NASA/TM-- 1998-208821

Electric Propulsion for

Low Earth Orbit Constellations

Steven R. Oleson

Dynacs Engineering Company, Brook Park, Ohio

John M. Sankovic

Lewis Research Center, Cleveland, Ohio

November 1998

https://ntrs.nasa.gov/search.jsp?R=19990008485 2018-05-11T16:59:06+00:00Z
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NASA/TM-- 1998-208821

Electric Propulsion for

Low Earth Orbit Constellations

Steven R. Oleson

Dynacs Engineering Company, Brook Park, Ohio

John M. Sankovic

Lewis Research Center, Cleveland, Ohio

Prepared for the

1998 JANNAF Propulsion Meeting

sponsored by the Chemical Propulsion Information Agency

Cleveland, Ohio, July 15-17, 1998

National Aeronautics and

Space Administration

Lewis Research Center

November 1998
Acknowledgments

Research for this paper was done at NASA Lewis Research Center's

On-Board Propulsion Branch under Contract NAS3-27186.

This report contains preliminary

findings, subject to revision as

analysis proceeds.

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Price Code: A03

Available from

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Price Code: A03
Electric Propulsion For Low Earth Orbit Constellations

Steven R. Oleson

Dynacs Engineering Co., Inc.

NASA Lewis Research Center Group

Brookpark, Ohio 44142

John M. Sankovic

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio, 44135

Hall Effect electric propulsion was evaluated for orbit insertion, satellite repositioning, orbit maintenance and deorbit applications for a sample low earth orbit satellite constellation. Since the low masses of these satellites enable

multiple spacecraft per launch, the ability to add spacecraft to a given launch was used as a figure of merit. When

compared to chemical propulsion, the Hall thruster system can add additional spacecraft per launch using

planned payload power levels. One satellite can be added to the assumed four satellite baseline chemical launch

without additional mission times. Two or three satellites may be added by providing part of the orbit insertion with

the Hall system. In these cases orbit insertion times were found to be 35 and 62 days. Depending on the electric

propulsion scenario, the resulting launch vehicle savings is nearly two, three or four Delta 7920 launch vehicles out

of the chemical baseline scenario's eight Delta 7920 launch vehicles.

INTRODUCTION

Many new, low earth orbit (LEO) constellations are

being planned or put into service. 1The lower altitudes

of these satellites necessitate many more satellites as

opposed to the only three or four geostationary

satellites, required for global coverage. Use of electric

propulsion for these LEO constellations can greatly

reduce propulsion system wet mass. This benefit can

be translated into longer lifetimes, larger payload

masses, or reduced launch masses. This final benefit

allows for either using smaller launch vehicles or

launching more satellites per launch vehicle. With the

many satellites required for LEO constellations this last

scenario can allow a substantial reduction in the

amount of launch vehicles required to place the

constellation into service.

Hall thrusters have been used on LEO spacecraft

since their introduction on the Russian Meteor series

• 2

of low earth orbit weather satellites. Many proposed

LEO satellite systems have relatively high power

payloads, 3 which are not in use during satellite

delivery, repositioning and disposal and could be

effectively used by an electric propulsion system to

increase payload mass or reduce launch mass.

A previous study dealt with electric propulsion

• 3

specifically for LEO communication satellites. In the

study described in this paper an assessment of the

benefits of Hall propulsion systems for a "generic"

LEO constellation are made. The performance

advantages were determined in terms of increased

number of satellites per launch vehicle. The sample

mission uses available information on launch

vehicles and a sample satellite constellation to create

the generic scenario. 2'4'5

MISSION ANALYSIS, OPTIONS AND

ASSUMPTIONS

Several mission tools were used in these analyses to

provide low thrust trajectory, atmospheric drag, earth

oblateness and shadow modeling. The numerical

optimization program Solar Electric Propulsion

Steering Program for Optimal Trajectory (SEPSPOT)

