl ·m attractiw. 18',; cdls made by today\

processes would not yield an economically attractive system.

The other systems require more tedmology advance: Brayton and silicon are the least risky.

Environmental Effects Differential factors-Env•ronmenlal effects di:ferences are summarized as

follows:

• No significant environmental effects associated with energy conversion were found.

• The principal factor is launch vehicle emissions. which are proportional to SPS mass.

• A launch pad fire with gallium arsenide ta~nic• appears kss of a toxicity problem than the

hydrochloric acid effluent from shuttle SRB's.

Arsenic was specifically investigated. Launch accident cloud analyses indicate that concentratiO!b

get below allowable levels quickly. Further. this is not a routine condition. but an exceptional

condition.

Materials Differential Factors-Materials factors are displayed on a very compressed logarithmic

scale in Figure 1-14. We have picked five materials that are used in sufficient quantity in SPS sys­

tems to present polential concerns. The figure shows how many SPS's can be built per year with

today's production rates and finally. how many SPS\ total could be built with the total known

reserves. Reserves are quantities available by today's recowry process at today·s costs. Silicon is

produced in large quaa.tities for mdallurgical reasons. The reserves arc on the order of half the cmst

of the earth. so there is no supply problem.

Presently the U.S. production of columbium is essentially zero. But the world production is suffi­

cient to build several SPS's per year. The reSc:rves in the United States are not large, but world

rc:serves are adequale. Aluminum is no problem. Gallium was the only material indicating a polen­

tial problem. The assumptions are very important for gallium. We show ~.000 tons of gallium per

SPS and no process improvement. There are known potentials for process improvements up to

about a factor of 4. Gallium today is produced as a byproduct of aluminum production. The yield

is about one-fourth of what it coul<l be. Akoa. for example. has stated that if more gallium is

needed from ahaninum. it could be obtained with more inveo;tment in reco\'ery l!quipm~nt. Gallium

production rate may he more of a problem than total reserv~s.

IS

1o8

1o5

i 1o4

1000 ... z ... -' 100 c > 5 10 0 ...

1

.1

.01

.001

D 180-20689-1

Table 1 ·S. Eneqy Convenion: Technology AdYancemeat Requirements

SIUCON

CONTINUOUS CELLIBLANKET MODUCTION fROCESS

-ANNEALING

PHOlOVOLTAIC THERMAL CYCLE

GALLIUM THIN-FILM BRAYTON RANKINE ARSENIDE

THIN-FILM THIN..flLM RELIABLE HIGHTEllP· GALLIUM ARSENIDE TECHNOLOGY FLUID ERATUrtE APPLICATION - CONTAINMENT 11£1' 1.L VAPOR PROCESS PRODl.'CT ION TECHNOLOGY

- PROCESSES -CONTINUOUS RtLIABLE CELL/BLANKET FLUID PROOUCllON CONTAINMENT PROCUS

-ANNEALING

CONCLUSION BRAYTON AND SILla>N LEAST RISK

GALLIUM WAS ONL V IDENTIFIED MATERIALS AVAILABILITY PROBLEM. VALUES SHOWN ASSUME 2000 TONS GALLIUM PER SPS AND r~ PRODUCTION PROCESS IMPROVEMENT

COLUMBl!JM SILICON (BRAYTON) ALUMINUM GALLIUM

SPSIYR SPS TOTAL SPSIVR SPS TOTAL SPS/VR SPS TOTAL SPSIVR1

SPS TOT Al

1D ,w WORLD!;

U.S. Cl. ~ Ow Uo t-ic Oc Z..1 en> Wcz: >"' ffi> :s a:m

WORLD I

WORLD

lu.s.

I WORLD

U.S.

WORLD

U.S. PRCD. IS -o

.s. ES. AVE

.. 0

U.S. PROO. NOT AVAIL.

