NASA Technical Memorandum 103211

• ....... Lunar Orbiting Microwave

..;.__:_ Beam Powe_ SyStem _

Edgar H. Fay

Sverdrup Technology, Inc.

Lewis Research Center Group

Brook Park, Ohio

and

Ronald C,_Cu!I

National Aeronautics and SpaceAdministration -_ _:_:_- _- _--_-_:--_ _-............ _

Lewis Research Center .......

Cleveland, Ohio

Prepared for the

' 25th Intersociety Energy Conversion Engineering Conference

=_, .... ._ cosponsored by the AIChE, ANS, SAE, ACS, AIAA, ASME, and IEEE -

Reno,- Nevada, Augtist_12_-17, 1990

(NASA-T,M-I03211} LUNAR ORBITING _ICROWAVE NgO-25IT3

I:}EAM POWER SYSIEM (NASA} '9 p CSCL 22B

Unc I a s

G311B 02'992033

LUNAR ORBITING MICROWAVE BEAM POWER SYSTEM

Edgar H. Fay

Sverdrup Technology, Inc.

Lewis Research Center Group

Brook Park, Ohio 44142

Ronald C. Cull

NASA Lewis Research Center

Cleveland, Ohio 44135

ABSTRACT

A microwave beam power system using lunar orbiting solar

powered satellite(s) and surface rectenna(s) was

investigated as a possible energy source tor the moon's

surface. The concept has the potential of reduc_ system

mass by placing the power source in orbit. This can greatly

reduce and/or eliminate the 14 day energy storage

requirement of a lunar surface solar system. Also

propellants required to de-orbit to the surface are greatly

reduced.

To determine the practicality of the concept and the most

important factors, a "zero4h order" feasibility analysis was

performed. Three different operational scenarios

employing state of the art technology and forecasts for two

different sets of advanced technologies were investigated.

To reduce the complexity of the problem, satellite(s) were

assumed in circular equatorial orbits around the moon,

supplying continuous power to a single equatorial base

through a fixed horizontal rectenna on the surface.

State of the art technology yielded specific masses greater

than 2500 kg/kw, well above projections for surface

systems. Using advanced technologies the specific

masses are on the order of 100 kg/kw which is within the

range of projections for surface nuclear (20 kg/kw) and

solar systems (500 kg/kw). Further studies examining

optimization of the scenarios, other technologies such is

lasers transmitters and nuclear sources, and operational

issues such as logistics, maintenance and support are

being carried out to support the Space Exploration

Initiative (SEI) to the moon and Mars.

INTRODUCTION

Operations on the surface of the moon will depend on a

reliable electrical energy source. Providing low cost

electrical power on the surface presents a significant

challenge. Energy storage requirements for the 14 day

eclipse period make surface solar power systems heavy.

Proposed nuclear power sources have masses highly

dependent on power level and may have political and

safety concerns. An alternative is to place the the power

source in orbit and beam the energy to the surface. The

concept has the potential of lowered system mass by

greatly reducing and/or eliminating energy storage and

also reducing propellents required to de-orbit to the

surface.

Specific masses for surface power plants are projected to

range from a low of approximately 20 kg/kw for high

powered nuclear up to approximately 500 kg/kw for

continuous solar power [1]. The objective of this study was

11odetermine, with reasonable technology projections, if a

beam power system was competitive with surface power

sources.

Multiple approaches including systems using nuclear

sources and/or laser beaming can be conceived end will be

addressed in future work. For this study, a simple system

where energy is beamed at microwave frequencies from a

solar powered satellite(s) to a fixed non-tracking rectenna

(an antenna that receives and converts the RF energy to

useful dc electrical energy) is investigated. To further

reduce the complexity, a single equatorial base and

satellite(s) in circular equatorial orbitwere assumed.

It is recognized that more optimum designs are possible

and will be explored in future work. Likewise, operational

issues such as logistics, installation, maintenance, etc.

must be addressed. (See Future Work and Conclusions

Section). This first assessment with limited scenarios and

technology is thus presented as an initial scoping

document.

