Affordable Spacecraft: Design and Launch

Alternatives

January 1990

OTA-BP-ISC-60

NTIS order #PB90-203225
Recommended Citation:

U.S. Congress, Office of Technology Assessment, Affordable Spacecraft: Design and Launch

Alternative --Background Paper, OTA-BP-ISC-60 (Washington, DC: U.S. Government

Printing Office, January 1990).

Library of Congress Catalog Card Number 89-600776

For sale by the Superintendent of Documents

U.S. Government Printing Office, Washington, DC 20402-9325

(order form can be found in the back of this report)
Foreword

Several major DOD and NASA programs are seeking ways to reduce the costs of

launching spacecraft. However, it typically costs much, much more to build a spacecraft than

to launch it into a low orbit. Therefore, unless the costs of building spacecraft are reduced, even

dramatic reductions in costs of launching to low orbit would reduce total spacecraft program

costs by only a few percent.

This background paper examines several proposals for reducing the costs of spacecraft

and other payloads and describes launch systems for implementing them. It is one of a series

of products of a broad assessment of space transportation technologies undertaken by OTA

at the request of the Senate Committee on Commerce, Science, and Transportation, and the

House Committee on Science, Space, and Technology. In 1988, OTA published the special

report Launch Options for the Future: A Buyer’s Guide and the technical memorandum

Reducing Launch Operations Costs: New Technologies and Practices. In 1989, OTA

published the background paper Big, Dumb Boosters: A Low-Cost Space Transportation

Option? and the special report Round Trip to Orbit: Human Spaceflight Alternatives. A

summary report on space transportation, entitled Access to Space to be published in the spring

of 1990, will be the final report in the space transportation series.

In undertaking this effort, OTA sought the contributions of a wide spectrum of

knowledgeable individuals and organizations. Some provided information; others reviewed

drafts. OTA gratefully acknowledges their contributions of time and intellectual effort. OTA

also appreciates the cooperation and assistance of the Air Force and NASA. However, OTA

is solely responsible for the content of this background paper and the other OTA publications.

u JOHN H.-GIBBONS

Director
Advisory Panel on Affordable Spacecraft:

Design and Launch Alternatives

M. Granger Morgan, Chair

Head, Department of Engineering and Public Policy

Carnegie-Mellon University

Michael A. Berta

Assistant Division Manager

Space and Communications Group

Hughes Aircraft Company

Richard E. Brackeen

President

Martin Marietta Commercial Titan, Inc.

Edward T Gerry

President

W. J. Schafer Associates, Inc.

Jerry Grey

Director, Science and Technology Policy

American Institute of Aeronautics and Astronautics

William H. Heiser

Consultant

Otto W. Hoernig, Jr.

Vice President

Contel/American Satellite Corporation

Donald B. Jacobs

Vice President, Space Systems Division

Boeing Aerospace Company

John Logsdon

Director, Space Policy Institute

George Washington University

Hugh F. Loweth

Consultant

Anthony J. Macina

Program Manager

IBM Federal Systems Division

George B. Merrick

Vice President

North American Space Operations

Rockwell International Corporation

Alan Parker

Senior Vice President

Technical Applications, Inc.

Gerard Piel

Chairman Emeritus

Scientific American

Bryce Poe, II

General, USAF (retired)

Consultant

Ben R. Rich

Vice President and General Manager

Lockheed Corporation

Sally K. Ride

California Space Institute

Tom Rogers

President

The Sophron Foundation

Richard G. Smith

Senior Vice President

JLC Aerospace Corporation

William Zersen

Consultant

NOTE: OTA appreciates the valuable assistance and thoughtful critiques provided by the advisory panel members. The views

expressed in this OTA report, however, are the sole responsibility of the Office of Technology Assessment. Participation on

the advisory panel does not imply endorsement of the report.

iv
OTA Project Staff on

AFFORDABLE SPACECRAFT—DESIGN AND LAUNCH ALTERNATIVES

Lionel S. Johns, Assistant Director, OTA

Energy, Materials, and International Security Division

Alan Shaw, International Security and Commerce Program Manager (from March 1989)

Peter Sharfman, International Security and Commerce Program Manager (through February 1989)

Ray A. Williamson, Project Director (from February 1989)

Richard DalBello, Project Director (through January 1989)

Michael B. Callaham, Principal Analyst

Eric O. Basques (through August 1989)

Gordon Law

Administrative Staff

Jackie Robinson Louise Staley
Acknowledgments

The following organizations generously provided OTA with information and suggestions:

American Rocket Corporation

Boeing Aerospace Company

California Institute of Technology,

Jet Propulsion Laboratory

Defense Advanced Research Projects Agency

Defense Systems, Inc.

Electromagnetic Launch Research, Inc.

NASA Headquarters

Naval Research Laboratory

Orbital Sciences Corporation

Radio Amateur Satellite Corp.

Strategic Defense Initiative Organization

TRW, Inc.

U.S. Air Force Systems Command, Space Systems

Division

U.S. Space Command

This paper has also benefited from the advice of many experts from the Government and the private

sector. OTA especially would like to thank the following indivuals for their assistance and support. The views

expressed in this paper, however, are the sole responsibility of the Office of Technology Assessment.

Steve Altes

Orbital Sciences Corporation

Ivan Bekey

NASA Headquarters

James Bennett

American Rocket Corporation

Carl Builder

The Rand Corp.

Paul Dergarabedian, Sr.

Aerospace Corp.

George Donohue

The Rand Corp.

Dani Eder

Boeing Aerospace Company

John Gaines

General Dynamics

Daniel Gregory

Boeing Aerospace Company

Col. Charles Heimach, USAF

USAF Systems Command, Space

Systems Division

Abraham Hertzberg

University of Washington

Ross M. Jones

California Institute of Technology

Jet Propulsion Laboratory

Jordin Kare

Strategic Defense Initiative

Organization

Clark Kirby

TRW, Inc.

Henry Kolm

Electromagnetic Launch Research,

Inc.

George Koopman

American Rocket Corporation

Robert Lindberg

Orbital Sciences Corporation

Andrew C. McAllister

Radio Amateur Satellite Corp.

Lt. Col. Tom O’Brien, USAF

U.S. Space Command

Scott Pace

U.S. Department of Commerce

Miles Palmer

Science Applications International

Corp.

David Rossi

Orbital Sciences Corporation

Arthur Schnitt

George Sebestyen

Defense Systems, Inc.

Larry Stern

George Mason University

Peter Wilhelm

Naval Research Laboratory

Col. John Wormington, USAF

USAF Systems Command, Space

Systems Division

vi
Contents

Page

Chapter 1. Summary ..o. ..o. ... o.. ... ... ... .*. ... *.. ... ... ... ... Q.*...**""""""*""""""""""""" 1

THE HIGH COST OF SPACECRAFT’ . . . . . . . . . . . . . . ........................................s 1

APPROACHES TO REDUCING PAYLOAD COSTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

OPTIONS FOR CONGRESS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

Options for Influencing Spacecraft Design . . . . . . . . . . . . . . . . . . . . . . . . . .. .. .. .. .. .. ... .... .... O. 4

Options for Promoting the Development of Launch Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

Chapter 2. Weight and Volume Margins ● ....0.0 . . . . . . . . . . . . . . . . . . . . . . . . . . . ● .00..00. ● ...0...0 7

Chapter 3. Fatsats .....000 .....000. . . . . . . . . . . . . . . . . . . ..0.0.0.0 .0.....00 .0..0.0.. .00.0.... .0 0 9

STANDARDIZING SUBSYSTEMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

UPPER STAGEHAND ORBITAL-TRANSFER VEHICLES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

ESTIMATING POTENTIAL COST REDUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

A Parametric Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

Bottom-Up Analyses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

A Comparison of Parametric and Bottom-Up Analyses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

ACHIEVING POTENTIAL COST REDUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

Chapter 4. Lightsats ......00 . . . . . . . . . . . . . . . . . . .... . . . . . . . ....000.0 ● 19

SPACECRAFT CONCEPTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

CONCEPTS FOR PHASED ARRAYS OF SPACECRAFT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

LAUNCH REQUIREMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

LAUNCH SYSTEM OPTIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

Existing Launch Vehicles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

Air-Launched Vehicles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

Standard Small Launch Vehicle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

Other Options.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

ISSUES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

Chapter 5. Microspacecraft .0.00..0 ● .00..... . . . . . . . . . ...00.0.. ● .....000 ● .......0 ● 0...000. . . . . 29

SPACECRAFT CONCEPTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

LAUNCH SYSTEM OPTIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

ISSUES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

Boxes

Box Page

3-A. Fatsats: The Good and Ugly, and the Bad . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

4-A. Brilliant Eyes? . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

5-A. Lasers for Rocket Propulsion: The State of the Art . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

Figures

Figure Page

2-1. Value of Weight Margin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

3-1. Average Launch Cost v. Payload Capability. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

3-2. Cost for Hypothetical Mission With Current Launch Cost Trend . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

3-3. Cost for Hypothetical Mission If Launch Costs Are Reduced 67 Percent . . . . . . . . . . . . . . . . . . . . . . . 13

3-4. Payload Cost v. Weight Trade-offs and Economic Launch Costs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

3-5. If a Satellite Were Allowed To Be Heavier It Could Cost LESS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3-6. The Effect of Reducing Launch Cost on the Optimal Weight and Cost of a Payload. . . . . . . . . . . . . 15

4-1. The Global Low-Orbit Message Relay (GLOMR) Satellite . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

4-2. Microsat: A 22-lb Communications Satellite for Amateur (“Ham’’)Radio Operators . . . . . . . . . . . . 21

4-3. Sparse Array of Satellites Could Provide Detailed Radar Maps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

4-4. Orbiting Low-Frequency Array of Radioastronomy Satellites . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

4-5. Pegasus Launch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

4-6. Projected Performance of Pegasus: Payload v. Orbital Altitude and Inclination . . . . . . . . . . . . . . . . . . 25

5-1. Laser-Powered Rockets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

5-2. Mars Observer Camera Microspacecraft Design Cutaway View.... . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

Tables

Table Page

3-1. Ratios of Nonrecurring to Recurring Cost of Spacecraft Subsystems . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 0

5-1. Estimated Cost of a 20-Megawatt Laser for Powering Rockets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

vii
Chapter 1

Summary

Several major efforts are aimed at finding

ways to reduce the cost of launching spacecraft.

However, it typically costs much, much more to

build a spacecraft than to launch it to low Earth

orbit (LEO). Unless spacecraft costs are reduced, even dramatic reductions in launch costs

will have only a small effect on total spacecraft

program costs.

This Background Paper reviews four possible

approaches to spacecraft design that have been

proposed to reduce total spacecraft program

costs. Adopting them could change the launch

rates, payload capacity, and reliability demanded

of conventional launch vehicles, or create a

demand for exotic launch systems to launch very

small spacecraft cheaply. Conversely, developing new, economical launch systems would

strengthen incentives to adopt these new approaches to spacecraft design.

This is one of several publications documenting OTA’s broad assessment of space transportation technologies requested by the House Committee on Science, Space, and Technology, and

the Senate Committee on Commerce, Science,

and Transportation. Previous publications in

this assessment examined a variety of future

launch options,l ways to reduce the costs of

launch operations,2 low-cost, low-technology

(“big, dumb") boosters,3 and options for transporting humans to and from orbit.4 A final report

will be published in 1990.

THE HIGH COST OF

SPACECRAFT

Because of the high cost of spacecraft, a

dramatic reduction in launch cost alone will

not substantially lower spacecraft program

costs. Although launching a pound of payload to

LEO currently costs about $3,000, procuring

that pound of payload typically costs much

more. For example, representative U.S. spacecraft bussess of types first launched between

1963 and 1978 cost between $130,000 and

$520,000 per pound dry,6

including amortized

program overhead costs. Procurement of the

mission payloads carried on those busses cost

about 50 percent more—about $200,000 to

$800,000 per pound.7 Reducing launch costs

from $3,000 to $300 per pound of payload, a

goal of the Advanced Launch System program,8

would reduce the total cost of procuring and

launching a dry spacecraft (half bus, half

mission payload) by less than 2 percent.

A spacecraft bound for a high orbit or another

planet requires an upper stage, which when

fueled is typically more than twice as heavy as

the spacecraft but costs less. Even so, a payload

consisting of a Centaur upper stage (about

$2,250 per pound) and a spacecraft weighing a

third as much (half bus, half mission payload)

might cost from $40,000 to $160,000 per pound.

Reducing launch costs to $300 per pound would

Iu.S. Conm55, office of l’whnolo~ Assessment, Launch Options for the Future: A Buyer’s Guide, OTA-ISC-383 (Washington, DC: U.S.

Government Printing Office, July 1988).

2u.s. Conge55, office of T~hnoloW Assessment, lled~ing hunch Operations Costs: New Technologies Und practices, OTA-TM-ISC-28

(Washington, DC: U.S. Government Printing Office, September 1988).

3u.s. conw~5, office of TWhnology As~ssment, Big, D@ Bo o sters: A ~w-cost space Tra~pO~~ion option? (WfaSMn~On, DC: U.S.

Congress, Office of Technology Assessment, International Security and Commerce program, February 1989).

4u.s. con=e55, office of T’ec~oloW A~ssment, Rou~ Trip ~ orbit: H~an spuceflightAfter~tives, OTA-ISC-419 (Washington, DC: U.S.

Government Printing Office, August 1989).

5~e c(bw~~ of a ~acar~t consists of the s~ctme, ~wer ~Wce5, and o~er subsys~ms r~uired to support the rnisslon payloads it carries.

%.Jnfueled.

7Space System ad Operatlom Cost Reduction and Cost Credibility Workshop, Executive S ummary (Washington, DC: National Security

Industrial Association, 1987), p. 3-10, fig. 3.7.3. These estimates were derived by amortizing nonrecurring costs (e.g., for development) over four

satellites; some programs procure more than four while others procure only one. OTA inflated the estimated costs, which were in 1982 dollars, to

1987 dollars using the GNP inflator from table B-3 of the ECOrWndC Report ofthe President, January 1989, and then to 1989 dollars by assuming 4.2

percent annual inflation in 1987 and 1988.

8101 s~t. 1067.

–l–
2 ● Affordable Spacecraft: Design and Launch Alternatives

reduce the total cost of procuring and launching

such a payload by only 2 to 6 percent.

APPROACHES TO REDUCING

PAYLOAD COSTS

To reduce payload costs, and for other

reasons, novel approaches to payload design and

fabrication have been proposed:

● Design payloads to fit launch vehicles

leaving size and weight margins of about

15 percent

. Allow payloads to be larger and heavier:

Fatsats

● Allow satellites to be simpler, and make

them lighter: Lightsats

● Design Microspacecraft to be launched like

artillery shells

Each type of spacecraft—fatsat, lightsat, or

microspacecraft-would impose unique launch

demands. New, large launch vehicles would be

needed to launch the heaviest satellites.

Lightsats could be launched on existing launch

vehicles, but new, smaller launch vehicles might

launch them more economically. In wartime,

small launch vehicles could be transported or

launched by trucks or aircraft to provide a

survivable means of space launch. Microspacecraft could be launched on existing launch

vehicles, but they might instead be launched by

more exotic means such as a ram cannon,

railgun, coilgun, or laser-powered rocket. Some

of these might be proven feasible in the next

decade.

Weight margin: Designing payloads to fit

launch vehicles while reserving ample size

and weight margins can reduce the risk of

incurring delay and expense after assembly

has begun.

It is often the case that satellites grow

substantially heavier than expected as they

proceed from design to construction. For example, dry weights of military spacecraft have

been about 25 percent greater, on the average,

than initially predicted. Growth in estimated

weight may be caused by underestimating the

weight of a spacecraft (especially one designed

to use the most advanced technology) or by

changing mission requirements during development (requiring hardware to be added). If a

payload grows so heavy during assembly that it

threatens to “gross out” its assigned launch

vehicle (i.e., cause its weight to equal or exceed

the maximum allowable gross lift-off weight),

the payload must be redesigned to cut its weight.

This causes delay and increases cost.