was used for determining optimal solar electric

propulsion starting orbits and optimal steering for

• 6

constant and shaded thrusting orbits. The numerical

orbit integration program Systems Evaluation of

Orbit Raising (SEOR), was used to test the use of

• 7

circumferential steering. The routine, Thrusting

Orbiter with Atmospheric Drag (TOAD) was used to

assess the impact of atmospheric drag on the transfer

time and velocity change (AV) required for the low

thrust transfer. 8 All chemical systems were assumed

to burn impulsively. Repositioning operations were

simulated using a scheme developed elsewhere. 9

Constant, Circumferential Thrusting

In operation LEO satellite systems have active,

relatively high power payloads which require power

• 3

in shade and sunlight. Because the payload is

usually not in use during satellite delivery,

repositioning and disposal, the power could be made

available to the propulsion system. Thus, in this

NASA/TM-- 1998-208821 1
study,theelectricpropulsion(EP)systemsdescribed

areassumedtooperatefromthesolararraysduring

thesunlitportionsof thetrajectoryandfromthe

batteriesintheshadowportion.Thisuseof payload

batterypowerfor electricpropulsionhasprecedent

with North-Southstationkeepingusing arcjet

thrustersongeostationary spacecraft.It isassumed

thattheadditionalcyclinganddifferentcharging

patternswill haveminimalimpactonthemulti-year

powersystems;a shortelectricpropulsionorbit

insertionandde-orbitaddsonlya fewextramonths

tomanyyearsofcycling.

Onebenefitofusingthepayload'spowersystemin

light andshadeis theavoidanceof non-thrusting

periodsduringshadow.This shouldallow for

simplified,circumferential steering.It hasbeen

shownwith SEPSPOTthattherequiredin-plane

steeringanglefor anothersampleLEOspacecraft

withoutshadowingis 0° or simplycircumferential

(i.e.,perpendicular totheradiusdirectionintheplane

• 4

of the orbit). The shaded optimal steering is more

complex and varies depending on shadow conditions

throughout the trajectory.

Other power/orbit/steering scenarios are possible.

For instance, using all the available, beginning-of-life

(BOL) solar array power, a higher power (but

heavier) electric thruster system could be used but

only during sunlit portions of the orbit. Such a

trajectory would require more complex steering as

mentioned above. In addition, the BOL power would

not be available at the end-of-life and thus would

require a throttleable thruster system. Another

possibility would be to use shorter electric propulsion

burns and start in an elliptical orbit; the electric

propulsion system imitates a chemical thruster. This

method, while reducing AV, would probably require

a longer trip time, as shown by Pollard and Janson. 1°

These options will be considered in further analyses.

SYSTEM ASSUMPTIONS

Two propulsion systems were assumed in the

analysis performed for this study. The candidate

systems were meant to be representative of

state-of-the-art and to show the benefit of electric

propulsion. The state-of-art chemical system assumed

a 5 kg dry mass (less tanks), an 8% tankage, and an

Ispof 235 seconds. 2 The candidate electric propulsion

system was a 1.2 kW xenon Hall thruster with

parameters shown in Table I. This candidate Hall

thruster technology is representative of the

SPT-100 u, the T-100 _2, and the D-55. _3 More

information concerning each, including availability,

can be found in the referenced texts.

Mission Phases

A generic surveillance system of three planes of

8 satellites each along with 8 spares at an altitude of

1850 km and 90 ° inclination was chosen as the

• 2

sample LEO constellation. Based on the reported

satellite mass the sample baseline satellite is

extrapolated to be approximately 608 kg at launch

with a non-propulsion mass of 440 kg. The payload

power is assumed to be 1.2 kW. A hydrazine

chemical propulsion system is baselined for this

sample mission. The Delta 7920 t'11which is assumed

to deliver four chemical satellites to the operating

orbit, was used in this analysis. A l0 year lifetime is

assumed including orbit maintenance and an

end-of-life de-orbit. 14 De-orbit for the chemical

system is assumed to consist of changing only the

orbit perigee to 185 kin and allowing atmospheric

drag lower the apogee and then de-orbit the

spacecraft. To simplify steering, the electric

propulsion systems are assumed to spiral down to a

300 km circular orbit and allow drag to de-orbit the

spacecraft. Analyses have shown that total disposal

times for both the chemical and electrical propulsion

are roughly equivalent for the sample system.

The spacecraft are also assumed to require a multiple,

quick reposition capability of one 45 ° orbit spacing

change in 3 days for each year of service. This

capability is roughly equivalent to a total 450 °

in-plane reposition at a rate of 15°/day reposition.

