Communications Satellite Technology [1 ed.] 9781600862878, 9780262021012

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Progress in Astronautics and Aeronautics

Martin Summerfield, Series Editor PRINCETON UNIVERSITY

VOLUMES

EDITORS

1. Solid Propellant Rocket Research. 1960

Martin Summerfield

2. Liquid Rockets and Propellants. 1960

PRINCETON UNIVERSITY

Loren E. Bollinger THE OHIO STATE UNIVERSITY

Martin Goldsmith THE RAND CORPORATION

Alexis W. Lemmon Jr. BATTELLE MEMORIAL INSTITUTE

3. Energy Conversion for Space Power. 1961

Hathan W. Snyder

4. Space Power Systems. 1961

Nathan W. Snyder

INSTITUTE FOR DEFENSE ANALYSES

INSTITUTE FOR DEFENSE ANALYSES

5. Electrostatic Propulsion. 1961

David B. Langmuir SPACE TECHNOLOGY LABORATORIES, INC.

Ernst Stuhlinger NASA GEORGE C. MARSHALL SPACE FLIGHT CENTER

J. M. Sellen Jr. SPACE TECHNOLOGY LABORATORIES

6. Detonation and Two-Phase Flow. 1962

S. S. Penner CALIFORNIA INSTITUTE OF TECHNOLOGY

F. A. Williams HARVARD UNIVERSITY

7. Hypersonic Flow Research. 1962

AVCO CORPORATION

8. Guidance and Control. 1962

Robert E. Roberson

Frederick R. Riddell

CONSULTANT

James S. Farrior LOCKHEED MISSILES AND SPACE COMPANY

9. Electric Propulsion Development. 1963

Ernst Stuhlinger

10. Technology of Lunar Exploration. 1963

Clifford I. Cummings and Harold R. Lawrence

NASA GEORGE C. MARSHALL SPACE FLIGHT CENTER

JET PROPULSION LABORATORY

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11. Power Systems for Space Flight. 1963

Morris A. Zipkin and Russell N. Edwards GENERAL ELECTRIC COMPANY

12. lonization in HighTemperature Gases. 1963

Kurt E. Shuler, Editor NATIONAL BUREAU OF STANDARDS

John B. Fenn, Associate Editor PRINCETON UNIVERSITY

13. Guidance and Control — II. 1964

Robert C. Langford GENERAL PRECISION INC.

Charles J. Mundo INSTITUTE OF NAVAL STUDIES

14. Celestial Mechanics and Astrodynamics. 1964

Victor G. Szebehely

15. Heterogeneous Combustion. 1964

Hans G. Wolfhard

YALE UNIVERSITY OBSERVATORY

INSTITUTE FOR DEFENSE ANALYSES

Irvin Glassman PRINCETON UNIVERSITY

Leo Green Jr. AIR FORCE SYSTEMS COMMAND

16. Space Power Systems Engineering. 1966

George C. Szego INSTITUTE FOR DEFENSE ANALYSES

J. Edward Taylor TRW INC.

17. Methods in Astrodynamics and Celestial Mechanics. 1966

Raynor L. Duncombe U.S. NAVAL OBSERVATORY

Victor G. Szebehely YALE UNIVERSITY OBSERVATORY

18. Thermophysics and Temperature Control of Spacecraft and Entry Vehicles. 1966 19. Communication Satellite Systems Technology. 1966 20. Thermophysics of Spacecraft and Planetary Bodies Radiation Properties of Solids and the Electromagnetic Radiation Environment in Space. 1967

Gerhard B. Heller NASA GEORGE C. MARSHALL SPACE FLIGHT CENTER

Richard B. Marsten RADIO CORPORATION OF AMERICA

Gerhard B.Heller NASA GEORGE C. MARSHALL SPACE FLIGHT CENTER

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21. Thermal Design Principles of Spacecraft and Entry Bodies. 1969 22. Stratospheric Circulation. 1969

Jerry T. Bevans TRW SYSTEMS

WillisL.Webb ATMOSPHERIC SCIENCES LABORATORY, WHITE SANDS, AND UNIVERSITY OF TEXAS AT EL PASO

23. Thermophysics: Applications to Thermal Design of Spacecraft. 1970

Jerry T. Bevans

24. Heat Transfer and Spacecraft Thermal Control. 1971

JohnW. Lucas

25. Communication Satellites for the 70's: Technology. 1971

TRW SYSTEMS

JET PROPULSION LABORATORY

Nathaniel E. Feldman THE RAND CORPORATION

Charles M. Kelly THE AEROSPACE CORPORATION

26. Communication Satellites for the 70's: Systems. 1971

Nathaniel E. Feldman THE RAND CORPORATION

Charles M. Kelly THE AEROSPACE CORPORATION

27. Thermospheric Circulation. 1972

WillisL.Webb ATMOSPHERIC SCIENCES LABORATORY, WHITE SANDS, AND UNIVERSITY OF TEXAS AT EL PASO

28. Thermal Characteristics of the Moon. 1972

John W. Lucas

29. Fundamentals of Spacecraft Thermal Design. 1972

John W. Lucas

30. Solar Activity Observations and Predictions. 1972

Patricks. Mclntosh and Murray Dryer

JET PROPULSION LABORATORY.

JET PROPULSION LABORATORY

ENVIRONMENTAL RESEARCH LABORATORIES NATIONAL OCEANIC AND ATMOSPHERIC ADMINISTRATION

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31. Thermal Control and Radiation. 1973

Chang-Lin Tien UNIVERSITY OF CALIFORNIA

32. Communications Satellite Systems. 1974

P. L. Bargeliini COMSAT LABORATORIES

33. Communications Satellite Technology. 1974

P. L. Bargeliini COMSAT LABORATORIES

(Other volumes are planned.)

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Communications Satellite Technology

Purchased from American Institute of Aeronautics and Astronautics

The MIT Press Cambridge, Massachusetts and London, England

Progress in Astronautics and Aeronautics

An American Institute of Aeronautics and Astronautics Series Martin Summerfield, Series Editor Volume 33

Purchased from American Institute of Aeronautics and Astronautics

Communications Satellite Technology

Edited by P. L Bargellini

Technical papers selected from the

CLARKSBURG, MARYLAND

COMSTAT LABORATORIES

AIAA 4th Communications Satellite Systems Conference, April 1972, subsequently revised for this volume. In addition, the volume contains specially invited papers to round out the coverage of the subject.

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Copyright ©1974 by The Massachusetts Institute of Technology

All rights reserved. No part of this book may be reproduced in any form or by any means, electronic or mechanical, including photocopying, recording, or by any information storage and retrieval system, without permission in writing from the

publisher. This book was printed by Alpine Press, Inc.

and bound by Colonial Press, Inc. in the United States of America. Library of Congress Cataloging in Publication Data

AIAA Communications Satellite Systems Conference, 4th, Washington, D.C., 1972. Communications satellite technology.

(Progress in astronautics and aeronautics, v. 33) "Technical papers selected from the AIAA 4th Communications Satellite Systems Conference, April 1972, subsequently revised for this volume. In addition, the volume contains specially invited papers to round out the coverage of the subject." 1. Artificial satellites in telecommunication—Congresses. I. Bargellini, P. L, ed.

II. Title. III. Series. TL507.P75 vol. 33 [TK5104] 629.1'08s ISBN 0-262-02-101-3 73-15612

[621.38'0422]

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Preface

xi

I Orbit and Attitude Control

1

Variation in Range, Range-Rate, Propagation Time Delay, and Doppler Shift for a Nearly Geostationary Satellite

3

VICTOR J. SLABINSKI

Active Attitude and Orbit Control of Body-Oriented Geostationary Communications Satellites

29

MARSHALL H. KAPLAN

INTELSAT IV Nutation Dynamics

57

J. T. NEER

Effect of an Autotracking Antenna on the Nutational Stability of a Dual-Spin Spacecraft

85

J. E. MclNTYRE AND M. J. GIANELLI

Attitude Determination for the SKYNET and NATO Synchronous Communications Satellites

127

CHARLES W. HANSON AND JOHN V. BROWN

Momentum Wheel Three-Axis Attitude Control for Synchronous Communication Satellites

141

J. E. KEIGLER, W. J. LINDORFER, AND L. MUHLFELDER

II Propulsion and Power

163

.Propulsion Requirements for Communications Satellites

165

WILLIAM C. ISLEY AND KENNETH I. DUCK

Survey of Auxiliary-Propulsion Systems for Communications Satellites

191

L. B. HOLCOMB AND D. H. LEE

Selected Comparisons among Propulsion Systems for Communications Satellites

245

BERNARD FREE AND GEORGE HUSON

Power Processing Systems for Ion Thrusters

261

B. G. HERRON, D. R. GARTH, R. C. FINKE, AND H. A. SHUMAKER

Low Thrust Propulsion System Effects on Communication Satellites D. F. HALL AND W. C. LYON

279

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Results from Tests of a Large Lightweight Solar Array Unit

307

K. L HANSON

III Communications

329

Polarization Isolation Characteristics of a Dual-Beam Reflector Antenna

331

J. W. DUNCAN, S. J. HAMADA, AND W. C. WONG

Measurements of Earth Station Antennas G/T Ratio by Radio Stars and Satellites

359

HENRY J. KOCHEVAR

Design of the INTELSAT IV Transponder

373

S. B. BENNETT AND I. DOSTIS ,

A Time Division Multiple-Access System for INTELSAT

393

D. J. WITHERS AND A. K. JEFFERIS

Synchronization of Earth Stations to Satellite-Switched Sequences

411

R. A. RAPUANO AND N. SHIMASAKI

Results of German TDMA-Experiments

431

G. ECKHARDT, H. HABERLE, B. REIDEL, AND H. RUPP

Performance of a Digital Adaptive Echo Canceller in a Simulated Satellite Circuit Environment

455

H. G. SUYDERHOUD AND M. ONUFRY

PSK and QPSK Modulation and Demodulation in Digital SatelHte-to-Satellite and Satellite-to-Ground Links

479

C. LOUIS CUCCIA AND WILFRED E. LEE

Characteristics and Applications of Multibeam Spacecraft Antennas

503

K. G. SCHROEDER

A Near-lsotropic Microwave Antenna for Communications Satellites

533

W. S. GREGORWICH AND C. W. WESTERMAN

Index to Contributors to Volume 33

541

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PREFACE

A proper perspective of communications satellite technology can be acquired through the identification of three distinct periods of its development. Starting the countdown with the remarkably precise forecast by Arthur C. Clarke in 1945, the art will be thirty years old in the next couple of years. In an epoch like ours of rapidly advancing technologies, one third of a century is a long enough time span to witness great changes. Yet, in the first period, which lasted some twelve years, nothing really happened except a few continued speculations and the use of the moon as a passive communications satellite in 1954-55. Only in the second period, from 1957 to 1964, were man-made satellites actually used for communications purposes. .This period was characterized by feverish activity; following Sputnik, numerous launches of space vehicles took place in the U. S. and the U.S.S.R., among which one finds a variety of communications satellite.programs, such as SCORE, COURIER, ECHO, TELSTAR, RELAY, and SYNCOM. The intensity of these efforts and the diversity of the approaches (e.g., passive vs active satellites and the widely different orbital and spacecraft characteristics) helped in the redefinition of certain fundamental problems. Thus, the third period, which can be called the operational phase, began in 1965; the launches of the INTELSAT I (Early Bird) and the first MOLNIYA satellites inaugurated the commercial era of satellite communications. The word "commercial11 is used here in a broad sense to cover civilian and/or military communications at the operational level rather than for experimental purposes. Curiously enough, while the Western world used satellites for international, world-wide communications, the U.S.S.R. was first in using them extensively and effectively for domestic purposes. I would like now to invite the reader to consider three recent major events. The first event is represented by the continued successful launchings of satellites of the INTELSAT IV series, which are now providing global services of an extent and variety undreamed of a few years ago.* The second event was the successful launching, .one in November 1972 and another in April 1973, of the Canadian Anik satellites, the *From the 240 telephone circuits of INTELSAT I launched in 1965, after only six years the communications capacity of a single satellite has jumped to 4000 telephone circuits (nominal) in INTELSAT IV.

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XII

first communications satellites in geosynchronous orbit specifically intended to provide domestic services. The third event is represented by the FCC long-awaited rulings on U.S. domestic satellites. The continued expansion of the INTELSAT international communications satellite consortium is well known. The man in the street is aware of the fact that the words, "via satellite" flashing on his home TV screen, imply that what he sees and hears comes to him from faraway places via one, or more, INTELSAT satellites. It is a fact that, although intercontinental telephone and telegraph traffic may be routed along communications paths other than satellites (e.g., submarine cables, and in some cases high-frequency circuits), television and other wideband, world-wide services are possible only by satellite. The success of the INTELSAT venture was accurately and humorously summarized three years ago by a well-known European space scientist, who stated that, among all organizations throwing money into space, only INTELSAT had shown the capability of getting it back with interest. It is now clear that other organizations have already joined, or are about to join, the club because in regard to domestic systems the vacillations of many years are about to be replaced by one of the most interesting races in the history of technology. Although some ruthlessness in one form or another is expected, it is now clear that satellites will greatly contribute to domestic communications. As a result of the success and future promise of commercial communications satellite systems for national and international communications, the present volumes are largely devoted to this central theme. Volume 32, which is devoted to Systems, is organized in three Chapters covering, respectively, International Applications, Advanced Concepts, and Special Topics. Volume 33, which is devoted to Technology, also is organized into three Chapters, treating Orbit and Attitude Control, Propulsion and Power, and Communications. In addition to contributions of a general nature, whose timeliness and significance cannot be missed and which will interest all readers, the specialist will find in-depth studies in specific areas of the highly complex combination of subsystems which form a communications satellite. Although it is certain that communications capacity per satellite will increase in the next five years by at least one order of magnitude, it has been pointed out that this result

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XIII

will be accomplished without recourse to brute force solutions in terms of spacecraft mass and power. The reader will find papers outlining the various approaches that will lead to the achievement of the above-mentioned goal. In addition to the advantages expected from higher efficiency solar cells and batteries for eclipse operation, the major step in spacecraft engineering will be the move toward body-stabilized structures capable of permitting the precise pointing of several very narrow beam antennas toward chosen areas of the Earth. Savings in the total satellite mass will be made possible by electric propulsion (ion engines) for North-South stationkeeping and orbital repositioning. - These spacecraft technologies, combined with advanced communications techniques such as the use of spectral regions below and above 10 GHz and the frequency reuse made possible by spatially orthogonal waves and by separate narrow antenna beams, will lead to communications capacities progressively higher than those presently available. The next most significant change in communications satellites will occur when their function will be upgraded from the present relatively simple one of repeaters to that of switchboards in orbit. Keeping in mind the interaction of the space and the Earth segments, it is clear that new systems can be designed with considerable freedom, although the expansion of present systems is often constrained by the existing installations and the huge investments represented by them. Aside from expanded telephone communications, television, and data communications, many interesting problems will arise from the interaction of satellites, conventional radio broadcast, and CATV systems. Again, the reader will find much food for thought in the volumes. As history has repeatedly shown, mistakes are made, especially by experts, in evaluation of potential new technologies. It took about three decades for the airplane to become commercially important and ultimately competitive against trains and ships; and in the beginning of aviation, not too many took it seriously. In a similar manner today, in certain conservative quarters of the electrical communications industry, satellites are still regarded as "just another means of communication." Some experts have gone so far as to state that satellites can provide only what other coiranunications systems already provide. This is true in a qualitative sense only; comparatively speaking, in a different field, a jet aircraft and a horse and buggy are both means for moving people! Yet, qualitative statements are inadequate for the engineer and the economist, and only when some measure of quantity is introduced can each technique be assessed in terms of a hierarchy of potential possibilities, values, and costs. m

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XIV

Then it also becomes clear, although not without considerable effort, how different technologies should interact to serve mankind in an optimum manner. Throughout several decades of this writer's activity in the field of electrical communications, one item of the utmost importance has become apparent: the need for greater coordination among the various forms of communication systems.' It is my belief that satellites are bound to play a major role towards this coordination. Although analog techniques were advancing in Earth-based and satellite systems, digital technologies have made their appearance and show great promise. Traditionally, frequency and/or time have been the classical domains for multiplexing and accessing. Onboard satellite switching, as previously mentioned, coupled with time division multiplexing and time domain multiple access, will result in further substantial increases of satellite capacity. Thus, the advances of satellite systems will stimulate the growth of surface communications and will accelerate the trend toward the coordinated progress of both. Eventually, all-digital communications systems, consisting of surface facilities integrated with space-borne switching centers, will come into being.

As soon as large capacity domestic and regional satellite systems are established, their emergence into the international network on one side, and their interfacing with the capillary local networks on the other, will provide for mankind's needs to communicate to an extent of scale, convenience, and economy well beyond present-day standards. It should be noted that the bulk of the present two companion volumes consists of papers given at the AIAA 4th* Communications Satellite Systems Conference held in April 1972 in Washington, D.C. Two previous similar Conferences resulted in the publication of one volume in 1966 and two volumes in 1971. After accepting the Editor's task for one or two books covering the 1972 Conference, I tried first to arrange the papers so as to provide some continuity with respect to the previous volumes and then to adopt a logical order in the presentation of the available contributions. Secondly, I made an attempt to obtain new contributions. In this regard, I hope that readers will conclude, as I have, that the limited quantity of the new submissions is compensated by their quality and significance. As a member of the AIAA Technical Committee on Communications Systems, which was responsible for the organization of the Conference, I would like to acknowledge the efforts and help of many colleagues and friends on the Committee. Special

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XV

recognition is due W. L. Pritchard, General Chairman of the Conference Committee, and Richard L. Booton Jr. and Charles M. Kelly, Technical Program Co-Chairmen. Acknowledgment also is due Dr. Martin Summerfield, Series Editor-in-Chief; Miss Ruth F. Bryans, Director of Scientific Publications at AIAA Headquarters; and my secretary, Mrs. Shirley H. Taylor, who provided coordination and secretarial assistance. .