To reduce the risk of exceeding vehicle

payload capacity, program managers could require designers to design each payload to fit its

assigned launch vehicle with room to spare and

to weigh substantially less than the maximum

weight the vehicle can launch to orbit. However,

this design philosophy would lead to more

stringent size and weight constraints than would

otherwise be imposed. If a mission simply could

not be performed by payloads predicted to be

small enough to fit the largest launch vehicles

with adequate size and weight margins, new,

larger launch vehicles would have to be developed to provide the desired margin. In many

cases, however, sufficient margin could be

provided by clever design, e.g., by designing

several smaller single-mission payloads instead

of a single multimission payload, or developing

and using an electric-powered or space-based

orbital transfer vehicle (OTV) instead of a

conventional OTV.9

Fatsats: If payloads were allowed to be

heavier for the same capability, some could

cost substantially less. For example, OTA

estimates that Titan-class payloads that cost

several hundred million dollars might cost

about $130 million less if allowed to be five

times as heavy.

9

To place a satellite in a high orbit, a launch vehicle must carry both the satellite and either an upper stage to take the satellite directly to the high

orbit, or an OTV to take the satellite from a low-altitude parking orbit to the high orbit. A conventional OTV weighs two or three times as much as the

satellite it carries. An electric-powered OTV could weigh less than a conventional OTV of comparable capability, creating more weight margin. A

space-based OTV could be launched separately from its payload, allowing the payload to weigh as much as its launch vehicle could carry while reserving

the desired weight margin. However, operation of space-based OTVs would be complex and require costly infrastructure.
Chapter I-Summary ● 3

If payloads were allowed to be much heavier,

a manufacturer could forego expensive processes for removing nonessential structural material, as well as expensive analyses and tests for

assuring the adequacy of the remaining structure. Standardized subsystems, which could be

produced economically in quantity, could be

used instead of customized subsystems designed to weigh less.

The savings that might be realized are uncertain. In principle, they could be estimated by

comparing the costs of a heavy payload and a

light payload that perform the same functions

with the same capability. However, the United

States has never designed and built two payloads, one heavy and the other light, that

perform the same functions equally well, in

order to compare actual costs. A few estimates

have been derived by comparing the cost and

weight of an actual spacecraft with the estimated

cost and weight of hypothetical heavier spacecraft of comparable capability. Designers have

also compared the estimated costs and weights

of hypothetical spacecraft of comparable capability and different weights. All such studies

predict payloads could cost less if allowed to

weigh more, but the estimates of savings differ.

An accurate estimate of potential savings

requires a detailed trade-off analysis for each

payload. Achieving these savings will probably

require giving spacecraft program managers,

and those who establish mission and spacecraft

requirements, incentives crafted specifically for

the purpose, and may require developing new

launch or orbital-transfer vehicles to carry the

spacecraft.

Lightsats: If allowed to be less capable,

reliable, or long-lived, payloads could be both

lighter and less expensive. Useful functions

such as communications and weather surveillance could be performed by payloads small

enough to be launched on small rockets from

airborne or transportable launchers.

Small, simple, and relatively inexpensive

civil and military satellites have been, and still

are, launched at relatively low cost on small

launch vehicles or at even lower cost, sometimes

for free, as “piggyback” payloads on larger

launch vehicles.

The Department of Defense is considering

whether the increased survivability and responsiveness such spacecraft could provide would

compensate for possibly decreased capability.

Some missions might be accomplished as well

by a swarm of several small satellites as by a

single large one. If so, a swarm would be less

expensive in many cases, because smaller satellites typically cost much less per pound than do

large ones. Even if the satellites were launched

individually, which would increase total launch

cost, total mission cost might be lower.

Microspacecraft: Spacecraft weighing only

a few pounds could perform useful space

science missions and might be uniquely economical for experiments requiring simultaneous measurements (e.g., of solar wind) at

many widely separated points about the

Earth, another planet, or the Sun.

These could be launched on existing launch

vehicles. Eventually, it may be possible to

launch them on laser-powered rockets or, if they

are as rugged as a cannon-launched guided

projectile, 10 with a ram cannon or an electromagnetic launcher. Within the next decade,

experiments now being planned may establish

the feasibility of some of these launch systems.

Their costs cannot be estimated confidently

until feasibility is proven, but at high launch

rates they might be more economical than

conventional rockets. An electromagnetic launcher

could also be constructed in orbit to launch

microspaceprobes to outer planets; they would

arrive years earlier than if they were propelled

by conventional rockets.

IOA camon-lauc~ guided projectile is an artillery shell equipped with a system (e.g. TV camera and computer) for recomztig a target ad

movable fm or other means for steering the projectile toward a target. The Army’s M712 Copperhead is an example.
4 ● Affordable Spacecraft: Design and Launch Alternatives

OPTIONS FOR CONGRESS

What can Congress do to promote spacecraft

cost reduction and, thereby, reduce the cost of

space programs? Some options deal directly

with spacecraft design; others would promote

the development of launch systems that could

launch small, inexpensive spacecraft at low cost

or heavy spacecraft with generous weight margins.

Options for Influencing Spacecraft Design

Option 1:

Congress could order a comprehensive study

of how much the Nation could save on space

programs by:

●

●

●

●

●

designing payloads to reserve more weight

and volume margin on a launch vehicle;

allowing payloads to be heavier, less

capable, shorter-lived, or less reliable;

designing standard subsystems and buses

for use in a variety of spacecraft;

designing spacecraft to perform single

rather than multiple missions; and

using several inexpensive satellites instead

of a single expensive one.

Lockheed completed such a study in 1972;11

anew one should consider current mission needs

and technology. It would complement the Space

Transportation Architecture Study (STAS) and

more recent and ongoing studies12 that compare

space transportation options but not payload

design options.

As noted above, to estimate potential savings

accurately, a detailed trade-off analysis must be

done for each payload, or more generally, for

each mission. So, for greater credibility:

Option 2:

Congress could require selected spacecraft

programs--for example, those that might require a new launch vehicle to be developed--to

award two design contracts, one to a contractor

who would consider the unconventional approaches mentioned above.

Option 3:

Congress could require both the Department

of Defense and NASA to refrain from developing

a spacecraft if the expected weight or size of the

spacecraft, together with its propellants, upper

stage, and support equipment, would exceed

some fraction of the maximum weight or size

that its intended launch vehicle can accommodate. Public Law 100-456 required the Department of Defense to require at least 15

percent weight margin in fiscal year 1989.13

New legislation could extend this restriction to

NASA and could require size margins in future

years.

Options for Promoting the Development of

Launch Systems

Option 4:

Congress could fund the development of the

Shuttle-C cargo launch vehicle, the Advanced

Launch System, liquid-fueled rocket boosters

for the Shuttle, or a larger Titan launch

vehicle. 14 Any of these vehicles could launch

payloads larger and thus less expensive (for

comparable performance) than payloads designed to fly on Shuttles or Titan IVs. Alternatively, if payload size and weight are not

increased, these proposed launch vehicles could

provide greater size and weight margins, thereby

reducing the risk of needing costly weightreduction efforts. However, their greater payload capacity would also enable payload pro- —

ll~ckh~~ssiles and Space CO., ]Wact of Low-Cost Rejivbishable and Standard Spacecraji Upon Future NASA Space Progmms, NTIS

N72-27913, Apr. 30, 1972.

%g.,tie Air Force’s Air Force-Foc~ed SEAS, NASA’s Next Manned Space Transportation System study, the Defense Science Board’s Natioml

Space Luunch Strategy study, and the Space Transportation Compankon study for the National Aero-Space Plane Program.

13s= S. Rcpt. 100-326, p. 36, and H. Rept. 100-989, P. 282.

Id’rhe Co$ts md ~nefits of these launch systems would differ; see U.S. Congress, op. cit., foomote 1, and the forthcoming find report of this

assessment.
Chapter 1--Summary ● 5

gram managers to forego these potential savings

and instead pursue greater payload performance

by increasing payload size or weight. If they do

so and weight overrun occurs, it would probably

cost more to trim the weight of a larger payload

than it would to reduce the weight of a lighter

payload by the

weight margins,

reduce this risk.

Option 5:

same percentage. Requiring

as described above, would

Congress could continue to fund the development of the Standard Small Launch Vehicle

(SSLV15) and the Sea Launch and Recovery

(SEALAR) system16 to provide survivable means

of launching military lightsats. The Department

of Defense probably will not allow operational

lightsats to be designed for such launch vehicles

until the vehicles are operational. Hybrid rockets, which can use liquid oxygen to burn

nonexplosive solid propellant similar to tire

rubber, could also be designed to launch

lightsats from transportable launchers. Such

hybrids would have some safety advantages and

might be allowed where conventional solid- or

liquid-fuel rockets are not. Later—perhaps by

2005-NASP-derived vehicles might be able to

launch 20,000-pound payloads in wartime.17

With continued funding, the National AeroSpace Plane Program would continue to develop

technology for NASP-derived vehicles.

Option 6:

Congress could fired the development of a

laser or a direct-launch system (e.g., a railgun,

coilgun, or ram cannon) for launching microspacecraft at high rates economically. Many

uses have been proposed, but to date only the

Strategic Defense Initiative Organization (SDIO)

has identified a plausible demand for high-rate

launches of microspacecraft. However, demand

for launches of scientific, commercial, and other

microspacecraft could increase, perhaps dramatically, if launch costs could be reduced to a few

hundred dollars per pound and payloads were

inexpensive. The SDIO estimates development

and construction of a laser for launching 44-

pound payloads would require about $550

million over 5 or 6 years. The SDIO estimates

it could launch up to 100 payloads per day (more

than 20 Shuttle loads per year) for about $200

per pound.

Railgun proponents predict a prototype

railgun capable of launching 1,100-pound projectiles carrying 550 pounds of payload could be

developed in about 9 years for between $900

million and $6 billion, including $50 million to

$5 billion for development of projectiles and

tracking technology. If produced and launched

at a rate of 10,000 per year, the projectiles (less

payload) would cost between $500 and $30,000

per pound (estimates differ). The cost of launching them might be as low as $20 per pound—i.e.,

$40 per pound of payload.

Is’’f’he SSLV is being develop~ by the Defense Advanced Research projeets Agency (D~A).

IGSEALAR is king developed by the U.S. Naval Research Laboratory.

ITS= U.S. Congress, op. cit., footnote 4, pp. 67 and 74.
Chapter 2

Weight and Volume Margins

As payloads progress from the drawing board (or

computer-aided design workstation) to the launch

pad, their volumes tend to expand to fill the payload

fairings of their betrothed launch vehicles, and their

weights tend to grow to the maximum that the launch

vehicles can launch. Dry weights of representative

U.S. military satellites have been about 25 percent

greater, on the average, than the estimates of dry

weight made when they were proposed.l On occasion, satellites grow so heavy that their assigned

launch vehicles would be unable to place them in the

desired orbits. When this happens, drastic and

expensive efforts are undertaken to reduce the

weight of the satellite—and sometimes of the launch

vehicle as well. A TRW executive has said, “We

have to spend numbers like $150,000 a pound trying

to get the last few pounds out of a spacecraft. ”2

A contract for spacecraft development and production typically specifies that the spacecraft perform certain functions and not weigh more than a

specified maximum weight. Usually this maximum

weight is chosen to be less than the maximum weight

a particular launch vehicle can place in the desired

orbit. The difference between these two values

provides a margin that allows for contingencies such

as less-than-expected launch vehicle thrust.

The spacecraft designer initially tries to design the

spacecraft to perform the specified functions and not

weigh more than the specified maximum weight,

minus

1. a “contingency” amount by which the estimated weight of the spacecraft is expected to

grow as the design matures, and

2. a “weight margin” representing greater-thanexpected growth in the estimated weight as the

design matures.

Based on its experience producing high-tech government spacecraft and spacecraft subsystems, TRW

expects weight to grow ultimately 15 percent larger

than initial estimates and hence budgets a contingency of 15 percent initially. As the design matures

and the estimated weight increases as expected, the

contingency budget is decreased.

TRW may also allow a margin of 15 percent or

more for greater-than-expected weight growth, initially. TRW estimates that the value of weight

margin increases with decreasing margin, from zero

at 15 percent margin, to $5,000 to $10,000 per pound

at 5 to 10 percent margin, to $40,000 per pound at O

percent margin, to as much as $100,000 per pound

at negative margin, depending on the time remaining

before launch (figure 2-l). Margins are seldom

negative at an early design stage; if margin becomes

negative shortly before the spacecraft is to be

launched, expensive redesign may be required to

reduce the spacecraft weight below the maximum

weight specified in the contract, and TRW estimates

that it maybe worth up to about $100,000 per pound

to avoid this. This is much greater than the cost of

transporting a typical spacecraft to its operational

orbit: about $15,000 to $30,000 per pound to

geosynchronous orbit, as estimated by TRW.3

TRW notes that spacecraft designed to incorporate advanced-technology subsystems can consume weight margins with unusual rapidity, leading

to redesign (which may require 3 or 4 months),

increased risk, and possibly additional redesign, as

well as downward revision of ‘requirements. Two

Figure 2-l—Value of Weight Margin

Marginal value of margin ($1,000 /lb.)

100

r

--- ~

80 -

I

I

\

60 j ‘,

I

-- GEO Iaunch (Titan IV + IUS)

4 0 I

‘“’d --’~-- LEO launch (Titan IV)

o

+ —--=k.—— —

— — -— — ~––- — ——T– —===+ —

- 5 0 5 10 15 2 0

Weight margin (percent)

SOURCE: TRW (plotted by OTA).

1P. Hillebrandtet al., space Diviswn Unmanned Space Vehicle Cost Model, Sixth Edition, SD TR-88-97 (Los Angeles Am, CA: Headquarters, SPace

Systems Division, U.S. Air Force Systems Command, Nove~r 1988), p. VIII-6; distribution limited to U.S. Government agencies only.

~.W, Elverum, Jr., “Reliability Up/Coats Down Through Simplicity,” Space Systems Productivity and Man@acturing Conference IV (El Segundo,

CA: Aerospace Corp., 1987), pp. 175-1%.

Sclmk ~rby, TRW, prsonal communication, Nov. 16, 1988.

-7–
8 ● Affordable Spacecraft: Design and Launch Alternatives

redesign cycles totaling 6 to 8 months could lead to

a 10 to 25 percent cost growth. TRW also notes that

volume constraints are also costly.

How much margin should be reserved? Reserving

too little increases the risk of delay and cost overrun,

should redesign become necessary, because weight

growth exceeds expectations late in a development

program. Yet reserving too much margin imposes an

opportunity cost: one either foregoes the opportunity

to add extra fuel or equipment, and hence capability,

to the payload, or one foregoes the opportunity to

make the payload less expensive by allowing it to be

heavier. In the latter case, the opportunity cost is

tangible and can be estimated.4 The optimal margin

will be that at which the marginal opportunity cost

of not making the payload a pound heavier equals

the marginal value that the additional margin would

provide by reducing risk of cost overrun and

schedule slippage.

If the marginal value of weight margin is as

estimated by TRW (see figure 2-1), then a margin of

11 to 12 percent would be optimal for a Titan IV

payload. 5 This is comparable to the margin (15

percent) that Public Law 100-456 required the

Department of Defense to reserve for satellites DoD

approves for development in fiscal year 1989. In

reporting on the National Defense Authorization Act

for fiscal year 1989, the Senate Committee on

Armed Services proposed requiring that

. . . the Under Secretary of Defense for Acquisition

shall not approve for development a new satellite if

the proposed payload weight exceeds 85 percent of

the lift capacity of the launch vehicle(s) identified

with the proposed satellite, and shall not approve for

development a block change if the proposed payload

weight exceeds the weight of the existing payload.b

This language was endorsed in conference and

signed into law.8

The expectation that increasing weight margins

would reduce cost risk does not imply that it would

save money to build a new launch vehicle large

enough to launch the heaviest payloads with increased weight margins. Predicting whether it would

save money would require comparing the cost of

developing the vehicle with the total benefits it

would provide—including the reduction in cost risk

that increased weight margins would provide. Ironically, building larger launch vehicles could increase

this risk: Without discipline, payloads might still be

designed to allow little margin, and margins could

still be tight or negative, but with greater consequence. To eliminate a l-percent negative margin,

one would have to trim 2,000 pounds from a

200,000-pound payload, compared to only 390

pounds in the case of a 39,000-pound payload. In

some cases this could be done by carrying less fuel

than planned and accepting reduced payload performance. If not, reducing the dry weight of a

200,000-pound payload by 1 percent would cost

$240 million more than reducing the dry weight of

a 39,000-pound payload by 1 percent, if weight

reduction costs $150,000 per pound. This illustrates

why allowing margin for, and controlling, weight

growth will be much more important for large

payloads (such as proposed heavy-lift launch vehicles could carry) than for smaller ones.