WORLD

•GRAPHITE FIBER PRODUCERS ARE TOOLING UP FOR THOUSANDS OF TONS PER VEAR

I I I I I I WORLD I I

THERMIONICS

THERMIONIC DIODE TECHNOLOGY

GRAPHITP

SPSIVR SPS TOT AL

ID w 15 I a.

:& 18w , ... ~ I ~:5

WORLDI ~ ~ PAOO >w NOT I ffi >

U.S. AVAIL. ~ !; I a: ID

I

Figure 1-14. Energy Conversion Comparison Materials Differential Facton

16

OJUGINAL PAGE 18 or POOR QUALITI

D 180-20689--1

Graphite today (graphite fibers of an Aerospac."\! quality. not total gr.tphite ). is a small production

item. But the producers are presently tooling up for very high production ratc:s: graphite tih<r is

becoming a commercial product.

Energy Conversion Evaluation Summary-Four ene~y conversion options were found that make

SPS look promising. Any one of them would work: they are not remarkably different in owrall

potential. Thennal engines are more complicated but require less technology ad\'ancc. The rt:oto­

voltaic."S require continuous production cell process development more than they need anything else.

As will be shown in the cost summary. there is not much difftarence in production cost projections

for the Silicon and Br.tyton systems Gallium arsenide looks slightly cheaper. but there is a huge

uncertainty in the data.

We propose to concentrate on the silicon concentr.ation ratio = I with annealing capability and

Brayton sysh~ms. We propose gallium-arsenide a~ an advanced technology option. showing one way

that the SPS system could ~row to achieve potentially lower costs with technology improvement.

There are certain things in the potassium Rankine system that need further evaluat:on. especially

machinery mass properties.

A problem was experienced with thin films in that the data base was just not sufficient to drdw

definitive conclusions. Finally. we recommend rejecting two systems, specifically thermionics and

skam Rankine. The recommendations stated apply to the Part II study and not necessarily to an

SPS progr.am. They relate to the fact that the study objectives are. in part. to minimize uncertainties

in mass and cost.

Construction Location Evaluation-The statement of work specified the evaluation factors listed in

Table 1-6 for construdion location. The construction locations to be evaluated are low Earth orbit

CLEO) and geosynchronous orbit (GEO). A significant proportion of the evaluation relates to selec­

tion of orbit-to-orbit transportation. because with LEO construction the satdlite (or satellite mod­

ules) must be moved with low thrust to prevent structural failure. The power output from the mod­

ules can be used to opernte an electric propulsion system for the transfer. With GEO construction

the satellite hardware is not cc1pable of self-powered transfer: conventional means must be used. For

the purposes of this study reusable L02/LH 2 orbit transfer vehicles were baselined. In summar)·.

LEO constmction impli.:s self-powered electric rocket transfer: GEO construction impli.:s chemicJl

vehicle transfor of HLLV-sized payloads to the GEO construction bas~.

17

D 180-20689-1

Table 1-6. Construction Location Evaluation Factors

l:.valuation Factor

a) Transportation Requirements

b) Construction Requirements

c) SPS Overall Design Req.iirements

d) SPS Performan'-~ and Degradation Potential

e) Launch Site Difforential Effects

0 System Startup Requirements

g > Operations Considerations

h > ( ollision Considerations

i) System Cost Differential factors

H Clrbital Transfer (omplexity Factors

Transportation Requirements-The principal difference in tr~ nsportation rel1uirements bdween LEO

and GEO construction location is the difference in total delivery to low Earth orbit. shown in Fig­

ure 1-15 in tenns of numbers of HLLV launches. The difference results primarily from the

differences in propellant ret1uircd for the transfer due to the great diffe:-ence in propulsion spedfo.:

impulse. typkally 5000 sec for ekctric rockets versus 470 seconds for LO-,/LH-, chemical rockds.