P. L. Bargellini May 1973

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VARIATION IN RANGE, RANGE-RATE, PROPAGATION TIME DELAY, AND DOPPLER SHIFT FOR A NEARLY GEOSTATIONARY SATELLITE Victor J. Slabinski* Communications Satellite Corp., Washington, D. C. Abstract Equations are derived and graphs are presented giving the satellite range and range-rate variations as a function of the orbital elements and Earth station location. These range and range-rate variations readily give the propagation time delay and Doppler shift for signals sent through the satellite. The graphs permit a simple determination of the worst case variation in time delay and Doppler shift for signals sent between two stations, given the orbit inclination, eccentricity, and longitude drift rate. The effect of orbit maneuvers on range-rate also may be determined with the graphs. 1. Introduction 1.1 Need for Range and Range-Rate Analysis The basic concept of satellite communication using timedivision multiple access (TDMA) requires that a certain phase synchronization be maintained between signals from different Earth stations to allow the signals to be interleaved as they pass through the satellite repeater. Television networks, and some proposed high-speed data networks, require that signals originating from different stations be kept in time synchronism at the receiving station to facilitate the interleaving This paper is based upon work performed at Comsat-Plaza under the sponsorship of the International Telecommunications Consortium (INTELSAT). Views expressed.are not necessarily those of INTELSAT. *Member of Technical Staff, Celestial Mechanics Division.

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4

V. J. SLABINSKI .

of signals from different stations. However, even small changes in the satellite range from Earth stations can affect the required synchronization of signals sent through a communications satellite: 1) The propagation time delay increases by 1 ysec for every 0.300 km increase in satellite range. 2) The fractional frequency change in the received signals (Doppler shift) is a 1 part in 109 decrease for every 0.300 m/sec increase in range-rate. Some variation in propagation time delay and some Doppler shift always occur because a satellite in synchronous orbit is never perfectly stationary with respect to the rotating Earth. Even if a satellite could be placed in synchronous orbit with zero inclination, zero eccentricity, and zero longitude drift rate, the satellite motion would soon depart from this geostationary condition because of perturbations of the orbit. For example, the gravitational pull of the sun and moon at times can change the orbit inclination by 0.005°/day, and Earth*s triaxiality can give a longitude acceleration of 0.0016°/day2. Lubowe1 considers the perturbations to the orbit of an initially geostationary satellite, and his graphs illustrate the resulting variation with time of the transmission path through the satellite for a particular station pair. This paper gives a graphical means for evaluating the path length variation for any station pair, given the orbit inclination, eccentricity, and longitude drift rate.

1.2 Time Delay and Fractional Doppler Shift The propagation time delay T for a radio signal, transmitted from Earth station 1 through a communications satellite to Earth station 2, is approximately

T = (p + p )/c 1

2

(1)

where p is the satellite range from the Earth station and c is the speed of light, 2.997 925 x 108 m/sec. In this approximation, each path (uplink or downlink) gives a path delay taken as p/c, a quantity which depends only on the relative location of station and satellite. Relativistic effects neglected in this approximation can give an error of 0.4 ysec in T for a nearly geostationary satellite, out of a total delay of ^0.3 sec. This error from neglecting relativity is much less than the error from this report's neglect of sun and moon pertur-

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RANGE VARIATION OF GEOSTATIONARY SATELLITES

5

bations to the orbit, which can give a 12-hr oscillation with an amplitude of ^6 ysec in T.

A signal transmitted from Earth station 1 with frequency f 1 will be received at Earth station 2 as a signal with a Doppler-shifted frequency f2. The fractional change in the frequency (which this report calls the fractional Doppler shift) is approximately

(f2 - fi)/fi = -[dpi/dt + dp2/dt]/c

(2)

where t is time. In this approximation, each path gives a contribution ~(dp/dt)/c to the total fractional Doppler shift, which contribution depends only on the relative motion of station and satellite. Relativistic effects (specifically, the clock paradox and the gravitational red shift) neglected in this approximation may give an error of order 1 part in 1011 in this fractional Doppler shift for a nearly geostationary satellite. This error from neglecting relativity is much less than the error from this reportT s neglect of sun and moon perturbations, which can give a 12-hr oscillation with amplitude ^0.9 x 10~9 in the fractional Doppler shift of a signal sent through the satellite. Equations (1) and (2) show that we may calculate separately each path T s contribution to the time delay variation and the Doppler shift, and then combine the separate contributions to find the total time delay variation and the total Doppler shift. The remainder of this paper is concerned with the graphical evaluation of the range and range-rate quantities to use in these two equations. 1.3

Restrictions on the Satellite Orbit

This paper considers nearly geostationary satellites with orbit inclinations i < 1°, orbit eccentricities e < 0.001, and mean drift rates D under l°/day. These restrictions allow the separation of effects resulting from inclination, eccentricity and drift rate. The results of this paper can be applied to satellites whose orbits are somewhat outside these limits with reduced accuracy. Operational communications satellites are usually within these limits. 2. Basic Range and Range-Rate from Graphs On the assumption that the satellite follows an unperturbed elliptic orbit, Appendix A derives the approximate expressions presented in Sec. 2.1 for the range and range-rate

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V. J. SLABINSKI

of a nearly geostationary satellite. These expressions are accurate to ^1.2 km in range and ^0.1 m/sec in range-rate, amounts that are comparable to the errors from neglecting the perturbations discussed in Sec. 2.5. The coefficients in the range and range-rate expressions can be evaluated readily for any station location from the set of graphs given in this report, as discussed in Sec. 2.2. Section 2.3 explains the determination of worst-case range variation, time-delay varition, range-rate, and Doppler shift for a given tolerance on orbit inclination, eccentricity, and drift rate, while Sec. 2.4 gives a numerical example illustrating the use of the graphs for this purpose. 2.1 Theoretical Basis of the Graphs Appendix A shows the range p of a nearly geostationary satellite from an Earth station at east longitude L , geocentric latitude i (north positive) and distance R from Earth's center is approximately p = p ' - A. sin nt - A sin (nt +•) m i e

(3)

The first term on the right is the mean range p to the satellite at a given time. This is the range to the satellite's mean position^ a point over the equator halfway between the perigee and apogee heights, at the satellite mean longitude L . The mean longitude here may be considered as a longitude center which moves linearly with time, about which the true satellite position undergoes various sinusoidal oscillations in the course of a day. Because L changes, p has a slow, nearly linear change with time, as discussed later in this section.

In the second and third terms of Eq.(3), n is the average angular frequency of orbital motion about Earth's center (the mean motion) for the satellite, approximately 2ir rad/1436 min for a nearly geostationary satellite. It is related to the satellite orbital period P by n = 27T/P

(4)

Here t is (approximately) the time since the satellite crossed the Earth's equator traveling northbound (satellite at the ascending node of its orbit). Within 0.11°, nt gives the geocentric angle (argument of latitude) along the orbit, from this equatorial crossing to the satellite position.

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RANGE VARIATION OF GEOSTATIONARY SATELLITES

LOCATION WITH MAXIMUM AMPLITUDE 8.12 011.3)f SI = ffl = ffffffl—1111 nTT-fflffl^T^wm^fff—pm—rrnT4-i rtt

'0

20

40

60

80

LONGITUDE DIFFERENCE |Lm - Lel BETWEEN SATELLITE AND STATION (DEG)

Fig. 1 Amplitudes for inclination part of variations, nA for range-rate and A for range. (Notes: For south latitudes, take negative of quantity for corresponding north latitude. For .other inclinations i, multiply quantity from graph by ± in degrees.) The second term on the right in Eq.(3) gives the oscillation in satellite range resulting from the north-south motion of the satellite caused by the orbit inclination to the equator. The amplitude A. is a function of station position and is linearly proportional to the inclination i.. This amplitude may be found from Fig. 1. The amplitude is positive for sta-r tions in the northern hemisphere and negative for stations in the southern hemisphere. This sign difference indicates a 180° phase difference between hemispheres for this part of the

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8

V. J. SLABINSKI

Table 1 Range variation resulting from the orbit inclination Sense of the range variation Satellite motion over the equator Southern nt Northern hemisphere hemisphere station station Northbound M3° Decreasing Increasing (ascending node) Southbound (descending node)

VL80°

Increasing

Decreasing

range variation. Table 1 gives the sense of the range change resulting from the inclination when the satellite crosses the equator (satellite at orbit node) . The third term on the right in Eq. (3) gives the oscillation in satellite range caused by the orbit eccentricity. The eccentricity gives a satellite oscillation in longitude about l^ which is 90° out of phase with the accompanying oscillation in geocentric radial distance, so that the satellite traces an oval path relative to its mean position over Earth1 s rotating surface. The amplitude Ae is a function of station position and is linearly proportional to the eccentricity e. This amplitude may be found from Fig. 2 and is always positive. The phase angle $ depends mainly on the perigee position along the orbit, with a slight dependence on station location, as explained in Sec. 3.2. In a worst-case analysis, $ may be treated as an orbital parameter independent of station location. Appendix A also shows the range-rate dp/dt of a nearly geostationary satellite may be given as dp/dt = pm - nA-£ cos nt - nAe cos (nt + cj>)

(5)

The first term on the right gives the range-rate resulting from the mean drift of the satellite eastward or westward in longitude. The quantity fon is a function of station position and is linearly proportional to the mean drift rate D (positive to the east) , defined by

D = dL/dt

(6)

This contribution to the range-rate stays constant through a day, with a slow change from day to day. pm may be found from Fig. 3. The second term on the right in Eq. (5) gives the

fD

rt O

P

fD P

CO •

"d oo

O Hl-h p

LO

>

PfD P . O

> 5 ro co

CO •

I-P

O W O

P

i-h P

Pa 00

fD

rt

p" P^

£

o

p

o

Pa

fD

Pa

i-t fD

fD

co

fD


+ M) + (v - M)]

- sin (A - ft) + 2e cos (A - ft) sin M

(A15)

where A is the mean orbital longitude, A E M + CD + ft

( A 1 6 )

Making such true anomaly expansions in Eq.(A9) and omitting terms in e2 gives p - -fa2 + R2 - 2 aR cos £ cos2 ~ cos (A - S) - 2 aR

sin £ sin i sin(A - ft)

+ cos & sin2 -| cos[2(A - ft) - (A - S)] - 2 a

2

e { l - — cos & cos2 4 cos (A - S) } cos (A - 0) -ft)

1 Jr - {2 ^ cos & cos2 j sin(A - S)} sin (A - w - ft) |f (A17)

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20

V. J. SLABINSKI

where terms with a factor of sin i and sin2(i/2) have been omitted from the last square bracket, since they give an error in the range oscillation due to the eccentricity of less than 0.6% of the amplitude (0.22 km worst-case range error). We also may take cos2(i/2) - 1 within this bracket with little additional error. The terms A and S are linear in time and go through one cycle in a day, so the quantity (A - S) is nearly constant through the day and may be treated as a phase angle in Eq. (A17). Appendix B shows this quantity may be replaced by the mean longitude of the satellite east of the station, (L - L ). m e

is p

m

We expand the last equation to show that the mean range 1

E l{a2 4- R2 - 2 aR cos £ cos2 ± cos (L - L ) }2 2 m e J

(A18)

and the range variations due to the orbit inclination and eccentricity are

i

i - (aR/p ) cos A sin2 ^ cos [2(A - fl) - (L - L )] m 2 m e (6p)

E -A

cos [(A - fl) + (6 - a))]

(A19)

(A20)

where A. E (aR/p ) sin & sin i A

2

1

(A21)

"T

E —— {l - 2K cos(L - L ) + K2[4 - 3 cos2(L - L )]}. (A22) e p m e m e m

tan 0 E [2K sin(L

K E (R/a) cos £

(A23)

- L ) ] / [ ! - K cos(L - L )] e m e

(A24)

m

-90° < 0 < +90°

We have as the maximum magnitudes of the ratios

( 6 p ) . / p * - sin i pitch, and yaw components of control angular momentum, respectively h^ ,hz = roll and yaw orbit rate decoupled momentum commands, respectively = I£,IY>IZ satellite principal moments of inertia J = wheel moment of inertia about its axis of symmetry K,K = roll and pitch autopilot gains, respectively k = yaw-to-roll gain ratio m = mass of wheel system = ^Xo^Zc roll and yaw control moments, respectively P = average wheel power required 7" = disturbance torque T^i = thruster misalignment torque T^ = motor torque magnitude 6,y = roll and yaw gimbal angles, respectively e = over-all beam-pointing error ^1'^2'^p = damping ratio of roll, yaw, and pitch dynamics, respectively T,T = roll and pitch time constants, respectively c)),9,^ = roll, pitch, and yaw attitude angles, respectively ^ = wheel speed fijj = nominal wheel speed oo0 = orbit rate = 7.8 x KT5 rad/sec a)^,0)2,Wp = natural frequencies of roll, yaw, and pitch dynamics, respectively COX,O)Y,(JOZ = satellite angular velocity components in bodyfixed coordinate system Introduction

Communications satellites of the mid-1970s and beyond are likely to be three-axis stabilized and actively controlled in both attitude and orbit. Of course, they will be placed in geostationary orbits and employ narrow-beam antennas for high effective radiated power with limited mass. Several such satellites are currently under development, e.g., Canadian Communications Technology Satellite, the European Telecommunications Satellite, and post-INTELSAT IV satellites. These

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ACTIVE ATTITUDE AND ORBIT CONTROL

31

vehicles have the same basic configuration as shown in Fig. 1. To take maximum advantage of this type of vehicle, the control system must maintain accurate position and orientation.

Automatic attitude control and commanded orbit corrections are achievable by several means. Attitude control torques can be produced by momentum exchange devices and mass expulsion techniques. At synchronous altitude gravity gradient and magnetic moment methods do not produce sufficient torques to permit satisfactory correction of orbit control thruster misalignments and execution of attitude acquisition maneuvers. Orbit control forces must be generated by mass expulsion, since solar sails and other exotic techniques are not Assumed satellite practical for this application.Fig. 1 configuration. Specific torque producers and

thrusters for this mission are several in type and have been the subject of much controversy. Selection criteria and performance considerations are presented here for both orbit and attitude devices. Current developments in technology and hardware are discussed,and comparisons of some system combinations offered. Promising concepts for future generation satellites are identified and des.cribed. In particular, double-gimbaled momentum wheels and all-electric thruster systems represent two developing areas which are competing for future applications on long-life communications satellites. An example configuration is used to illustrate design considerations for these two types of control systems.

Control Requirements

Control requirements are the necessary responses to perturbations and commanded deviations from an existing situation

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M. H. KAPLAN

to fulfill mission objectives. This definition applies to both orbit and attitude maneuvers, each of which has varying requirements depending upon the situation, e.g., acquisition, reacquisition, station and attitude maintenance, etc. Of primary concern here is nominal position and orientation control. Associated torque requirements include responding to perturbing moments from outside as well as internal sources. Significant perturbing torques result from solar pressure and inclination control thruster misalignment. Solar pressure torques are sinusoidal in nature with period equal to that of the orbit. Inclination control thruster torques are produced by uncertainties in center-of-mass location and thrust vector direction. Such torques may not be steady but are limited to discrete time intervals during which inclination errors are corrected. Occasional momentum dumping is necessary if a momentum exchange device is incorporated in the control system. Stationkeeping requirements include only responses to outside forces. The longitude holding requirement is a function of longitude only and is the result of equatorial asphericity (triaxiality), which is the cause of the only in-plane secular perturbation of significance to geostationary orbits. Other in-plane disturbing forces are periodic and of lower amplitude. Inclination corrections require the major portion of propellant and are the result of lunar-solar attraction on the satellite. The secular effect can be approximated in terms of velocity increment required to maintain zero inclination. This varies between 40 and 51 m/sec/yr over an 18.6-yr cycle because of precession of the lunar orbit. Limits on orientation and position errors are defined through considerations on available power, communications requirements, and contractual ground station tracking and coverage agreements. Narrow-beam antennas require high pointing accuracy, especially for ground stations far away from the boresight location. Attitude limit components can be determined somewhat arbitrarily from the over-all beam pointing requirement. max

\/ max

max

max

For example, a 60° great circle arc away from the subsatellite point will produce an off-nadir angle of E = 8.27°. If e max = 0.1°, an acceptable combination of attitude angles is *max = emax = 0.07°, i(Jmax =0.2° (typical value if no yaw sensor is used).

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Attitude Control Techniques There are two basic methods of producing attitude control torques that are considered here: momentum exchange and mass expulsion. Stabilization systems employing one, two, three, and more wheels have b.een proposed recently. Pitch axis control usually is provided by a reaction wheel through variation of speed. One- and two-wheel systems do not use yaw sensors because roll/yaw coupling provides yaw information if the pitch wheel momentum is high. The double-gimbaled momentum wheel system is of particular interest because of its unique combination of advantages over other momentum devices. An example control system is presented to illustrate design and parameter selection procedures for this type of system.

Mass expulsion attitude control has not been widely used except where a momentum device was the primary attitude actuator. However, long-life, nonspinning satellites which employ narrow-beam antennas may offer a situation in which attitude thrusters with low impulse bits and high reliability are preferred over momentum systems. An example all-electric thruster system is presented to illustrate design procedures and selection of components. Double-Gimbaled Momentum Wheel System Design Considerations Specifications

The double-gimbaled momentum wheel offers control torques about all three vehicle axes through wheel speed control and gyrotorquing. If the wheel size and speed are correctly selected, momentum exchange permits cancellation of cyclic torques without employing attitude jets, and only periodic momentum dumping is necessary due to secular disturbance torques. This is in contrast to mass expulsion control systems, in that continual thrusting and roll, pitch, and yaw sensors are required for accurate attitude maintenance. Only a roll/yaw (Earth) sensor is needed for the double-gimbaled wheel system described here.

Wheel systems of this type have yet to be proven for long life applications. Reliability and useful life properties may indicate that redundancy is needed. Weight advantages over mass expulsion devices are not yet clear, even for a 10-yr lifetime. However, hardware properties are left for a later development phase, and only preliminary design considerations are presented based on nominal mode control requirements.