4sw CCA p~~e~c Analysis” in ch. 3.

Sviz. a paylo~ that would cost between $100 million and $5 billion if built to weigh 39,000 pounds-i.e., to reserve no mwgin.

6s. Rept. 100-326, P. 36.

TH. Rept. I(K)-989, P. 282.

8S. ReptO 100-326” did not Swify ~he~er~e [es~ate o~ propo~ Satel]ite weight should inc]ude ex~t~ weight growth in addition tO the IIOIIlllld

weight estimated from the design for the satellite. If so, Public Law 100-456 would require at least a 15 percent weight margin in addition to whatever

weight growth is expected. If the payload consisted solely of an unfueled satellite, 25 percent weight growth would be expected [P. Hillebrandt et al.,

op. cit., footnote 1, pp. VIII-6], so Public Law 100-456 would require a 40 percent weight margin in addition to the nominal weight estimated from the

design for the satellite.
Chapter 3

Fatsats

Many experts find it plausible that a payload

could be designed to perform a function at lower cost

if it were allowed to be heavier. However, there have

been few attempts to estimate just how much cheaper

payloads could be, if allowed to be heavier, or how

much they should weigh in order to minimize the

total cost of producing and launching a payload.

The first section of this chapter discusses one of

several ways in which payloads could be made less

expensive if allowed to be heavier: standard subsystems could be used in lieu of customized subsystems

designed to minimize weight. The next section

describes how a high-altitude satellite could be

allowed to be heavier (and hence less expensive)

without using a larger launch vehicle: by using an

orbital transfer vehicle with electric engines (e.g.,

arcjets). The third section discusses estimation of

cost versus weight trade-offs, with subsections

describing parametric and bottom-up methods and

comparing parametric with bottom-up estimates.

The final section discusses some organizational

obstacles to reducing cost by allowing payload

weight growth.

STANDARDIZING SUBSYSTEMS

One of several ways to trade off cost for weight is

to use standard spacecraft subsystems or busses.2

The use of a standard or previously developed

subsystem may result in a heavier spacecraft but

allow a satellite program to avoid paying part or all

of the substantial nonrecurring costs of developing

a custom subsystem. In addition, because of learning

and production-rate effects, it helps reduce the

recurring cost of producing the standard subsystem.

Using a standard subsystem could reduce subsystem cost by a factor roughly equal to 1 plus the ratio

(expressed as a Ii-action) of nonrecurring to recurring

cost. For example, the ratio of nonrecurring cost to

recurring cost is typically 2/1 for a spacecraft bus

(see table 3-l), so the nonrecurring cost of developing a spacecraft bus is about twice the recurring cost

of producing a bus. If a mission requires one

spacecraft, the cost of developing and producing a

custom bus would be about three times the cost of

producing a suitable previously developed bus. By

using a previously developed bus, one could therefore save about two-thirds of the cost of a customized bus.3 The cost would be reduced by a factor of

3, i.e., by 1 plus the ratio of nonrecurring cost to

recurring cost (2/1).

More could be saved on subsystems with higher

ratios. For example, the cost of structure, which has

ratios ranging from 5/1 to 8/1, could be reduced by

a factor of at least 6, and possibly 9.

The amount saved could be a small percentage of

total spacecraft program costs, which also include

the costs of mission-peculiar payloads and program

overhead, etc. A 1972 Lockheed study4 concluded

that development and use of standard subsystems

could save only about 4 percent of the cost of 91

payload programs, when used in addition to low-cost

design methods and payload refurbishment. The

savings attributable to standardization would be

about 6 percent of the program costs already reduced

by low-cost design and refurbishment.5

Manufacturers have estimated that 95 percent

learning might be achieved, i.e., every time the

cumulative number of units produced is doubled, the

incremental unit cost would decrease 5 percent.6 The

Air Force has assumed 95 percent learning in

estimating first-unit production costs from lot sizes

ISW, e.g., LoC~~MiSSileS and Space Co., t~acrOfhW-CoStRefwbiStile and Standard Spacecraft Upon Future NASA S@zce Progmms, NTIS

N72-27913, Apr. 30, 1972.

2A ~ac=r~’’bm’ consists of hose spacecr~t subsystem~.g. structure, thermal control, telemetry, attitude control, Power, and PmP~sion—that

are not peculiar to a particular mission, as cameras or radio relays would be.

3The cost of ~~~at~g ~ Off..the-shelf subsystem into a spacecr& is sm~l. The Boeing comp~y’s p~~etric Cost Model predicts that the cost

of integrating an off-the-shelf subsystem into a spacecraft would be about 3 percent of the cost of designing anew subsystem for the spacecraft. [Boeing

Aerospace Co., May 1989]

o~c~~ Missiles and Space Co., op. cit., footnote 1.

5~ his study, ~c~eed ~med a ‘ ‘mission m~el’ now recowi~ w highly i~ated; this probably 1~ to overestimation of the potential savings

from standardization. On the other hand, the percentage savings atrnbutable to standardization might have been greater if refurbishment had not been

assumed.

6F.K. Fong et & um~~ed spacecr@ cost Model, s~ ~., sD.’l’’slAsAs (~ Angeles AF’B, CA: He~qu~ers, space Systems Division, U.S.

Air Force Systems Command, June 1981); NTIS accession number AD-B060 824L; disrnbution limited to U.S. Government agencies only.

-9-
10 ● Affordable Spacecraft: Design and Launch Alternatives

Table 3-l-Ratios of Nonrecurring to Recurring Cost

of Spacecraft Subsystems

If a previously developed subsystem can be used in a new spacecraft

in lieu of a custom-designed subsystem, subsystem cost could be

reduced by a factor roughly equal to one plus the ratio of nonrecurring

to recurring cost-e.g., threefold for a spacecraft bus (a spacecraft without its mission payload).

Ranges of ratios

nonrecurring Cost/

Subsystem recurring cost

Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5/1 to 8/1

Propulsion (apogee kick) . . . . . . . . . . . . . . . 2/1 to 6/1

Thermal control . . . . . . . . . . . . . . . . . . . . . . . 4/1 to 40/1

Attitude control . . . . . . . . . . . . . . . . . . . . . . . . 1/1 to 2/1

Electrical power . . . . . . . . . . . . . . . . . . . . . . . 1/1 to 2/1

Telemetry, tracking, & command . . . . . . . . . 2/1 to 3/1

Spacecraft (less mission payload) . . . . . . . . about 2/1

Communications mission payload . . . . . . . . 2/1 to 3/1

SOURCE: U.S. Air Force Systems Command, Space Systems Division.

and prices,7 but there have been too few buys of each

type of U.S. spacecraft to demonstrate learning

conclusively.

UPPER STAGES AND ORBITALTRANSFER VEHICLES

Often a payload has an upper stage or an

orbital-transfer vehicle (OTV) in addition to a

spacecraft. Some analysts have considered options

for reducing the costs of upper stages by allowing

them to be heavier.8 Others are considering the

opposite approach: making OTVs smaller-and

perhaps more expensive--in order to save money by

using a smaller launch vehicle, allowing the spacecraft to be larger, or providing more margin for

spacecraft weight growth. Space-based OTVs have

also been proposed;9 they could be reused and would

not be launched together with the spacecraft, which

could therefore be larger. However, refueling and

maintaining them could be complicated and might

require the development and maintenance of costly

space- or ground-based infrastructure.

Electric propulsion could be used to make OTVs

smaller and lighter while at the same time increasing .

the mass they could deliver to a high orbit. They

could use photovoltaic (“solar”) cells to generate

electricity10 to power an electrostatic ion thruster, an

arcjet thruster, or an electric engine of some other

type (many are possible) that has an exhaust velocity

much greater than that of a chemical rocket. This

would reduce the mass of fuel required for orbital

transfer, increasing the payload that could be transferred. The “dry weight” of an electric OTV

(EOTV) could also be smaller than that of a chemical

OTV of comparable total impulse, further increasing

the payload that could be transferred.

There would be drawbacks. An EOTV would

produce little thrust, so transfer of a payload from a

low-altitude parking orbit to geostationary orbit

might take 3 to 6 months. Before reaching its

destination orbit, the payload would age a few

months and, more important, might degrade because

of its longer transit through the Van Allen radiation

belts. An EOTV would be designed to tolerate such

a transit, but some satellites might not be. A

near-term solution would be to shield sensitive

satellites against the radiation, but this would reduce

the maximum satellite mass that could be carried.

The longer transfer time could also be detrimental

to security. A military satellite on an EOTV would

remain longer at low altitude and within range of

low-altitude anti-satellite weapons than if it rode a

conventional OTV to its destination orbit. If a

critical satellite fails or is damaged, an EOTV might

not be able to replace it without a serious lapse in

mission performance.ll These drawbacks could be

mitigated by launching military satellites on schedule rather than on demand. That is, near the end of

its projected useful life, each satellite would be

deactivated, maintained as a spare, and replaced by

a new satellite, which could have been launched

months earlier.

The Air Force Systems Command Space Systems

Division estimates that EOTVs would be more

economical than conventional OTVs for selected

7p. Hillebrandtet d., space Division unrnan~dspace Vehicle Cost Model, Sixth Edition, SD TR-88-97 (Los Angeles A.FB, CA: Headquarters, SPace

Systems Division, U.S. Air Force Systems Command, November 1988); distribution limited to U.S. Government agencies only.

SD~i ~er, “why Spacwraft Should Get Less Expensive If Launch Costs Decrease, ” unpublished, undated, marked ‘‘ D582-1(X)03-1. ” Boeing

presented results of this analysis at the ALS Phase 1 System Requirements Review, and to OTA in a briefing, “ALS program Development. ” See also

Hughes Aircraft Co., Space and Communications Group, Design Guide for ALS Payloads, October 1988.

gE.g., see ~ckh~d Missiles ~d Space CO., Fiwl Report---payload Eflects Analysis Study, LMSC-A990556 (Sunnyvale, CA: June 1971)! NTIS

accession number N71-3749630; and J.M. Sponable and J.P. Penn, ‘‘Electric Propulsion for Orbit Transfer: A Case Study, ’ Journal Of PrOPU/SIOn ad

Power, vol. 5, No. 4, July-August 1989, pp. 445451,

IONucle~ Wwer was once considered more promising; see Imckheed Missiles and Space CO., Op. Cit., fOOmOte 9.

llHowever, in ~me scenarios, neither could a conventional OTV.
Chapter 3-Fatsats ● 11

missions. For example, a 5,250-pound spacecraft

could be launched to geostationary orbit either on a

Titan IV launch vehicle with an Inertial Upper Stage

for about $250 million or on a Delta II launch vehicle

with a solar-electric OTV for about $124 million,12

saving about $126 million. A Titan IV with an

EOTV could launch a 30,000-pound spacecraft to

geostationary orbit for an estimated cost of $269

million. The Space Systems Division is planning to

demonstrate an expendable solar-powered EOTV,

probably with an experimental payload, sometime

between 1993 and 1995.

ESTIMATING POTENTIAL

COST REDUCTION

There have been few attempts to estimate how

much cheaper spacecraft could be, if allowed to be

heavier, or optimal weights for minimal production

and launch costs. Analyses of historical data show

that heavier spacecraft are typically costlier than

lighter spacecraft;l3 usually, however, heavier spacecraft are also more capable than lighter spacecraft.

They perform more functions, more difficult functions, or similar functions better.14 So these analyses

do not answer the important questions:

1.

2.

How heavy should a payload15 be in order to

minimize the combined costs of payload

production and launch on a currently operational launch vehicle? Are payload weights

actually optimized for current launch vehicles?

How heavy should a payload be in order to

minimize ‘combined production and launch

cost, if the launch cost per pound of payload

were reduced, or the maximum payload weight

that could be launched were increased, by

some factor? By what factor would total

payload production and launch cost be reduced?

Answering these questions requires comparing

the costs of heavy payloads and light payloads that

perform the same function equally well. Unfortunately, such data do not exist; there are no two

payloads, one large and the other small, designed at

the same time (and hence with comparable technology available) to perform the same set of functions

equally well. The few analyses that have estimated

how much cheaper a payload could be if allowed to

be heavier have been hypothetical. They are based

on both “bottom-up” and parametric estimates.

Bottom-up estimates are obtained by designing

two or more versions of a payload to perform the

same functions at minimum cost without exceeding

weight or size limits, which differ from version to

version. For example, two versions of a communications satellite could be designed: one to be launched

on a Scout launch vehicle, the other on an AtlasCentaur. Each version would be designed to minimize production and launch cost. Comparing the

costs of the two different versions would indicate

how much less expensive the larger version would

be.

Bottom-up estimates are time-consuming and

expensive to derive, and there may be no basis for

assuming the cost-versus-weight trade-offs derived

would apply to versions larger or smaller than those

designed or to payloads that must perform different

functions. For example, there is no rationale for

expecting that bottom-up estimates of costs and

weights of communications satellites of comparable

capability could be used to illustrate cost-versusweight trade-offs for remote-sensing satellites of

comparable capability.

Parametric estimates are obtained by assuming

that if the weight of a payload were allowed to

increase, the minimum cost at which it could be built

would vary in some qualitative way-e. g., approach

a limit, or decrease exponentially. The parameters of

the relationship-e.g., the minimum costs for particular weight limits—are chosen to make the hypothetical relationship fit historical cost and weight

data, bottom-up cost and weight estimates, or both,

as well as possible.

The fit will not be exact, however;l6 it may be a

good fit on the average, with some payloads costing — - - —

lz~cluding $3 million for RDT&E.

13P. Hillebrandt et al., op. cit., footnote 7, and figure 4-1 of Lockheed Missiles and Space CO., Op. cit., foomote 9.

14s= ~ckh~ Missiles and Space Co., op. cit., footnote 9.

15Thepq/o~ of a lawch vehicle may ~ a ~acar~t (e.g., a satellite or plane~ probe) toge~er wi~ an upper stage (to propel it to a transfer orbit

or escape trajectory) and support equipment for attaching them to and releasing them from the launch vehicle, Some launch vehicles (e.g., the Space

Shuttle and sounding rockets) sometimes carry payloads (e.g., scientific instruments) that remain attached to the launch vehicle.

16u~e= mere ~s. many Pwmeters (i.e., statistic~ degrees of fi~om) in tie model mat tie av~lable data ~e too few to estimate them with

statistical significance.
12 ● Affordable Spacecraft: Design and Launch Alternatives

more, and others less, than predicted by the relationship (the “model”). The fact that the model

represents only average costs and cost-versusweight trends may be an advantage in studies, such

as this one, which focus on such averages and trends.

However, it would be a limitation if accurate

cost-versus-weight trade-offs for a specific mission

were required. If the functions to be performed by

the payload were specified in detail, and if resources

permitted detailed engineering design of several

alternative versions of different weights or sizes,

then bottom-up estimation could be used and should

provide greater accuracy.

Thus bottom-up estimation and parametric estimation are complementary approaches to cost estimation.17 Neither approach by itself would be of

general value: a bottom-up estimate is applicable

only to a specific payload, and parametric estimates

are abstract and hence useless unless fitted to

bottom-up estimates of cost and weight.18

The bottom-up and parametric approaches are

exemplified by two analyses produced almost two

decades ago: a parametric model developed by Carl

Builder at the Rand Corp.,19 and a bottom-up

analysis by Lockheed Missiles and Space Co.20 Both

were cited in congressional debate on the merits of

the Space Shuttle.21

A Parametric Analysis

Builder did not estimate cost reductions for

particular payloads; he described a procedure for

doing so using data or assumptions about payload

cost and launch cost. He assumed average launch

cost would vary as the payload capability of the

launch vehicle raised to some power A (this is called

a “power-law” relationship) and payload cost

would vary as the payload weight raised to some

power B. To use Builder’s model, one must specify

the exponent A and the initial cost and weight of a

payload designed to minimize payload plus current

launch cost. One need not specify what the payload

does; in theory, it should make no difference. If

launch costs are reduced by some factor,22 the

optimum weight and the cost of a functionally

equivalent payload designed to minimize payload

plus reduced launch cost may be calculated using

formulas derived by Builder. The difference between the old and new minimum total costs is the

savings obtainable if launch costs are reduced by the

specified factor, assuming payloads are reoptimized

to take advantage of the new launch costs.23

OTA derived an estimate of the exponent A used

in Builder’s model by fitting a straight line to points

on a log-log plot of the payload capabilities24 and

average launch costs25 of Delta, Titan, and the Space

Shuttle. Figure 3-1 shows the three points and the

line obtained as a least-squares fit to the points. The

slope of the line corresponds to a value of 0.74 for

the exponent A, implying that average launch cost

would increase by two-thirds if the payload weight

were doubled.