Construction Requirements-(onstruction n!t1uirements in low Earth orbit involve sever.ti nuisance

factors. As noted in Figure 1-16 atmosphere drag results in an aver.tge propellant consumption (to

keep the construction facility orbit trimmed) of about 800 kilograms per day. The construction

approaches that we have developed do not appear sensitive to light/dark cycling on crew producli\;­

ity. For gravity gradient effects. the only practical thing to do is to select a stable attitude and con­

struct in that attitude. Ont' simply car, .10t afford to expend enough propellant to hold a non-stable

attitude.

Thermal t:fti!cts may have some influence on construction. but with low,o\."ffo.:ient-of~xpansion

graphite epoxy. Hhe baseline structure l that does not appear to be a strong consideration. There an:

similar thermal dfrcts in geosynduonous orbit. although less frequent.

In geosynchronous orbit. radiation environment may be an issue if massiw shielding is required.

With the Apollo/Skylab cn:w radiation exposure standards. construction bases can provide most of

the shielding required without mass penalties. Thl' solar tlare contribution can b\.' avoided with a

"storm cellar.'" If it is ncccssar} to go to a lower level radiation standard. and add shielding. there is

not mud1 differen~e in the amount of shielding rl.'quircJ. but there is quite a Jiffercn(I.' in thi: trans­

portation cost I a one-time CO!>t for each facility 1.

18

32

28

24

20

LAUNCHES PER DAV 16

12

8

4

2

D 180-20689· 1

CHEM OTV r-\_ GEO CONST '\.

t AVG

4 6

SATELLITES PER VEAR

.. MAX

8

• JSC SCENARIO B •BALL/BALL Hl..LV

P/L•4od< KG

Figure 1-1 S. Construction Location Evaluation

SPS810

• DRAG-AVERAGE PROPELLANT CONSUMPTION (Oz - Hz> 700 TO 800 Kg /DAY TO KEEP ORBIT TRIMMED

• LIGHT/DARK - NOT A MAJOR PROBLEM WITH MECHANIZED CONSTRUCTION

•GRAVITY GRADIENT-CONST!1UCT IN STABLE ATTITUDE

~

•RADIATION ENVIRONMENT IS ISSUE IF MASSIVE SHIELDING REQUIRED

• CONSTRUCTION BASE LIKELY TO PROVIDE &TO 15glcm2

APOLLO/ SKYLAB GEO

_....-(SHADED REGION IS SOLAR PROTON CONTRIBUTION)

INDUSTRIAL -~

1--~~~.._~~~--~.a.-~

' 10 100 SHIELDING, g/an2 AL

Figure I· 16. Construction Location Evaluation Construction Requirements

19 ()IUGINAL PAGE 1S or rooR QUIJj'rl

D 180-20689-1

SPS Design Requirements-A number of design requirements on the satdlite were found for LEO

construction. First is modularization, Cwhich may be a blessing in disguise). The satdlitc must be

modular, because otherwise it is not controllable during the orbit tr.msfer. The photovoltaic system

must have the capability to operate the proportion of the array used for electric propulsion at

reJuceJ voltage: and must incorporate additional power distribution provisions. Further requin:d i~

the capability to accept propuls;on installations and then provide for their removal at GEO. All of

these were reflected as satellite impact costs as part of the LEO/GEO construction cost differential

factors. In most cases. the costs were trivial. but thesl' additional design re{1uirements represent

additional <lesig.l c:>mplexities.

SPS Performance and Degradation Potential-An SPS module being transported from LEO to GlO

must pass through the Van Allen trapped radiation belts. If the transfer is accomplished by a high­

thrust system in a total time of roughly six hours the radiation dose received is minimal. even with

the scantie~t of shielding. Electric propelled transfers (low thrust). however. arl! likl!ly to re{1uirl! 2

months to a year. and substantial doses will be receivl!d. (The electric-powered trip time can be var­

ied over a wide range by selection of power level and specitk impulse. Radiation dl!gradation is a

significant factor in trip time selection.)

Significant radiation dl!gradation phenomena were identified for solar cells and for plastic film

retlectors. Representative solar cdl degradation data are shown in Figurl! 1-17 and plastic film

reflector estimates are shown in Figure 1-18. Additional potential degradation concerns include

plastic matrix compositl! structural materials: no data werl! found on radiation degradation cf these

materials.