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Linearized equations of motion are presented, and specific satellite properties are assumed for an example vehicle. Appropriate control laws and system parameters were selected to counteract disturbance torques resulting from solar pressure and limited station change thruster misalignment. Elimination of thrust misalignment torques arising from electric north-south engines has been considered separately. Stability of control systems was determined and response obtained for realistic disturbance torques. An appropriate wheel momentum was selected and weight estimates made. The vehicle considered was selected on the basis of current launch technology and expected spacecraft constraints. Assumed specifications and expected disturbance torque expressions are listed in Table 1. Consider the satellite initially placed and deployed in geostationary orbit at the proper attitude by some other system. The control system initiates operation with nominal wheel momentum, H^, and zero gimbal deflections and attitude errors. A rigid body model is used for this preliminary design. Flexibility effects are left for a more comprehensive investigation. Beam pointing requirements are consistent with the use of only horizon sensors. The lifetime requirement is somewhat academic here, because hardware reliability and mean time between failures are not of direct concern. However, comparisons with mass expulsion systems will depend on useful satellite lifetime. Equations of Motion

Development of the equations of motion for this satellite is based on a classical formulation-* which assumes a rigid

Table 1 Assumed satellite specifications ________________and disturbance torques_______________

Initial in-orbit mass Moments of inertia

716 kg (1574 Ibm) Ix = Iz = 2000 kg-m2 (1475 slug-ft2) IY = 400 kg-m2 (325 slug-ft2) Useful life 7 yrs Attitude accuracy ^Pitch and roll = 0.05° requirements \Yaw = 0.40° Solar pressure torques (TX = 2 x 10~5 ( 1 - 2 sin 030t) N-m (t = 0 at 6 AM or

in Eq. (28) from WQH^ to K, the roll autopilot gain, resulting in

T =

x V

+ KT + K +

*

* V

This represents a classic second -order system with driving functions TX and -H^ip, the gyroscopic coupling term which permits yaw errors to appear in roll sensor outputs. The desired yaw response would be well damped and decoupled from roll. However, the yaw control law cannot provide direct damping because yaw angles are not measurable in this case. Therefore, the control law uses roll angle and a pseudorate modulator to form the complement of Eq. (18): M

Zc

= H

Z

+

V^X ~ V

=

- kK(T * + *>

(20)

Equation (19) can be written as -kK(T + < ( > ) = kiyj; + kl^ - kTx

(21)

Then Eqs . (17) and (21) give a new yaw equation,

kl

- kT

(22)

This is interpreted as a damped second-order system with forcing functions T^ , kT^, and -kl^, the roll coupling term. This latter term is not desirable because it represents yaw errors resulting from roll transients. Direct compensation for this in the control law is not practical, because it requires the second derivative of (f), which would introduce excessive noise into the system. It is convenient to define "orbit rate decoupled momentum commands" at this point as 5 h

Xc & \c

(23a)

(23b)

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M. H. KAPLAN

The control laws, now written as Xc

Xc

o Zc

Zc

0) v 0 Xc

(24a)

*>

(24b)

-kK(T +

represent an undamped oscillator at orbital frequency, 0) . Thus, oscillation of momentum commands at this rate has the effect of decoupling roll and yaw dynamics which arise from the orbital frequency. In other words, commands on gimbal motion are referred to an inertial frame as is the angular momentum vector, even though the satellite is pitching at the rate 0)0. In terms of gimbal angle deflections, definitions (23) become Xc

By using these forms with Eqs. (19), (22), and (24), a control system block diagram was formulated as shown in Fig. 6. Roll sensor input is processed through a pseudorate modulator, and deflection commands y and 6 are produced according to the selected control laws.

To investigate system stability and responses the roll/ yaw equations are rewritten in Laplace form as Tx(s)

S^ + K(TS

Tz(s)

-kK(Ts +

V

*2 2 or L.,

L

L

engines, may be fired in sequence so that power drain is equivalent to only a portion of the units. Duty cycles for attitude control need not be constrained otherwise. The large N-S and station-change thrusters operate within some restrictions. Station change engines will require most of the power available in order to maneuver the satellite in reasonable time. Thus, no high-density communications functions can be performed while major longitude changes are occurring. This is consistent with operational philosophy. However, N-S thrusting must maintain a rather strict duty cycle because of restraints on firing positions around the orbit. Thus, there is a power penalty in wieght to run these thrusters. The level of thrust is, however, high enough so that on a 50% daily duty cycle, yearly N-S corrections can be accomplished in only 155 days. Thus, no thrusting should be required during shadow periods. In any case, batteries will not be used to run large engines, because corresponding maneuvers are performed after panels are deployed and in sunlight.

In summary, many areas related to the application of electric thrusters to a specified geostationary communications satellite without momentum devices were covered partially or in depth. Considerations have included mission and control requirements, sensor usage, orbit determination, electric thruster technology, attitude and orbit maneuvers and maintenance, and system interactions. Although many questions could not be answered primary problems were identified. Based on these considerations the following conclusions have been made: 1) An all-electric thruster control system can provide complete control impulses for the geostationary satellite studied. 2) Restrictions on exhaust impingement and interference resulted in some nonoptimum ttiruster site selections. However, full redundancy is still possible, and adverse E-W coupling effects of N-S thrusters can be overcome easily by the E-W engines. 3) A specific satellite configuration was used here. However,

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limited variations in initial mass and shape would not affect the selection of thruster types or general conclusions. Conclusions Promising attitude and orbit control concepts for bodyoriented communications satellites have been presented and specific examples discussed in detail. Design considerations for double-gimbaled wheel attitude control and all-electric thruster control have been included because these techniques represent two currently developing systems competing for application to the next generation of communications satellites. Momentum wheels permit yaw control without direct yaw sensing and provide optimum propellant usage and inherent attitude stiffness. All-electric thruster configurations eliminate all rotating components which have limited life in orbit and permit efficient use of propellant with full redundancy. The same system may provide position as well as orientation control. There is no clearcut weight or power advantage between these two systems. The final choice will depend on reliability, life, simplicity, and, finally, weight. References

^•Pritchard, W. L. and Bargellini, P. L., "Trends in Technology for Communications Satellites," Astronautics & Aeronautics, Vol. 10, No. 4, April 1972, pp. 36-42. 2

Dougherty, H. J , , Lebsock, K. L . , and Rodden, J. J . , "Attitude Stabilization of Synchronous Communications Satellites Employing Narrow-Beam Antennas," Journal of Spacecraft and Rockets, Vol. 8, No. 8, August 1971, pp. 834-841. 3

Kaplan, M. H., "Estimation and Correction of Electric Thruster Misalignment Effects on an Oriented Geostationary Satellite," COMSAT Technical Review, Vol. 3, No. 1, Spring 1973. ^Kaplan, M. H. and Edwards, T. L., "Double-Gimbaled Momentum Wheel Control System for a Body-Stabilized Spacecraft Configuration," TM CL-45-72, August 31, 1972, COMSAT Laboratories, Clarksburg, Md. 5

Lyons, M. G., Lebsock, K. L., and Scott, E. D., "Double Gimballed Reaction Wheel Attitude Control System for High Altitude Communications Satellites," AIAA Paper No. 71-949, Hempstead, N.Y. , August 1971.

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Cannon, R. H. , "Some Basic Response Relations for ReactionWheel Attitude Control," ARS Journal, Vol. 32, No. 1, January 1962, pp. 61-74. Nicklas, J. C. and Vivian, H. C. , "Derived-Rate Increment Stabilization: Its application to the Attitude Control Problem," Transactions of ASME Ser. D; Journal of Basic Engineering, Vol. 84, March 1962, pp. 54-60. o

Scott, E. D. , "Pseudorate Sawtooth-Pulse-Reset Control System Analysis and Design," Journal of Spacecraft and Rockets, Vol. 4, No. 6, June 1967, pp. 781-785. 9

Kaplan, M. H. , "All-Electric Thruster Control of a Geostationary Communications Satellite," Journal of Spacecraft and Rockets, Vol. 10, No. 2, February 1973, pp. 119-125. l^Holcomb, L. B. , "Survey of Auxiliary Electric Propulsion Systems," Addendum to "Satellite Auxiliary -Propulsion Selection Techniques," Tech. Rept. No. 32-1505, July 15, 1971, Jet Propulsion Lab., Pasadena, Calif.

, B. , "Chemical and Electric Propulsion Tradeoffs for Communications Satellites," COMSAT Technical Review, Vol. 2, No. 1, Spring 1972, pp. 123-145.

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INTELSAT' IV NUTATION DYNAMICS

J. T. Neer* Hughes Aircraft Co., Culver City, Calif.

Abstract INTELSAT IV has successfully "become the second generation Gyrostat stabilized dual-spin vehicle. Although presently carrying the bulk of the international telecommunication traffic, the satellite has demonstrated the applicability of the Gyrostat stabilization technique for precision pointing missions. The nutational design of INTELSAT IV was altered after an extensive empirical investigation of the destabilizing effects of fuel motion in a spinning satellite. Excellent correlation between flight and laboratory test results was obtained. Highly efficient passive eddy current nutation dampers were flown for the first time and provided the required damping over the range of mission operations.

Presented as Paper 72-537 a"t "the AIAA kfh Communications Satellite Systems Conference, Washington, D.C., April 2^-26, 1972. In part, this paper is based upon work performed under the sponsorship'of the International Telecommunications Consortium (INTELSAT). Any views expressed here are not necessarily those of INTELSAT. The author wishes to acknowledge, particularly, the technical efforts of J.O. Salvatore, G.J. Adams, D.B. Krimgold, R.H. Bernard, and G.J. Cloutier in supporting the design and analysis of the INTELSAT IV nutation control system. Special thanks also are due COMSAT for not only permitting the in-orbit nutation tests and measurements to be made but also providing early nonspinning fuel slosh test results. The later tests were conducted at COMSAT Laboratories and the special efforts (Ref. 7) of E. Martin are acknowledged. ^Senior Project Engineer. 57

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J.T. NEER

Nomenclature a d E I m N R

= = = = = = =

r £> 6 X cr T CD

= = = = = = =

acceleration due to nutation, fuel tank diameter, ft energy dissipation, ft-lb/sec transverse moment of inertia, slug ft equivalent simple pendulum mass, slugs frequency ratio; dimensionless = Ai/con distance "between vehicle e.g. and pendulum radius of gyration, ft distance "between, spin axis and accelerometer axis, ft traction of critical damping nutation angle, deg or rad nutation frequency, rad/sec inertia ratio; dimensionless time constant, sec frequency or spin rate, r/sec or r/min

Subscripts D DD i L n P res T U

= = = = = = = = =

damper dedamper inertial lower natural; referring to damper natural frequency platform resonance; transverse upper

Superscript 1

= prime notation; denotes perturbed valve (see text)

I.

Spacecraft Description

INTELSAT IV, built by Hughes Aircraft Company for the Communication Satellite Corporation, is the latest in a series of commercial communication satellites owned and operated by the International Telecommunications Satellite (INTELSAT) Consortium. The two main elements of the spacecraft are the spinning rotor, comprising 60% of the on-station spacecraft weight of l600 Ib, and the despun Earth-oriented platform, containing the communication repeater and its antennas. A rotating interface, consisting of conventional ball bearings,

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despin motor, a rotary transformer, and slip rings, sustains the relative motion between the two bodies, permits signal transfers to take place, and affords an electrical path over which power from the solar panels and "batteries can flow to the repeater payload. The spinning rotor provides a basic gyroscopic stability to the spacecraft. As seen in Fig. 1, the appearance of the antenna mast is dominated by the presence of large, 50-in.-diam, spot-beam antennas. These high-gain, narrow-beam antennas will be used for transoceanic communications between areas of maximum service, e.g., United States and Europe. The spot-beam reflectors are independently steerable over the full Earth disk in both elevation and azimuth by antenna positioner mechanisms. Located behind the spot-beam reflectors are two Earth coverage communication receive antenna horns and two transmit antenna horns. These are horn antennas oriented vertically with reflector "mirrors" used to direct the beams toward Earth.

The position and orientation (reaction control) subsystem is mounted on the rotor so as to benefit from the centrifugal 1 2 3 3a 4 4a 5 6

T E L E M E T R Y AND COMMAND ANTENNA A N T E N N A MAST S T R U C T U R E RECEIVE ANTENNA, GLOBAL COVERAGE TRANSMIT A N T E N N A , G L O B A L C O V E R A G E T R A N S M I T A N T E N N A , SPOT-BEAM DISH FEED HORN ANTENNA POSITIONING MECHANISM TELEMETRY HORN ANTENNA NUTATION DAMPERS (2) FOR WARD SUN SHIELD, QUARTZ MIRROR MATERIAL WAVEGUIDE OUTPUT MULTIPLEXERS DE-SPUN E L E C T R O N I C EQUIPMENT P L A T F O R M T R A V E L I N G W A V E T U B E S - T W T ( 1 2 x 2) TWT POWER SUPPLY C O N V E R T E R S TELEMETRY AND COMMAND REPEATER ELECTRONICS B E A R I N G A N D P O W E R T R A N S F E R A S S E M B L Y (Bapta) ELECTRICAL POWER CONTROL ELECTRONICS BATTERY PACK (2) A T T I T U D E C O N T R O L T H R U S T E R - R A Dl A L ( 2 ) 18a S P I N - U P JET (2) 19 H Y D R A Z I N E M O N O P R O P E L L A N T T A N K S ( 4 ) 20 SUN S E N S O R (3) 21 A C C E L E R O M E T E R (2) 22 E A R T H S E N S O R (3) 23 SPUN W I R I N G H A R N E S S C O N N E C T I O N S 24 DE-SPIN C O N T R O L E L E C T R O N I C S 25 N O Z Z L E - A E R O J E T G E N E R A L SO L I D - P R O P E L L A N T APOGEE MOTOR A T T I T U D E C O N T R O L T H R U S T E R - A X I A L (2) AFT THERMAL BARRIER BOOSTER ADAPTER F O R W A R D SOLAR PANEL AFT SOLAR PANEL R E S I N - B O N D E D A L U M I N U M H O N E Y CO M B D R U M SOLARCELLSBONDEDTOSURFACEOFDRUM

Fig. 1 INTELSAT IV spacecraft.

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J.T. NEER

field. The INTELSAT IV satellite will be maintained continuously in its synchronous equatorial orbit at its designated longitude with the spin axis pointing north. Propulsive thrust is by catalytic decomposition of anydrous hydrazine; 300 Ib of hydrazine are carried in the four propellant tanks. Roughly 35 Ib of propellant are allocated for initial injection errors5 with the remainder being expended at a rate of about 35 Ib per year (90% of which is for north-south stationkeeping). Two jets, one axial and one radial, are required to maintain both attitude and station. For increased reliability the radial and axial jets are redundant. Two additional jets are provided for initial spinup and spin trim control once in synchronous orbit. After apogee boost the spin jets can be operated in a pulsed mode to

provide a backup technique for attitude precession and inplane velocity maneuvers. The spinup jets in combination with the axial jets (which are biased 0.5° in the despin direction) provide two-way spin speed control. Additional functional redundancy for north-south stationkeeping is provided by the radial jets which are canted about 25° out of the equatorial plane so that the thrust is applied through the e.g. The cant results in a kofo coupling into out of plane or inclination maneuvers.

The 1562-lb apogee motor is carried in the aft half of the rotor section. This solid rocket provides the needed impulse to place the spacecraft in its final synchronous equatorial orbit. By virtue of this maneuver, INTELSAT IV becomes the first Gyrostat stabilized vehicle to be apogee-boosted into synchronous orbit.

Continuous Earth orientation of the despun platform is required for the directional communication antennas. The despin control system (DCS) provides a means of autonomous inertial pointing of the platform. Sun or Earth sensor information (three Earth sensors, two sun sensors - all rotor mounted) is utilized to establish the instantaneous inertial attitude of the rotor. A magnet/pipper coil pair located in the bearing and power transfer assembly (BAPTA) establishes the relative phase between the rotor and the platform. A brushless d.c. motor provides the necessary torque to compensate continuously for friction torques produced by the bearing and slip rings. Because of the importance of despin control, not only to communication service but also to nutational stability, an automatic rate control logic has been included. If, for any

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INTELSAT IV NUTATION DYNAMICS

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reason, the despin electronics is deprived of inertial reference information, the control logic automatically will

switch to a rate hold mode wherein the spacecraft is controlled to maintain the most recent platform-rotor relative rate. The rate hold mode can toe open loop commanded from the ground if desired. An ancillary despin mode is provided by ground commanding an artificial pulse train to which the spacecraft despins. This pseudo Earth mode is used for portions of the transfer orbit and apogee boost sequence and, if needed, for certain failure modes of operation. Two pendulous eddy current nutation dampers are located on the antenna mast. Identical in design but varying in performance, these dampers provide the basic nutation stability for the satellite. Redundant active nutation dampers (AND) provide increased stabilizing torques during transient periods when the dampers are operating in a saturated (amplitude limited) mode, e.g., following apogee motor burn.

II.

Mission Summary

A typical launch and injection sequence for INTELSAT IV consists of the key events shown in Fig. 2. The Atlas/Centaur launch vehicle is used to inject the 3100-lb spacecraft into an elliptical transfer orbit having an apogee at synchronous altitude and a perigee of 380 naut miles. The orbit inclination is 28.h°. Prior to separation, the Centaur reorients the vehicle into an orbit normal attitude. Concurrently at separation a clamp, which holds the platform and rotor together during launch, is released and the spinup sequencer is initialized. The spinup commences about 2 sec after separa.tion with the firing of . both hydrazine spinup jets. (This allows time for the adapter Fig. 2 INTELSAT IV mission sequence.

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'

J.T. NEER

to clear the aft end of the spacecraft.) The spin jets fire for about 8.5 sec, sufficient to stabilize the spacecraft gyroscopically against the separation transient. The jets are then commanded off for ko sec, during -which time the hydrazine and nitrogen separate under the centrifugal force field created by the initial burn. The jets are subsequently refired for approximately U6 sec. The resultant spin rate is nominally 51 r/min. At the completion of spinup, the sequencer closes the latching valve upstream of the spin jets, commands on the despin control electronics and the redundant motor drivers, and selects the Earth sensor despin mode. The active nutation damping system also is energized at this time. About 15 min after separation, the spacecraft is visible to the Carnarvon, Australia tracking, telemetry, and command (TT&C) station. A spacecraft status check is made via PCM and FM telemetry. A rotor mounted accelerometer provides the basic information to assess spacecraft nutation stability. Subsequent to the checkout, the spacecraft is maneuvered into the attitude required for apogee motor burn. The spin axis precession angle between the separation and apogee injection attitude is about 36.5°. Orbit and attitude data are then collected in preparation for a touchup maneuver prior to injection into synchronous orbit. Apogee motor burn typically occurs on second or fourth apogee for an Atlantic satellite and third or fifth orbit for a Pacific satellite. The transfer orbit period is nominally 10 hr 50 min. Prior to apogee motor burn, the despun platform is spun at a small rate — a few revolutions/min — sufficient to average out any transverse torques produced during thrusting from a statically imbalanced platform. The active nutation damping system operates at the end of burn to reduce the nutation to its threshold value — nominally 0.^°. The platform is then despun. The resultant

spin speed is close to the resonant condition for the dampers, whereupon the system passively damps with a net time constant less than 25 sec. Approximately 30 min after injection into synchronous orbit, the satellite is maneuvered into the nominal orbit normal attitude. The spacecraft then drifts to its intended longitudinal station where it is stopped via spin synchronous radial jet pulsing.