Figure 3-1 also shows a point representing the

predicted payload capability of the Pegasus airlaunched vehicle (see figure 4-6) and the price

charged by its operator, Orbital Sciences Corp.

(OSC), for the launches and launch options purchased by the Defense Advanced Research Projects

Agency. The line fitted to Delta, Titan, and Shuttle

costs and payload capabilities accurately predicted

the cost (to the government) of a Pegasus launch.

If launch cost is assumed to vary with payload

weight in the way described above,26 Builder’s

model predicts that a payload that would cost $1

billion if designed to weigh 39,000 pounds could be

17u.s. @gess, office of Technolo~ Assessment, Reducing Launch Operations Costs: New Technologies and Practices, OTA-TM-ISC-*8

(Washington, DC: U.S. Government Printing Office, September 1988), app. A.

18As no~ a~ve, hem is seldom more than one historical “data point” for the cost and weight of fictionally identical SPaceraft.

19cwl H. BUil&, Are ~~h Vehicfe c’05f5 a Bottle~~k t. ECO~rniCa/spuce operatio~.? Rand Working document D-19482-PR, December 1969.

%ckheed Missiles and Space Co., op. cit., footnote 9.

21s= ~erem~s of Senatm Mond~e in tie CongresswMl Recor#e~te, May 26, 1971, pp. 171~-]7106, and tie rem~ks of Senator Anderson

in the Congressional Record-Senate, Apr. 20, 1972, pp. 13786-13790.

22BY tie me factor, regardless of payload weight.

~More P=isely, “mw~g paylo~ we optim~]y si~ for minimum total cost in both cX.WS.” Builder, OP. cit.! foomote 19.

24T0 a ~oonemi.-hi~ Ofiit inc~ined 28.5 de~s. U.S. Congess, Office of T~hno]ogy As~ssment, ~nch Optzonsfor the Future: A &4yer’s Gw”de,

OTA-ISC-383 (Washington, DC: U.S. Government Printing Office, July 1988), table 2-1.

2S~cluding ~m~ fix~ amu~ ]aMch co~t, OTA ~~~ launch at tie m~imum rate es~at~ in table A-1 of ibid. and usd the nominal

cost-estimating relationships in that table.

~Ioee, in ~oP~m t. tie paylo~ capability of the launch vehicle raked to tie Power 0-74.
Chapter 3—Fatsats ● 13

Figure 3-l—Average Launch Cost v.

Payload Capability

Average launch cost (millions of dollars) $l,ooo ~– ---- -- -

~

$100 ~ , ,p - ‘

~

I

,,.F’ “

$10- ,-

1

●-

+ ,,”

$1 ‘“:’-’ “- - ——T–——” — -–-—---

0.1 10 100

Payload’ capability (1,000s of Ibs.)

Power-law fit ● Delta + Titan

* Shuttle ■ Pegasus (not fitted)

SOURCE: Office of Technology Assessment and Orbital Sciences Corp.,

1988.

made for 5.8 percent less if it were allowed to be

twice as heavy. More generally, if allowed to be

heavier or designed to be lighter, the cost of such a

payload would be proportional to its weight raised to

the power –0.086 (the exponent B in Builder’s

model). Figure 3-2 shows how the payload cost and

the launch cost would vary with the payload weight.

Designing the payload to weigh 39,000 pounds

would minimize the total cost.

Figure 3-3 shows how the payload cost and the

launch cost would vary with the payload weight if

launch cost were reduced by a factor of 3—i.e., by 67

percent. It illustrates that total cost is insensitive to

weight for weights between 80,000 pounds and (at

least) 200,000 pounds. The optimal weight would be

about 150,000 pounds. Reducing the launch cost by

67 percent would reduce the total cost by only 11

percent. It should be emphasized that this is the

estimated cost reduction achievable by allowing

payload weight to grow without changing payload

performance. It assumes that the baseline payload

was designed to minimize total cost. If a baseline

payload was not designed to minimize total cost,

redesigning it (possibly to weigh more) could save

money even if launch costs are not reduced. If launch

costs are reduced, additional savings could be

obtained by allowing weight growth; these additional savings are the savings estimated by Builder’s

model.

Figure 3-4 shows these results in a different form

along with results of similar analyses of other

hypothetical payloads initially weighing 39,000

pounds but costing $2 billion, $3 billion, and $4

Figure 3-2-Cost for Hypothetical Mission With

Current Launch Cost Trend

Cost (millions of dollars) $1,400 ;–

$1,200 “

$1,000-,

$8001

$600

$400

$200 ~

$0 , -— — —

10 30 50 70 90 110 130 150 170 190

Payload weight (1,000s of Ibs.)

r- —- Payload m Launch

SOURCE: Office of Technology Assessment, 1990.

Figure 3-3-Cost for Hypothetical Mission if Launch

Costs Are Reduced 67 Percent

Cost (millions of dollars) $1,400 ~-- --—–-1

$1,200 

$1,000

$800 -

$600 -

$400 -

$200 -

$ 0 --—~

10 30 50 70 90 110 130 150 170 190

Payload weight (1,000s of Ibs.)

~ Payload _ Launch

SOURCE: Office of Technology Assessment, 1990.

billion. Figure 3-4a shows how these costs could be

reduced, according to Builder’s model, if the

weights were allowed to grow. For each weight,

figure 3-4b shows the (“economic’ launch cost at

which that weight would be optimal. These estimates predict that allowing the weight of a Titan-IVclass payload to increase by 400 percent (for

example) would reduce payload cost by an amount

that is nearly the same for a $1-billion payload as for

a $4-billion payload. The estimates also predict that

the lower the initial cost of a payload, the more the

cost per launch must be reduced to justify increasing

its weight by a large factor.

Builder’s assumption about the relationship between payload weight and launch cost would lose

. .
14 ● Affordable Spacecraft: Design and Launch Alternatives

Figure 34-Payload Cost v. Weight Trade-offs and —

Payload $5

$4

$3

~

$2

$1-

$0 ~. ——

0

Economic Launch Costs

cost (billions of dollars) A — —. — —

——. . —. I

-–- — — ..7——.—— —

50 100 150 200

Payload weight (1,000s of Ibs.)

Economic launch cost (millions of dollars) B

$120,

$115-

$110-

-

$105- - d

+ –

$100 L—

0

‘ \ -—–-..+

Payload 1 ‘\~

Payload 2

Payload 3

‘ .l ,

= . -.

Payload 4 -=., ----- -.. .

. —= -— —-r–— ~

50 100 150 200

LV payload capability (1,000s of Ibs.)

SOURCE: Office of Technology Assessment, 1990.

validity if extrapolated to extremely heavy payloads. 27 It would imply that the average cost of

launching a pound of payload could be made as low

as desired--even lower than the cost of the fuel

required to launch a pound of payload—by building

a launch vehicle of sufficiently large payload

capability. But his assumption fits the estimates of

Pegasus, Delta, Titan, and Shuttle launch costs in

figure 3-1 very well. The cost-versus-payload curve

fitted to Delta, Titan, and Shuttle launch costs

predicted the cost of Pegasus accurately, even

though the payload capability of Pegasus is eightfold

smaller than that of the smallest vehicle (Delta) on

which the curve is based. The curve could probably

be extrapolated with comparable validity to a

payload capability of 200,000 pounds.

Similarly, Builder’s assumption about the relationship between payload weight and payload cost

would also lose validity if extrapolated to extremely

heavy payloads.28 It would imply that they could be

built at a lower cost per pound than that of the bulk

structural material (e.g., aluminum) from which they

are made. However, this is not a problem for the

ranges of payload weights and costs-shown in figure

3-4.

Bottom-Up Analyses

Lockheed used bottom-up analysis to estimate

how much the cost of building, launching, and

operating selected payloads could be reduced by

making them larger29 and by other measures.30 Lockheed considered three payloads, selected to span a

range of costs, that had been built and launched and

for which design and cost data were available. The

least expensive was the Lockheed P-1 1 subsatellite,

which could be modified for use as a Small Research

Satellite (SRS). The most expensive was the Orbiting Astronomical Observatory, the redesigned version of which was designated OAO-B. The other

was the Lunar Orbiter, which could be modified for

use as a Synchronous Equatorial Orbiter (SEO), four

of which could perform Earth resources observation.

Lockheed estimated how much the costs31 of the

OAO-B and SEO payload programs could be

reduced if the payloads were redesigned to be

launched on unmanned expendable launch vehicles

or to be launched on a (then) proposed version of the

Space Shuttle.32 The savings estimated for the first

 A is less than one.

 B is less than one.

 t. for 

trade-off curve for each subsystem was a which, for a minimum cost per pound and, at the other

extreme, approached the minimum weight achievable.

 refinishing in orbit or retrieving to repaired or on 

 development, testing, if the Shuttle 

used, refurbishment) of satellites.

 considered two versions.
Chapter 3—Fatsats ● 15

case were attributed to payload growth.33 However,

Lockheed did not redesign the baseline versions of

OAO-B and SEO to minimize cost without exceeding the baseline weight, hence Lockheed attributed

to weight growth some cost reduction that should

have been attributed to improved design. The

amount of cost reduction misattributed to weight

growth cannot be determined from Lockheed’s

report, so the cost reductions Lockheed attributed to

weight growth should be considered upper bounds

on cost reductions achievable by allowing weight

growth.

Figure 3-5 compares Lockheed’s estimates of the

weights and the average unit costs34 of the baseline

OAO-B and SEO with Lockheed’s estimates of the

weights and costs of “low-cost” versions designed

to be launched on expendable launch vehicles. The

potential savings in fiscal year 1988 dollars would

be $10.1 million (21.3 percent) and $43 million

(15.2 percent) for SEO and OAO-B, respectively.

The estimated weight growth required to achieve

such savings would be 170 percent and 69 percent,

respectively.

A more recent bottom-up analysis by Boeing

Aerospace Co. estimated the cost of a “typical”

payload could be reduced by a large percentage if

weight growth by a modest percentage were allowed, and that it would save money to allow such

weight growth if launch cost per pound were

reduced.35 For example, Boeing estimated the cost

could be halved if the weight were allowed to grow

30 percent (see figure 3-6).36

Boeing actually considered a payload consisting

of an upper stage and a hypothetical spacecraft using

specific subsystem technologies and with subsystem

weights in an assumed ratio.37 Boeing claimed the

hypothetical spacecraft was typical, implying that

similar cost reductions could be expected for other

types of spacecraft. However, some of the subsystem

technologies Boeing assumed for the spacecraft

were atypical. For example, the analysis estimated

Figure 3-5-if a Satellite Were Allowed To Be Heavier,

It Could Cost Less

Average cost (FY 1988, millions of dollars) 1,000 ‘- —

I

—

t Baseline

1 * — “Low-cost ”

1

Orbiting *

100 Astronomical

Baseline Observatory-B . .Low-Cost ”

➤ .

I Synchronous

t Equatorial

Orbiter

1

10 ‘——-- ‘

-

”

– ,.--~

1,000 10,000

Weight (Ibs.)

SOURCE: Estimates from Lockheed Missiles and Space Co., 1971, plotted

by Office of Technology Assessment, 1990.

Figure 3-6-The Effect of Reducing Launch Cost on

the Optimal Weight and Cost of a Payload

Optimal growth ~ _ —–

1

factor

‘—--.l

--t-+”

0“ - ——— - 1

0 1,000 2,000 3,000

Launch cost (dollars per pound)

4,000

Boeing Aerospace Co. estimated that, if average launch cost per

pound were reduced from $3,600 per pound, it would be

economical to redesign a hypothetical payload to allow it to be

heavier by the weight-growth factor indicated and less costly by

the cost-reduction factor indicated. Boeing assumed that the

original design minimized the sum of payload and launch costs at

a launch cost of $3,600 per pound, and that the average launch

cost per pound does not depend on the payload weight.

SOURCE: Boeing Aerospace Co., 1988.

 savings estimated for attributed to payload growth; the rest was attributed to reduced cost and to 

that only a reusable vehicle such the Shuttle could provide: intact abort capability, on-orbit checkout, repair, and refurbishment. Additional savings

in both cases were attributed to use of improved technology.

 number of satellites launched (6 to 

by OTA.

 Op. cit., footnote 8.

 have misquoted would 

growth. See, e.g., Thomas M. ‘Status of the ALS Program,’ Space Systems Productivity and Conference-V (El CA: The

Aerospace Corp., 1988), p. 30.

 to change when the payload was redesigned with relaxed weight 
16 ● Affordable Spacecraft: Design and Launch Alternatives

the cost and weight of the satellite’s electric power

subsystem by assuming it consisted entirely of solar

cells (which typically dominate the cost of a

satellite’s power subsystem), with no batteries

(which typically dominate the weight of a satellite’s

power subsystem).

Some of the proposed cost-reducing and weightincreasing substitutions Boeing proposed--e. g., substitution of commercial-grade solar cells for spacecraftgrade (S-class) solar cells—would decrease payload

reliability and expected lifetime to an extent not

estimated by Boeing. Thus Boeing did not estimate

weight-versus-cost trade-offs for equal reliability;

some of the savings Boeing attributed to weight

growth should have been attributed to reduced

reliability .38 Therefore, the savings Boeing attributed to weight growth may be upper bounds for

spacecraft of the type Boeing considered.

Another recent bottom-up analysis estimated

cost-versus-weight trade-offs for some subsystems

but not for complete payloads.39 Like earlier studies

by Lockheed40 and the Aerospace Corp.,41 it identified payload “cost drivers"—i.e., costly payload

components or testing-and recommended changes

in payload design, components, testing, or operations that might reduce space program cost. Many of

these changes would require increasing the weight of

the payload (the “fatsat” approach discussed here)

or specifying a simpler or easier mission, or fewer

missions (the “lightsat" approach discussed in

chapter 4).

A Comparison of Parametric and

Bottom-Up Analyses

Estimates differ on how much cheaper payloads

could be if they were allowed to grow to a specified

weight, and how much they should grow if cost per

launch were reduced to a specified amount. One

parametric estimate by OTA predicts that a hypothetical expensive payload as heavy as a Titan IV

could launch could cost about $130 million less if

allowed to be five times as heavy. It would be

economical to design the payload to be so heavy if

it could be launched for less than about $100

million. 42 Less would be saved if the baseline

payload cost were comparable to or less than the

average Titan IV launch cost, estimated hereas$117

million. The only bottom-up estimate that could be

compared to these is one by Boeing, which predicts

much greater savings (at least 76 percent, or $760

million for a billion-dollar payload) but is based on

a conceptual design for a payload that is atypical in

important respects; moreover, the redesigned payload was allowed to use less reliable components,

and launch cost per pound was assumed to be

independent of payload weight.

OTA also derived parametric estimates to compare to detailed bottom-up estimates by Lockheed

for two spacecraft. Lockheed estimated 66 percent

greater savings for one spacecraft (SEO), and 360

percent greater savings for the other (OAO-B), but

attributed to weight growth some savings that should

have been attributed to optimization of the baseline

designs. However, these discrepancies are comparable to the unexplained statistical variations often

encountered in spacecraft cost estimation. The less

detailed Boeing estimate, if applicable, would predict much greater savings than predicted by OTA:

530 percent more for OAO-B, and at least 630

percent more for SEO, at the weight growth factor

proposed by Lockheed.43

ACHIEVING POTENTIAL COST

REDUCTION

Realizing most of the potential savings predicted

by these estimates will probably require creation of

incentives to dissuade satellite program managers

and designers from adding capability, and thereby

weight, until launch vehicle lift margin, engine-out

sg~ckh~dids. and estimated how many satellites would be required to provide comparable mission performance for 10 years. bckh~d concluded

that use of S-class components, which would maximize expected satellite life and minimize the number of replacements required, would be most

cost-effective.

39Hughes Aircraft Co., op. cit., footnote 8.