Moves smaller than 45 ° may take longer than

15°/day rate. Also, a greater reposition can be made

with the same fuel but at a slower rate. Repositions

are achieved by maneuvering up or down to a 'coast'

orbit where the difference in satellite speed allows a

differential drift. At the appropriate time the

spacecraft is returned to its original orbit. While use

of chemical propulsion for repositioning requires

practically all the drift to be made at the coast orbit,

the electric propulsion system takes longer to transfer

to the coast orbit and allows for some drift during

transit. A complete discussion of repositioning can

be found in Free 15and Pollard. 16

In summary, three mission tasks are assumed for the

constellation.

-- Ten Orbit Repositions: 45 degree in-plane orbit

maneuver in three days - 17m/s/reposition

(chemical), -24 m/s/reposition (electric)

NASA/TM-- 1998-208821 2
OrbitMaintenance: -5 m/s/year(includes

margin)+ 7m/sforinjectionerrors

Deorbit:

• Chemical Case:lowerperigee from

1850 km to 185 km

• Hall Cases: Lower orbit from 1850 km

circular to 300 km circular

For some of the electric propulsion scenarios the

launch vehicle is allowed to leave the spacecraft off at

lower orbits to allow the electric propulsion system to

complete the orbit insertion. The lower starting orbit is

chosen to allow additional satellites to be launched.

-- Orbit Insertion

• Electric only: raise orbit from lower

circular parking orbit to the 1850 km

operating orbit

Mission task / propulsion system combinations are

also shown in Table II.

Results

Baseline Chemical Scenario

The baseline system was assumed to use a hydrazine

monopropellant chemical system (225 s Isp, 8%

tankage fraction) for the repositioning, orbit

maintenance and the de-orbit. The Delta 7920 was

assumed to deliver four, 608 kg satellites to the

required 1850 km orbit. In all cases the separated

Delta dispensing adapter, margin and penalty for the

larger, 10 ft. fairing was assumed to be 160 kg. 5 The

yearly 45 °, 3 day reposition requires a AV of

approximately 17 m/s. Orbit maintenance is assumed

to be 5 m/s/year. After the 10-year lifetime the

satellite must be de-orbited. A disposal orbit perigee

of 185 km was assumed which would produce a

de-orbit time of roughly three months. The velocity

change needed to lower the orbit perigee to 185 km is

403 m/s. Thus, the total AV required was 628 m/s.

The chemical hydrazine system mass required to

perform these maneuvers, assuming a 608 kg initial

mass, was 168 kg. Thus, the non-propulsive

spacecraft mass required for performing the mission

was found to be 440 kg. The spacecraft mass

breakouts and the number of spacecraft per launch

vehicle for the chemical and electrical propulsion

options are shown in Figures 1 and 2, respectively.

Electric Propulsion Scenarios

The approximately 440 kg non-propulsive mass

found in the Baseline Chemical Scenario was also

assumed for the electric propulsion scenarios. The

chemical propulsion system was replaced by an

electric propulsion (EP) system. A 1.2 kW Hall

propulsion system was considered based on past

work (see Table I). 4 Because the payload power is

assumed to be 1.2 kW in sunlight and shadow, the

EP system was assumed to run off the solar arrays in

sunlit portions of the trajectory and the batteries in

the shadow portion. This use of payload battery

power for electric propulsion was described in the

mission analysis section. It was assumed that the

additional cycling and different charging pattern will

have minimal impact on the assumed 10 year system.

Five Satellites per Launcher

In this scenario the Hall thruster system directly

replaced the chemical system. The spacecraft are

launched directly to the 1850 km orbit by the Delta

launch vehicle. Repositions would be the same

45 ° in 3 day rate as the chemical system also. While

the AV to perform the reposition is larger than the

chemical reposition AV, due to the relative low thrust

of the Hall system, the much larger Isp of the Hall

system more than offsets the penalty. The de-orbit

altitude was changed to 300 km circular to allow for

simplified constant thrusting from the Hall system.

Once the spacecraft reaches the 300 km orbit, drag

could be counted on to complete the de-orbit.