The communication equipment is put through an extensive checkout test. The satellite is then ready for service. Through the 7 yr of normal operation, the spacecraft is maintained at the desired station to within +0.5°- Inclination maneuvers occur every 12 to 15 months whereas the east-west maneuvers occur about every 3 to h months. The spacecraft attitude is maintained to ±0.5° of the orbit normal. Solar

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INTELSAT IV NUTATION DYNAMICS

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torques precess the vehicle at a variable rate throughout the year; the rate varies from -0.003° to +0.025°/day. III. Nutation Design INTELSAT IV belongs to the class of vehicles that are spin stabilized. More specifically,, the design incorporates a dual spin configuration with a large spinning section, the rotor, and an Earth-oriented despun platform. The vehicle is stabilized about its axis of minimum moment of inertia. Although this condition is a stable one for a classical rigid body," small dissipative effects (such as structural flexing, fuel slosh, etc.) will cause the minimum inertial spin axis to cone (nutate) about the angular momentum vector. If the nutation is uncontrolled, the amplitude will increase, ultimately reaching a new stable equilibrium condition characterized by a pure spin about the axis of maximum moment of inertia. By providing dissipation devices on the platform and restricting the dissipative effects on the rotor, the dual spin configuration can be passively stabilized about the desired axis of minimum moment of inertia. This specific application of spin stabilization is referred to as the Hughes Gyrostat stabilization technique. TACSAT I, launched in February 1969, was the first flight demonstration of the Gyrostat principle. INTELSAT IV has become the second generation satellite series stabilized by this method. The hardware implementation of the Gyrostat principle into the INTELSAT IV design has taken the form of a pair of platform mounted nutation dampers and redundant rotor mounted active nutation dampers. The despun dampers render the configuration stable throughout the normal mission sequence. As an auxiliary backup to the despun dampers, the active nutation control system utilizes the hydrazine propulsion system as a means of imparting large transverse torques to the spacecraft. Flight data have established the fact that the actual stability margins are entirely consistent with both the prelaunch analyses and dedamper tests. Furthermore, within the nominal spin speed operating range, the nutation dampers by themselves guarantee nutational stability. The active nutation dampers serve nominally in a backup capacity. However, because of the possibility of subsystem failures occurring in either the despin control system or the command system in conjunction with anticipated large nutation transients, such as that following apogee motor burn, the AND system serves in a protective capacity.

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Requirements The preceding description of the spacecraft and its associated operational sequence should give the reader an appreciation of the complexity of the INTELSAT IV mission. In view of the mission plan, the spacecraft represents an ambitious application of the Gyrostat stabilization technique. The collective implications of not only the basic mission requirements but also the earlier flights of ATS-V and TACSAT I and an extensive fuel slosh test program had a major impact on the final design of the nutation control system flown on INTELSAT IV. The following discussion reviews the significant system design considerations that influenced the nutation control system. The nutation control system must satisfy two fundamental objectives: l) provide basic vehicle stability consistent with the mission pointing requirements of ±0.35° and, equally important, insure stability during the transfer orbit phase of the mission; and 2) rapidly damp nutation transients associated with normal stationkeeping operations. To satisfy these objectives, it is necessary to consider the following major vehicle design and mission related factors: l) variable inertia configuration, 2) transfer orbit operations, 3) apogee boost dynamics, U) allowable steady-state nutation, 5) damper performance, and 6) dedamper magnitude.

The requirement to carry the apogee kick motor in the satellite itself makes the INTELSAT IV fundamentally different from its pioneering predecessor TACSAT I. The presence of the apogee motor severely alters the vehicle mass properties between the ascent phase and the synchronous orbit phase. As a consequence of the mass properties variations, the inertia parameters, which characterize the spacecraft nutation dynamics, exhibit appreciable variations. Table 1 summarizes the critical inertia properties of the spacecraft at certain periods in the mission and shows that, in the transfer orbit, the vehicle has an inertia ratio (rotor spin moment of inertia (MOl) to total transverse MOl) of 0.278, while in synchronous orbit the inertia ratio is initially about 0.351 and subsequently decreases to 0.292 as fuel is depleted. As a direct consequence of this inertia change, the inertial nutation frequency (damper excitation frequency) varies over a 30% range. A further complication is the range of spin speeds over which the spacecraft could be expected to operate. To limit the amount of spin change resulting from fuel expulsion (300 Ib over the mission) and potential jet misalignments, INTELSAT IV has the capability of controlling spin speed via its reaction

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Table 1 Characteristic INTELSAT IV dynamic parameters Synchronous orbit Transfer orbit

Parameter

Start of life End of life

Transverse MOI, slug-ft2

830

5lU

kQ6

2

Rotor spin MOI, slug-ft

230

182

1^2

Inertia ratio, rotor spin MOI/trans. MOI

0.278

0.351 0.292

Rotor spin rate, r/min

50.8

Inertial nutation frequency, rad/sec

I.k8

Rotor nutation frequency,

rad/sec

•

Composite damper time constant, sec

^50

1.83

3.88

~50

1.53

3-U

150 to 200 18 to 60

'

3-75

30 to 90

control system. Without spin speed control, the nutation dampers would have to be moderately broadband low Q devices to span the possible operating frequency range. Since INTELSAT IV must survive the transfer orbit phase of the mission on its own, several critical operational events must be anticipated. They are initial stabilization, perigee pass (out-of-view periods), and apogee boost. As discussed earlier the spacecraft is separated from the booster, autonomously spun up, and stabilized out of command visibility. Since the spacecraft is out of view for nearly 15 min, the system must be capable of performing the critical initial stabilization unaided. To maximize the reliability of the dampers, they are not mechanically restrained at launch. By eliminating an electromechanical caging mechanism, the dampers remain operationally passive, subject only to the despun condition of the platform. As a consequence of the design, any nutation transient induced at separation will excite the dampers, which in turn act to damp out the motion. Since the dampers are mechanically limited to a physical displacement of ±k°, a nutation angle in excess of about 1° will cause the dampers to impact structural stops. Once the dampers begin to operate in an impacting mode, the net damping time constant is reduced. To insure normal linear operation of the dampers, two identical active nutation damper systems are commanded on

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at the end of spinup "by the sequencer. Any nutation in excess of 0.5° "will result in a firing of one or both of the axial jets, which act to reduce the amplitude below a threshold level. The active nutation damper system plays another key role at the end of apogee boost. The AND system will be discussed later.

During the perigee phase of the transfer orbit, the spacecraft is out of view of the TT&C stations and must therefore operate autonomously. To insure that the vehicle is nutationally stable throughout this period, the pseudo Earth circuit of the DCS is ground commanded at the desired rate. The desired rate may be either at the nominal despun value or at some other rate, typically a counterrotating rate, which results' in a tuning of the platform excitation frequency to the resonant operating point of the dampers. Further stability protection is afforded by the AND system. When the spacecraft ascends to the injection apogee, it is prepared for the most demanding dynamical transient of its mission — apogee boost. Although apogee-boosting a Gyrostat stabilized vehicle is similar in principle to a single spinner, the requirements to minimize the injection errors and insure spacecraft stability simultaneously significantly impact not only the vehicle design but also the operational procedures pre- and post-boost. As a result of detailed analysis and digital simulation of the INTELSAT IV apogee boost dynamics,2 two key requirements developed: l) The despun platform must be statically balanced. Furthermore, the platform must be rotating at a small rate during the firing period so as to average out any transverse torques resulting from residual platform imbalance. 2) The post-boost nutation transient would most likely result in nonlinear damper operation; therefore the AND system should be available to augment the dampers at this time. The first requirement has been operationally implemented on the three spacecraft launched to date. In the case of F-2 (the first spacecraft launched in the INTELSAT IV series) and F-3, the platform was counterrotated at 4 r/min. On F-^4 the platform was forward spun to about 2.5 r/min. The later rate was selected to improve the passive stability immediately following boost. During the thrusting phase itself, the friction torque increases to about 1 ft-lb as the vehicle accelerates. The DCS, operating in the relative rate mode, tracks this variable friction and maintains the desired relative rate during burn. Any attendant spin change associated with apogee

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INTELSAT IV NUTATION DYNAMICS

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motor firing -will result in an equal spin change in the platform. The largest spin change observed of the first three spacecraft launched (April 1972) has been 0.5 r/min spin down on F-2. F-3 spun down 0.3 r/min whereas F-U spun up 0.1 r/min. The second requirement to operate with the AND system on during and following burn insures nutational stability for amplitudes exceeding the linear operating range of the dampers . On ~F-k the AND system was observed in operation immediately after boost. The nominal injections of the first three satellites (April 1972) indicate that the foregoing operational and system requirements are compatible with the apogee boost dynamics of the INTELSAT IV satellite. Once successfully in orbit and in service, the satellite must be stabilized so that the basic mission pointing objective of ±.0.35° can be met. To that end it was concluded that the residual steady-state nutation should be no greater than ±0.02°. This requirement manifested itself in a damper design that had no inherent mechanical stiction. Furthermore, the dampers are required to attenuate the nutation with a time constant no greater than 5 min. This requirement influenced the selection of the damper performance parameters, specifically the natural frequency, damping factor, pendulum mass, and physical location on the spacecraft. The ultimate fulfillment of the above objective fell on the predictability of the net magnitude of all dedamping factors on the rotor. The rotor of INTELSAT IV contains a number of readily identifiable (and some not so obvious) nonrigid elements that could provide destabilizing energy losses. It was the objective of an extensive dedamper investigation to identify, categorize, and analyze potential dedampers. The results of that investigation will be discussed in some detail later. The successful performance of the three INTELSAT IV satellites launched to date (April 1972) is sufficient demonstration that all the foregoing requirements have been met and exceeded by substantial margins. Nutation Dampers The INTELSAT IV nutation dampers (Fig. 3) are of the pendulous eddy current type. The operation of the damper is simultaneously initiated by the presence of spacecraft nutation. The attendant motion causes the pendulous magnetic tip mass to move relative to the fixed conducting vane. The relative velocity of the magnet swinging over the aluminum vane induces small electrical eddy currents which circulate

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•within the "body of the aluminum. These currents in turn produce a countermagnetic field to the permanent magnet. A resistive force proportional to the relative velocity is generated -which acts to reduce the motion between the swinging pendulum and stationary vane. The circulating eddy currents generate resistive heat losses in the conducting material. Thus a fraction of the nutational energy is dissipated as heat. The remainder is converted (through the damper reaction torques) to rotor spin energy via the despin control system which acts to keep the platform rate zero.

Fig. 3 INTELSAT IV

nutation dampers.

The mechanical construction of the damper is dictated by the pendulum and its attachment to the structural frame. The damper pendulum is supported on a cantilevered torsion -wire. By dimensionally controlling the beryllium copper torsion wire, the natural frequency of the compound pendulum can be accurately established. At the free end of the wire journal bearings support the pendulum loads induced by "bottoming" during launch and apogee boost. In a 0-g environment, the wire has sufficient bending stiffness to center itself within the journal bearings. This centering action eliminates any nutation deadband resulting from stiction forces. The harplike structure, which supports the magnet pendulum assembly, contains the impact stops mentioned earlier. These spring loaded stops are preloaded to 5 lb and have sufficient stiffness to absorb transient impact velocities on the order of 10 ft/sec.

The basic performance of a damper of this type can be evaluated either by numerical integration of the vehicle equations of motion or by the more expedient energy sink approximation method. By use of the latter approach, it can be shown^ that the effective time constant of the nutation damper for a despun platform is given by (l - N 2 2

mR

(1)

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From Eq. (l) it is seen that the time constant depends on three spacecraft parameters Irp, X^9 R and three damper parameters m , ^ , and con. Tables 1 and 2 give all the critical parameters required to evaluate Eq. (l).

The two dampers on INTELSAT IV have "been designed with different performance characteristics. The reason for desiring variable damper performance was to optimize the performance for synchronous orbit operation without compromising transfer orbit stability. The lower nutation damper is tuned to the synchronous orbit inertia configuration and, since spin speed can be controlled, the Q of the oscillator is increased so as to improve its performance. This improvement was deemed necessary following the fuel slosh test program. The upper damper is tuned to a compromise frequency between transfer and synchronous orbit inertia conditions. The Q is lowered to cover a broader frequency range, which results in a reduction of performance. To offset some of the reduced performance, the upper damper is located at the maximum radial distance, R. As a consequence of the lower damper tuning and the broader bandwidth of the upper damper, the latter provides the primary stabilizing time constants in the transfer orbit. Performance evaluation tests are conducted on each damper after unit final assembly, and following spacecraft acceptance testing, to insure that each unit performs to specification. A rather simple laboratory test was devised which has proven to be quite adequate for predicting in-orbit time constants to within a 10$ accuracy. The procedure consists of a force free deflection test wherein the transient pendulum motion is recorded on a strip chart. From this test it is possible to determine both the percent overshoot and therefore the damping Table 2 INTELSAT IV nutation damper performance parameters

Parameter

Upper damper

Lower damper

Pendulum mass, slugs

O.OU65

0.01+65

Natural frequency, rad/sec

1.63 ±0.02

1.89 ±0.02

Fraction of critical damping

0.38 ±0.12

0.08 ±0.02

Distance between pendulum eg and spacecraft eg, ft Synchronous orbit Transfer orbit

8.37 10.0

6.55 Q.h

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ratio, £,, and the damped natural frequency from which, given £>, the undamped natural frequency can "be determined. To minimize gravity effects on the pendulum frequency, the damper is accurately mounted perpendicular to the local vertical. The pendulum is vertically off-loaded by a low inertia suspension system which lifts the pendulum through its eg. The data are then adjusted for a deterministic coulomb friction force associated with the test setup. The predictability of the damper performance is influenced by the following parameters: magnetic pole piece dimensions, flux density, air gap width, mass, pendulum frequency, and conducting vane resistivity. The latter parameter, which is temperature dependent, accounts for the majority of the uncertainty in the value of £,. The dampers are positioned on the mast such that they are openly exposed to the space environment. Subsequently the dampers are expected to operate over a temperature range of -100° to +250°F. The resulting variation in £,is ±30% about the ambient value.

Given the foregoing parameters, the composite damper curves for the synchronous and transfer orbit are shown in

Figs, k and 5. Figure k shows that the optimal performance at resonance yields a composite time constant of about l8 sec, but because of temperature uncertainties the time constant could be 36 sec. In synchronous orbit, resonance occurs at a rotor spin speed which is given by the lower damper natural frequency divided by the inertia ratio. It should be pointed out that at resonance the energy sink model is conservative and somewhat inaccurate. The difficulty with the model is that it assumes the nutation angle is not changing to first order over one nutation cycle. With time constants on the order of 20 sec and nutation periods on the order of 3 sec,

the assumption is not valid. In this regime numerical integration provides a more accurate damper time constant. When the excitation frequency exceeds the resonant frequency by about 2%, the energy sink model is quite adequate.

As seen in Fig. 5 in the transfer orbit at about 50 r/min, the composite damper time constant varies between 150 and 200 sec. As noted earlier, this is a consequence of the off resonant condition brought about by the transfer orbit inertia condition as well as the increased transverse MOI.

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INTELSAT IV NUTATION DYNAMICS

Fig. h Synchronous orbit damping time constant vs rotor rate (despun platform).

71

Fig. 5 Transfer orbit damper time constant vs rotor spin rate (despun platform).

Active Nutation Damping

The INTELSAT IV active nutation damping system is similar in design and implementation to the one flown on ATS-V. The main elements of the control loop are the following:

ACCELEROMETER

CONTROL ELECTRONICS

SPACECRAFT DYNAMICS

REAC TION CON!•ROL SYST EM (AXI>XL JET)

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The AM) system consists of two independent systems mounted on the spinning rotor. Each system consists of an accelerometer, a signal conditioning unit (active nutation damper electronics, ANDE), and the utilization of an axial jet. The methodology of actively controlling nutation is the following: l) A sensor detects the presence of nutation and establishes the phase and amplitude of the motion with respect to a rotor-fixed coordinate system. 2) The sensor output signal is threshold detected, amplified, and converted to a jet command. 3) The axial jet fires once per nutation cycle for some fraction of the cycle which results in the application of a transverse torque in opposition to the nutational motion. The basic technique used for active nutation control is illustrated in Fig. 6. As the transverse rate vector, corp, rotates clockwise at rotor nutation frequency with respect to a rotor fixed x-y plane, a rotor-fixed jet develops a torque impulse which, on the average, opposes the sense of the coijj vector; thus the residual transverse angular momentum is removed and the nutation angle decreases.

Although the AND system has not "been required to maintain nutation stability during normal mission operation, the design implementation has been verified in orbit. Following the apogee motor burn of the F-U spacecraft, the AND system was observed in operation reducing the 1° nutation transient to threshold, 0.^2°. On the first two spaceDIRECTION OF U ROTATION @ DIRECTION OF ROTOR SPIN craft , the AND opera12 = ( < r - 1 ) c o = ROTOR NUTATION FREQUENCY tion was not observed, but the accelerometer signals received shortly after thrust 'ACCELEROMETER termination strongly suggest that the nutation was actively reduced to threshold. The feasibility of manually controlling APPLIED JET TORQUE nutation was demonstrated on F-2 as part 0 AXIAL JET of a checkout of both the passive and active control system. Fig. 6 Active nutation control concept. T

s

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Prelaunch Dedamper Investigation Prior to the launch of INTELSAT IV, an extensive search was conducted for major sources of energy dissipation in the rotor. Much of the effort was tailored after the TACSAT I and ATS V postlaunch investigations and utilized similar analyses and test techniques where applicable. Primary use was made of the energy sink analysis technique for approximating the energy dissipation rates of a nonrigid element on a spinning nutating body. When either the analysis yielded a result of significant concern or proved to be too dependent on critical unjustifiable assumptions, hardware tests were conducted. These tests fall into two basic categories: those that were qualitative and those that were quantitative. A qualitative dedamper test typically consisted of verifying a critical assumption, i.e., giving factual credibility to a reasonable engineering judgment. An example of a test in this category is a test to measure the frequency response and/ or damping factor, £, of some complex piece of rotor mounted hardware such as a conical sun shield, flexible thermal barrier, or a solid propellant engine. Since the natural frequency and damping factors are critical parameters required in the energy sink calculation, direct measurement often served to rule out certain candidates as not significant and, on the other hand, elevate any suspect dedamper to the next phase of test verification — the quantitative test. At one point in the INTELSAT IV study, the list of potential dedamping devices exceeded a dozen. It was possible, however, to narrow the list to about a half dozen potentially significant dedampers. Table 3 lists these major dedampers, gives an order of magnitude comparison with the dampers, and describes the method used to evaluate their respective dissipation rates. The structural dedampers, including the apogee motor, were all found to be several orders of magnitude below the dampers. The energy sink approach was used exclusively in evaluating these dedampers. In certain cases it was a difficult analytical task to evaluate the lowest natural frequency of some complicated piece of hardware like a sun shield. At such a point a test was conducted on the subject piece to determine its frequency response. The test also provided the damping factor of the mode excited. In principle, the test is analogous to the damper test discussed earlier. The conclusion of all the structural dedamper studies was that the rotor was an essentially rigid body and whatever energy dissipation was present it was negligible, relative to the damper capabilities.