APPROACH

Three different scenarios of beaming power from orbit to

the lunar surface were investigated. Figure I shows the

block diagrams for each scenario.

The first scenario, Case 1, consists of 8 single satellite in

orbit that transmits power to the surface when in the field of

view of the receiver (rectenna). Energy storage is needed

on the satellite as well as the ground. Batteries on the

satellite provide power for transmission when the satellite is

in the eclipse while within the beaming field of view (see

figure 2) The energy storage on the surface provides

power to the user while the satellite is out of view.

Case 2 is a constellation of satellites providing continuous

coverage of the reclenna (see figure 2). Energy storage

on the surface is therefore not necessary. Batteries are

provided on each satellite is the energy source for

transmission to the surface when that satellite is in the

eclipse.

Case 3 is also I constellation of satellites providing

continuous coverage, but with the energy storage located

only on the surface to provide power to the user when the

satellites are in the eclipse.
Figure1.SystemBlockDiagrams

Figure2.Lunar Orbiting Satellites

but uses SOA components available in the 1990's (see

figure 3). The structure uses Space Station Freedom

(SSF) trusses. (An "average" specific mass is shown in

Table 1, but structure mass is not a linear function of area).

The solar array employs single crystal silicon cells, and

nickel hydrogen batteries are chosen for energy storage

(SSF designs). The nickel hydrogen batteries are sized

assuming a depth of discharge (DOD) of 50%. Power

Management and Distribution (PMAD) uses components

developed for SSF. The thermal management radiators'

specific masses assume a rejection temperature of 25° C.

Six percent of the masses for solar array drives, batteries,

transmitter, and PMAD is added to account for integration

masses.

Table 1 indicates the specific masses and officiencies

assumptions for the technologies used in this study,

Three levels of technologies were considered:

1. State of Art (SOA) : This design is similar to the Solar

Power Satellite (SPS) studied considerably in the lg7o's,

The aluminum slotted waveguide anlenna array uses

parameters from the SPS [2]. No space qualified

microwave tubes of the power level required were known,

therefore parameters for 35 GHz terrestrial gyrotron were

assumed for the transmitter. (As for all technology levels,

the transmitter efficiencies were assumed constant for all

frequencies. Efficiency is expected to decrease with

higher frequency, and this effect will need to be factored__in

later analyses, but it is expected thai the efficiency at

higher frequency will be more a function of time and

technology evolution than basic physics.)

2. Advanced: Advanced technology represents the

evolution of state of the art technologies to levels believed

1o be possible by the year 2000. Advanced lightweight

solar cells and batteries are used for the power source.

Satellite batteries in this case are sodium sulfur operating at

2

ORIGINAL PAGE IS

OF POOR QUALITY
a temperature of 350 ° C which leads to a correspondingly

lower specific mass for the radiators. The sodium sulfur

batteries are sized assuming 50% DOD. PMAD

components are improved and their operating temperature

is also raised, lowering the thermal specific mass.

Regenerative Fuel Cells (RFC's) were chosen for surface

energy storage since the charge and discharge cycles are

of relatively long durations. The RFC's are sized assuming

an 80% DOD. The transmitter is an improved efficiency

tube.

F_ure 3 shows schematically a beam power satellite using

advanced technology. Four of these Case 2 satellites

beaming power at a frequency of 100 GHz from an altitude

of 2459 km provide I Mw of continuous power to a user(s)

on the surface of the moon. tl is of the same general size

as SSF. The SSF transverse boom length is 145 m and its

solar array "tip-to-tip" length is 72 m. The beam power

utellite's overall dimensions are 148 x 58 m. As will be

discussed later, lower power levels or altitudes would

reduce each satellite's size.