As a design practice. the idl!ntifil!d modl!s of degradation were compensakd by oversizing or. in the

case of solar cells. in certain cases by annealing. Consequently. radiation degradation effrcts were

reflected in cost tradl!s for LEO wrsus GEO ..:onstruction.

Launch Site Differential Effects-The principal effect on launch site operations was due to the

aforementioned difference in launch rates. This diffrrence was also retleded in costs of providing

facilities capable.: of supporting the requisite launch rates. with estimated values of 10.6 billion for

LEO construction and 15.8 billion for GEO constmction. These facilitks would be incrementally

procured over a period of years as the launch rate increased along with the rate of ~PS capacity

addition.

System Startup Requirements-Certain startup factors Wl!re identified that are unique to the LLO

con~truction option:

20

~ t Ill _, ... Ill s s s a • _,

i ... 0

i u s r

D 180-20689-1

0.9

38Y£ARS

0.7

PORTION OF SILICON

INITIAL POWER

0.6 AVAILABLE

0.6

OA

0.3

0 60 100 150 200 2SO 300 350 400 .e&O

THICKNESS OF CELL COVER, pm

Figure 1-17. Photovoltaic Array Radiation Degradation Characteristics

100

eo

ID

40 • • •

20

e THUS All FACETS Will 3E INSTALLED AT LEO e GEO-ASSEMBLED SPS IS LESS THAN 1% LIGHTER

TRANSFER EFFECTS IALL FACETS OUTI --._____

DEGRADATION CURVE FOR AlUMINIZED KAPTON CFROM PROJECT ABLEI

...E:!!... • 1.0505 69.3'.

CONCEN I RA TOR MASS IS 8.706% OF SPS

l5Jl5%) X c• .7011%1 • 0.44% of SPS

45DAY1•8 YRI 900AY 1•18 YRI

180 DAY 1=28 YRI

1 1800AY TRANSFER +38YEARS

72.l'Jli

19.3"

0---~~~---~~-'------------~.....1~--~---~-----------~--------...._--..___._ 0.01 0.1 1.0 10 100

YEARS EQUIVALENT GEOSYNCHRONOUS ORBIT EXPOSURE

Figure 1-18. f.elf Power Transfer Has Utde Oversizing Impact

:!I ()lUGINAL PA:; or pOOR QU

D 180-20689-1

(I) The SPS moduks must have chemical propulsion attitude control capability with suftident

control authority to establish a sun-pointing orientation to activate th•: power generation sys­

tem. Th!s capability is al~o required to successfully ..:xecute the orbit transfer, in view of fre­

quent passage through the Earth's shadow.

(2) Orbit transl er simulations by the FSTSA study indicated that a typical 180-day transfer would

experience 800 to 1000 occultations by the Earth's shadow. In GEO service the satellite will

experience about 80 occultations per year.

(3) Satellih: modules arriving at GEO must be joined together. \I hcreas '.f constructed at GEO the

satellite may be of a rnon1Jlithic dc!>ign. or :f modular can be construded with moduks joilll·d.

One concept for module joining is illustrated in Figure 1-19. This concept b db~.-u~sl.'d in morl.'

detail in the body of the report.

Operations Considerations-No strong disaiminators in or::rations were found. LEO '-·onstn11.:tion

has more distinct kinds of operations. notably chemical and electric orbit transfer operations (chem­

ical for crew rotation and resupply at GEO) and SPS module assembly operations at GEO. The

number of orbit transfor vehicles in flight at one time all requring control is ab0ut double for LEO

construction. 13 :! vs 16 at an SPS addition rate of 4 per year) but the number of HLL V operations

is less, as discussed above.

Collision Considerations-LEO construction operations result in an increased risk of collisions. The

situation for the photovoltaic SPS is summarw:d in Figure 1-:!0. The collision analysis is described

in ti.e body of the report.