Preliminary simulations have shown that the Hall

system de-orbit time, a combination of the powered

spiral down and the coasting deorbit, would be

roughly three months - the same as the chemical

baseline. Adding in the orbit maintenance to the

repositioning and deorbit, the AV budget comes to

1058 m/s. The wet Hall system mass to perform the

mission is only 54 kg. Thus the launch mass to

provide the 440 kg non-propulsive spacecraft mass is

494 kg. With a launch capability of 2470 kg to the

1850 km, 90 ° inclination orbit the Delta 7920 should

be able to launch five spacecraft. No assessment of

volume constraints have been made, but it seems

reasonable that an additional satellite could fit in the

10 ft. fairing.

NASA/TM-- 1998-208821 3
Six Satellites per Launcher

Instead of using the Delta launcher to place the

satellites directly into the final orbit like the chemical

baseline mission and the five satellite electric

propulsion scenario, this option begins from a lower

circular orbit. At this starting orbit the Delta can

place more mass. The EP system was then tasked

with raising the spacecraft to the final 1850 km

circular orbit and then maintaining, repositioning and

de-orbiting the spacecraft. The starting circular orbit

was chosen so that six spacecraft could be launched

on the Delta launcher.

By using a Hall system the required EP circular

starting orbit was 880 km with a trip time of 35 days.

De-orbit time was still around three months total. The

total mission AV was approximately 1500 m/s.

Spacecraft launch masses were 510 kg with a Hall wet

propulsion mass of 70 kg. With a launch capability of

3060 kg to the 880 kin, 90 ° inclination orbit the Delta

7920 should be able to launch six spacecraft.

Seven Satellites per Launcher

The higher starting orbits could be lowered even

further to 400 km (any lower may encounter

excessive drag) and a higher Isp Hall system used to

allow seven spacecraft to be launched. The Hall

system is assumed to be modified to a 1850 second

Iv, 55% efficient system. The Iv is increased to

minimize the required fuel mass at the expense of trip

time. Starting from the lower 400 km starting orbit

would also contribute to a longer insertion time

(62 days). The AV budget for the orbit insertion,

repositioning, maintenance, and de-orbit comes to

1800 m/s. This budget is higher due to the larger

orbit insertion (714 m/s). It is also higher due to a

higher reposition AV (27 m/s) due to the lower thrust

of the higher Iv Hall system. The Hall system wet

mass for this option is 68 kg. The launch mass

required for the 440 kg non-propulsive mass is

508 kg. Thus seven spacecraft would total 3556 kg,

within the Delta 7920 capability for the 400 km, 90 °

starting orbit.

Volume Considerations

Packaging additional satellites into the Delta 7420

fairing was not considered in this analysis due to lack

of packaging and dispenser information.

Launches Saved

For the assumed system, a total constellation of

32 satellites including 8 spares must be launched to

provide complete service. Assuming all the satellites

were to be launched on Delta 7920s, eight launch

vehicles would be required: 32 satellites / 4 per

launch = 8 Deltas. With electric propulsion adding

one satellite per launch almost two Delta launch

vehicles could be saved: 32 satellites / 5 per launch =

6 Deltas plus two satellites. These two satellites

could be spares and perhaps piggy backed on

another launch or launched on a smaller launch

vehicle when needed. With electric propulsion

adding two satellites per launch almost three Delta

launch vehicles could be saved: 32 satellites / 6 per

launch = 5 Deltas plus two satellites. Again the two

satellites should be cheaper to launch. Finally, with

electric propulsion adding three satellites per launch

the required number of Delta launch vehicles can be

almost halved: 32 satellites / 7 per launch = 4 Deltas

plus four satellites. Again, these four satellites can be

considered spares and could be launched on smaller

launch vehicles.

CONCLUSIONS

It was shown that the mass of up to three satellites

can be added to multiple LEO constellation

spacecraft launches by using electric propulsion for

orbit insertion, maintenance, repositioning, and

de-orbit. The result is almost halving the number of

launch vehicles required, from eight to four and a

half. If only one satellite is added per launch

practically no extra time is required for orbit

insertion of other mission operations. To add two or

three extra satellites, orbit insertion times of 35 and

62 days would be needed, respectively. A simple

circumferential steering method was assumed which

relies on the payload's solar array and battery power

and eliminates the more complex steering required

when shading of the solar arrays must be considered.