%ckheed Missiles and Space Co., op. cit., footnote 9.

blspacecra~ Cost Drivers Study--+inal Report: Phase 1 (El Segundo, CA: The Aerospace COW., October 1983).

bz’’f’he Advanced Launch System is intended to launch such payloads fO1 much less.

b3However, Boe~g estimated that it would not be economical to ~ek the ex~eme cost reduction predicted for SEO unless launch cost per Pourtd were

reduced by a factor of 36.
Chapter 3—Fatsats ● 17

Box 3-A—Fatsats: The Good and Ugly, and the Bad

HEAO—During a budget crunch, NASA took steps to repackage its High-Energy Astronomical Observatory

instruments, designed for two Titan payloads, into three Atlas-Centaur payloads. This reduced launch costs by about

2 x $125K -3 x $60K = $70K in today’s dollars, by TRW’s calculation. It also created extra weight margin. This

allowed designers to design the spacecraft with high safety factors (strength margin etc.), so that they could dispense

with the costly construction and testing of model and qualification spacecraft.l 

Phobos 1 & 2-In July 1989, two science spacecraft, Phobos 1 and Phobos 2, were launched from the Soviet

Union toward Phobos, the larger of the two moons of Mars. Their busses were designed and built in the Soviet

Union, as were some of their instruments; other instruments were designed in Austria, France, Sweden, Switzerland,

West Germany, and several East European nations participating in Project Phobos. Some of the Soviet instruments

were designed with generous weight margins. Jochen Kissel, a West German member of the project’s scientific

council, said, “We could use standard printed circuit boards rather than ultraminiaturized parts . . . It made

everything cheaper and simpler.

Nevertheless, because of greater-than-expected weight growth, some instruments were removed from one or

the other spacecraft. The two spacecraft were originally intended to carry identical suites of instruments, so that

Phobos 2 could perform all functions of Phobos 1 in case Phobos 1 failed, or vice versa.

As it happened, Phobos 1 did fail. More accurately, contact with Phobos 1 was lost late in 1988, because an

erroneous command was transmitted to Phobos 1 from the ground. To compensate for the loss, mission directors

planned to command Phobos 2 to rendezvous with Phobos rather than proceeding to the smaller Martian moon,

Deimos, as it would have had Phobos 1 succeeded. Phobos 2 lacks the radar mapper, neutron spectrometer, and solar

x-ray and ultraviolet telescopes of Phobos 1, but carries an infrared spectrometer and a hopping lander which Phobos

1 lacks.3

Ironically, contact with Phobos 2 was also lost on March 27, 1989, about two weeks before the planned

encounter with Phobos was to occur (on April 9-10).4 Nevertheless, Phobos 2 gathered a significant amount of data

before this failure. This anecdote illustrates the value of generous weight margins on some instruments, the cost of

negative weight margins on other instruments and on the spacecraft as a whole, the value of redundant spacecraft,

and the risks of human errors and compound failures.

Milstar-Milstar is an advanced communications satellite being built for the Department of Defense. A few

will be built—fewer than originally planned—allowing nonrecurring program costs to be amortized over a few

satellites. Aside from this economy, Milstar appears to exemplify the antithesis of the fatsat philosophy: it is

designed to be large in order to cram it with capability, not to reduce its cost. It has had to be redesigned at least

once to reduce its estimated weight and add margin. Costly edge-of-the-art technologies have been adopted to reduce

weight, and some are so risky that additional greater-than-expected cost and weight growth could occur.

Milstar was made fatter to be better, not cheaper, and its gross weight kept growing. A subsystem designer

quipped, “Milstar is going to gross everybody out. ”

ITRW, briefig, NOV. 16, 1989. See also Science, vol. 199, Feb. 24, 1978, p. 869, and Astrophysical/Journal, vol. 230, June 1, 1%’9, P. 540.

2EI-iC J. hrner, “Mission to Phobos,” Aerospace America, September 1988, pp. 34-39.

31bid., and “Phobos 1 Loss to Change Mars MiSSiOn,” Aviation Week and Space Technology, Oct. 3, 1988, p. 29.

4C6Soviets IJJSe Contact With Phobos 2 Spacecraft,” Aviation Week and Space Technology, Apr. 3, 1989, p. 22.

capability, and reliability are reduced, requiring spend less than the amount appropriated. When

expensive weight reduction programs for redress. funding allows, spacecraft purchasers, with rare

The managers of satellite programs funded by exceptions, opt to increase performance rather than

line-item appropriations have little incentive to reduce cost.
Chapter 4

Lightsats

Lightsats are satellites that are light enough to be

launched by small launch vehicles such as a Scout or

Pegasus, or others now in development. Military

lightsats could be designed for wartime deployment

or replenishment from survivable transportable launchers to support theater commanders. Civil lightsats,

and some military ones, would not require transportable launchers; they could be launched by a wider

variety of launch vehicles, including larger ones.

The first Soviet and U.S. satellites were lightsats,

according to this definition. Explorer I, the first U.S.

satellite, weighed only 31 lb but collected data that

led to the discovery of the Van Allen radiation belt.

But in another sense, the first satellites were fat—as

heavy as early launch vehicles could launch. As

larger launch vehicles were developed, larger, more

capable satellites were developed to ride them.

Nevertheless, small satellites continue to be

launched for civil and military applications that

require only simple functions.

Interest in lightsats has grown recently,l partly in

anticipation of new rockets designed to launch them

at low cost, and, in the case of military lightsats,

because of a desire for a survivable means of

launching satellites-e.g., transportable launch vehicles too small to launch large satellites.2

A few years ago the Defense Advanced Research

Projects Agency (DARPA) began to examine

lightsats, initially to demonstrate the ability of

simple and inexpensive satellites to perform simple

but useful tasks, and, more recently, to demonstrate

the utility of satellites small enough to be launched

from transportable launchers to support theater

commanders during a war. DARPA is considering

concepts for several types of lightsats--for communicaions, navigation, radar mapping, and targeting. The

Army, Navy, or Air Force may choose to procure

similar lightsats for operational use. They would be

designed to be affordable as well as small, because

many might be needed for replenishment after

attrition. “With lightsat, we can undoubtedly put up

satellites for less money than it will cost the Soviets

to shoot them down,’ said Dr. John Mansfield,

while Director of DARPA’s Aerospace and Strategic Technology Office.3

SPACECRAFT CONCEPTS

The first lightsat developed by DARPA’s Advanced Satellite Technology Program was a communications satellite, the Global Low-Orbit Message

Relay (GLOMR; see figure 4-l). Weighing only 150

lb, GLOMR was launched by the Space Shuttle on

October 31, 1985, into a 200-mile-high orbit inclined 57 by degrees.

DARPA has ordered nine more UHF communication satellites from Defense Systems, Inc., the

contractor that built the GLOMR. Seven of these,

called Microsats, will weigh approximately 50 lb

each. These satellites will be launched together into

a polar orbit 400 nautical miles high by the Pegasus

launch vehicle. Once deployed, they will spread out

around the orbit. They will carry ‘‘bent-pipe’ radio

repeaters—i.e., the messages they receive will be

retransmitted instantaneously to ground stations.

The other two satellites will be larger “store/

foreward" satellites called MACSATs (forMultipleAccess Communication Satellite). They will weigh

about 150 lb each and will be launched together on

a Scout launch vehicle. They will store messages

received from ground stations and forward (or

“dump”) them when within range of the ground

stations to which the messages are addressed. The

bent-pipe satellites and the first store/foreward

satellites will cost DARPA about $8 million excluding launch costs.

Amateur (“ham”) radio operators have built a

series of small satellites carrying radio beacons or

repeaters, the first of which, OSCAR I, was launched

in 1961. Since 1969, the nonprofit Radio Amateur

Satellite Corp. (AMSAT) and its sister organizations

worldwide 4 have built or designed several of these

for scientific, educational, humanitarian, and recreational use by hams. AMSAT, which has only one

paid employee, is now building four 22-pound

IS=, e.g., A.E. Fuhs and M.R. Mosier. ‘‘A Niche for Lightweight Satellites,’ Aerospace America, April 1988, pp. 14-16., and Theresa M. Foley,

“U.S. Will Increase Light.sat Launch Rate to Demonstrate Military, Scientific Uses, ’ Aviation Week and Space Technology, Sept. 26, 1988, pp. 19-20.

z~ere is no consensus on whether this is the best approach to assuring continued mission performance in w~ime.

@ot~ by James W. Rawles, “LIGHTSAT: All Systems Are Go,” Defense Electronics, vol. 20, No. 5, May 1988, p. 64 ff.

4AMSAT-Nofi America hm sister organizations in &gentina, Aus~alia, Br~il, Britain, @-many, Italy, Japan, Mexico, and the Netherlands. The

Soviet Union has also launched satellites for amateur radio operators.

–19–
20 ● Affordable Spacecraft: Design and Launch Alternatives

Figure 4-l-The Global Low-Orbit Message Relay

(GLOMR) Satellite

SOURCE: Department of Defense.

(l0-kilogram) communications satellites called Microsats5 (figure 4-2). (These Microsats are unrelated

to the above-mentioned Microsats developed for

DARPA and to the “microspacecraft” discussed in

the next part of this report.) All four satellites are

designed to receive, store, and forward digital

messages using a technique called packet communications. All four use a standard bus, the Microsat

bus. One of the satellites, called PACSAT, is being

built for AMSAT. An almost identical satellite

called LUSAT is being built for a sister organization,

AMSAT-LU, in Argentina. A third satellite, nicknamed Webersat, will carry, in addition to its packet

radio repeater, a low-resolution color TV camera

designed at the Center for Aerospace Technology at

Weber State College in Logan, UT. The fourth

satellite, Digital Orbiting Voice Encoder (DOVE),

was built for BRAMSAT (AMSAT-Brazil). It will

carry a digital voice synthesizer to generate voice

messages that can be received by students using

inexpensive “scanner’ radios.6

AMSAT contracted with Arianespace to launch

these four satellites together for $100,000. The

Ariane 4 launch vehicle will also deploy the

UoSAT-D and UoSAT-E amateur-radio satellites

built at the University of Surrey in England in

addition to the four Microsats and its primary

payload, the SPOT 2 photomapping satellite.7 The

launch, originally scheduled for June 1989, has been

postponed until January 1990, at the earliest.

Microsats are among the smallest communications satellites ever built. They are lightsats, because

they are designed to perform relatively simple

functions. But they are also fatsats, because they are

heavier and larger than they would be if built by

methods usually used for more conventional (and

more expensive) satellites. Assembly of some Microsat subsystems is literally a cottage industry.

Although some printed circuits are being built for

Microsats by a contractor using high-tech methods,

others are being built in the homes of Amateur radio

operators all over the country. When they volunteer

to assemble printed circuits, AMSAT sends them the

instructions.

CONCEPTS FOR PHASED

ARRAYS OF SPACECRAFT

Groups of small satellites could collectively

provide communications or radar capabilities that

could otherwise be provided only by a large satellite.

To do so they would have to operate coherently as

elements of a phased array: all satellites must relay

the signals they receive to a satellite or ground

station that can combine them in a way that depends

on the relative positions of the satellites, which must

be measured extremely accurately. When transmitting, the satellites must all transmit the same signal,

“The AMSAT-NA 73 Amateur Radio, May 1989, 83-84, and Doug and Bob 

 The Next Generation of OSCAR Satellites—Part l,” May 1989, pp. 37-40.

 digital voice synthesizer is similar to the developed at the University of Surrey in England for use on the 

and OSCAR-11 amateur-radio satellites built there.

 Ward, “Experimental OSCARS,” May P. 62 
Chapter 4--Lightsats ● 21

Figure 4-2--Microsat: A 22-lb Communications

(1) Flight model beside Leonid Labutin, UA3CR.

(r) Close-up view. (Copyright Radio Amateur Satellite Corp.)

Photos: Andrew C. MacAIlister, WA5ZIB.

but each satellite must delay its transmission by a

period that depends on its relative position.

The Air Force has considered “placing large

phased arrays in space with major components of the

arrays not rigidly connected to each other” (see

figure 4-3), because “If we can achieve coherence

among these components, phased arrays can be

spread out over very large volumes in space, giving

them an unprecedented degree of survivability. It

therefore may be possible to create a phased-array

device (e.g., a space-based radar) that we can place

into space and enhance simply by adding more

relatively inexpensive elements whenever the threat

increases and budget pressures permit.”8 If small

enough, each element could be launched by a small

launch vehicle.

If a large radar or communications satellite were

divided into several modules, “crosslink” equipment for communications among the modules would

have to be provided, which would add some weight

and cost. Aside from this, historical cost data

indicate that a communications mission payload

might cost less if divided into several smaller

payloads of equal aggregate weight and power.9 The

same might be true of radar equipment. Each module

would need its own bus, but the cost data also

indicate that several small busses would probably

cost less, or no more, than a single bus of equal

aggregate weight.10 Learning and production-rate

effects could make the small modules and busses

even less expensive. Economies of scale in launching might make it economical to launch as many as

possible on a large launch vehicle, but they could be

launched individually on small launch vehicles if

desired-e. g., in wartime, if peacetime launch facilities have been damaged.

Coherent operation of several satellites requires

relative positions to be measured with errors no

greater than a fraction of a wavelength of the

radiation to be sensed or transmitted. The accuracy

required for coherent operation of several satellites

as a microwave radar or radiotelescope has already

been demonstrated.11 Coherent operation of several

 Command, Project Forecast Executive Air Force Base, MD: Headquarters, Air Force

Systems Command, undated). This concept is described in greater detail in the classified report. see PP. to 

of Annex D of vol. IV, which authorized readers may request from the Defense Technical Information Center (accession number 642).

 Division Edition, Angeles AFB, CA: Headquarters, 

Systems Division, U.S. Air Force Systems Command, November 1988); distribution limited to U.S. Government agencies only.

 Ground Antennas Link for Radio Astronomy Observations, ” Aviation Week and Space Technology, Dec. 1, 1986, pp.

32-33.
22 ● Affordable Spacecraft: Design and Launch Alternatives

SOURCE: U.S. Air Force.

satellites as a ladar or optical telescope is beyond the

state of the art. However, someday it may be feasible

to launch several telescope modules, each smaller

than the Hubble Space Telescope, and assemble

them in space (or allow them to assemble themselves) into a rigidly connected phased array12 that

would operate as an optical telescope with better

light-collecting capability and resolution than the

Hubble Space Telescope. If technology advances

further, two or more such arrays, not rigidly connected to one another, could operate coherently to

further increase light-collecting capability and, especially, resolution.13

A scientist at the Jet Propulsion Laboratory (JPL)

of the California Institute of Technology (CalTech)

has proposed a less ambitious phased-may radiotelescope that could be begun today: an Orbiting

Low-Frequency Array of 6 or 7 satellites in a

formation 200 km across (see figure 4-4). It could

map astronomical sources of radio signals with

wavelengths longer than 15 meters; such signals

cannot penetrate the Earth’s ionosphere to reach

ground-based radiotelescopes. The angular resolution of the proposed array could be comparable to

that of a dish antenna 200 km across. JPL estimates

that each satellite would cost about $1.5 million and

weigh less than 90 kg (200 lb) if cylindrical, or 45 kg

(100 lb) if spherical14 They must be launched to a

circular orbit at least 10,000 km high, so the

equivalent weight to low orbit would be about 170

lb for the spherical satellite.

LAUNCH REQUIREMENTS

Operational launchers for military lightsats must

meet several requirements, the most distinctive

being survivability in high-intensity (if not nuclear)

conflict. Such survivability is required of U.S.

strategic and theater missile launchers, and operational lightsat launchers might adopt some of their

features. Like the rail-mobile launcher for the Soviet

SS-24 intercontinental ballistic missile (ICBM), the

similar launcher being developed for U.S. Peacekeeper

(M-X) ICBMs, and the Hard Mobile Launcher being

developed for the Small ICBM (“Midgetman”),

lightsat launchers could pursue survivability

through mobility. They could also employ concealment, as do the Pershing 2 launcher and submarines

(e.g. Trident) that launch ballistic missiles.

Operational military lightsat launchers would

probably be required to launch on short notice and to

sustain higher launch rates than typical space launch

facilities do. On the other hand, lightsat launch

vehicles would be useful with less lift capability than

most launch vehicles have, although DARPA has

said the Scout launch vehicle lifts too little to be

useful as an interim launch vehicle for launching

developmental lightsats, and the Scout does not have

the survivability and launch rate desired for wartime

use. The Pegasus launch vehicle may provide an

innovative means of improving launch flexibility

and survivability.

 Coherent of Modular Imaging Collectors (COSMIC) described by the National Research Council in Space Technology

to Meet Future Needs (Washington, DC: National Academy Press, 1987), p. 39.