System Cost Differential Factors-The~e dre discussed in conjunction with energy comersion co~t

differentials. LEO construction consistently shows Iowa overall cost as a result of rt!duced cost for

Earth-to-orbit transportation.

Orbit Transfer Complexity-The sdf-powen:d operations associated with LEO construchon arc

more compkx as r~gards propulsion systems. flight control. guidance and navigation. and software.

Orbit transfer systems arc discussed in some dl.'pth in the body of the report.

Figure I-~ I summaritcs the construction loi.:ation evaluation. A hullet shows the preferred option

for each evaluation factor. Th·.:re an~ b bullets for GEO and 3 for LEO which indicates GEO con­

struction. However. the co~.t of LEO construi.:tion has consistently been found to be cheaper than

GlO. Either option is workable. LEO is cheaper, but more i.:emplex. In a '-·ommercial environment

i.:osts will eventually dn\'c the decision. Much mon: interesting is the question. not where satellite

number I 0 or 50 will bl.' built but where will the first one he built. The condusions of <.-. .. r anJlysi!>

may not appl} to thi~ ':ucstion. Thi.' rci.:ommcndation!> in the Figure arc aimed .1t rt!Jucing Sl'lhiti\·

ity of the continuing analy~i~ to thl.' LEO v~ c;Eu is.,ue.

22

0180-20{>89- l

SPHJI NEW MODULE ARRIVAL

-

e CLOSE TO WITHIN 1000.2000 M WITH PROPULSION

INITIAL DOCKING Iii ALIGNMENT

MOVEABLE~ DOCKING \ CRANES

No. t

/;-

l>OCK1NG MODULES

OOCKlhG I RECEPTACLES ' ...

e MODULES HAVE DIFFERENT e 111nTUDE eALTITUOE

e MA5S"'8Mk~

FINAL POSIT!( ~ING

CONTROlS ATTITUDE ANO CLJSING SPEED

MODULE No.2

r;JLLSINTO POSITION

Figure 1-19. Satellite Module Docking at GEO

... 711

;; Ill ~ :> 0

~i II!~ ...

" .. zU o< 1:: og !! 0 ~ ... 8i

2ID

15

10

6

REGIONIKMI

TIMEIN REGION IDAYS

CONSTRUCTIQN

500

ORIGINAL PAGB I! OF POOR QUALl'l'Y

500

TO

1000

1C.I

·:·:·::·:·:: - GROWTH ::·:·:·:·:·:· OBJECT

......

1000 TO

2000

22.2

MODEL (YEAR 20001

-CJRRENT OBJECT MODEL IYEAR 1975)

2:llOO TO

3000

11.0

3000

TO 5000

22.3

~

5000 TO

10000

29.9

figure t-20. Number of Collisions

• 3X 3M OBJECTS • NOOBJECTClEAN\JP e NO AVOIDANCE DURING TRANSFER e 100%ARRAY DEPLOYED AT GEO e ~OF ARRAY DEPLOYED

DURING CONST. & TRANSFER

0 ......

10000

TO

2!000

34.2

0

20000 TO

35788

39.8

GEO

oPS

10,950

D 180-20689-1

For a dewelopment program. these kinds of questions arc important:

• When must manned GEO operations be-gin'!

• What are relative transportation costs with dewlopmental Earth launch system (q; .. shuttk­

derived HLLV)'!

• Can the: dcwlopmental SPS confaguntion (such as 1-GW module) be tr.ansfem:d to GEO with

transmitter inSlalkd or stowed'!

• Can the developmental SPS configuration survive the transfer radiation environment?

• What is the program funding vs time compari- n~

Program funding vcrst1S time is always an important factor in deftlopment decisions. For LEO con­

struction. ekctric propulsion must be developed. while for GEO. more in,-esunent may be needed in

~ ~tion~ and in earlier manned GEO operations.