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Table 3 INTELSAT IV dedampers Dedamping source

Order of magnitude relative to dampers

Evaluation technique/ comment

Fuel slosh

Laboratory tests on scaled vehicle. Time constants directly measured and scaled to flight. Corroborative flight data validate test techniques.

BAPTA

BAPTA compliance test, similar to TACSAT test.1 Indicated INTELSAT IV BAPTA was either neutrally or positively stabilizing. Assumed negligible.

Structural damping

10-3

Apogee motor

10-3

Modal analysis supported by frequency and/or damping tests •where necessary to verify analysis. Analysis supported by measurement or propellant loss tangent.

In the case of the BAPTA, the analytical difficulties in estimating internal energy dissipation were numerous. Since it was known that the TACSAT I BAPTA exhibited a mechanical anomaly that ultimately led to a 1° steady nutation motion, the INTELSAT IV BAPTA was given special attention to assure that it did not contain a significant destabilizing effect. To this end a quantitative dynamic test was conducted on a special fixture. The test, similar to the one used to evaluate the TACSAT BAPTA,1 consisted of applying a rotating (at inertial nutation frequency) moment on the BAPTA despun section. Two orthogonal proximitors provided a measure of the BAPTA deflection. The resulting data were interpreted and reduced using plots of angular deflection vs applied moment. The area enclosed by the plot is the energy transferred between the BAPTA and the test fixture. The direction of motion determines the sign of the energy transfer. It was concluded to first order that the BAPTA was neutrally stabilizing, i.e., had no effect on the growth or decay of the

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INTELSAT IV NUTATION DYNAMICS

nutation. To second order the BAPTA had a small stabilizing influence on the nutation. Since this effect is beneficial, it was treated as a bonus damper and not included in any of the subsequent stability margin calculations. After the BAPTA had been cleared as a major dedamper on INTELSAT IV, attention focused on the potential dissipation resulting from the hydrazine sloshing in the four conispherical tanks. A test was set up at Hughes to quantitatively measure the energy dissipation in conispherical tanks mounted to a test vehicle. Figure 7 shows the vehicle on an air bear ing in a high altitude chamber. The energy dissipation rate, E, can be determined by measuring the time rate of change of the nutation angle, 9 (i.e., the diverging time constant T) , and using the relationship

- near the resonant value ELECTRONICS for maximum damper perWINDOW formance while the fuel is being depleted and throughout this maximum energy dissipation region. FREEDOM AIR B E A R I N G

It is worth noting that the fuel slosh tests indicate that

Fig. 7 Fuel slosh test fixture and instrumentation.

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spherical tanks of comparable dimen4 sions do not hold 2. 4 TANKS, d = 1.44 FT, EQUAL LOADirJG TANKS ASSUMED significant dedamp2 3. CURVE GOOD OVER SPIN RANGE ing advantages over OF 40 TO 60 R/MIN 103 — the conispherical SPHE -SP^HE RF 8 — K -/ tank. Where the 6 f ( / / 4 conisphere peaks / r out near 65% frac2 tion fill, the \ 7c ONIS PHER E 102 — rw sphere peaks out at BEST ESTIMATE OF FLIGHT 8 — ' DEDAMPER TIME CONSTAN 6 50$ fraction fill. 4 At its worst the SPACECRAFT sphere is about a A F-2 2 V F-3 factor of 3 better I I I | 1 I | 1 1 I I I ml than the conisphere 0 20 40 60 80 100 at its worst value. ff, PERCENT FRACTION FILL Presently only limited data exist on inertia ratios Fig. 8 INTELSAT IV fuel slosh other than INTELdedamping time SAT IV, but the constant. data that are available indicate that the fluid dissipation is very sensitive to this dimensionless parameter which controls the nutation excitation frequency. On future spacecraft designs incorporating spin stabilization about an axis of minimum moment of inertia, fuel sloshing effects remain a potentially significant source of dissipation and as such must be accounted for in the control system design. » 6

1.

SYNCHRUIMUUS UKBI 1 UUIMMLiUKAl IUIM

a-= 0.355, IT = 515 SLUG-FT2

I

From the INTELSAT IV experience, it has been firmly established that simple fluid mechanical models are wholly inadequate when attempting to describe the rotating fluid dynamical motion occurring in propellant tanks. Experimentation and empirical extrapolation provide the only meaningful engineering solution to bound the problem for a specific design. The three INTELSAT IVs launched to date have corroborated the above conclusions. Fuel slosh is the major dedamper on the spacecraft. The laboratory test results agree well with the flight data. The successful operational performance of satellites is in itself a conclusive demonstration that the nutation control system is functioning as designed.

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Flight Nutation Data Description of Instrumentation Information on the spacecraft nutation is provided "by an analog sensor mounted to the rotor. The accelerometer is a force "balance closed loop transducer in which an acceleration applied to a seismic mass is counter"balanced "by an electromagnetic force generated by a current from the servo loop. The current passed through a suitable resistor provides a voltage proportional to the acceleration. The seismic element contains a capacitive sensor and the force restoring coil. The seismic element is supported by three equally spaced sets of flexures forming a kinematic parallelogram. This permits a translational linear motion in the sensitive axis with virtually no mechanical hysteresis.

To understand the following data better it is necessary to appreciate the telemetry data link between the accelerometer and the strip chart recorder.on the ground. Figure 9 indicates the flow of the accelerometer output through the entire telemetry circuit. The accelerometer output is coupled to a signal conditioner in the telemetry system via a lead filter. This filter acts to remove the d.c. bias resulting from misalignments. In the signal conditioner the voltage is amplified by a commandable factor of 1 or 20 and used to modulate an FM subcarrier oscillator which in turn phase modulates the telemetry beacon transmitter. The signal is then transmitted to the ground where it is demodulated and put through another signal conditioner which has three gain

Fig. 9 FM accelerometer signal flow.

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settings: XI, XU, and X10. Finally the data are displayed on a strip chart recorder. The data may be passed through a low pass filter at this final point in order to suppress the spacecraft and link induced background noise. Since the nutation sinusoid of interest is around 0.5 Hz, a three-pole filter with a 2-Hz corner and 20-db/octave roll-off sufficiently suppresses the high-frequency noise that the nutation sinusoid can be detected down to the arcsec regime. However such resolution is only possible after an amplification of 200 times in combination with the low pass filter.

Time Constant Data The nutation time constant data from the F-2, F-3, and F-lj. spacecraft have been assembled into one graph, Fig. 10. Since the rotor spin speed at resonate nutation frequency differs slightly for each spacecraft because of differences in the vehicle inertias, the abscissa is dimensioned in rotor r/min from the resonant value. The data points have been adjusted to eliminate dedamping effects. The dedamper time constant assumed was 175 sec. This value is derived later. It is seen that the data all fall within the expected bound based on damper performance predictions. The data scatter is primarily associated with temperature variations which influence the damping factors.

-2

-1

0

+1

+2

o R E S DEVIATION OF ROTOR SPIN RATE FROM DAMPER RESONANCE, R/MIN S/C

DATA

ITSFT 2

F-2 F-3 F-4

O D A

515 515 515 0.38 0.27 0.50

0.355 0.351 0.355 0.08 0.06 0.1

u)RES 50.8 R/MIN 51.5 50.8

FRACTION FILL 73-75% 73-75% 73-75%

175 s 175s 175 s

1.89rad/s

Fig. 10 Synchronous orbit damper time constant.

The transfer orbit time constants are given in Fig. 11. Here the spin range is enlarged to encompass the expected operating range. Since the three inertia ratios are nearly equal, the abscissa is dimensioned in r/min. This is a consequence of the off resonance performance being insensitive to cr variations. Again the dedamper has been subtracted out of the data. The value assumed was 600 sec. This time constant is

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determined in the next section. Dedamper Time Constant

= = = = =

825 SFT2 0.277 10' 8.4' 0 R/MIN

Normally the composite damper time constant dominates and masks the dedamper. The only method of establishing the magnitude of the dedamper is to effectively degrade the damper performance until it is roughly comparable to the dedamper. When this condition is established, the net time constant becomes quite long. As mentioned earlier, the

FLIGHT DATA OF-2 DF-3 AF-4

0.27 0.08 0.38 0.08 0.50 0.08 NOTE: THE'LONGESt TIME CONSTANT IS THE NET VALUE MEASURED, Tn. THE LOWER TIME CONSTANT IS THE ESTIMATED DAMPER TIME CONSTANT, To, AFTER THE 600 S DEDAMPER IS REMOVED.

W s , R/MIN

time constants add reciprocally, so if the composite damper

Fig. 11 Transfer orbit net damper time constant vs rotor rate.

time constant is T^ and the dedamper time constant is TDTJ, then the net time constant, T is given by T

DTDD

!_ = _JL_ T

TT^

DD

(3)

T

DD - TD

Equation (3) can obviously be solved for one gets

if desired and

Typically TD is known to about 10$ accuracy and T is measured experimentally to a ^% accuracy, However the error on T-pp, as can be established by differentiating, is strongly dependent on the difference T - T^, If this is a small difference, the error in TTJJJ is large. The only -way to maximize the difference between T - TJJ is to find a condition where Tp is slightly less than T ,

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The operational technique used to measure the dedamper is to spin the platform. With the platform spinning, the excitation frequency as seen by the passive dampers can be adjusted to achieve near neutral vehicle stability. Equation (5) is used to determine the damper time constant when the platform rate is nonzero. * -

[d - N» 2 r + HC, n - j ~

(5)

mR"

where N ' = X /u>^ 2

2

1 - co /co p n

9,

b£

f

= bL

The two dampers can be individually evaluated and summed as before. Thus, given the same parameters as before and knowing the platform rates, the damper performance can be predicted.

By incrementally varying the platform rate, starting with the despun condition and inducing nutation, the net time constant can be measured directly from the accelerometer. The composite damper time constant is computed and then the dedamper estimated from Eq. (k) . As the damper performance degrades and the dedamper becomes more noticeable, the net time constant will rapidly start to deviate from the predicted damper curve. The results of a dedamper test on F-3 are presented in Fig. 12. Here the platform was counterspun to h r/min. Intermediate data points were -2 and -3 r/min. The results of

the test show the trend described above. The net time constant despun was hO sec, whereas at -h r/min, the net time constant was almost 1200 sec. From these data the best estimate for the dedamper is 175 +15 sec. As noted in Fig. 10, these data points correspond to a fraction fill of 75% and agree quite well with the fuel slosh test results shown in Fig. 8. The test results were conclusive in isolating the magnitude of the dedamper to a high accuracy and simultaneously identifying the major source of dedamping as fuel sloshing. A similar dedamper test was -conducted on F-U in the transfer orbit. However, to effectively degrade the damper

Purchased from American Institute of Aeronautics and Astronautics

o t* NJ

" IT

ivj

o

( P/ kRAMETER

*> O

00

QCUg

"Cu ^L

>

= 515 S Fl[2

\ rDC)

^H

REDK :TED

i

- 0.351

; T DC

= 49.05 R/ MIN

Hx andu>y the corresponding expressions for the axisymmetric vehicle in Eq. (3), that is, a;

x

=

u>

o

COS At

u>

, y

=

w

o

sin At, a; = X 9

o

(8)

Thus , to lowest order, Eq. (7) "becomes

lx

r

I i //

(cos At 6h

^

y

-- sin At 6h ) dt

^

(9)

The equivalent expression for the transverse rate change

O

=

A.

27T

rJ W^ n

c^ I /~X7cos At 6h__

/ Jr\

\ J X. \\ ^

B

sin At 6h,, A

Note that Eq. (lOa) reduces to Eq. (10) in the event that the two inertias, A and B, are equal. 3. Autotracking Antenna

For a dual-spin vehicle with a tracking antenna, the angular momentum about the system e.g. can be written fe =

I

I

(tf + fl£) + n-^ x ^ + I2« + m^2 x r^

+ I (w + ^J + m r x r - mr x r

(17)

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EFFECT OF AN AUTOTRACKING ANTENNA

97

Now r is the vector from the coordinate system origin to the system e.g. But the origin is chosen to coincide with the system e.g. when the antenna is correctly positioned in azimuth and elevation relative to the platform, that is, if the vehicle is not nutating and the antenna is pointed in the desired direction, r = 0. Thus r and r take on nonzero values only as the antenna moves about its nominal position, relative to the platform, in an attempt to track out nutation. Hence, to first order mr x r = 0 £,

= u) x £-, £2

=

£ x £o' ^

=

£3

+

g

~

( 1 8 )

'-, ) X U

Combining (17) and (18) yields

h = ^

where

(19)

Ju> + J w 1 + C 1On ~ -L^ - ^

J = I

I+ ml \ ~l ' ~1 I "^1-1 / + i2+ m2

+ I 3 + m 3 ( r 3 • r^-^^ J

^

f

l = X3 + m 3 {^3 ' tt'-^

and where all the matrices and vectors are resolved in the platform xyz system depicted in Fig. 2. The xyz coordinate system is fixed relative to the platform with origin at the frozen system e.g. and with the z axis parallel to the bearing axis. In addition, it will be required that the x and y directions be selected so that the J matrix in Eqs. (19) and (20) has the approximate representation / 0 J

=| 0

B

I_

1

(21)

I

that is, the xy axes are rotated so that the 1, 2 term of J is zero (also, the 2, 1 term is zero, since J is symmetric).

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98

J.E. McINTYRE AND M.J. GIANELLI

This zero condition will hold only for one antenna orientation; as the antenna orientation changes, the 1, 2 term will develop some nonzero value. However, it is required only that the J matrix approximate the form in Eq. (21), i.e., that the 1, 2 term be much smaller than either the 1, 1 or 2, 2 term. Hence, in what follows it will be assumed that the coordinate axis has been chosen so that the J matrix is approximated by Eq. (21) with J(l, 2) and J(2, l) very nearly zero over the antenna pointing envelope. If we follow the method of Sec. III.2, the h vector is written as

h

= h + 6h

where ho is a nominal nutating condition about which the averaging is conducted and 6h is a perturbation which causes the nutation decay. The nominal condition is taken to be the nutational motion of the vehicle (with no motion of the antenna relative to the platform) and with the platform itself completely despun. Thus, from Eqs. (19) and (21)

(22)

where ho is characterized by the steady nutational motion x

=

cu

o

cos At

J(X2A1) sin At X =

Subtracting (22) from (19) and noting that in xyz coordinates takes the form

Purchased from American Institute of Aeronautics and Astronautics

T

yields 6h as

e + J (1,

xz z 6h

=|

I yz z

I

xzU)x

2,2 +

I

yzU> y

" ^ " Cw u)

z

e +

"*"

where jj ^ T 12 denotes the 1, 2 element of the matrix [j]_ ^^J and where ^ is evaluated at any particular antenna setting for which the stability is to be analyzed.

Finally, substituting (25) into (lOa) yields the nutation time constant as I/T

=

(l/7s)

7257'

+ (l/T^T

where 27T/\

' dt

cos

27TU,

z

^^

(27a)

cos \t 2 7TW

2

sin \t| Ve dt A (2?b)

27T

X"

^(2,3) cos At

^(1,3) sin At

dt

(27c]

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100

J.E. McINTYRE AND M.J. GIANELLI

The determination of the rates u>z, e, and f requires the three control equations governing the platform azimuth motion and the antenna azimuth and elevation motion. Under the assumption that these motions are small, the control equations can be linearized to yield

L[X3gL1 p

|~ETI

L

S

"1

L 3g J

(28b)

e + [X El

J3,2

e + ["ETI

L

row

s

E]

-13,3

V+^J/l

L

-I 3rd' row

| w . . l =

T,(28c)

where J]_ is evaluated in platform coordinates and I3g in antenna coordinates. (The matrix Igg is very nearly equal to the antenna inertia about the gimbal center. It differs slightly from this condition because of the motion of the system center of mass relative to the coordinate system origin. For the averaging method it suffices to use for I3g the antenna inertia about the gimbal center.) By using the zeroth order expressions forwx and v given in Eq. (23), Eqs. (28) can be solved for the quasi-steady-state values of e, ^, and

(30a)

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EFFECT OF AN AUTOTRACKING ANTENNA

antenna elevation. error rate

101

= - fe + A_(2,lL JL x + A J_ (2,2)w ~y + A_(2,3)u> J_ z 1I (30b) L J

antenna azimuth error rate

e +A 1 (3,l)u> x +A^ 3,2)0^^3,3)0/1

(30c

•while, for the RR mode ,

platform error rate

=

- [ * eos e + ^(3.1)^+ ^(3,2)^+

Ola)

antenna elevatic elevation = - fe + A,(2,l)u> + A (2,2)o, + A (2,3)w 1 , I J_ X JL V JL Z I error ra.te »^ _j rate antenna azimuth rate

= -^

(31b)

(31c)

Since the rates u>x, o>y, e, V 3 and u>z vary sinusoidally at nutation frequency, the applied torques may be represented as a gain and phase shift operating on the appropriate error signal. Hence, m T

s

T 1

e

= I I^ K 1 s

; v4 \ s '

—

' ve /

I IV PT 1

' e

-,

A (h

\ }€

where the gain and phase are evaluated torque response at nutation frequency. spin torque is the despin error signal by gain Ks and shifted in phase by the

e

from the open loop For example, the deincreased in amplitude angle s.

Equation (28) now may be solved by the method of undetermined coefficients, and the results substituted into Eq. (27) to produce the desired time constants. For the variables u>x andu> v , the zeroth order expressions in Eq. (23) are used. The procedure one would follow to determine the stability contribution of an autotracking antenna would involve the

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102

J.E. McINTYRE AND M.J. GIANELLI

repeated solution of Eqs. (27) and (28) for the different antenna orientations within the tracking envelope. This solution can be accomplished analytically, although a computer evaluation of the time constants vould be more efficient, since the algebra is messy and many computations are required. Furthermore, since the time constant has been resolved into the three components, rs , r e , and r^5 it vould not be too difficult to determine what type of a fix was needed in the event a stability problem arose. Unfortunately, Eqs. (27) and (28) provide no immediate design information or physical insight regarding the stability of different system configurations. However, some information of this sort is obtained under the idealized condition in which the antenna tracks perfectly.

k.