E

ID

Figure 3. Beam Power Satellite Concept (sized using

advanced technology)

:',I!I

II I

11]

3. Th_ Film: Thin film solar cells and sor¢l state

MMIC (Microwave Monolithic Integrated Circuit) devices are

presently available for terrestrial applications and

projections were made if space versions of these products

could be developed around 2000. This technology

assumes an inflatable structure to support an integrated

solar array and antenna, see figure 4 [4]. Amorphous

silicon solar cells and RFC's are chosen for power

components. The transmitters are solid state MMIC

devices in a phased array. This technology level is used

only for Case 3 since no thin film storage technology was

known to exist for the satellites,

Figure 4. Thin Film Technology Beam Power Satellite

Concept

n

Incid,nt

Ught

Beam •

Table 1. Technology Performance Parameters

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Reels.n= (k=,_)

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1

3

- Advsnced ThiR Film

% .or. % ._,. %

14 $ _S ,? S

1 0

3 0

21 10

6 3

70 15

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14 9s 7 is 1 |S

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$ 40 $ eO .1 so

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1 IS 1 iS 1 |S

A spreadsheet was developed to perform energy balances

and calculate the beam power system mass for each

scenario and technology level. Power components are

sized as a function of user power. End-to-end system

masses are calculated including all support such as thereat

management and structural subsystems. By using three

different technology levels under three scenarios, the

sensitivity of system specific mass to various subsystem

and component performances was obtained. Also, the

ORIGINAL PAGE IS

OF POOR QLIALITY
effect of technology on achieving a specific

mass competitive with surface options was determined.

The modular nature of the spreadsheet can accommodate

different power sources (such as a nuclear reactor) and

transmission types (such as laser) for future studies.

Orbit parameters calculated for the three cases are

presented in Table 2. The beam times, battery discharge

and recharge times, and the number of satellites required

for continuous coverage assume a 135° rectenna field of

view.

ASSUMPTIONS

Transmission frequencies of 2.45, 35, 100, and 300 GHz

were investigated in this study. The receivers were taken

to be placed flat horizontally on the surface with an

assumed field of view of 135 ° .

A near field approximation equation was used to size the

antenna and rectenna:

Pr/Pp, 1-exp{-AtAr/(LR)2} Equation I.

Where:

R- Separation Distance (m)

L- Wavelength (m)

At,=Antenna Area (m2)

Ar=Rectenna Area (m2)

Pt=Transmitted Power (w)

Pr,,Received Power (w)

When the satellite is directly overhead (i.e.. normal to the

rectenna), a 15% space transmission loss was assumed in

the calculations (i.e. Pr/Pt is set equal to 0.85). The overall

transmitter to rectenna link efficiency will be reduced from

the 85% space transmission efficiency due to off-normal

pointing losses which are a complex function of many

parameters such as antenna/rectenna design, orbit

parameters, antenna pointing, etc. A simple numerical

integration for an ideal tracking antenna at variable distance

yielded a 54% link efficiency, which was used for all cases.

(See Discussion of Results and Future Work sections for

further discussion of this assumption.)

The propulsion requirements to land a surface power

system from a low lunar orbit will require propellent at least

equal to the mass of the power system, effectively

doubling the entire system mass. The orbiting portions of

the beam power system, of course, requires no de-orbiting

capability. To provide a common scale for comparisons

with surface-only systems, masses presented are surface

equivalent masses (is orbiting masses are added as half of

what they would weigh on the surface).

For altitudes tess than 1336 kin, the eclipse time is greater

than the beam time (see Fig. 2). Thus, for periods of the

lunar night, the satellite will be entirely in eclipse while

within the rec_enna field of view. For Case 1 and 2, this

means that the satellite batteries' must be sized to provide

the entire energy for transmission since the solar arrays will

be unable to produce power. The satellite batteries

discharge time thus equals beam time. For altitudes at or

above 1336 kin, the solar array can supply some power

directly for transmission even in the middle of the lunar

night. The discharge time thus transitions from beam time

toeclipse time above 1336 kin. Recharge time in both

cases is when the satellite is in the sun but not transmitting.

The solar array is thus sized to charge the batteries during

this recharge time.