Cost Diffeteatial Factors-This discussion begins with energy payback considerations. Solar cells arc

very energy intensive. Pr.:senkd in figure 1-22 are energy costs in kilowatt hours per kilogr.am of

"'"dk. The e.•crgy payback for solar cells as a function of this energy cost is also shown on two

5alcs These seal~ show SPS and ground applications. Pricing the energy at 40 mills per kilowatt

hour. the actual cost of 4he energy is shown on the outside scale.

The main reason toda) 's \.'l:lls are so intensiw ts that yields are very poor. Most of the silicon. in

which a great deal of energy is invested. ends up as waste (saw filings and grindings•. Continuous

proccSSt"s can probably reach a yield ran~ of 60'K to 800( makino= the payback very attradive.

Eni.:rg)· c~t is a basic factor in the cost of solar cells. like materials cost in building hardware. If the

energy cost is below l ~/watt one might be reasonably confident that "'-ells in the 1~/watt r.ange,

made by a continuous production process. would be possible.

Why is energy cost important? The reason is that energy payback tim~ is economically significant.

Typical economics equations uSc:d. for example. by Caputo and Truscello in the JPL report. predict

the cost of energy from an energy system. One can dose the loop in these equations b) setti:'lg the

capital cost of the eneri,'}· in\·esteJ in the system equal to the capital cost of cnern that the system

produces. In other words one does not borrow from a cht>ap energy system to create an expcnsi-..-e

energy system. Wirn typical economic factors. ene~y payback less than about 8 years is essential.

Othen1;ise. enc:rgy cost will spiral upward ever more rapidly. as inJicated in Figure 1-~3. With con·

tinuous pro<ludion proceSSt":. for solar cells cor with thennal engines•. SPS payback (for the total

s)·)tem • is I 1; to 3 yeah where this curve is not very stei.:p. With its short lyba1:k. SPS should have:

economic ad"anta~es when the: system tedmology is m.iture.