Perfect Track Situation

In the perfect track case, the antenna control loops perform with zero error even while the vehicle is nutating. This is an idealized situation, but one which should be nearly satisfied for a good autotracking system. Both tracking configurations described in Section II will be analyzed using this assumption. The CAR mode will be treated first. The perfect track condition requires that the inertial motion of the boresight axis be zero. Thus,

from which it follows that

e

=

s ^u>

- c fu>

(32a)

- tan e fcfwx + s

Substituting Eqs (32) into Eq. (28c) and using the error signal of Eqs. (30a) provides

w z = w o T sin (\t + 17 )

(33)

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103

EFFECT OF AN AUTOTRACKING ANTENNA

where

r

+ k

= \c

p

\2

P

p

A kj + K - 2k0\ K cos 3/ s 3 s

r -

cos

r

sin 77

= 3

cos

-K sin

=

=

k

sin

cos

=

I

xz

+

" Jl(3'3)

| Jl*

- ^(3,3) tan e s f

I yz - I 3^ C

tan

2 ~ Ji(3'3)

Finally, substituting Eqs. (32) and (33) into (27) yields the time constant expressions as -AT

cos 17

2V AB

1 1

1

- X

(35a)

(35b) 2,2

,/O r c 77 + tan e s ^ (35c)

- ^(2,3)

F 377 + tan e c

- c fa; ,

^ = 0

(47)

Thus the elevation time constant is again governed by Eq. (38) with l/re =

(-A/2^AB) ^'(1,2)

( 4 8 )

while the azimuth expression becomes

I/T,7 = 0 V

( 4 9 )

The elevation time constant is the same as for the CAR mode since it operates off the same error signal. The azimuth expression is zero, because with the azimuth loop holding perfectly there is no motion of the antenna in azimuth relative to the platform.

The platform azimuth motion can again be represented as u>

z

= cj

o

F sin (\t + rj )

where F and 77 are determined from Eqs. (28a) using the ideal rates of Eqs. (47) and the despin error rate from Eq. (31a). This evaluation yields

Purchased from American Institute of Aeronautics and Astronautics

108

J.E. McINTYRE AND M.J. GIANELLI

22 2 + KS c e - 2Cc~0\ K0c.cec(A cosrj

= _

sin 77 =

-

- KS (se) c ^ s < i >S + KS se sy ^1 * / v / \ c.0 ' / v -_Li ' ^S

- KS (se) c^c* - SK se s^ S

^j

\/U A ) s6 i_ 0_L 1 S

For either the CAR or the RR tracking modes, the perfect

track approximation allovs for a reasonable assessment of the antenna's effect on system stability. Because of this, the perfect track assumption can be used during a preliminary design stage to estimate the time constants associated -with the antenna azimuth and elevation motion [i.e., Eqs. (38) and (^5) or (^9)]. The small tracking errors resulting from imperfect antenna servo loop performance play only a minor role. The dominant effect comes from the vehicle mass geometry arrangement. To complete a preliminary design stability assessment, some knowledge of the platform1s azimuth motion is required. The constants T and 77 can be evaluated using either Eqs. (3*0 or (50), provided all the parameters are known — an unlikely situation during a preliminary design stage. However, a crude estimate can be made by assuming the antenna is locked and treating the platform/antenna system as a rigid body. This

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EFFECT OF AN AUTOTRACKING ANTENNA

109

estimate, together with the perfect track time constant expressions, will allow an early indication of potential problem areas.

5.

Stability Dependence on Mass Geometry

So far, the analysis has focused on the development of mathematical expressions governing the nutation time constant. In this section, two examples of stable and unstable configurations are given in an attempt to illustrate the strong dependence of system stability on vehicle mass geometry. Consider the vehicle shown in Fig. 5a containing two identical antennas symmetrically positioned with respect to the x-z plane and with both antennas at zero azimuth and zero elevation relative to the platform such that the coordinate system of each antenna is parallel to the platform system. Let the right antenna (marked ® in Fig. 5a) have position vectors £3 and u, given by.

; antenna (5) u =

(u

Ji.

(5l)

0, 0)

Then the left antenna (because of the symmetrical arrangement about the xz plane) will have vectors

; antenna (E) u =

(52)

(u , 0, 0) Jv

In addition, require that l) each antenna is locked in azimuth and tracks perfectly in elevation, and 2) the antenna coordinate system is an antenna principal axis system (i.e., the matrix 13 is diagonal). Now in the presence of a nutational transient, the perfect track condition requires that both antennas move in identical fashion relative to the platform. Because of this identical motion, one would expect the two antennas to contribute in similar manner to system stability. That this is not the case follows from Eq. (38) which, using Eqs. (51) and (52), becomes {l/T

(ly

e>right antenna = U/2V1S)m3V3

e'left antenna =-(\/2 VAB)m u^ = -(l/re)right antenna

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J.E. McINTYRE AND M.J. GIANELLI

110

GIMBAL CENTERS

Thus, the elevation motion of the antenna on the right is stabilizing, while that of the antenna on the left is destabilizing.

The above example illustrates that the stability behavior of an autotracking antenna can depend rather strongly on which side of the platform the antenna is mounted. The phys-

©

a)

0, u

y

= 0

'ENNA (T)

ical reason for this dependence can be explained in terms of the reaction torque arguments of Sec. III.l. If one works through the analysis, it is a simple matter to show that the x component of the reaction torque which antenna ® exerts on the vehicle is equal and opposite to the x component exerted by antenna ©. It is this difference which is responsible for the differing stability effect of the two antennas. Consider next the slightly different geometrical arrangement shown in Fig. 5"b. The two antennas are again identical, are both at zero azimuth and zero elevation and symmetrically arranged relative to the xz plane. The vectors and u in this case become

D) ux = 0, uy

right antenna

0

Fig. 5 Top view of platform antenna geometry for simplified case.

u = (0,u0) j

(53)

Purchased from American Institute of Aeronautics and Astronautics

EFFECT OF M AUTOTRACKING ANTENNA

u

=

(0, -u

«y

111

; left antenna

0)

In addition, assume that l) each antenna is locked in elevation and tracks perfectly in azimuth, 2) the antenna coordinate system is an antenna principal system (i.e., I_ is diagonal). Again, each antenna's autotracking motion is identical, yet the stability contribution of each is different. From Eq. (1*3),

m ~juy z ~j cos ( ? ? + « ) =

(- \T/2'\&B) m u z sin 77 •~) j ~)

=

-90°

right antenna

cos

(+ M/2^AB) m u z sin 77 -3 j

or =

O

left antenna

+90°

Thus

right antenna

antenna

and the stability properties of the antenna on the right are equal and opposite to those on the left. As before, a physical explanation for this behavior can be obtained from the reaction torque arguments of Sec. III.l.

The foregoing two examples indicate a strong dependence of stability on vehicle mass geometry. In Fig. 5a5 the stability contribution changes sign as the antenna is moved from the right to the left side of the platform, while in Fig. 5b, a second sign change occurs as the antenna e.g. is moved from outboard to inboard of the gimbal center along the platform y axis. Note also in Fig. 5a that a l80° rotation in azimuth

Purchased from American Institute of Aeronautics and Astronautics

J.E. McINTYRE AND M.J. GIANELLI

112

or elevation of either antenna changes the sign of the time constant. Thus the stability is also strongly dependent on the relative orientation of the antenna with respect to the platform.

IV.

Example Problem

To illustrate the previously described dynamical effects, consider the example dual-spin configuration shown in Fig. 6 with mass properties and control loop data given in Table 1. While the example problem parameters do not pertain to any specific dual-spin BEARING A X I S vehicle, they are representative of the present generation of gyrostat vehicles. There are , however, several seemingly contrived features of ANTENNA e.g. this example which require explanation. GIMBAL CENTER 1) The platform

GIMBAL CENTER

PLATFORM e.g.'

ROTOR e.g.

Fig. 6

Geometry for example problem.

without the antenna has a sizable product of inertia in the xz plane. Although this mass geometry is not typical, it is included here to insure azimuth motion of the platform and antenna in the presence of nutation. Such a platform product of inertia is present on the TACSAT 1 communication satellite. 2) The azimuth gimbal of the antenna is not coincident with the bearing axis. This geometry is representative of a complex communications

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EFFECT OF AN AUTOTRACKING ANTENNA

113

Table 1 Example problem parameters /200.

i,.! o. \0.

(

-0.0909 >

-0.0909 | m

0.

200. 0.

(

o.\

0.

250. / /

/

-3.227 y \

300.

0.

i+o. \

0.

300.

0.)

Uo.

0.

\ 1.772 y

Ao.

250 y

0.

o.\

7-0.0909^

15-

0.

!3 - 1 0. \

0.

^ '

15. y

£2

r_

^le" ^6

= I

\

1 m2 = 25 slugs

-0.0909

= I1 1.1*09 l

= 2 5 slugs

)••(' • )

5-772 y'

\1.5/

m

= 5 slugs

n= 60 RPM Antenna shaping network open loop transfer function (both azimuth and elevation) GA(s) = .A.

[3000. (0.2 s + l)/(0.025 s + l)]

Despin shaping network open loop transfer function

G™Q(S) = rJote:

[150. (2.5 s + 1)/(0.2 s + l)]

Vectors given for antenna at zero azimuth and elevation condition.

satellite -with several antennas; the single antenna shown in Fig. 6 can be viewed as one element on such a system. 3) The control loop models contain neither motor dynamics nor friction. In a tracking application, the motor time constant must be very small compared with the nutation period. Hence, in evaluating nutational behavior, the motor drive can be modeled as a pure gain. Since the control loop gains are large, friction will not appreciably affect the control loop performance and can be neglected. Since the control loops and the mass properties of the individual bodies (i.e., platform, rotor, and antenna) are

Purchased from American Institute of Aeronautics and Astronautics

J.E. McINTYRE AND M.J. GIANELLI

114

fixed, the example problem illustrates the dependence of system stability on antenna orientation and tracking mode. The stability is evaluated using three distinct methods: averaging, eigenvalue, and digital simulation. For the eigenvalue analysis, the system Euler equations and the equations governing the relative motion of the rotor and the antenna were combined with the actual control loop representation to form the relevant set of differential equations. These equations "were then linearized and the associated eigenvalues determined. The nutation time constant is the reciprocal of the real part of the eigenvalue corresponding to the transverse rate vector. The accuracy of this time constant was verified at several points with a complete nonlinear digital simulation which is essentially an upgraded version of the one described

in Ref. 9.

'o

cc

0.0

3.75

7.5

11.25 15.0 18.75 22.5 . TIME (SEC)

26.25

30.0

o

MO

0.0

o

EC) UTH

O

b

AT^

2 §

Data for the CAR tracking mode are presented in Figs. 7 — 9 . Simulation results for the zero azimuth-zero

0.0

4.0

8.0

12.0 16.0 20.0 24.0 28.0 TIME (SEC)

a) Antenna tracking errors

"b) Despin control loop performance

c) Nutational response

d) Spin axis inertial motion

Fig.

7 Digital simulation results for CAR mode.

32.0

Q ME CONSTANT, r g , SECONDS

TIME CONSTANT, r .SECONDS

O 1-3

O 1-3

Purchased from American Institute of Aeronautics and Astronautics

i-b 3 O CD 4

O CO O *

1

V

\

P

H

I

p

_l

>

\\

C

s

\ \ V Ii

T I M E C O N S T A N T . T .SECONDS

X

/

_6

> 1^ i

\

£

(D H' OQ

CO P

8b

3d

I

OQ Q

P b

0)

H- MJ P 4

O O 1=5 CO dP

d- 0)

HO1 £ dHO

H3 HO O

CO O

.1

0.0

-0.1 -0.2 ELEVATION ANGLE (HAD)

-0.3 -O/

Time constant contribution from despin control loop for CAR mode — averaging method.

Purchased from American Institute of Aeronautics and Astronautics

116

J.E. McINTYRE AND M.J. GIANELLI

Fig. 9& System time constant for CAR mode from averaging and eigenvalue methods.

Fig. 9b

0.1

Time constant contribution from elevation tracking for CAR mode — averaging method.

0.0

-0.1 -0.2

-0.3

E L E V A T I O N ANGLE ( R A D )

Fig. 9c

Time constant contribution from azimuth tracking for CAR mode — averaging method.

Fig. 9d

Time constant contribution from despin control loop for CAR mode — averaging method.

Purchased from American Institute of Aeronautics and Astronautics

EFFECT OF AN AUTOTRACKING ANTENNA elevation case are shown in Fig. 7»

117

Observe that the control

loops track rather well, effectively cancelling 95% of the platform motion. The nutation angle decays fairly rapidly with the time constant evaluated from the x component of the transverse rate to be approximately 65 sec.

The averaging and eigenvalue results are shown in Fig. 8a

for different antenna orientations about the zero azimuth-zero elevation condition and in Fig. 9a for orientations about the l80° azimuth-zero elevation condition. The eigenvalue and simulation time constants agree perfectly at the one compari-

son point, while the averaged time constants are more conservative than those computed by the eigenvalue method. This might tempt one to discard the averaging in favor of the eigenvalue approach. However, the advantage of the averaging analysis is that the total system time constant can be broken down into antenna elevation and azimuth components and a platform azimuth component. This breakdown is shown in Figs. 8b-d and 9~b-d. The antenna elevation component will be discussed first. The elevation time constant behavior is best explained in terms of Eq. (38). Because 13 in antenna coordinates is diagonal, the 13T(l,2) term in Eq. (38) is always zero. Hence for perfect track l/re

= (\/2\fxB)

m.

Now if the shift in system e.g. location resulting from changing antenna orientation is neglected, then the mass properties in Table 1 indicate that ux' varies only with elevation, and y3* varies only with azimuth. Furthermore, ux? increases as elevation is increased from 0 to ^5° and decreases as elevation is decreased from 0 to -^5°. The variable y3 , on the other hand, decreases monotonically as ^ varies from 0 to l80° with yx at l80° equal and opposite in sign to y3 at 0°. Hence, two antenna settings which differ only in azimuth and be a l80° angle should yield elevation time constants which are negatives of each other. Such a trend is indicated in the plots of Figs. 8b and 9"b with slight deviations resulting from elevation tracking errors and the fact that the vehicle e.g. shifts as the antenna orientation is changed.

The antenna azimuth time constant for the CAR mode is shown in Figs. 8c and 9c. A detailed explanation for this time constant behavior is difficult, since even under the perfect track assumption of Eq. (^5) the parameter dependence is complex. However, the perfect track model can be used to

Purchased from American Institute of Aeronautics and Astronautics

118

J.E. McINTYRE AND M.J. GIANELLI

show why the antenna positions about the zero azimuth condition are stabilizing, whereas those about the l80° azimuth condition are not. From Eq. 2

I F cos ( "H + oi ) - tan e sin

(2,3)

( ot -

where a is defined in Eq. (kk) . Now from the mass properties data in Table 1, it is a straightforward matter to show that

^(1,3)

~-cf

J1(2,3)

~-s^

provided the system e.g. shift with antenna orientation is neglected. Thus, for \^_ and A 2 approximately equal

a - $ ^TT

sin (a - $ ) -

Q

and the azimuth time constant expression becomes =

A/2VAB

V ( V X 2 } J d . S ) + J 2 ( 2 , 3 ) Tcos (17+*)

From this expression it follows that if $ is changed by l80° and all other parameters are fixed, the time constant Ty will change sign. Of course, all other parameters are not fixed. In particular, n varies withf . However, for all the antenna orientations shown, *i falls in either the first or fourth quadrant. For this reason T^ is positive about the zero azimuth position of Fig. 8c and negative about the l80° azimuth position of Fig. 9c. For a physical explanation of the sign change in r^ , one must resort to a reaction torque argument. The relative azimuth motion of the antenna with respect to the platform is much the same for the antenna at either the zero or l80° azimuth setting. However, the change in azimuth does change the sign of the inertia terms J]_(l,3) and J1(2,3), and it is through these terms that the antenna loop reacts on the platform. Thus the reaction torque and, therefore, the time constant change sign with a l80° rotation of the antenna. The final data set for the CAR mode are the time constants caused by platform azimuth motion. These data are shown in Fig. 8d for the antenna in the vicinity of the zero azimuth position and in Fig. 9& for the 180° position. Note that both conditions are stable, but that the effect is more

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119

EFFECT OF AN AUTOTRACKING ANTENNA pronounced at the l80° condition.

To partially explain this

behavior, neglect for the moment the coupling of the antenna motion into the platform despin loop. Under this condition and assuming A]_ and \ 2 are nearly equal, Eqs. (3*0 and (35a) indicate that

yz

xz

•where Ixz and IyZ are defined in Eq. (2l). Now from Table 1 and Fig. 6 it is a simple matter to show that the quantity (lxz2 + Iyz2) is three to four times larger at the l80° azimuth condition than at the zero azimuth condition. For this reason, the time constants in Fig. 9d are smaller than those in 8d. Data for the RR mode is presented in Figs. 10 — 12. Recall that in the RR mode the azimuth track is accomplished /> -f\J V \7^A /^s-Systems, Vol. 26, edited by N.E. Feldman and C.M. Kelly, The MIT Press, Cambridge, Mass., 1971, Chap. IX, pp. 605-653. 4

Russell, W. and Dahl, P., "An Evaluation of the Comparative Merits of Dual Spin and Momentum Wheel Stabilized Spacecraft for a Synchronous Communications Satellite," Aerospace Corp. Rept. TOS-0059 (6141-02)-2f March 1971.

Purchased from American Institute of Aeronautics and Astronautics

MOMENTUM WHEEL THREE-AXIS ATTITUDE CONTROL

1 61

^Dougherty, J.; Lebsock, K.L., and Rodden, J.J., "Attitude Stabilization of Synchronous Communications Satellites Employing Narrow Beam Antennas," Journal of Spacecraft and Rockets, Vol. 8, No. 8, Aug. 1971, pp. 834-841. 6

Perkel, H., "Stabilite - A Three Axis Attitude Control System Using a Single Reaction Wheel* ' AIAA Progress in Astronautics and Aeronautics: Communication Satellite Systems Technology, Vol. 19, Academic Press, New York,

1966, pp. 375-400. ^Cummings, W.D., Barfield, J.N., and Coleman, F.J., Magnetic Substorms Observed at the Synchronous Orbit," Journal of Geophysical Research, Space Physics, Vol. 73, No. 21, Nov. 1968. ^Coleman, W.D. and Cummings, P.J., "Stormtime Disturbance Fields at ATS-1," Journal of Geophysical Research, Space Physics, Vol. 76, No. 1, Jan. 1971, pp. 51-62. , A, "Nutation Damping Dynamics of Axisymmetric Rotor Stabilized Satellites," ASME Winter Meeting, Chicago, Nov. 1965. , V. and Stewart, B., "Nutational Stability of an Axisymmetric Body Containing a Rotor," Journal Spacecraft and Rockets, Vol. 1, No. 6,

Nov-Dec.1964, pp. 682-684. ^Martin, E., "Fuel Slosh and Dynamic Stability of Intelsat IV," AIAA Paper 71-954, Aug. 1971, Hempstead, N.Y.