For Case 1, the surface batteries provide power when the

satellite is out of the rectenna's view. The surface batteries

are recharged during the beam time and discharged during

the remainder of the orbit. For Case 3, the surface storage

is charged when the satellite is in the sun and transmitting

power. During periods of the lunar night when the satellite

is eclipsed, the storage provides power for the surface. In

all cases, the energy storage requirements are much less

than a surface solar power system.

DISCUSSION OF RESULTS

Table 3 shows the Cases 1,2 and 3 specific masses for a 1

Mw system at 2.45, 35, 100 and 300 GHz, using state of

the art, advanced and thin film technologies. All specific

masses are shown at their optimum altitude, but no system

optimization was performed except trades of antenna and

rectenna sizes. The specific masses range from 2920

kg/kw using state of the art technology at 2.45 GHz to 84

kg/kw using thin film technology at 300 GHz. For any of the

three cases using state of the art technologies, and Case 1

advanced, the resulting specific masses are greater than

the surface solar option. Therefore, they will not be

discussed further.

Table 2. Orbit Parameters

all times In hour=;

A/tttude Orbtt E¢llpee No. Beam

(kin) I_dod Time _'e Time

9O 1.95 ,78 30 .07

1443 2.04 .76 20 .10

211 2.14 ,75 15 .14

374 2,42 .74 10 .24

533 2.70 ,75 | .34

9OO 3.38 .77 6 .S6

1336 4.31 .82 S ,86

24 50 6.78 .92 4 1.7

105G_ 34 1.53 3 11,3

Cue1&2

Dt=eMq;e I_¢haq_

Time

.07 1.17

.10 1

,14 1.3g

.24 1.68

.34 1.95

_56 2.61

.82 3.45

•g2 5.08

I.s3 727.

Case 3

_u rfa,_4 Rtnrnnm

Dlsc:l_rge lUdmharue

Tim=

403

251 420

235 436

_5

186 485

153 518

128 543

91 580

30 641

ORIGINAL PAGE IS

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Frequency

(GHz)

2.45

35

100

30O

Table 3. Specific Mass Results

Case 1 Case 2 Case 3

(k_/kw) (k_/kw_ (kg/k wl

SOA Adv SOA Adv SO& Adv ThinFilm

2920 887 2474 901 8863 1286 367

2482 811 1448 387 3043 415 170

2460 589 1328 327 2606 295 107

2452 582 1269 282 2447 200 84

For Case 3, the efficiencles assumed for the thin film

devices represent todays terrestrial versions (and should

go higher with time) but specific masses for the

COml_l'i_nfs quoted and structures designed with them

have a large error band since there are still many

unknowns. There is a higher confidence value placed on

the performance of the advanced technologies since it is

an extension of large space structure design such as SSF

end a more established space qualified technology

database exists.

One of the first parameters of interest is transmission

frequency. Figure 5 shows the antenna/rectenna areas,

calculated using equation 1., versus altitude for 2.45, 35,

100 and 300 GHz. Altitude and frequency have a large

effect on antenna/rectenna areas and tend to drive the

satellites to a lower orbi_to minimize the antenna]rectenna

size and mass.

Figure 5. Antenna/Rectenna Area vs Altilude

2o0o(

The system mass is composed of power dependent

components (which scale lineady with power level), and the

antanna/reclenna which are not a function of power, but of

altitude and frequency. The antenna]rectenna will

contribute most of the mass to very small power systems.

As power level is increased, the antenna/rectenna sizes

remain constant and further mass increases come only

from the power system. The total system specific mass

thus will decrease with increasing power.

One Mw continuous supplied power at 100 GHz

transmission frequency will be used in further discussions

of Case 2 Advanced, Case 3 Advanced e.nd Case 3 Thin

Film. The specific mass results can be scaled (to a first

order) to other power levels and frequencies by

proportional increases (decrease) of the plotted

antenna/rectenna specific mass for decreased (increased)

power and/or frequency.