24

MEHRREO CONITRUC-

EVALUATIOll FACIOR TION LOCATION

LEO GEO

el 'l'RAlllPORTATION • RIGUIRPllMS .. a.nucTIOll lllO

B£QUIREllEll1'S sn.ECTIOll

cl IPIOVIAAU. • -llEGUIRBIBlfS

• IPIPEllFORMAllC£ MD • 0£GRADATION l'OTEllT1AL

tlUl.:Hlfll DlfffRUmAl EFRCIS •

• SYSTEMST...-... • REOUIRDEIJ1'S

fl Ol'ERATIC*S • COllSIDERATICMG

NCOLLISIOll COllSIDERATIOICS •

I SYSTEM COST • DIFFERENTIAL FA.."JOAS

J OR&ITAL TRMSFER CDm'UJUTY FACTORS •

D 180-20689-1

D&:ISlCJllORIVER

UO REQUIRES LESS FLIGHlS AlllO LESS CNt-ORaT PROPELLANT TRANSFER

•uo DRAG 6 l>Altlt P£RIOOS VS. GEO RAOIATIOel 6 OISTAM:f

UORlOUIRU~lATIOll

6 OTHER SPECIAUZATION

DEGRADATION DUE 10 VAiii ALLEN RADIATIOllCM BE ~TED

ffWER lAUlllCHES fOR LEO

LEO STARnJPMORf COMPLEX

UO HAS llORE DISTINCT KllmS OF Of'£RA TIONS

A80UT 15 COLUSIOlllSISPS FOR LEO VS 1 FOR GEO

UO AllOUT 2" CHEAPER OV£RAU. TRANSPORTATION

ELECTRIC PROl'UlSIOll llEOt llORE COMl'l.EX

HIGHLIGHTS

• EllNER LEO OR GEO CONSTRUCTION IS A VIABt.E OPflON

• LEO IS LOW£R COST BUT-E a...EX FLIGHT CONTROL; PAOl'UUION

e •A COl&WRCIAL ElllVIAOMIENT COST Will DRIVE THE OECISIOll

e llllTIAl DtCISIONS Dfl'ENDEllT Olll DEVEll)PMENT PflOGRAll

• lmCUAR ANSl'ftR-eont llDDES PROBA8l Y HAVE THEIR PlACE

• SELECI' llOOl!I~ SPS FOR a>NSTRUCTIOfl BASE ANALYSIS

F1g111e 1-21. Consaruction Locaaioa Evaluation Summary

' D

a: z ~

i Ir 0 a

~ c ?: K > I 6

I • > z f • 4

2

• Sll.ANE PROCESS

• ENERGY OOST l.04IKWlf

• SPACE CELLS• 111' BLANKET EFFICIENCY 6 GS WTS EFFICIENCY. 30 K1ftfMl-OAY

"' • GROUND CELLS• 141' BLANKET ~ ~

EFFICIENCY 6 9°" POW£R .. u COLLECTION EFFICIENCY. 5 l(WHIM2.oAy .. • 0 0 • " c %

?: 2 :c K " u i ~ > c > .. 0 POTPmAL c .. FOR z t CIONTllllUOUS

' ..

PROCESS

TYPICAL TODAY

• 0 •o .2 A

YIELD KG SOLAR CELLS PEA KG SEMICONDUCTt'A GRADE SILICON

Fiple 1-22. Energy Costs and Payback for Silicon Solar CeDs

25

ORIGJNAL PAGI • or POOR QUALITY

1.8

DI 80-20689-1

Cost differential factors reflect cost impacts of all the evaluation factors. Thi! data shown in Figure

1-24 are for a 112-SPS program. for our preferred SPS designs. The number I unit SPS will be a sig­

nificantly higher cost than the program average. Also. note that there are some things t'tat were not

'-'"OSted in Part I. Our ROM estimate of the range for the uncosted items is shown.

We did the costing in two ways. first was in a mature industry fashion. a projection of things in a

commercial environment with high quantities of mass production. The second method was aero­

spa~ cost prediction techniques, using our parametric cost model. For satellite production. the

mature industry projection and the higher aerospa"-e prediction are shown. Transportation costs

used only the aerospace methods. The results were mature industry system costs trending LI less

than S2.000 a kilowatt for the I 12-SPS. No significant differences were sc:en between the silkon

photovoltaic and Brayton thermal engine. Typkal LEO versus GEO differences are also shown.

Shown in Figure 1-25 are two alternate systems costs to provide an idea of cost r-.mges. The silicon

array addition system is more massive than the annealable system and suffers a significant satellik

imract if constructed in LEO. The gallium arsenide system costs are a rough-order-of-magnitude

projection because of the relatively great technology extrapolation. Because it is low in mass . .,·ery

low potential costs are projected for the future. We did not know how t" .:stimate an uncertainty.

Also in the gallium a~nide case. the U:.0-(;l:.0 diUerence 1s less. because the satdlite is low in mass

and the transportation cost contribution is less significant.

Transportation Evaluation Summary--The results of the transportation add-on t?'ik are summariz.:d

in Figure 1-26.

Two Earth launch options w.:re analyzed: ( 2) A ballist,.:. two-stage sea reco.,·ery \'dtide with a

retractable payload shroud that was 1 O<Y1c recoverabk. ( ~) A two-stage wing-wing whide that was

also )()()';( recoverable. No significant diffaences Wt!r.: found in cost pt!r flight or pl!rformance. For

the ballistic system tt.~ main technical concern is sea recowl). It appears feasible. but there is not

much data base. For the! winged system. the lowest achievable payload density is considerably

higher. There are conc.:ms about launch and recovery siting because the booster is a down range!

lander and a suitable place to launch must haw a down range ,·ecovery site. The wing-wing vc!hicle

also has a somewhat higher DDT &E cost.