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PROPULSION REQUIREMENTS FOR COMMUNICATIONS SATELLITES William C. Isley* and Kenneth I. Duckt NASA Goddard Space Flight Center, Greenbelt, Md.

Abstract The concept of characteristic thrust is introduced herein as a means of classifying propulsion system tasks related particularly to geosynchronous communications spacecraft. Approximate analytical models are developed to permit estimation of characteristic thrust for injection error corrections, orbit angle relocation, north-south station keeping, east-west station keeping, spin axis precession control, attitude rate damping, and orbit raising applications. Performance assessment factors are then outlined in terms of characteristic power, characteristic weight, and characteristic volume envelope, which are related to the characteristic thrust. Finally, selected performance curves are shown for power as a function of spacecraft weight, including the influence of duty cycle on north-south station keeping, a 90° orbit angle relocation in 14 days, and finally comparison of orbit raising tasks from low and intermediate orbits to a final geosynchronous station. Power requirements range from less than 75 w for north-south station keeping on small payloads up to greater than 15 kw for a 180-day orbit raising mission including a 28.5°-plane change.

Presented as Paper 72-514 at the AIAA 4th Communications Satellite Systems Conference, Washington, D.C., April 24-26, 1972. The authors gratefully acknowledge the assistance of D. L. Endres of NASA Goddard Space Flight Center in review of the manuscript and verification of the analytical expressions contained herein. *Head, Systems Analysis Section. fAerospace Engineer, Systems Analysis Section.

165

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Nomenclature RS Aa Wd Vs M T Fc 6 e At w0 Ae0 Ae7 N Ft Fr AI 0 AI t Fn AX tc 00 T

= geosynchronous radius (1.38335619 x 108 ft) = orbit semimajor axis minus geosynchronous radius (ft) = mean daily satellite longitudinal drift (deg/day) = geosynchronous satellite velocity (10,087.405 ft/sec) = satellite mass (slugs) = allowable maneuver time (sec) = characteristic thrust (Ib) = true anomaly angle at midpoint of thrust application (rad) = one-half thrusting arc (rad) = thruster on time per correction pulse (sec) = geosynchronous orbital rate (0.72921979 x 10"4 rad per sec) = average eccentricity change per orbit = total eccentricity change per maneuver = total number of orbits required for correction = tangential thrust magnitude (Ib) = radial thrust magnitude (Ib) = average inclination change per orbit (rad) = total inclination change per maneuver (rad) = normal thrust magnitude (Ib) = longitudinal displacement angle (rad) = coast time (sec) = satellite longitude taken with respect to Earth minor axis (rad) = normalized time

J22 // Re A0p Iz wz d

=

T^

= disturbance torque level (f t-lb)

77 I * Ad Wf

= thruster operational duty cycle = moment of inertia (slugs-ft2) = magnitude of rate damping per maneuver (deg/sec) = final payload weight (Ib)

harmonic of the Earth triaxial gravity field (-1.68 x 10"6) = Earth gravity constant (1.40764433 x 1016 ft 3 /sec 2 ) = Earth radius (2.092564 x 107 ft) = total precession angle per maneuver (deg) = spin axis moment of inertia (slugs-ft2) = spin axis angular rate (rad/sec) = thruster moment arm (ft)

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PROPULSION REQUIREMENTS I sp AV r V0 V Pc ^pc a Wp Wt Wc Ec ]3 Dp Vse

167

= propellant specific impluse (Ibf-sec/lbm) = effective required velocity increment (ft/sec) = initial parking orbit circular velocity (ft/sec) = instantaneous orbit circular velocity (ft/sec) = characteristic power (kw) = Power conditioning efficiency = tankage weight factor = propellant weight (Ib) = system tare weight (Ib) = characteristic system weight (Ib) = characteristic volume envelope (ft 3 ) = tankage volume factor = propellant storage density (lb/ft 3 ) = thruster modules volume envelope (ft 3 ) I. Introduction

It has been well established from past experience in flight programs that optimum choice of a propulsion system is closely linked with specific mission application requirements as well as consideration of technology readiness. The diverse number of propulsion tasks, in addition to the complex (and often subjective) nature of the selection process for flight applications, has made it extremely difficult to identify hardware development goals (particularly for electric propulsion systems). The development of new concepts to the point of prototype hardware availability involves- considerable cost and schedule lead time in order to factor potential use into future missions. However, the seeming paradox is that hardware development must reach this state before any serious consideration can be expected from flight programs. Since only limited funding is now available for such purposes, it is especially important that technology goals be clearly identified where most likely benefits will accrue in near-term missions. This problem is approached herein by a proposed new method of classifying propulsion tasks and relating them to projected flight applications. Estimation techniques are presented for selected propulsion tasks which have been identified for geosynchronous communications satellites. Systems performance assessment factors are then described and related to current system concepts.

II. Historical Background Early satellite propulsion applications employed cold gas systems, since they 1) required little or no electrical power from a satellite that had severely limited maximum power capability, 2) met minimum operational lifetime total impulse requirements of one year or less, and 3) offered no major environmental or interface problems with the spacecraft. Such uses generally involved

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unloading of momentum wheels, attitude control of spinning satellites, and vernier orbit adjustment. As more stringent missions evolved, it became evident that cold gas would lead to unacceptable system size and weight in order to satisfy requirements. This was a direct result of the low propellant efficiency (specific impulse) nominally in the range of 60 sec for nitrogen and from the large tankage volume needed at maximum permissible storage presures. Hydrogen peroxide was subsequently used for similar reasons to achieve higher total impulse capability within reasonable tankage size and weight but was still limited by a specific impulse less than 130 sec. Also, hydrogen peroxide was found to be quite sensitive to chemical breakdown and produced a number of operational problems for long life service. With the development of Shell 405 catalyst, hydrazine became a major candidate for extended mission requirements, since the propellant is stored as a liquid, requires little electrical power, and produces a specific impulse ranging from 150 sec in pulse mode up to 225 sec at equilibrium operation. Extensive development was pursued at thrust levels in the range of 5 Ib and, more recently, at thrust levels down to 0.5 Ib and 0.1 Ib. There are also potential applications of bipropellant systems that could produce in excess of 300 sec specific impulse over a thrust range similar to that of hydrazine.

Consideration of electric propulsion systems can bring a new dimension into the tradeoff requirements for flight applications. This results from their ability to produce substantially higher propellant efficiency (specific impulse greater than 3000 sec) at the expense of additional electrical power consumption. It is also possible to somewhat adjust specific impulse and thrust level as a function of available power. The major problem has been that past flight programs have severely limited power for onboard propulsion to less than 50 w. On such a basis, one could not expect to realize any significant performance gain over competing chemical approaches since the total impulse requirements were low. This meant that development goals had to be established from the standpoint of unique capabilities of low thrust systems rather than performance gains. Typical examples were the ammonia resistojet systems flown on Application Technology Satellites (ATS) 1, 3, 4, and 5 for eastwest station keeping, and the cesium contact ion system flown on ATS 4 and 5 for a similar function. In the latter case the satellites were gravity gradient stabilized and low thrust was considered essential to eliminate the possibility of tumbling the satellite due to a misaligned thrust vector. Also, for the Synchronous Meteorological Satellite (SMS-C), a teflon propellant pulsed plasma system is now under development for east-west station keeping and precision precession control. The unique application here is the requirement to maintain spin axis pointing to within 5 arcsec of a given reference. The impulse bit capability of the pulsed plasma is well suited to such a thrusting requirement. Once again, the total impulse requirement is quite low and little performance gain is expected. The first real application of electric propulsion that offers a significant performance improvement is the cesium bombardment ion system experiment under development for ATS-F. In this case a 1 mlb thrust engine at 2500 sec specific impulse will be used to demonstrate northsouth station keeping of the spacecraft. For extended operation the total

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impulse requirement in this case becomes large, meaning that competing chemical systems would require substantial propellant weight to perform the same task. This weight advantage can be factored into cost benefits assuming availability of operational hardware. With the advent of lightweight and compact rollout or foldout solar arrays in the 1- to 15-kw range, it is now possible to consider electric propulsion applications at higher thrust levels and/or at the higher specific impulse values. Prospects for competitive use of electric propulsion as a "single onboard propulsion unit" is no longer a farfetched dream. As power availability increases for propulsion use during various parts of the mission sequence, it becomes possible to consider smaller launch vehicles to perform the same mission and even deliver a larger payload. By borrowing power from the spacecraft during maneuvering phases, electric propulsion can serve to cover the full range of propulsion tasks without reducing substantially power availability to a communications system during normal operation. III. Classification of Propulsion Tasks

Propulsion requirements have previously been categorized by either total impulse or by design thrust level on a system basis. This approach does not

work effectively for low thrust application, since it does not take into account the time allowed to perform each particular operation. On this basis, the term, characteristic thrust, is defined, and is the total impulse required for a given maneuver divided by the allowable time to perform the maneuver. Note that maneuvering time is not necessarily the thrusting time, as this definition permits combinations of thrusting and coasting phases. Therefore, Characteristic Thrust is a quantity tied to particular propulsion tasks and is not necessarily specified on a system basis. It is now possible to classify the various tasks performed by onboard propulsion systems during the entire mission sequence. In order to arrive at the various categories, it is necessary first to apply certain assumptions as to the mission applications and availability of supporting technology in power systems. The technology base is herein related to a maximum launch vehicle delivery capability of approximately 3000 Ib into a geosynchronous orbit (equivalent to the Titan IIIC capability). Payloads with capabilities of up-rated Delta and smaller also apply to this classification. This should be interpreted as primarily an economic constraint applicable over the next decade, which in effect eliminates present consideration of larger launch vehicles than Titan III-C. On the basis of the foregoing considerations, mission application requirements can be classified (somewhat arbitrarily) according to the levels of characteristic thrust. Bracketing has been achieved by establishing maximum values of allowable maneuver time, which in turn are dictated either by ground station requirements, mission sequence requirements, or arbitrary specification by projects. Four classification categories have been established to meet the current projection of mission application requirements. These are described in the following sections.

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Class I Systems (Characteristic Thrust > 1Q-2 Ib)

Typical mission requirements under this class include:

1) Correction of injection errors into geosynchronous orbit for payloads of the Titan IIIC capability. 2) Large orbit angle relocation for geosynchronous spacecraft for payloads of the Titan IIIC capability. 3) Gross satellite orbit raising missions covering a range from low altitude through geosynchronous altitude for the full spectrum of available launch vehicles. 4) Near-geosynchronous in-orbit network rapid replacement satellites for the full spectrum of available launch vehicles.

5) Near-geosynchronous orbit maneuvering spacecraft for remote maintenance and satellite inspection missions for the full spectrum of applicable launch vehicles. 6) Inclination adjustment and nodal drift compensation for sunsynchronous Earth resources mapping satellites. 7) Nutation damping control for spinning spacecraft having unfavorable inertia ratios for the full spectrum of available launch vehicles.

8) Spin-up and despin control of spacecraft of size within pay loads of the Delta and Titan IIIC. 9) Rapid spin axis precession maneuvers over large angles for pay loads covering the full spectrum of launch vehicles.

10) Attitude reference acquisition maneuvers for 3-axis stabilized spacecraft for pay loads of the Delta and Titan IIIC class. 11) Long-term orbit correction for sun-synchronous spacecraft having pay load capability of up-rated Delta. Class II Systems (Characteristic Thrust IP"2 - IP"3 Ib) Typical mission requirements under this class include:

1) Correction of injection errors into geosynchronous orbit for payloads of the Delta and Delta (up-rated) capability. 2) Small orbit angle relocation and drift cancellation for geosynchronous spacecraft having pay loads of the Titan IIIC capability; large orbit angle relocation for Delta Class payloads.

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3) Vernier orbit raising or lowering missions in nearby geosynchronous orbit for the full spectrum of available launch vehicles.

4) Orbit period adjustment for sun-synchronous spacecraft having requirements for variable ground track using Delta vehicle capabilities.

5) Drag cancellation for low altitude satellites for payloads up to capabilities of the Delta. 6) Attitude maneuvering and tracking control for three axis stabilized geosynchronous satellites for payloads of the Titan IIIC capability.

7) Coarse dumping of momentum storage devices for spinning and nonspinning spacecraft. Class III Systems (Characteristic Thrust IP" 3 - IP"4 Ib) 1) North-South station keeping of geosynchronous spacecraft for payloads of up-rated Delta and Titan IIIC capabilities.

2) Drag cancelation for high-altitude satellites for payloads up to capabilities of the Delta.

3) Inversion of gravity gradient stabilized geosynchronous spacecraft for up-rated Delta through Titan IIIC payload, capabilities. 4) Precision attitude stabilization for geosynchronous satellites for full spectrum of payload capabilities. 5) Vernier dumping of stored momentum in spinning and nonspinning spacecraft.

Class IV Systems (Characteristic Thrust < IP"4 Ib) 1) East-West station keeping of spinning geosynchronous spacecraft for the full spectrum of payload capabilities.

2) Active damping of librations for gravity gradient stabilized spacecraft. 3) Precision spin axis orientation and precession control for spinning spacecraft.

4) East-West station keeping of three axis stabilized geosynchronous spacecraft for the full spectrum of payload capabilities.

Propulsion tasks can furthermore be subclassified under each of the above categories into 1) continuous thrusting mode or 2) pulsed thrusting mode wherein requirements for thrust vector deflection can be imposed in

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either case. The approach for hardware configuration and operational modes is herein considered to be a part of the optimization process for particular mission applications. IV. Assessment Criteria .for Selected Propulsion Tasks

This section presents analytical approximations useful in sizing characteristic thrust for selected propulsion tasks. Specific forms for the analytical models were developed by the authors, primarily in unpublished work. However, the mathematical foundations are in most cases an application of the Gauss or Euler equations that can be found in numerous texts and technical journals.2"9 For this reason, there was no attempt to cite references in the body of this work, but rather to list typical sources for further study by those interested. Also, physical units of feet, pounds, seconds, and degrees were primarily employed in this treatment rather than the metric system, because it was felt that project engineering personnel are more familiar with use of such units. 1) Injection Error Corrections

All-chemical ascent propulsion trajectories normally employ a low altitude, near-circular initial parking orbit with subsequent burn that produces a transfer ellipse having its apogee at near-synchronous radius. The transfer orbit is usually inclined to the equator at an angle of around 28.5° (which is compatible with a Cape Kennedy launch). Apogee burn is normally timed so that the geosynchronous longitude at injection will fall as close as possible to the desired station location. Apogee burn removes all but a small inclination error and provides a near-circular final orbit with a period close to 24 hr. Injections can be deliberately designed to produce orbit periods less than, or greater than, the synchronous Earth period so as to generate a mean drift rate eastward or westward to achieve final station longitude. Terminal maneuvers therefore are required for the following functions: 1) drift cancellation and/ or period adjustment, 2) orbit angle relocation, 3) eccentricity removal, and 4) inclination adjustment. These requirements stem from the dispersions (errors) produced by the launch vehicle/kick stage guidance and propulsion systems. It is generally of prime importance to stop unwanted drift of the satellite, since it could move out of a safe communications and tracking region relative to given ground stations. In addition, unwanted large angle drift would necessitate either an extended recovery time or significant onboard propellant consumption to reach terminal station.

Because of an injection error in semimajor axis, a satellite will experience a mean daily drift given by wd = - 540 (Aa/R s ) A positive A a will generate a westward drift. The characteristic thrust required to cancel an unwanted drift can be expressed by

(1)

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Fc = Vs M wd/1080 T

(2)

The second step is to remove orbit eccentricity. If one assumes a maximum allowable correction time of less than 15 days, external disturbances from gravity field harmonics and solar/lunar sources can be neglected.' Under such conditions, eccentricity change per orbit can be written as Ae^ = (4 Ft cos 0 sin e)/M w0 Vs

(3)

The angle 6 represents the true anomaly at midpoint of thrust application and angle e is the arc moved through by the satellite during one-half a thrust cycle. For near-circular orbits, one can approximate the thrusting arc by the equation

2e « w0 At

(4)

The total eccentricity change requirement can be written as Ae~t = N Ae^

(5)

Also, characteristic thrust can be expressed as Fc = N Ft At/T

(6)

Combining Eqs. (3-6) produces Fc = (M Vs Ae^/2T cos 0) (e/sin e)

(7)

Thus, characteristic thrust is found to be a function not only of the expected total eccentricity change requirement, spacecraft mass, velocity, true anomaly angle at midpoint of thrust application, and allowable correction time, but also of the length of the angular arc during each thrusting cycle. Removal of inclination errors is the final step of the correction sequence and would involve for low-thrust systems the use of alternately fired (two correction pulses per orbit), paired nozzles directed north and south to permit essentially continuous operation over an orbit. The change in inclination per orbit which can be produced in the absence of external gravitational fields is Al^ = (2 Fn At/M Vs) (sin e/e)

(8)

Since the inclination corrections are assumed to be made in a relatively short period, one can essentially ignore the influence of external forces. By using similar expressions, as for Eqs. (4-6), characteristic thrust can be expressed for inclination adjustment as Fc = (M Vs AVT) (e/sin e)

(9)

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W. C. ISLEY AND K. I. DUCK

Figure 1 presents characteristic thrust vs allowable maneuver time for one typ-

ical eccentricity and inclination adjustment assuming a 2500-lb spacecraft in

near-geosynchronous orbit. 2) Orbit Angle Relocation

This task involves relocation of the operating station longitude starting

with a geosynchronous orbit and ending in a geosynchronous orbit. Tangential and combined tangential/radial thrusting schemes can both accomplish such maneuvers. Figure 2 presents a phase plane diagram for such a propulsion task. It provides a trace of subsatellite longitude angular rate and dis-

placement. Consider a station change (case A) from 100° west long to 40° east long. This can be performed by a tangential thrust application opposing the satellite orbit velocity direction, so as to produce a smaller period and yield an eastward drift. At thrust termination, one has the option of a coast phase for accumulation of longitude displacement. At a predetermined time,

thrust in the direction of satellite velocity would cancel the drift rate at the desired new station. A similar maneuver is shown for generating a westward drift to 160° west long. Case B illustrates a shorter thrusting arc with longer 0.1

1 DAY

D

O ^

INCLINATION CORRECTION

'

0.01

ECCENTRICITY CORRECTION

D

QC I

O

cc

111 o

0.001

M 0 6

= 77.7 SLUGS (2,500 LBS) = 0 7T

" T A^ = 0.01

O

= 0.00872 RADIANS (0.5 DEGREES)

0.00001 104

i 105

106

i

i i i i 11 107

ALLOWABLE MANEUVER TIME (T)

Fig. 1 Characteristic thrust requirement for eccentricity and inclination adjustment maneuvers.