“Sub-millimeter Waves, “ in & Micro Spacecraft for Space Science Workshop-Presentations, California

Institute of Technology Jet Propulsion Laboratory, July 6-7, 1988; pp. 136-142.
Chapter 4--Lightsats ● 23

Figure 4-4-Orbiting Low-Frequency Array of Radioastronomy Satellites (Artist’s Concept)

~“loo ‘ M

SOURCE: California Institute of Technology, Jet Propulsion Laboratory.

Civil lightsats and developmental military

lightsats would not require transportable launchers.

LAUNCH SYSTEM OPTIONS

Existing Launch Vehicles

Civil lightsats, or developmental military lightsats,

could be launched-alone or co-manifested with

other payloads-on currently operational launch

vehicles larger than the Scout. If they are small

enough, they could be launched by the Scout.

Air-Launched Vehicles

The communications satellites now being developed by DARPA are very light; several can be

launched on the Pegasus air-launched vehicle (ALV).

The Pegasus ALV is being developed by Orbital

Sciences Corp. (OSC) and Hercules Aerospace

Corp. as a $50 million privately funded joint
24 ● Affordable Spacecraft: Design and Launch Alternatives

Figure 4-5-Pegasus Launch

The Pegasus air-launched vehicle is released from a modified B-52 aircraft operated by NASA’s Dryden Flight Research Facility (artist’s

concept).

SOURCE: Orbital Sciences Corp.

venture. DARPA will pay OSC $6.3 million to

provide the launch vehicle for a government demonstration launch. This price includes neither the cost

of using NASA’s B-52 as an ALV carrier (see figure

4-5) nor the cost of safety support from the Air

Force’s Western Test Range. Pegasus is expected to

be able to launch a 335-kg (738-lb) payload into an

orbit 500 km (270 nmi) high and inclined 25 degrees,

or a 244-kg (537-lb) payload into a polar orbit 500

km high (see figure 4-6).

DARPA originally intended to launch the Microsats on the first Pegasus launch, but has decided to

launch them on the second launch, perhaps late in

1989. On its first flight, now scheduled for January

1990, Pegasus will carry:

1. a 150-lb Navy communication satellite,

2. a NASA scientific experiment payload, and

3. instrumentation to evaluate the performance of

the ALV.

DARPA’s contract has four launch options remaining, and OSC has offered to add six additional

launch options to the contract.

OSC and Hercules also expect Pegasus can launch

lightsats into geostationary orbit; they recently

signed an agreement with Ball Aerospace to launch

two BGS-1OO Ball geostationary satellites in late

1990 or early 1991. OSC estimates each launch will

cost between $6 million and $8 million, and that

each of the satellites, which will weigh about 400 lb

and carry a 400-watt single-channel transponder,

will cost between $5 million and $8 million. OSC

estimates each mission (satellite, launch, and support) will cost about $20 million. The satellite design
Chapter 4--Lightsats ● 25

Figure 4-6-Projected Performance of Pegasus: Payload v. Orbital Altitude and Inclination

1 Perigee Altitude, nmi

450

\

350

\

250 150

\\/\ ● ● ●

\ .

\ \

\ \ \ 1. *

%%% ] \ \ 450 \ 3. 53, 23. ,), nm

500

9

Circular Orbits

r t

) I I I I I I I I

o 300 400 500 600 700 800 900 1000

Payload to Orbit, lb

Equatorial orbits — “ — O“ inclination

SOURCE: Orbital Sciences Corp.

is modular; Ball is developing larger versions with

more transponders or more power per transponder. 15

Standard Small Launch Vehicle

Last year, DARPA issued a Request for Proposals

to develop a transportable ground-launched Standard Small Launch Vehicle (SSLV) capable of

launching 1,000 lb of payload to a polar orbit 400

nmi high. DARPA recently awarded a contract to

Space Data Corp. (a subsidiary of OSC) for development, one launch (from Vandenberg Air Force Base

in the fall of 1991), and options for four more. The

first stage of the vehicle proposed by Space Data

Corp. will use a solid rocket motor developed for the

Peacekeeper ICBM by Morton Thiokol. The second,

third, and fourth stages will be the first, second, and

third stages of Pegasus, without the wings. OSC is

Polar orbits

90” inclination

also developing a commercial version of the SSLV,

called Taurus.

Other small launch vehicles have been developed

or proposed by companies that performed Phase 1

SSLV studies for DARPA. Space Services, Inc.

(SSI) developed a launch vehicle called Conestoga,

which uses clustered Castor solid rocket motors. The

first Conestoga was successful on a sub-orbital

flight, but the second failed shortly after launch on

November 15, 1989. LTV Aerospace, which produces the Scout, could produce an upgraded version.

Lockheed Missiles and Space Co. proposed a launch

vehicle that would use the first- and second-stage

motors from Poseidon C3 fleet ballistic missiles and

Morton Thiokol Star 48 motors for the third stage.

Lockheed estimated that the vehicle could be

available in two years and could launch a 770-lb

Is’’ Pegasus, Ball to Launch Communication Satellites Into Geosynchronous Orbit,” Aviation Week and Space Technology, June 12,1989, p. 64, and

Military Space, June 19, 1989, p. 8.
26 ● Affordable Spacecraft: Design and Launch Alternatives

payload from a land-based launcher to a 250-mile

high orbit inclined 28 degrees.16

Other Options

The U.S. Naval Research Laboratory (NRL) is

developing a concept for a sea-launch system to

launch a partially submerged launch vehicle from a

platform towed out to sea.17 This system, called

SEALAR (SEA Launch And Recovery) might

provide the survivability required by lightsats.

Sea-launch systems have been tested, to different

degrees, by the U.S. Navy (Project Hydra), Truax

Engineering 18 (SEA DRAGON, SUBCALIBER),

and StarStruck (now American Rocket Co.). DARPA

is not known to be considering a sea-launch system

for its Advanced Satellite Technology Program, but

such a system might prove attractive to the Navy in

the future.

Storage, shipment, and mobile basing of small

launch vehicles could be made safer by using hybrid

rocket motors—rocket motors that use liquid oxygen to burn solid fuel, which can be inert (nonexplosive). American Rocket Co. (AMROC) of Camarillo, CA, has developed a throttleable, restartable,

70,000-lb thrust hybrid rocket motor, the H-500. On

its first launch attempt (October 5, 1989), the motor

failed and the prototype sounding rocket it was to

power collapsed and burned at Vandenberg Air

Force Base, CA. It is noteworthy that it did not

explode, and did very little damage to the pad.

The sounding rocket, a prototype of AMROC’S

planned Industrial Research Rocket, carried a payload designed by AMROC for a Strategic Defense

Initiative experiment and a prototype reentry vehicle

developed by the Massachusetts Institute of Technology (MIT) Space Systems Laboratory. The reentry vehicle was to deploy an umbrella-like structure

made of space-suit material and decelerate to a soft

landing in the Pacific Ocean, where it was to be

recovered.

AMROC is developing a larger sounding rocket

and an even larger Industrial Launch Vehicle, the

largest version of which is being designed to launch

a 4,000-lb payload into low Earth orbit. AMROC is

not specifically designing the launch vehicle for

survivable basing (although this is not precluded)

and has not entered DARPA’s SSLV competition.

Before the sounding rocket failure, AMROC expected a frost launch late in 1990.

AMROC is also developing a larger hybrid motor

for a larger launch vehicle, as well as a smaller

hybrid motor for various applications, possibly

including use on projectiles launched from electromagnetic launchers (discussed below).

General Technology Systems (GTS) is developing a small launch vehicle called LittLEO to launch

lightsats. It is expected to be able to launch almost

a tonne (2,200 lb) of payload into a polar orbit 300

kilometers (162 nautical miles) high. First launch is

planned for 1992, probably from Andoya, Norway.

GTS quotes a price of nine million pounds sterling

per launch, which is equivalent to roughly $6,400

per pound to LE0.19

E’Prime Aerospace Corp. (EPAC) is developing

a series of launch vehicles for launching payloads

weighing from 1,000 to 20,000 lb into LEO and up

to 8,000 lb into geostationary Earth orbit (GEO).

EPAC quotes a prices of $12 million per launch for

the smallest launch vehicle and $80 million for the

largest. 20 EPAC plans to launch from Cape Canaveral Air Force Station, FL, and from Vandenberg Air

Force Base, CA. A first launch is planned for 1992.

The Soviet Union is developing a launch vehicle

called “Start” for launching lightsats. Start would

use guidance and propulsion systems developed for

the SS-20 ballistic missile and could be launched

from a mobile launcher, carrying 300-lb payloads to

orbits 500 km high. Space Commerce Corp., in

Houston, is seeking customers for Technopribor,

which is developing Start, and quoting a price of

about $5 million to $6 million per launch. Technopribor estimates a test launch could be conducted

in 1991.

Lightsats can also be launched as “piggyback”

payloads on launch vehicles carrying larger primary

payloads. Many U.S. and foreign launch vehicles

have done this for years. Arianespace has developed

lbt+bctieed will ~velop Small Milit~ Booster Using Poseidon C3 H~dw~e! “ Aviation Week and Space Technology, Sept. 26, 1988, p. 18.

ITNRL, Nav~ Center for Space Technology, briefing for OTA staff, Feb. 17, 1989.

lgcapt, Robefl C. Truax, USN (ret.), “Commercial View on Launch Vehicles, ” Space Systems Productivity and Manufacturing Conference [V (El

Segundo, CA: Aerospace Corp., Aug. 11-12, 1987); pp. 55-69.

lgIan p~ker, “Getting There Cheaply, ” Space, vol. 5, No. 4, July-August 1989, pp. 45-48.

zOIbid.
Chapter 4--Lightsats ● 27

procedures to do so routinely with the Ariane 4

launch vehicle, which, as noted above, is scheduled

to launch six amateur-radio satellites in addition to

the SPOT 2 photomapping satellite on January 19,

1990. General Dynamics is planning to offer a

similar service using its commercial Atlas launch

vehicle, which could launch, in addition to a primary

payload, one 3,000-lb satellite or several smaller

lightsats to LEO, or a 2,000-lb payload to geostationary transfer orbit.21 This service could be offered

in late 1991 for about $6,000 per pound to LEO; the

primary payload owner may reserve the right to

approve the price offered.22

Someday, small lightsats might be launched on

laser-powered rockets (discussed below). Lightsats

could also be launched on vehicles proposed for

launching larger payloads or crews--e.g., the Advanced Launch System, the Advanced Manned

Launch System, and NASP-derived vehicles.23 Of

these, only National Aero-Space Plane (NASP)-

derived vehicles (NDVs) are intended to provide a

survivable capability for wartime launch.

ISSUES

Are lightsats the most economical answer to the

problem of satellite vulnerability? Replenishing

satellites in wartime is only one of several partial

solutions; others include hardening satellites and

stockpiling spare satellites in orbit during peacetime,24 as well as arms control, actively defending

satellites, and reducing reliance on satellites for

support of military operations.25

What military requirements could lightsats satisfy? An Air Force officer responsible for space

system planning said, “The challenge to the small

satellite community has been to get out of the mold

of a solution looking for a problem; that is, what

missions will a small satellite support. "26 According

to the previous Secretary of the Air Force, “The

decision on whether a system is ‘small’ depends on

such things as orbit, mission, requirements, and

technology capabilities. When these factors properly

converge, we have built Smallsats . . . . What we

want are a realistic set of requirements and concepts

for smaller systems.”27 Several concepts have

already been proposed, the most grandiose of which

are being considered by the Strategic Defense

Initiative Organization: ‘‘BrilliantPebbles” or larger

space-based missile interceptors, “Brilliant Eyes”

(space-based space-surveillance satellites that would

demonstrate Brilliant Pebbles technology—see box

4-A), decoys for Brilliant Pebbles, and “Small

Dumb Boosters” (orbital transfer stages with which

Brilliant Pebbles could rendezvous and mate).

Could lightsat technology and launch vehicles

benefit civil applications? They already have. For

example, for two decades amateur radio operators

have built and used lightsats for recreational, educational, and public-service communications. Conceivably, networks of tens or hundreds of lightsats

could provide continuous global communications or

navigation services commercially. There is some

commercial interest in concepts that would require

only a few lightsats.2829

Since Sputnik I and Explorer I, science has

benefitted from lightsats and will continue to

benefit-more so if launch costs are reduced.

Medium-sized multi-mission remote-sensing satellites have used some instruments, such as the

Scanning Multichannel Microwave Radiometer on

Seasat-1, 30 that are light enough to be mated to a

zlTtJ reach g~~tionq orbit (GEO) from geostation~ transfer orbit (GTO), the payload must have an orbital wansfer stage.

zzIan pinker, op. cit., footnote 19.

23s= U.S. Cmgess, ~fice of T~hnology Aswssment, Round Trip to Orbit: Human SpaceflightAlternatives, OTA-lSC-419 (W*in$On, DC: U-s.

Government Printing Office, August 1989).

24sW, for exmple, tie sp~ch hat tie Honorable Edward C. AMridge, Jr., Secretary of the Air Force, prepared for a luncheon at the Aviation club>

Crystal City, VA, Sept. 15,1988, and Col. Charles Heimach, USAF, Speech to Second Annual AIAA/USU Conference on Small Satellites, Utah State

University, Logan, UT September 1988.

2S’S= U.S. Cm=ss, office of T~hnology AsSssment, Antisatellite Weapons, Countermeasures, and Am Cotirol, OTA-ISC-281 (w’sshin~on~

DC: U.S. Government Printing Office, September 1985); reprinted in Office of Technology Assessment, Strategic Defenses (Princeton, NJ: Princeton

University Press, 1986).

Zwol. Chmlm Heimach, USAF, op. cit., footnote 24.

Z7Hm. Edwmd C. Aldridge, Jr., op. cit., footnote 24.

zgF~ and Mosier, op. cit., footnote 1.

2~e ~n~r for ~ovative TWhnoloa of (he Comonwedth of Virginia h~ commissioned ~onomi~s at G~rge Mason University to ~StXS

potential markets for small satellites. This study is nearing completion.

30A.R. Hibbs and W.S. Wilson, “Satellites Map the Oceans, ” IEEE Spectrum, October 1983, pp. 46-53.
28 ● Affordable Spacecraft: Design and Launch Alternatives

Box 4-A—Brilliant Eyes?

Lawrence Livermore National Laboratory (LLNL) is developing a new class of electronic high-resolution

wide-angle TV cameras that, from an altitude of 1,000 km (610 mi), could image a land area the size of the state

of Virginia and show individual buildings. The first prototype camera, completed in 1987, has optics that are about

1 ft in diameter and 16 in long, excluding electronics. With improved electro-optical components, its resolution

would be comparable to or slightly better than that of the French SPOT satellite (about 10 m) from a comparable

altitude (832 km). At a lower altitude it could show greater detail but would have a smaller field of view. On Earth,

it could be used as a telescopic TV camera to record the tracks of meteors and low-altitude satellites against the

night-time sky.

SPOT Imagery of the Pentagon and LLNL has also developed a preliminary the White House

Lightsats Could Produce Comparable Imagery design for a miniature version of this camera

compact enough for use as a satellite navigation system. The system is designed to get

periodic position updates by viewing many

stars at the same time. The total mass of the

system is expected to be less than 250 grams

(about half a pound). LLNL expects that “this

[wide-field-of-view] system, with its combination of high resolution and high light collection capability, will also find applications in

SOURCE: U.S. Geodetic Survey. robot vision and smart munitions.

1

LLNL, Energy and Technology Review, July-August 1988, pp. 88-89.

lightsat bus to become a lightsat for meteorology and mapping or “mediasat” applications.31 This might

oceanography. Arrays of lightsats could someday be more economical than using a large, monolithic

use interferometric (phased-array) and aperture- satellite; predicting an arrays’ relative economy

synthesis techniques to provide high-resolution would require comparing cost estimates based on

radar or microwave imagery for Earth-resources detailed designs.

 U.S. Congress, Office of Technology Assessment, Commercial From Space, (Springfield, National

 Information Service, May 1987).
Chapter 5

Microspacecraft

Microspacecraft would be satellites or deep-space

probes weighing no more than about 10 kilograms

(22 pounds).1 Tens or hundreds could be usedto

measure magnetism, gravity, or solar wind at widely

separated points simultaneously. A swarm of different microspacecraft could obtain detailed radio

images of galaxies, while others could be used for

communications, gamma-ray astronomy, or planetary photoreconnaissance.

They would not require development of new

launch systems; they could be launched like buckshot on existing small launch vehicles. However, if

there were a demand for launching thousands per

year, it might be cheaper to launch them on

laser-powered rockets (see figure 5-1 ), if these prove

feasible.