,-I /

/i

- I'--y . i--"--

_ 70( • C.,me3 ThinFilm

1

= |

01- . I • I • • I • •

o soo _ooc 1_oc =ooo _t_oo .. _oo 3o( _

This effect of frequency can be seen in figure 6 which 10(

shows the specific mass of the beam power system versus

frequency using advanced and thin film technologies. The 0

beam power system is sized to deliver 1 Mw of continuous 0 s0 100 ls0 2oo

power at the surface. The points plotted here relate to the Fmquwcy(OHm)

aJtitude yielding the lowest system mass. At 2.45 GHz the

antenna and rectenna masses are so large that they

dominate the total system mass. Overall though, the

transmission system does not contribute significantly to

the system mass at greater than 35 GHz and little is gained

at frequencies above 100 GHz for the cases investigated.

The small dependence on transmission system above 100

GHz reduces the importance of the 54% efficiency

assumption used for the transmission link. If the actual link

efficiency is less, the antenna/rectenna sizes can be

increased without major effect on total mass (assuming

larger sizes can be built).

Figure 6. Specific Mass Advanced & Thin Film Technology

250 300

Figure 7 shows the specific mass of the beam power

system under Case 2 scenario using advanced

technology. The PMAD and transmitter on each satellite

are sized to provide the total continuous power to the user.

As seen previously from Table 2, the beam time increases

with altitude which leads to a lower number of satellites

required. Thus, as the altitude increases, the reduction in

satellites, in turn, decreases the total satellite PMAD and

transmitter mass.

5 ORIGINAL PAGE IS

OF POOR QUALITY
The energy storage mass increases due to an increase in

discharge time. Notice tha_ the storage mass is increasing

up to 1336 kin, decreases up to 2459 kin, then continues

to increase as the altitude increases. At 1336 kin, the

discharge time changes from the beam time to the eclipse

time which causes the storage mass trend to become

discontinuous and decrease due to the array being able to

provide some power during transmission. Above 2459 ks,

the storage mass continues to increase from an increase in

eclipse time. (If the total discharge time from the

constellation is plotted versus altitude, the trends will be

the same as the storage mass shown in figure 7).

The arrays and storage are not a function of the number of

satellites but rather the total energy needed from the entire

constellation. On the other hand, the antenna needs to be

replicated on each satellite. Even so, the antenna (and

rectenna) are not significant mass drivers until altitudes

above 1000 ks.

Figure 7. Case 2 Specific Mass Advanced Technology

100¢

[] 100 GHz Arnenr_Reclonr_

] Transmiller

I _MAO

[] Slorl_e

[] Array

AIIItude (kin)

The specific mass is plotted for Case 3 advanced

technology in figure 8. All elements except the

antenna/rectenna decrease with altitude due to the

increasing view time with the satellite in the sun. The

RFC's mass is decreasing with altitude because of the

Figure 8. Case 3 Specific Mass Advanced Technology

lg0( c_,

• i_3111_% I _ =,-;,

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IR_-':':-i:i::l_ _::_ " l'rrrrrr111Tllll IIIIIIIIII

========================== =_==|.......'.._!==|||, =: .... =_......

tO0 1000 1000(

ARItld¢ (lun)

corresponding decreasing discharge time. The satellite

and surface PMAD and the transmitter masses decrease as

the altitude increases due to a lower peak power resulting

from increasing transmission times. The antenna/rectenna

do not contribute significantly below 2500 ks.

Plotted in figure g is the specific mass using the thin film

technologies at 1O0 GHz. The mass trends of the thin film

technologies are the same as in the advanced technology

case. However, the array has become much less of a mass

driver than the energy storage by virtue of its lower specific

mass. The energy storage total mass has not decreased

since ii assumes the same RFC's as in the advanced

technology. The transmitter mass is larger than, in th e

advanced technology due to the thermal support arising

from the assumed 50% efficiency. Finally, by using

inflatable structures for the integrated thin film antenna and

Iolar array, the significance of antenna weight has been

further reduced.