Orbit transfer options included a space-based :md a ground-based OTV. and self-power. Self-power

le~ns transportation costs about 25'1. The space based OTV showed 15' (better performance than

the ground based OTV. Thi: space-based orbit transkr vehicle requires on-orbit propdlant transfer

but based on work Joni." by General Dynanucs. it appears possible to transfer the propellant without

rotating the staging ha~. It may be sufficient maely to rotak th.: propellant by using electric

pumps to withdraw the propdlant and inject it into the OTV tanks in such a \\ay that a rotation is

set up within the tanks.

26

, ..

•

••

SAME CELLS USED IN SPACE

2

TYPICAL RANGE FOR TOTALSPS

•

D 180-20689-1

6 8

ENERGY PAYBACK TIME. YR

• CAPITAL COST IN ADDITION TO ENERGY • S1CIOOIKWe

• CAPITAL COST OF ENERGY 911V£STED • CAPITAL a>ST OF ENERGY PAOOUCED

• SOLAACELLOATA FOR CURRENT CELL PAOOUCTIO!ll natNOLOGY

18

TYPICAL SILICON SOI.AR CELLS USED ON GROUND

12 14

F"....,e 1-23. Energy Payback Tune Is Signif"acant!

MATURE SYSTEM COSTS TRENDING TO LESS THAN $2000/KWe

SILICON CR1/ANNEALABLE

LEO

r--, I I I I I I I I

GEO ..--, I I : I I I I I

l'RAVTON 1650K

LEO

r-, I I I I I I I a I I

GEO

r--1 AEROSPACE 1 : COSTING : 1 DELTAON I I SATELLITE I I PRODUCTION

SATELLITE PRODUCTION (MATURE INDUSTRY)

CRE\Y ROTATION & RESUPPLY

FREIGHT TRANSPORT

NOT COSTED IN PART I

• CONSTRUCTION BASE AMORTIZATION

• CREW SALARIES

•OPERATIONS& SUPPORT COSTS OTHER THA'4 TRANSPORTATION

• GROUND RECEIVING STATION CUSED JSC VALUE FOR MPrSI

ROM RANGE IS $300-$500 KWe

Figure 1-24. Cost Differential Facton Recommended Reference Systems (112 SPS Program)

27 ORIGINAL PAGE IS or POOR QUAUIY

I !

i 100D

SATELLITE IMPACT

D 180-20689-1

SILICON ARRAY ADDITION

LEO

r-, I I I I I I

GEO

r - 1 AEROSPACE I !COSTING I 1DELTACN

1SATELLITE 1 PRODUCTION

GALLIUM ARSENIDE :R•1 ANNEALABLE

LEO

1

GEO

I 1 I I I I I

I

SATELLITE PRODUCTION CROM PROJECTIOl'.ll

NOTCOSTEDINPARTI

• CONSTRUCTION BASE AMORTIZATION

• CREW SALARIES •OPERATIONS& SUPl'ORT

COSTS OTHER THAN TRANS?>ORT A TION

• GROUND RECEIVING ST A TION IUSED JSC VALUE FOR MPTSI

CREW ROTATION & RESUPPLY

FREIGHT TRANSPORT

Figure l-2S. Cost Differential Factors (112 SPS Program)

FREIGHT LAUNCH OPTIONS ORBIT TRANSFER OPTIONS

' ! I

.L

• PERFORMANCE AND CDSTP£R FLIGHT ABOUT EOUAL. <S'IO/LB

e BALLISTIC CONCERN: SEA RECOVERY"

e WING/WING CONCERNS: • PAYLOAD BAY DENSITY • LAUNCH & REC.OVEllY SITING • HIGHER DDT&E COST

CHEMICAL

SPACE GROUND BASED BASED

ELECTRIC

SELF POWER

e SELF.f'OWER REDUCES NET TRANSPORTATION COST ABOUT 25"'°

e SPACE-8ASED CHEMICAL OTV ABOUT 15'4 BEnER PERFORMANCE

e SELF.VOWER CONCLRNS

• RADIATION DEGRADATION • COMPLEXITY •COLLISION

• SPACE·BASED OTV cor~CERN. PROPELLANT TRANSFER

Figure 1-26. Transportation Evaluation Summary

28