Extremely rugged microspacecraft, constructed

like the Lightweight Exe-Atmospheric Projectile

(LEAP) being developed for the Strategic Defense

Initiative,2 could be launched to orbit by an electromagnetic launcher (railgun or coilgun) or a ram

cannon. An electromagnetic launcher in orbit could

launch them toward outer planets at muzzle velocities that would allow them to reach their destinations

and return data to Earth within a few years. This

might allow a graduate student to design a mission

and then receive mission data in time to use it in a

Ph.D. dissertation.

SPACECRAFT CONCEPTS

Several concepts for microspacecraft have been

proposed. One example is the Mars Observer

Camera (MOC) microspacecraft proposed by the Jet

Propulsion Laboratory at the California Institute of

Technology (CalTech). It would be a generic imaging microspacecraft; dozens could be launched on

each of several missions-to the Moon, the planets,

their moons, comets, and asteroids. A MOC microspacecraft would be shaped like an oversized hockey

puck, about 15 centimeters (cm) in diameter and 4

cm thick (see figure 5-2). It would weigh about 800

grams (g). A version could be designed to withstand

the accelerations to which electromagnetic launch

would subject them.

Placed in different orbits or trajectories, they

could trade off field of view for resolution, or vice

versa. For example, one MOC microspacecraft in a

polar orbit about Mars could serve as a Martian

weather satellite, providing two-color images with a

resolution of 5 to 10 kilometers (km )-sufficient to

resolve Martian clouds. A similar MOC microspacecraft in a lower orbit could serve as a mapper,

providing two-color images of a smaller field of

view with better resolution—100 meters (m). In time

it could map the entire planet. A similar MOC

microspacecraft in an even lower orbit about the

Moon could provide a two-color global map of the

Moon with 10 m resolution. Existing global maps of

the Moon currently show no features smaller than

several hundred meters.

LAUNCH SYSTEM OPTIONS

New, specialized launch systems need not be

developed to launch microspacecraft, because they

could be launched on existing launch vehicles—by

the dozens, if appropriate. However, some proposed

unconventional launch systems might prove to be

better or cheaper than conventional launch vehicles

for launching microspacecraft.

One example of such a system is a laser-powered

rocket that would use a laser beam, instead of

combustion, to heat the propellant, which could be

inert (i.e., nonreactive). If feasibility is proven, a

10-megawatt (MW) laser may be able to launch a

l-kg payload of one or more microspacecraft; a

gigawatt laser might launch a l-tonne (t) payload

consisting of several microspacecraft or a larger

spacecraft.

The smallness of microspacecraft has another

potential advantage: some microspacecraft could be

built to withstand high accelerations comparable to

those endured by cannon-launched guided projectiles such as Copperhead.3 Such “g-hardened”

IS= ROSS M. Jones, “Coffee-can-sized spacecraft,” Aerospace America, October 1988, pp. 36-38, and “Think small-in large numbers, ”

Aerospace America, October 1989, pp. 14-17.

Zsee U.S. Congess, Office of TwhnoIogy Assessment, SD1: Techncdogy, Survivubiliry, and So@are, OTA-l$C-353 (Washingon, DC: U.S.

Government Printing Office, May 1988), pp. 120-121.

sThe U.S. tiy’s M712 Copperhead cannon-launchd guided projectile is an artillery shell fired from a 155mm howitzer. It hm a %nsor and

electronics for detecting a spot of laser light on a target illuminated by a low-power, pulse-coded, target-designating laser aimed by a soldier or pilot.

When Copperhead detects such a spot, it steers toward it using fins deployed after launch. Copperhead rounds cost about $35,000 each.

–29–
30 ● Affordable Spacecraft: Design and Launch Alternatives

Pitch-over

Figure 5-1—Laser-Powered Rockets

\ \

‘-’.

— —

Payload

Ascent

Fairing (nose cone)

— Base late

Propellant(plastic, l

~ Laser beam

ice, or metal)

.A–Laser(nOtiOnal

)

\\

\ .

‘-’--..,.

SOURCE: Office of Technology Assessment, 1989.

microspacecraft could be launched by “direct

launch” systems4 such as railguns,s coilguns, 6 and

Orbital

insertion

A simple laser-powered rocket would have four parts: the

propellant (a block of plastic, metal, or ice), the payload,

the payload fairing (nose cone), and a base plate to which

the propellant is bonded and the payload and fairing are

attached. A large laser built on a mountain would beam

power to the rocket, vaporizing the propellant and

thereby producing thrust perpendicular to the surface of

the propellant.

The figure at left illustrates operation of the system

shortly after launch, when the rocket ascends vertically.

After the rocket rises above the densest part of the

atmosphere, the laser beam would be aimed off-center to

produce thrust asymmetrically, causing the rocket to pitch

(tilt) over, as shown above left. The rocket then begins to

accelerate downrange while ascending. When the rocket

reaches orbital altitude, it continues to accelerate

horizontally, as shown above, until orbital velocity is

attained.

ram cannons.7 In the near term, microspacecraft

could be launched by chemical rockets, such as

 (or a mission payload), no 

after they leave the muzzle of the launch system, if launched into solar orbit or an interplanetary trajectory. A projectile would require a small rocket

motor or some other kind of motor to enter Earth orbit.

 R. and Ali E. Dabiri, “Electromagnetic Space Launch: A in Light of Current Technology and Launch Needs and

Feasibility of a Near-Term Demonstration,” IEEE Transactions on vol. No. 1, January 1989, p. 393ff.

 and “Basic Principles of Coaxial Launch Technology,’ IEEE Transactions on vol. No. 2, March

1984, pp. 227-230.

 for of IAF-87-211, presented at the 38th

Congress of the International Astronautical Federation, Brighton, England, Oct. 10-17, 1987.
Chapter 5--Microspacecraft ● 31

Figure 5-2-Mars Observer Camera microspacecraft

Design Cutaway View-Actual Size

SOURCE: Jet Propulsion Laboratory, 1988

Scout and Pegasus, that are designed to launch

payloads of a few hundred kilograms for a fraction

of the cost of launching them on larger launch

vehicles.

In the remainder of this section we focus on

unconventional launch technologies. A surprisingly

large number of them have been proposed; one

recent review8 lists 60 propulsion technologies—in

addition to conventional chemical rocket technology—

that are potentially applicable to space transportation. About half are applicable to Earth-to-orbit

transportation; most are applicable to in-space

transportation (e.g., orbital transfer or escape),

which demands less thrust and power than does

Earth-to-orbit transportation. For brevity, we discuss

only two unconventional launch technologies here:

railguns and two-pulse laser-supported-detonation

(LSD) thrusters, which are the simplest of several

proposed laser-powered rockets. We discuss only

their application to Earth-to-orbit transportation

here, although both are also applicable to in-space

transportation. 9 In fact, orbital transferor reboost of

low-altitude satellites could be done with much

smaller lasers than would be required for launching

projectiles from Earth to orbit.10

Direct launch systems would subject payloads to

high accelerations. For a specified muzzle velocity,

the barrel length of any type of direct-launch system

must grow as the reciprocal of acceleration. To

achieve a muzzle velocity of 8 kilometers per second

with an acceleration of 1,000 gs,ll a direct-launch

system must have a barrel more than 3 km long. It

would be impractical for a launcher to be much

longer, or to subject a payload to correspondingly

lower acceleration.

To launch a projectile vertically at an acceleration

of 1,000 gs, the projectile must be subjected to a

force 1,001 times its weight.12 Hence exotic design,

fabrication, and testing processes are required—

especially for electronic and optical components—

and there are constraints on the shape of the

projectile, and, in practice, limits on its size.

Proposed projectiles have weights ranging from a

few kilograms to a tonne. They could carry payloads

such as fuel, food, water, structural components for

space assembly, and specially designed electronic

and optical systems such as those used in the Army’s

Copperhead cannon-launched guided projectile and

SADARM cannon-launched sensor-fuzed weapon,

and the Lightweight Exe-Atmospheric Projectile

being developed for the SDI.

 Eder, “Technological Progress and Space Development,” draft, Apr. 30, 1989.

 e.g., Ross M. ‘‘Electromagnetically Launched Micro Spacecraft for Space Science Missions,’ AIAApaper 88-0068, Jan. 11, 1988, and

Arthur Kantrowitz, ‘‘Laser Propulsion to Earth Orbit: Has Its Time Come?’ (cd.), Proceedings Workshop on Laser Propulsion,

 CA: Lawrence National Laboratory, July 7-18, 1986), vol. 2, pp. 1-12.

 op. cit., foomote 9, and Chapman, ‘‘Strategic Defense Applications of Ground-Based Laser Propulsion, ’ (cd.), op. cit.,

footnote 9.

 is ‘g”) of acceleration. acceleration of g, an object’s 

increases by 9.8 m (32 per second per second.

 force, it accelerate downward; its velocity would increase by

9.8 m per second each second. If sitting on the ground, it would be subjected to a force to its weight; this keeps it from falling into the ground-it

accelerates at zero If subjected to a force twice its weight, it would accelerate upward at 9.8 m per second per second, If subjected to a force thrice

its weight, it would accelerate upward at 19.6 m per second per second, and on.
32 ● Affordable Spacecraft: Design and Launch Alternatives

In a previous report, OTA described two proposed

direct launch systems: ram cannons and coilguns.l3

Coilguns could have important advantages over

railguns, which are simpler and more familiar

electromagnetic launchers (EMLs). For example,

coilguns can be designed so that the projectile does

not contact the barrel, avoiding barrel erosion, and

they can be scaled to launch large masses efficiently

at high velocities. Until recently, railguns were

expected to be very inefficient at launching multikilogram projectiles at the muzzle velocities (more

than eight kilometers per second) required to reach

orbit with minimal assistance from rockets. Low

efficiency would cause barrel heating and melting,

as well as a high electric bill. However, in recent

tests, railguns demonstrated unexpectedly high efficiencies in accelerating small projectiles to muzzle

velocities of 3 to 4 km per second,14 raising hopes

that they might be able to accelerate half-tonne

projectiles to more than 8 km per second.

Another cause for increased interest in railguns

for direct launch is the realization that ordinary

automobile batteries could be used for energy

storage and would cost much less than alternatives

previously considered. Automobile batteries could

also power coilguns.

The Air Force recently decided to demonstrate

suborbital launching of a microspacecraft using a

railgun at Eglin Air Force Base, Florida, after

augmenting its battery power system and adding a

barrel with a 30-cm bore, similar to one used

recently at Maxwell Laboratories. According to one

estimate, the upgraded gun could launch 5 to 10 kg

at 4 km per second three years from program start for

only about $10 million. A l-kg projectile launched

at 3 km per second with a l-kg sabot is expected to

reach an altitude of 200 km if the projectile’s nosetip

is allowed to ablate, or 400 km if the nosetip is

cooled by transpiration. *S

Proponents predict that a prototype operational

EML capable of launching 500-kg projectiles each

carrying about 250 kg of payload could be developed

in about six more years for an additional $900

million to $6 billion, including $50 million to $5

billion for development of vehicles and tracking

technology. l6 17

If produced and launched at a rate of 10,000 per

year, projectiles (less payloads) would cost as little

as $1,000 per kg according to one estimate, but over

60 times this, according to another estimate. An

EML projectile would require (besides its mission

payload) guidance, navigation, and control systems,

as well as a rocket kick motor to inject it into Earth

orbit. 18 Just as allowing a payload of specified

function to be larger allows it to be cheaper,

miniaturizing it makes it more costly, and ggardening it would make it still more costly. On the

other hand, if a mission required many projectiles,

high-rate production and learning effects could

reduce unit costs. Other launch costs might be as low

as $50 per kg at this rate.

If batteries are used, the limit would be about

10,000 launches (2,500 t) per year.19 Because of the

brief launch windows for rendezvous, very little of

this tonnage could go to a space station,20 but most

or all could be used for other applications, e.g.,

distributed low-altitude networks of tiny satellites

for communications, space surveillance, ballisticmissile defense, or space defense.

If the payload were reduced and the projectiles’

chemical rockets enlarged, the projectile would have

more cross-range capability; the launch windows for

rendezvous with a space station would be longer,

and more payload could be delivered to a space

station per year. However, the launch cost per pound

would increase. Chemical rockets could also be used

to reduce the muzzle velocity required, so that a

13u.s. Congess, Office of TechrIoIogy Aswssment, Launch Options for the Future: A Buyer’s Guide, OTA-ISC-383 (Washington, DC: U.S.

Government Printing Office, July 1988), p. 50.

~q~les R. palmer, “Electromagnetic Space Launch, ’ SA88091 (Mchan, VA: Science Applications International Corp., May 6, 1988).

IsIbid.

161bid,

ITMiles R. p~mer, perWn~ communication, Oct. 12, 1989.

ISA space pro~ bo~d for SOlaI orbit or a fly-by of another planet would not rwuire ~is.

19A ~eater la~ch rate could be achieved at greater cost if batteries are not u~d.

z(tp~mer ~d Dabiri, op. cit., fOOtnOte 5.
Chapter 5--Microspacecraft ● 33

smaller 21 or more conventiona122 direct-launch system could be used.

Laser propulsion would have important advantages over direct launch. Acceleration would be

much lower—about 6 or 7 gs on typical trajectories

to low orbit—so payloads would not have to be

designed to withstand gun-like stresses. Moreover,

no expensive device to store and quickly discharge

gigajoules of energy would be required, as it would

be for a railgun or coilgun, because power would be

beamed to a rocket continuously during ascent.

Perhaps the most important advantage of laser

propulsion is that a simple laser-powered rocket,

unlike an EML projectile, would not require guidance, navigation, and control systems, or a separate

kick motor for injection into orbit.

But laser propulsion would require a powerful,

expensive laser and a large, expensive adaptive

mirror. For efficient utilization and low average cost

per launch, both must operate reliably without

maintenance for longer periods than do existing

lasers. And laser launch operations would be halted

by overcast that would not impede direct launch.

Moreover, laser propulsion technology is less developed than EML technology and is predicated on

unproven theories of thermal blooming suppression

and thruster plasmadynamics. Validation of these

theories may require construction of a full-scale

launch system.

An SDIO official has estimated that a 20-MW

carbon-dioxide laser with a 10m-diameter beam

director telescope could launch rockets carrying 20

kg of payload for an incremental cost of about $120

per lb, assuming the laser efficiency is 15 percent,

the rocket efficiency is 40 percent, electricity can be

generated for four cents per kilowatt-hour, and the

structure and propellant each cost about as much as

the electricity. According to the SDIO, the 2mdiameter rocket structure would weigh only a few kg

and would require only about 120 to 150 kg of inert

propellant such as ice or polyformaldehyde plastic;

launching would require 30 to 40 megawatt-hours

(MW-h) of electric power.23

If the system could operate continuously without

downtime for maintenance or overcast, the launch

rate could be almost 60,000 per year. In practice,

occasional overcast would make full utilization

unachievable, and approaching it would probably

require at least two lasers and mirrors, so that one

pair could operate while the other is being serviced.

Even this would not assure operation most of the

time, unless the duty cycle of the lasers (i.e., the

fraction of time they are lasing) is much greater than

the duty cycles demonstrated by industrial and other

high-power lasers (see box 5-A). However, if a

launch rate of 100 per day could be maintained, over

1,600,000 pounds of payload could be launched into

orbit each year. This would be almost twice the

estimated combined capability of current U.S. space

launch systems,24 almost three times the annual

tonnage launched in 1984 and 1985, and about four

times the annual tonnage launched from 1980 to

1985.25

The SDIO postulates that, in practice, 100 payloads could be launched per day, on the average, and

the average cost could be as low as about $200 per

pound, if capital cost (table 5-1) were depreciated

over 5 years and if annual operating cost (excluding

rocket cost) were comparable to the annualized

capital cost of $90 million. The SDIO estimates that

launching only one or two payloads a day (500 per

year) would be sufficient to reduce average cost to

about $4,500 per pound and make laser-powered

rockets competitive with conventional rockets for

small payloads. Some users might be willing to pay

a premium for the speed with which a laser-powered

rocket could be prepared to launch a payload.

The SDIO estimates that a first launch to orbit

could be attempted about 5 or 6 years after program

start and expects to demonstrate a rocket efficiency

of 20 percent or more in experiments now being

planned. However, the highest efficiency demonzlp~mer and Dabiri, op. cit., footnote 5, and Henry Kolm and Peter Mongeau, “An Alternative Launching Medium,” IEEE Spectrum, vol. 19, No.