_. Case 3 Specific Mass Thin Film Techn01o_y

Im I

_::_-:<_ _ .. .... , ........ ,

,oo ....... ,oo_ ...... :_oc

AltluJUCh_ (kin)

FUTURE WORK

In all scenarios of interest, storage and PMAD are the

dominate mass. To reduce the system specific mass

without using very optimistic projections for energy

storage, means must be developed to further reduce or

eliminate the energy storage. One method of reducing the

energy storage requirements is to place the satellites into

inclined elliptical orbits such that they experience shorter

eclipses. High enough inclination could eliminate the

need for surface or orbiting energy storage except for

periods of earth eclipse (possibly twice a year) and for

emergency back-up.

Tho antenna end/or roctenna could be articulated with a

large gimbalted mechanism to point directly at the satellite,

capturing more incident RF energy and allowing inclined

orbits. The increased transmission efficiency would result

in lower power dependent masses. Additionally, the field

of view would be greatly increased resulting in fewer

satellites at a given altitude. Further, the beam time to

ORIGINAL PAGE IS

OF POOR QUALITY
eclipse ratio would be greatly increased resulting in

reduced storage requirements. More analysis needs to be

performed to understand the issues involved and to

determine the ultimate payback.

The antenna and rectenna masses dominate the total

system at 2.45 GHz, leading to the conclusion that higher

frequencies are necessary to lower the system mass. For

the cases investigated other considerations such as

structural needs or the ability to articulate the

antenna/rectenna may also justify the higher frequendes.

Conversely, transmltler efficiency versus frequency must

be determined. The transmitter can become a major mass

ddver if the actual afficiencles are considerably less than

assumed for this study. The mass could increase due to

the larger thermal management needed to reject the

increased waste energy. Also, lower efficiency would have

I ripple effect through the entire system by increasing the

size of the power dependent masses.

High magnetic fields are needed to operate gyrotron tube

microwave transmitters. These high magnetic field levels

require that for high frequencies, the gyrotron operates

either with superconducting magnets, or at harmonics of

the resonant frequency. Operation at a harmonic

frequency reduces the required magnetic field by a factor

approximately equal to the harmonic number, thereby

overcoming the need for superconducting magnets.

Beyond the second harmonic, however, the efficiency

rapidly degrades [5]. Further analysis needs to be done to

determine the mass of the thermal rejection system

needed to cool the superconducting magnet to cryo-level

temperatures.

In this study several assumptions were made that will need

to be addressed in future studies. The actual

transmitter/rectenna link efficiency should be calculated,

although it was seen that the antenna/rectenna size was

not the mass driver in the scenarios investigated.

Parameters such as antenna size and array/storage size

though should be traded to optimize system performance.

Operations issues such as installation, maintenance,

station keeping and reliability must be considered. Also,

the assumption of a single base,located on the equator,

may not be realistic and inclined orbits may need be

investigated to accommodate high latitude base locations.

The only figure of merit truly considered in this study was

apecific mass. Other criteria such as volume, cost and

operation issues must be addressed.

CONCLUSIONS

The energy storage for the 14 day "lunar night" is the major

mass driver for a lunar surface power system with a solar

source. Using an orbiting beam power system, the eclipse

times are greatly reduced, and therefore, the energy

storage mass is greatly reduced.

To realize this gain though, requires the additional

elements of a beam power transmission link consisting of

satellite(s), structure, PMAD, transmitter, antenna,

microwave link and rectenna. These add not only mass but

also efficiency losses to the power train that must be made

up with additional solar array size. The net mass change will

depend on many factors.

The study has demonstrated that the specific mass of a

beam power under certain technology and scenario

msump(ions falls between surface solar and nuclear power

options (is 20 to 500 kg/kw). More favorable system

scenarios then those examined and system optimization

could bring down the specific mass further. However,

more work needs to be accomplished to understand the

actual feasibility of microwave beam transmission.

In order to consider a beam power system as a lunar power

option, advanced and/or thin film technologies are

required. The costs and risks involved in developing these

technologies needs to be examined. Likewise,

operational issues such is satellite control, station

keeping, installation and maintenance need to be

addressed.