Assumed reposition rates were the same for the

chemical and Hall systems: 45 ° in 3 days. Deorbit

times were also the same for the chemical and Hall

systems- roughly three months.

NASA/TM-- 1998-208821 4
REFERENCES

1. Ruzicka, Milan, "The Race to LEO",

Launchspace, August/September, 1997.

2. Wilson, A., Jane's Space Directory, Eleventh

Edition 1996-97, 1996 Jane's Information Group

Ltd., Sentinel House, Surrey, UK.

3. Mukund, P., "Power Systems for LEO Satellites",

Launchspace, August/September, 1997.

4. Oleson, S.R., "Electric Propulsion for

LEO Communication Satellites", IEPC-97-148,

Aug. 1997.

5. Isakowitz, S.J., Samella, J., "International

Reference Guide to Space Launch Systems",

2nd edition, American Institute of Aeronautics and

Astronautics, Washington, DC.

6. Sackett, L.L., et al., "Solar Electric Geocentric

Transfer with Attitude Constraints: Analysis",

NASA CR-134927, Aug., 1975.

7. Horsewood, J.L., Flanagan, P.F., Maresca, P.A.,

"Program Manual for the Systems Evaluation of

Orbit Raising (SEOR) Computer Program",

Report No. 75-23, Contract NAS3-19087, Revised

March 1976.

8. Myers, R.M., et al., "Small Satellite Propulsion

Options", NASA TM-106701, AIAA Paper 94-2997,

June 1994.

9. Riehl, J., Morales, N., Oleson, S., "Benefits of Ion

Electric Propulsion for Earth Orbital Applications

Benefits Study", Unpublished Lab Report,

May 1995.

10. Pollard, J.E., Janson, S.W., "Spacecraft Electric

Propulsion Applications", Aerospace Report

ATR-96 (8201)-1, February 1, 1996.

11. Day, M. , Maslennikov, N. , Randolph, T.;

Rogers, W. "SPT- 100 Subsystem Qualification

Status", AIAA Paper 96-2713, Joint Propulsion

Conference, 32nd, Lake Buena Vista, FL,

July 1-3, 1996;

12. Petrosov, V.A., et.al., "A 2000 Hours Lifetime

Test Results of 1.3 kW T-100 Electric Thruster",

24 thIEPC, Moscow, Russia, Sept. 1995.

13. Sankovic, John M., Haag, Thomas W., Manzella,

David H., "Operating Characteristics Of The

Russian D-55 Thruster With Anode Layer",

AIAA Paper 94-3011.

14. Anon., "NASA Orbital Debris Assessment

Handbook", Goddard Space Fight Center,

Rev. 4/20/93.

15. Pollard, J.E., Janson, S.W., "Spacecraft Electric

Propulsion Applications", Aerospace Report

ATR-96 (8201)-1, Feb. 1996.

16. Free, Bernard, Babuska, Vit, "Repositioning Of

Geostationary Spacecraft - Chemical And Electric

Propulsion Options", AIAA Paper 96-1144,

International Communications Satellite Systems

Conference, 16th, Washington, DC, Feb. 1996.

NASA/TM-- 1998-208821 5
TableI--Propulsion Systems for LEO Constellation

Propulsion System Xenon Hall Thruster Modified Xenon Hall

Parameters Thruster

Desired PPU Input Power 1.2 kW 1.2 kW

Level

1500 s 1850 s Isp

Overall Efficiency (PPU

& Thruster)

Tankage

Masses:

Thruste_

0.45

10%

5 kg

34 % of Thruster

0.51

10%

Gimbals

Supporl 31% of Gimbals 31% of Gimbals

& Thrusters & Thrusters

Controlle_

Total Thruster + Gimbal ÷

Support + Controllei

Feed System

PPU

Cablin_

0.55 kg/Thruster

9.3 kg/thruster

1.5 kg/kWe

4.7 kg/kWe

0.4 kg/kWe

31 kg/kWt-disp.

9.1 kg/kWe

Thermal Sys. (92% PPU)

Total PPU + Feed ÷

Cabling + Thermal

5k_

34 % of Thruster

0.55 kg/Thruster

9.3 kg/thruster

1.5 kg/kWe

4.7 kg/kWe

0.4 kg/kWe

31 kg/kWt-disp.