4, April 1982, pp. 30-36.

Z?L~ge~amOn~have&aen ~@ tO ]a~~h~ubo~bltal~rOjectlle~and sounding rockets to altitudes of 4oo,0~ feet; sw C.H. M~hy and G.V. Bu1l,

“AReviewof Project HARP,” Annah of theiVew York Academy of Science, vol. 140, No. 1, 1966, pp. 337-357, and R.G.V. Bull and Charles Murphy,

Paris CannonetiThe Paris Guns and Project HARP (Springer-Verlag, November 1988).

23 Jordin T. K=e, ~~~ls~ La~r ~opulsion for ~w Cost, High Volume La~ch to mbit,” preprint UCRL-101 139, Lawrence Livermore National

Laboratory, Livermore, CA, June 2, 1989. A different launch simulation program (Kantrowitz, op. cit., footnote 9.) predicts each launch would require

about nine minutes (hence about 20 MW-h of electric power) and about 200 kg of propellant.

XOTA, hunch optw~ for tk Future: A Buyer’s Guide, op. cit., footnote 13, p. 20.

~bid., p. 5.
34 . Affordable Spacecraft: Design and Launch Alternatives

Box 5-A—Lasers for Rocket Propulsion: The State of the Art

A laser-powered rocket would use a laser beam, instead of combustion, to heat the propellant, which could be inert

(i.e., noncombustible). The beam could come from a ground-based laser; the rocket could be extremely simple and weigh

only 10 times its payload. For comparison, the Scout launch vehicle weighs 1,300 times its payload.

Studies of Earth-to-orbit laser propulsion postulate the use of infrared lasers, which could be carbon-dioxide or

deuterium-fluoride electric-discharge lasers or free-electron lasers. The most mature of these is the carbon-dioxide

electric-discharge laser, but free-electron lasers are more efficient. If a carbon-dioxide laser or a free-electron laser

operating at the same wavelength (0.01 mm) were used, the economical laser power would be about 1 megawatt (MW:

1 million watts) per kilogram (kg) of payload, if the laser-powered rockets can achieve the 40-percent energy-conversion

efficiency once predicted by laser-propulsion proponents. To date, only 10-percent efficiency has been achieved.

Laser-propulsion experts now predict that at least 20-percent efficiency can been achieved. At this efficiency, a laser power

of about 25 MW might be required to launch a 20-kg payload. If, pessimistically, no more than 10-percent efficiency is

achieved, about 50 MW might be required to launch a 20-kg payload.

It appears to be feasible to build such a laser; the U.S. has built a gigawatt (billion-watt) free-electron maser and

electric-discharge lasers of much greater peak power, but none that could produce even 10 MW average power for ten

minutes, the boost duration required to reach low orbit. Almost a decade ago, the Antares carbon-dioxide electric-discharge

laser at the Los Alamos National Laboratory produced brief pulses with a peak power of 40 terawatts (40 trillion watts).

But a very different design, similar to that of industrial lasers used for welding, would be required for prolonged operation

at high average power. Free-electron lasers have, to date, produced less peak power. A free-electron laser developed by

Los Alamos and Boeing has produced pulses of ten megawatts peak power, but only six kilowatts average power, at a

wavelength of 0.01 mm. In early experiments, the partially completed Paladin free-electron laser at the Lawrence

Livermore National Laboratory amplified five-megawatt pulses from a carbon-dioxide laser 500-fold, presumably

producing pulses of about 2.5 gigawatts peak power. The carbon-dioxide laser power is being increased to a gigawatt, and

the free-electron laser has now been extended. If its electron accelerator operates at the average power for which it was

designed (at least 25 megawatts), and if 40 percent of the electron-beam power is converted to laser beam power

(comparable to the efficiency demonstrated by a similar free-electron laser at a wavelength of 8.8 mm), the Paladin

free-electron laser would produce a laser beam of at least 10 megawatts average power.

Neither carbon-dioxide lasers nor free-electron lasers have demonstrated the duty cycle (the fraction of time a device

operates) that would be required for an operational launch system. The duty cycle of a free-electron laser designed for a

high duty cycle would be limited primarily by the lifetime of the cathode used by the electron accelerator. Loosely

speaking, the cathode is like the filament of a light bulb, and more closely resembles the cathode of a cathode-ray tube such

as a TV picture tube. Several cathodes designed for long life are being tested. Alternatively, an electron storage ring (an

arrangement of magnets) could be used to recirculate the electron beam, as was done in the first free-electron laser and

others.

Focusing a multimegawatt laser beam on a small rocket hundreds of kilometers away is another serious technological

problem; in particular, control of beam-degrading nonlinear optical effects, such as thermal blooming, has not yet been

demonstrated at any average power and beam diameter of interest. Some research sponsored by the Strategic Defense

Initiative Organization (SDIO) is aimed at demonstrating high-power beam control for ballistic missile defense; the beam

control required for propulsion would be more difficult in some respects.

Nevertheless, participants at a 1986 workshop on laser propulsion sponsored by SDIO and the Defense Advanced

Research Projects Agency expressed optimism that a free-electron laser and beam director then planned for other purposes

“should be capable of launching test payloads to low [Earth] orbit in the early 1990s. ” SDIO subsequently established

a laser propulsion program and considered using a free-electron laser and a beam director to be developed for the SDI

Free-Electron Laser Technology Integration Experiment (FEL TIE) to experiment with laser propulsion, even though the

FEL TIE laser would be designed to operate at a wavelength shorter than optimal for laser propulsion.

Subsequent budget cutbacks postponed by at least two or three years the date by which the FEL TIE laser and beam

director could be operating. More recently, SDIO decided the FEL TIE laser should use a radio-frequency linear accelerator

(RF linac) similar to the one developed for the Los Alamos-Boeing free-electron laser instead of an induction linac similar

to the one used in the Paladin free-electron laser. The Los Alamos-Boeing RF linac produced an electron beam of higher

quality than that produced by the induction linac used by the Paladin laser; however, use of an RF linac may cause the

FEL-TIE laser to produce laser pulses with a waveform that is far from optimal for laser propulsion. This, together with

the nonoptimality of the FEL TIE wavelength, may lead SDIO to abandon hope of using the FEL TIE laser for laser

propulsion experiments and force SDIO, perhaps teamed with other sponsors, to develop a laser and beam director

specifically for laser propulsion experiments.
Chapter 5--Microspacecraft ● 35

Table 5-l-Estimated Cost of a 20-Megawatt Laser

for Powering Rockets

Development. . . . .. .. .. .. .. .. .. ... ... ......$ 75 million

Laser . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .“$185million a

Telescope . . . . . . . . . . . . . . . . . . . . . . . . .. ... ..$lO0 million

Adaptive optics .. .. .. .. .. .. .. ... ... ... ....$ 15 million

Tracking .. .. .. .. .. .. .. .. .. .. ... ... ......$ 50 million

Power plant.... .. .. .. .. .. .. .. ... ... ......$ 50 million

Structure .. .. .. .. .. .. .. .. .. .. ... ... ......$ 50 million

Total capital cost . . . . . . . . . . . . . . . . . .. ... ... $450 million

Total nonrecurring cost . . . . . . . . . . . . . . . . . $525 million

There is a class of science and exploration

missions that can be enabled by microspacecraft

(i.e., infeasible with larger spacecraft). This class of

missions requires many simultaneous measurements

displaced imposition .. .. Examples . . . include: 1)

a global network of surface or atmospheric sensors

on planets such as Mars. . ., 2) measuring the spatial

and temporal structure of magnetospheres about the

Earth, Sun, or other regions of space, and 3) using

microspacecraftas distributed arrays for either radio

or optical signals.

estimated as $25 miliion +$8 per watt.

SOURCE: Strategic Defense Initiative Organization.

strated to date is about 10 percent. If only 10 or 20

percent efficiency could be attained, an 80- or

40-megawatt laser would be needed, and average

cost per pound would be greater than indicated

above. If the cost of the power plant increases in

proportion to its power, average cost would be about

$490 or $275 per pound for 10 or 20 percent rocket

efficiency, respectively, at a launch rate of 100 per

day. 26

Because of the brief launch windows for rendezvous only 2 payloads per day could be launched

directly to a rendezvous with the space station.

Payloads launched at other times would take longer

and require more fuel to rendezvous.27 The SDIO

considers 8 payloads per day to be a conservative

estimate of the number of payloads that could be

launched to rendezvous with the space station each

day. With additional investment, the laser and

rockets could be given more crossrange capability.

This could be done by making the beam director and

rockets larger or by adding a conventional chemical

rocket to the laser-powered rocket.

ISSUES

What could microspacecraft do that conventional

spacecraft couldn’t? The consensus of the NASA/

SDIO microspacecraft for Space Science Workshop

Panel 28 was that:

They would have another advantage: they could

be launched from Earth orbit toward outer planets

by space-based electromagnetic launchers (railguns

or coilguns) at muzzle velocities that would allow

them to reach their destinations and return data to

Earth years earlier than could spacecraft launched by

conventional rockets. This would accelerate the

cycle of acquiring knowledge.

What is the market for such services? How much

is now spent on conventional spacecraft for space

science which microspacecraft could do? The 1988

NASA budget was about $9 billion, of which about

$1.6 billion was for “space science and applications." 29 30 Much of this is for NASA’s ‘‘great

.

observatories,’ such as the Hubble Space Telescope

and the Advanced X-ray Astrophysics Facility, and

for planetary probes such as Galileo. The consensus

of the NASA/SDIO microspacecraft for Space

Science Workshop Panel was that:

microspacecraft cannot achieve the science objectives of the great observatory missions such as the

Hubble Space Telescope or the Advanced X-ray

Astrophysics Facility. Also, intensive, multi-faceted

science investigations such as those of Galileo at

Jupiter cannot be supported by the microspacecraft

concept. . . . many space science missions will have

to continue to use established technology. Microspacecraft, if they are to be used in deep-space

missions, must establish a new inheritance chain, for

example by being used in near-Earth scientific or

non-scientific missions.

z~alc~ated by OTA using the launch simulation program of Kantrowitz, oP. cit., footnote 9. The launch simulation program used by Kare,

‘‘Trajectory Simulation for Laser Launching,’ Kare, op. cit. (footnote 10), pp. 61-77, predicts a 50 to 100 percent longer ascent than does Kantrowitz’s

program (for the case of 40 percent thruster efficiency) and hence 50 to 100 percent higher electric power usage and incremental cost.

zTSee p.K. Chapman, op. cit., footnote 10.

Z8NAS~0AST & SD1OflST, M1cro~Pacecra~for Space Science work.s~~eport O$rhe wor~~p Panel, Cdifomia Institute of Technology Jet

Propulsion Laboratory, Oct. 6, 1988.

29UCS0 Congess, Conwessiond Budget Office, The NASA progr~ in rhe 1990’s aria’ Beyo~ (W~in@on, DC: Congressional Budget OffiCe, May

1988); figure 1; see also figure 4 and box 3.

So’r’he ~p~ment of Defense space program also includes some focused space science projecw.
36 ● Affordable Spacecraft: Design and Launch Alternatives

An EML, ram cannon, or laser (to power rockets)

may permit microspacecraft to be launched from

Earth to orbit at low average cost, but only if utilized

efficiently. Maximum efficiency would require launching on the order of 10,000 microspacecraft per year.

How much would these microspacecraft cost? Could

space science budgets pay for so many microspacecraft? If not, what other types of microspacecraft

might be launched by such a system to maintain an

efficient launch rate?

Possibilities include:31

●

●

●

●

low-altitude comsat

tary);

networks (civil or mili-

“Brilliant Eyes, “ “Brilliant Pebbles,’ or “Small

Dumb Boosters” for a strategic defense system; 32

logistics for a space station or other space

operations. Payloads could include structural

components, fuel, armor, etc.; and

intercontinental artillery.

The utility of these applications has not been

established. All require further analysis before they

can be used to justify developing a direct-launch

system or a laser and laser-powered rockets. Some

proposed logistics schemes appear more promising

than others. For example, it is probably feasible to

launch Small Dumb Boosters (orbital transfer stages)

with which Brilliant Pebbles could rendezvous and

mate. Some have proposed launching projectiles

loaded with water, liquid oxygen, and liquid hydrogen toward the Space Station Freedom, but the costs

of collecting and decanting them have not yet been

estimated.

The risk of satellite collisions would increase

greatly if tens of thousands of microspacecraft were

placed in orbit, unless a means of collision warning

and avoidance is developed. Existing space surveillance systems may be inadequate for tracking tens of

thousands of microspacecraft, although Brilliant

Eyes or Brilliant Pebbles could help with this.

Ground-based lasers could be used to change the

orbits of satellites equipped with slabs of inert

propellant, whether launched by laser or not.33

However, this may not be adequate for collision

avoidance, because such satellites may pass over

propulsion lasers only infrequently, so advanced

warning of a collision hazard would be required, but

might be costly and subject to false alarms.

Brilliant Pebbles would not require advanced

warning of a collision; they could be programmed to

avert collisions by dodging approaching spacecraft.

They could also be commanded to ram a nonmaneuverable satellite (e.g., a failed Brilliant Pebble) that

posed a threat to more valuable U.S. and foreign

satellites. But a successful intercept might generate

debris and increase the long-term risk to spacecraft.

Collision avoidance schemes based on other technologies developed for antisatellite or ballistic

missile defense applications have been proposed;

some would not generate debris.34

sl~les R. p~mer, op. cit. footnote 14.

gz’rhese were describ~ above in tie ~tion on Iightsats. Brilliant Pebbles would weigh tens of kilograms-more than the lightest microspacwrtit,

and more than some lightsats, but still light enough to be launched by laser-powered rocket.

33A laser could be used to maneuver satellites much heavier than those it could launch into orbit.

340TA hm jmt bp an assessment of technologies for controhtg space debris and protecting satellites from it. The assessment Wm r~uest~ by

the Senate Committee on Commerce, Science, and Transportation, its Subcommittee on Science, Technology, and Space, and the House Committee on

Science, Space, and Technology.
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Advanced Space Transportation Technologies Assessment

Round-Trip to Orbit: Alternatives for Human Spaceflight. OTA-ISC-419, August 1989, GPO stock #052-003-01 155-7;

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Launch Options for the Future: A Buyer’s Guide. OTA-ISC-383, July 1988. NTIS order #PB 89-1 142681AS.

Reducing Launch Operations Costs: New Technologies and Practices. OTA-TM-ISC-28, September 1988. GPO stock

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Big Dumb Boosters: A Low-Cost Space Transportation Option? OTA Background Paper, February 1989. (available

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Space Stations and the Law: Selected Legal Issues. OTA-BP-ISC-41, September 1986. NTIS order #PB 87-118

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International Cooperation and Competition in Civilian Space Activities. OTA-ISC-239, .July 1985. NTIS order #PB

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U.S.-Soviet Cooperation in Space. OTA-TM-STI-27, July 1985. NTIS order #PB 86-114 758/AS.

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Remote Sensing and the Private Sector: Issues for Discussion. OTA-TM-ISC-20, March 1984. NTIS order #PB 84-180

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UNISPACE ’82: A Context for International Cooperation and Competition. OTA-TM-ISC-26, March 1983. NTIS

order #PB 83-201848.

Space Science Research in the United States. OTA-TM-STI-19, September 1982. NTIS order #PB 83-166512.

Civilian Space Policy and Applications. OTA-STI-177, June 1982. NTIS order #PB 82-234444.

Radiofrequency Use and Management: Impacts From the World Administrative Radio Conference of 1979.

OTA-CIT-163, January 1982. NTIS order #PB 82-177536.

Solar Power Satellite Systems and Issues. OTA-E-144, August 1981. NTIS order #PB 82-108846.

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SDI: Technology, Survivability, and Software. OTA-ISC-353, May 1988. GPO stock #052-O03-O1084-4; $12.00.

Anti-Satellite Weapons, Countermeasures, and Arms Control. OTA-ISC-281, September 1985. NTIS order #PB

86-182 953/AS.

Ballistic Missile Defense Technologies. OTA-ISC-254, September, 1985. NTIS order #PB 86-182 961/AS.

Arms Control in Space. OTA-BP-ISC-28, May 1984. NTIS order #PB 84-198 209/AS.

Directed Energy Missile Defense in Space. OTA-BP-ISC-26, April 1984. NTIS order #PB 84-21011 l/AS.

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