Presently we are commencing a study to determine the

feasibility of beam power transmission concepts for the

moon and mars as part of the Space Exploration Initiative

(SEI). The system level studies described here will be

expanded to include other options such as laser

transmission and nuclear power sources. Also, more

intensive investigations into the performances, costs, risks

and developments of the technologies needed in a beam

power transmission system will be accomplished with

particular attention paid to system integration issues.

REFERENCES

[1] R.J. Sovie, "Nuclear Power Systems for Lunar and Mars

Exploration', to be published in 41st International

Astronautical Conference.

[2] W.C. Brown, "Design Study for a Ground Microwave

Power Transmission System For Use With a High Altitude

Powered Platform," NASA Wallops Flight Center, Contract

NAS 6-3200, pp. 2.15-2.23.

[3] Nils J. Diaz, et al., "UTVRs for Space Power

Generation', Prec. Space Power Vol. 8 Numbers 1/2, pp.

230, lg89

] G.A. Landis and R.C. Cull, *Applications of Thin Film

hnology Toward a Low-Mass Solar Power Satellite', to

be published in Vision-21 :Space Proceedings at Travel for

the Next Millennium.

[5] Kenwyn Long, NASA Lewis Research Center,

personal conversation, April 19, 1990.

This work was supported by the NASA Lewis Research

Center under contract NAS3-25266 with Ronald C. Cull as

monitor.

7
National Aeronautics and Report Documentation Page

Space Adminislration

1. Report No. 2. Government Accession No.

NASA TM-103211

4. Title and Subtitle

Lunar Orbiting Microwave Beam Power System

7. Author(s)

Edgar H. Fay and Ronald C. Cull

9. Performing Organization Name and Address

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

3. Recipient's Catalog No.

5. Report Date

6. Performing Organization Code

8. Performing Organization Report No.

E-5620

10. Work Unit No.

506-41-41

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

Prepared for the 25th Intersociety Energy Conversion Engineering Conference cosponsored by the AIChE, ANS,

SAE, ACS, AIAA, ASME, and IEEE, Rend, Nevada, August 12-17, 1990. Edgar H. Fay, Sverdrup

Technology, Inc., Lewis Research Center Group, 2001 Aerospace Parkway, Brook Park, Ohio 44142.

Ronald C. Cull, NASA Lewis Research Center.

16. Abstract

A microwave bean] power system using lunar orbiting solar powered satellite(s) and surface rectenna(s) was

investigated as a possible energy source for the moon's surface. The concept has the potential of reduced system

mass by placing the power source in orbit. This Can greatly reduce and/or eliminate the 14 day energy storage

requirement of a lunar surface solar system. Also propellants required to de-orbit to the surface are greatly

reduced. To determine the practicality of the concept and the most important factors, a "zero-th order "_'

feasibility analysis was performed. Three different operational scenarios employing state of the art technology and

forecasts for two different sets of advanced technologies were investigated. To reduce the complexity of the

problem, satellite(s) were assumed in circular equatorial orbits around the moon, supplying continuous power to a

single equatorial base through a fixed horizontal rectenna on the surface. State of the art technology yielded

specific masses greater than 2500 kg/kw, well above projections for surface systems. Using advanced

technologies the specific masses_on the order of 100 kg/kw which is within the range of projections for

surface nuclear (20 kg/kw) and solar systems (500 kg/kw). Further studies examining optimization of the

scenarios, other technologies such as lasers transmitters and nuclear sources, and operational issues such as

logistics, maintenance and support are being carried out to support the Space Exploration Initiative (SEI) to the

moon and Mars.

17. Key Words (Suggested by Author(s))

Microwave

Beam

Power

18. Distribution Statement

Unclassified- Unlimited

Subject Category 18

19. Security Ctassif. (of this report) 20. Security Classif. (of this page) 21. No. of pages 22. Price*

Unclassified Unclassified 8 A02

NASAFORM1626OCT86 *For sale by the National Technical Information Service, Springfield, Virginia 22161