9.1 kg/kWe

Table II--Mission phases and propulsion options for LEO spacecraft

Mission Phase Baseline: EP 5 EP 6 Sats/ EP 7 Sats/

Monoprop 4 Sats/Launch Launch Launch

Sats/Launch Scenario Scenario Scenario

•Orbit Insertion

Raise orbit from

880 km to 1850 km

-Raise orbit from

400 km to 1850 km

.Repositions

45 ° in orbit in 3

days (10 total)

•Orbit maintenance

5 m/s/y

.Deorbit

-Lower perigee

from 1850 km to

185 km

-Lower orbit from

1850 km to 300 km

Monoprop

(225 s Isp)

Monoprop

(225 s Isp)

Monoprop

(225 s Isp)

Hall

(1500 s Isp)

Hall

(1500 s Isp)

Hall

(1500 s Isp)

Hall

(1500 s Isp)

Hall

(1500 s Isp)

Hall

(1500 s Isp)

Hall

(1500 s Isp)

Hall

(1850 s Isp)

Hall

(1850 s Isp)

Hall

(1850 s Isp)

Hall

(1850 s Isp)

NASA/TM-- 1998-208821 6
Satellite Mass Breakouts

7OO

600

500

400

300

200

100

0

4 Sats /

Launch:

Monoprop

Required Bus Mass

.... i.... i.... ......... i...

5 Sats. / 6 Sats. / 7 Sats. /

Launch: Launch: Launch:

Hall Hall Orbit High Isp

Insertion Hall Orbit

reBus & Payload Mass nFuel Mass nDry Propulsion Sys. I

Figure lmSatellite Mass Breakouts

LEO Constellation Launch Masses

4000

3500

3000

2500

2000

1500

1000

500

0

Delta 7920* to 400 km / 90° 61 day insertion

Delta 7920* to 880 km / 90° 35 day insertion ::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::

508

::::::::::::::::::::::::::::::::::::::::::::::::::::

508

::::::::::::::::::::::::::::::::::::::::::::::::::::::

508

Delta 7920* to 1850 km / 90 °

608

494

608

494

I

4 Sats /

Launch:

Monoprop

510

510

510

5 Sats. / 6 Sats. / 7 Sats. /

Launch: Launch: Launch:

Hall Hall Orbit High Isp

Insertion Hall Orbit

* Does not include lOft. Fairing, Adapter, or contingency Masses

Figure 2mLaunch Masses for Each LEO Constellation Propulsion Scenario

NASA/TM-- 1998-208821 7
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4. TITLE AND SUBTITLE

Electric Propulsion for Low Earth Orbit Constellations

6. AUTHOR(S)

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7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Lewis Research Center

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NASA TM--1998-208821

11. SUPPLEMENTARY NOTES

Prepared for the 1998 JANNAF Propulsion Meeting sponsored by the Chemical Propulsion Information Agency,

Cleveland, Ohio, July 15-17, 1998. Steven R. Oleson, Dynacs Engineering Company, Brook Park, Ohio 44142;

and John M. Sankovic, NASA Lewis Research Center. Responsible person, Steven R. Oleson, organization code 5430,

(216) 977-7426.

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13. ABSTRACT (Maximum 200 words)

Hall Effect electric propulsion was evaluated for orbit insertion, satellite repositioning, orbit maintenance and de-orbit

applications for a sample low earth orbit satellite constellation. Since the low masses of these satellites enable multiple

spacecraft per launch, the ability to add spacecraft to a given launch was used as a figure of merit. When compared to

chemical propulsion, the Hall thruster system can add additional spacecraft per launch using planned payload power

levels. One satellite can be added to the assumed four satellite baseline chemical launch without additional mission times.

Two or three satellites may be added by providing part of the orbit insertion with the Hall system. In these cases orbit

insertion times were found to be 35 and 62 days. Depending on the electric propulsion scenario, the resulting launch

vehicle savings is nearly two, three or four Delta 7920 launch vehicles out of the chemical baseline scenario's eight Delta

7920 launch vehicles.

14. SUBJECT TERMS

Electric propulsion; LEO satellite; Hall thruster

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

15. NUMBER OF PAGES

13

16. PRICE CODE

A03

18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT

OF THIS PAGE OF ABSTRACT

Unclassified Unclassified

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-1B

298-102