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PROPULSION REQUIREMENTS

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A < o COAST CASE B /

WESTWARD 160°W

y

I 100°W

F t < oV

\

COAST CASE A Ft = 0

F t > O/

COAST

175

^ \Ft>0 \

40° E

N

X

EASTWARD

/

Ft = (J22 ACCELERATION NOT INCLUDED)

Fig. 2 Geosynchronous station relocation phase plane. coast interval to perform the orbit angle relocation task. As the coast phase increases, time to perform the maneuver will increase, but characteristic thrust required will decrease. Therefore, it is a matter of specifying an allowable

operating time to achieve the new station. In polar coordinates such a

maneuver would be represented by a spiral trajectory which produces inter-

mediate orbits having an average period slightly different from the geosyn-

chronous period. For such small changes in radius, one does not have to be concerned with phasing of thruster operation to offset large eccentricity buildup.

For low thrust acceleration equal to or less than 10~2 ft/sec 2 , one can

approximate large angle changes by

3 T (T + tc) 4 RSM

T2

4 R s Vs M 2

(10)

Equation (10) can also be rearranged to provide a solution for tc. Therefore,

given spacecraft mass (M), allowable maneuver time (T), and the required longitude change (AX), one can establish the coast time as a function of Fc. If there is only interest in the minimum allowable time to perform a relocation (i.e. no coast phase), Eq. (10) reduces to

|FC| = 4 |AX| Rs M/3T2

(11)

Figure 3 plots the minimum time required for station relocation against the magnitude of the angle change for selected values of characteristic acceleration

O

> ^

5 H
cr x Q. LU LU ^

10.0

0.25 1.0

< CO

0.1

2000

4000

I 6000

8000

10,000

SPACECRAFT WEIGHT (IBS)

Fig. 7 Characteristic thrust requirement for east-west station keeping.

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W. C. ISLEY AND K. I. DUCK

plasma systems, operating in the characteristic thrust regime from 25 - 50 pdb, are well suited to this propulsion task. 6) Limit Cycle 3-Axis Attitude Control Three-axis attitude stabilization can be performed using either fixed nozzle pulsed mode operation as for resistojets or by beam deflection techniques as for the ion engine. Fixed nozzles can also be canted in pairs so that effective beam deflection is accomplished by varying the pulse rate of each nozzle. For soft limit cycle operation (SLC), an impulse bit is provided by the thruster at one deadband extremity to offset the disturbance impulse. Figure 8 presents a phase plane representation of SLC operation. Characteristic thrust level associated with SLC mode is given by Fc = Td/r?d

(19)

In Fig. 8, three cases are shown to illustrate the influence of 77 on phase plane trajectories. A value of 77 greater than 0.5 is generally not acceptable for control purposes. Smaller values than 0.5 are acceptable, and the selected value normally depends upon the margin of safety which one desires in overcoming disturbances. Generally, a value of 0.1 should be more than adequate for sizing purposes. 7) Attitude Rate Damping Attitude acquisition following orbit injection and possibly a despin operation normally require attitude rate despin damping. Characteristic thrust for this maneuver can be expressed as Fc = IA0/57.3 Td

Fig. 8 Soft limit cycle phase plane trajectories.

(20)

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Generally, the allowable maneuver time is determined by the field of view of attitude sensors. 8) Orbit Raising

An orbit raising maneuver primarily relates to transfer of a satellite from an intermediate or low-altitude near-circular, inclined orbit to a near-circular, equatorial synchronous orbit. Conventional all-chemical high thrust methods were briefly described under Part 1 of this section, wherein each firing occurs over only a small arc of the orbital path. Low thrust orbit raising, however, requires essentially continuous thrusting over many orbit revolutions, thereby generating a transfer spiral. One has the options of coplanar spiral, followed by inclination adjustment or combined spiral and inclination adjustment. More advanced studies also have examined elliptical initial parking orbits to maximize payload gain or minimize transfer time. For characteristic accelerations of 10 ft/sec 2 or less, and assuming near-circular initial parking orbits, one can estimate the characteristic thrust level required for such a maneuver by the expression FC = (Wf I S p/T) (eAVg!sp _ l)

(21)

For a coplanar transfer without inclination change, eccentricity correction, or compensation for nonthrusting arcs due to solar shadowing, AV r is equal to the difference in circular orbital velocities at initial and final conditions. Edelbaum1 has developed the AV r associated with combined spiral and inclination removal, which has the form AV r = \/vo -

2V v

o

cos

(* AI t/ 2 >

+ y2

(22)

Figure 9 presents a plot of AV r against initial circular orbit altitude in nautical miles utilizing Eq. (22). This assumes continuous thrust application of constant total magnitude with no shadow regions, and no compensation for

eccentricity adjustment. Passage of large solar arrays through the Van Alien radiation belts during repeated orbit revolutions creates a degradation in the

solar cells, resulting in a decrease of power that is estimated to range from 15% to 35%, depending primarily upon choice of initial parking orbit and thrust level. To utilize the higher initial power capability properly, an electric stage would have to be designed for incremental thrust level operation, probably combined with vernier thrust adjustment for power matching. A beam deflection capability (now in development status) would permit attitude control during the extended thrusting phase. Specific applications studies are required to properly assess the guidance, shadowing, eccentricity buildup and correction, and solar array degradation in the radiation belts. Such activity is now in progress at NASA Goddard Space Flight Center and elsewhere.

V. Performance Assessment Factors

The characteristic thrust technique developed in Sec. IV for sizing typical propulsion tasks can now be applied to performance assessment of selected

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W. C. ISLEY AND K. I. DUCK

184 20000 C/5 Q_

18000

o EJ 16000

o C/5

cc 14000

I O

INCLUDES 28.5° PLANE CHANGE

12000

100

1000

5000

INITIAL CIRCULAR ORBIT ALTITUDE (NM)

Fig. 9 Characteristic velocity increment requirement for geosynchronous orbit raising fron circular parking orbits. hardware systems. This section will describe briefly the major tradeoff parameters of characteristic power, weight, and size of equipment. In order to properly evaluate competing systems, one should consider both hardware performance capabilities and applications factors which are mission peculiar. Optimization categories for hardware performance are: 1) maximize operational life expectancy within requirements set by the propulsion tasks; 2)minimize system weight, power, and volume; and 3)maximize the number of combined uses of the system in terms of characteristic thrust requirements (operational flexibility and .complexity). Applications factors relate primarily to spacecraft interface and environmental considerations (e.g., EMI/RFI, special thermal control, special structural requirements, impingement, deposition on surfaces, and field interactions). Interaction with other spacecraft systems could create operational problems that would eliminate a given propulsion candidate from further investigation on any given mission. For this reason, it has not been possible to develop a cookbook procedure that will identify best systems for new flight applications prior to a detail design stage for the composite spacecraft. It is, however, possible to trade off relative merits in hardware performance so that, applications factors being acceptable, a given approach could be recommended for use. It is proposed herein that the performance assessment factors include the following: 1) characteristic system power, 2) characteristic system weight, and 3) characteristic volume envelope. Characteristic power generally is expressed in the form (23)

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Both Kp and 7?pC are quite sensitive to hardware system type, propellant specific impulse and operating thrust level. The characteristic system weight is generally expressed in the form

Wc = Wp (1 + a) + Wt

(24)

Characteristic system volume envelope is generally expressed in the form EC = (W p /D p ) (1 + 0) + Et

(25)

It is readily apparent that the three performance factors given by Eqs. (23-25) are specific application dependent. System hardware technology status is obviously an essential part of any meaningful performance assessment. The propulsion analyst must match demonstrated system capabilities against requirements set for each propulsion task. The previous discussion herein has dealt with methods of sizing a selected task and establishing on a composite basis the characteristic thrust range for a given mission. Table 1 presents a summary of low thrust system technology status which can be used to develop comparative assessments of performance. Since weight and volume envelope considerations are highly configuration dependent, they are not included in the table. As a first attempt, one should establish comparisons based upon total existing system capabilities. As a general rule, thrust level and/or propellant storage capability will not match a given new task, so that the analyst must evaluate the extent of design and/or performance modifications necessary to meet acceptance standards of the project. Many of the systems shown in Table 1 have in the past or will be in the near future flown on a given mission application. Other remaining systems have been developed and tested in ground facilities to the point where a known level of confidence exists in operational performance and reliability. Typical examples of this are the SMS pulsed plasma units14 that have been tested beyond flight requirements and the EOS DG-2 cesium bombardment ion system which completed a 8149-hr continuous endurance test. It is evident that thrust levels up to the 1-mlb operating region are available at specific impulse capability up to 6700 sec. At higher thrust levels resistojets offer new capabilities, but at specific impulse no greater than 3 times existing all-chemical approaches. The ion systems are attractive for applications where kilowatts of onboard power can be made available. The SERT II flight demonstrated the operational capability of mercury bombardment systems at 6.3-mlb thrust level and 4200 sec specific impulse. In-flight operation of 5 1/4 months was achieved on one unit and 3 months on the second unit. More recent activity has involved large-scale assembly of a breadboard system utilizing SERT II type thrusters, but including three thruster modules, two power conditioning units, gimballing mechanisms, and related control hardware. This work is in progress at JPL.25 To solve the life-limiting mechanisms and extend thruster operation beyond one year, NASA Lewis Research Center has initiated an endurance test for a 30-cm electron bombardment ion thruster (nominal 27mlb) at Hughes Research Laboratory.26 The operating life goal for this test has

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186

Table 1 Low thrust system technology statusa Ref.

System designation

Solid teflon pulsed plasma (Fairchild Ind.)

10

Thrust level, mlb

0.0038 (0.667pps)

System power input, w 2.1

Specific impulse, sec 220

Cesium contact button ion (EOS)

11

0.020

35

6500

Cesium contact button ion (HRL)

12

0.020

35

6700

Ammonia resistojet (AVCO)

13

0.050

11

130 (S.S.)

Solid teflon pulsed plasma (Fairchild Ind.)

14

0.0507 (1.835pps)

23

405

100-julb cesium bombardment ion (EOS)

15

0.102

43

3900

Solid teflon pulsed plasma (Fairchild Ind.)

16

0.14 (2.0PPs)

50

1000

17

0.37

64

1920

8-cm cesium strip ion (HRL)

18

0.50

125

5000

1-mlb MESC cesium bombardment ion (EOS)

19

1.00

141

2500

20

6.3

977

4200

DG-2 cesium bombardment ion (EOS)

21

7.07

1217

5000

High temperature resistojet

22

10.00

200

23

20.00

5-cm mercury bombardment ion

(HRL)

15-cm mercury bombardment

ion (SERT II)

(MARQUARDT) Hydrazine resistojet (TRW)

8

750 (S.S.-H 2 ) 350 (S.S.-NH 3 ) 230 (S.S.) 175 (P.)

^S.S. = steady state; P. = pulsed, minimum impulse bit, low duty cycle; pps = pulses per second.

Refer to Ref. 24 for JPL study on additional candidate auxiliary propulsion systems.

been set at 6000 hr, and, if completed satisfactorily, will exceed most life requirements for the orbit raising applications associated with geosynchronous satellites. Since ion propulsion systems operating in the kilowatt power regime appear to offer attractive capabilities for future geosynchronous communications satellites, a characteristic power function was developed by the authors for estimating payload performance. This expression is given by 0.0403 Fr (2.05 x 106 +

1 + 26.3 Fc - 227

(26)

The power conditioning efficiency is actually a function of the power input, but generally ranges between 0.82 for small thrusters up to 0.90 for thrusters in the millipound region. Equation (26) does not take into account the differences between thruster efficiency for cesium engines and mercury engines, since available test data indicates that uncertainty in power requirement for

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PROPULSION REQUIREMENTS

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larger engines is of greater magnitude than variations caused by the propellants. Care must be taken to use values of Isp which meet operating life requirements for each thruster type and thrust level considered. One can modify Eqs. (24) and (25) so that they are functions of Fc rather than Wp, yielding

Wc = (Fc T/Isp) (1 + a) + Wt

(27)

EC = (Fc T/Isp D p ) (1 + 0) + Et

(28)

It is now possible to evaluate ion propulsion systems for specific tasks using known system capabilities. Figure 10 illustrates characteristic power requirement for ion propulsion applications as a function of satellite weight in geosynchronous orbit. East-West station keeping for ±0.1° tolerance is shown for continuous thrust applied at the strongest triaxial field acceleration point. North-south station keeping is presented as a function of thruster duty cycle to maintain inclination within ±0.1°. A similar plot could readily be generated to show the increased power requirement for a canted thruster configuration (such as on ATS-F) where there is a significant radial component. A 90° station relocation in 14 days is included to illustrate typical power requirements. Four cases of orbit raising, are included to show that it is possible to spiral from a low or intermediate altitude circular parking orbit to geosynchronous orbit in a period of time ranging from 90 to 180 days if power is available in the range frqm 3 - 1 5 kw. Thrust level requirement for this application would range from about 20 - 200 mlb. The performance calculation assumed an operating specific ORBIT RAISING MANEUVERS (CASE A: 200 NAUTICAL MILE PARKING ORBIT WITH PLANE CHANGE OF 28.5 DEGREES: CASE B: 3,000 NAUTICAL MILE PARKING ORBIT WITH SIMILAR PLANE CHANGE) 90 DEGREE STATION RELOCATION IN 1

2,000

NORTH-SOUTH STATION KEEPING

EAST-WE STATION KE

1,500 3,000

1,000 2,000
K)

g

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L. B. HOLCOMB MD D. H. LEE

these data, a thruster lifetime (space environment) of 3000 to 5000 hr can be expected -without large accelerator electric current drains. Extended life tests of tantalum and tantalum (50 ppm yttrium) neutralizer filaments -were conducted. Four tantalumyttrium filaments operated in excess of 10,000 hr, while two pure tantalum filaments failed at less than 10,000 hr. All six were tested at a pressure of 8 X 10~6 N/m2 or lower, to avoid hydrocarbon vapors that could embrittle the filaments. The neutralizer performed well during the space tests of the ATS-D experiment. Emission-limited neutralizer operation was observed at one point during the test. It was later postulated that one of the gravity gradient booms passing through the ion thruster exhaust beam caused this anomaly. 5-cm Mercury-Bombardment Thruster Low-thrust mercury-bombardment thruster research has been underway at NASA Lewis Research Center (LeRC) for-several years. The most recent low-thrust mercury electronbombardment thruster22,23 employs a 5-cm-diam discharge chamber and a glass-coated grid. A series of preliminary thruster tests were conducted at LeRC with a variable-geometry discharge chamber. Hughes Research Laboratories (HRL) is presently developing this experimental thruster into a flightprototype system capable of providing north-south stationkeeping of a synchronous satellite. The system includes ion engine, zero-g feed system, and neutralizer. Thrust vectoring (±10°) is accomplished by electrostatic beam deflection. The ion engine utilizes a plasma-br-idge neutralizer. The HRL 5-cm mercury electron-bombardment thruster system is depicted in Fig. 2k. The thruster employs a 5-cm-diam permanent-magnet discharge chamber (Fig. 25). A hollow enclosed cathode and a conical cathode pole piece are used to improve discharge chamber performance at the low thrust. Glass-coated grids are incorporated to reduce discharge power losses at low specific impulses. All of the propellant flow goes through the hollow enclosed cathode. The method of thrust vectoring is not firmly established; however, an electrostatic thrust vectoring scheme is being extensively investigated at HRL. Glass grids capable of electrostatic beam vectoring are not presently developed; however, experimental two-grid systems have demonstrated the technique.^ The 5-cm mercury-bombardment thruster could rely on a standard glass grid and engine gimbaling, or a two-grid system and grid translation to provide thrust vector control. The feed system is similar to the

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AUXILIARY-PROPULSION SYSTEMS

227

SERT-II mercury feed system. The

propellant is stored in a reservoir and expelled by a pressurant gas. An elastomeric bladder is used for propellant-pressurant separation. The feed system of the HRL design has a mass of 1 kg (2.2 Ibm) and a capacity for up to 6.2 kg (13.6 Ibm) of mercury. An isolator is provided so that Fig. 2i| Five-cm mercurythe main cathode and bombardment thruster the neutral!zer can (photo courtesy of HRL), share the same propellant tank. The plasma-bridge neutralizer is designed to provide sufficient electrons to neutralize a 30-mA beam with 6.3% of the main propellant flow required for the neutralizer. A power conditioning system mass of 3.6 kg (8.0 Ibm) is estimated,^5 vith an efficiency of 85% expected. The total system has an estimated mass of 2.3 kg (5-1 Ibm) with a maximum propellant capacity of 6.2 kg (13.6 Ibm) of mercury. The HRL 5-cm

mercury-bombardment thruster develops a l61*0-|iN (370-M-lbf) thrust at a nominal 18,800-N-s/kg (1920-lbf-s/lbm) specific impulse with a discharge chamber mass efficiency of 75% with glass grids. Discharge chamber mass efficiency includes neutralizer propellant consumption. Detailed thruster performance data taken from Ref. 22

l/2-cm-DIA

ROD MAGNETS

BORON NITRIDE GLASS-COATED ACCELERATOR GRID CERAMIC SLEEVE

ENGINE BODY

Fig. 25

Five-cm mercury-bombardment thruster discharge chamber (representative of HRL and LeRC designs).

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L. B. HOLCOMB AND D. H. LEE

are presented in Table 7« The plasma-bridge neutralizer requires a maximum of 12.0 w. The second line of data is the present thruster performance with two grids.26 Life test data of this thruster system with a glass grid is not available, since it is still in the developmental stage. Grid lifetime seems to limit thruster life at present. During a 1000-hr life test of a 15-cm glass-coated grid,2? weight loss measurements indicated that the grid would be 100% eroded in 22,000 hr. This life test was divided into two parts: 500 hr of normal operation followed by 500 hr of grid-edgetermination studies. During the second 500 hr of testing, this grid was modified with masks over the edges of the grid. Although lifetimes of 10,000 hr have been projected for glasscoated grids, a great number of glass grids have failed after a few hours of test.

Long-life glass grids seem to depend heavily on extremely tight quality control, since irregularities in glass coating induce premature failure.27 It has also been speculated that interactions with the test facility may be responsible for such failures, making ground test of these devices very difficult.28 Two-grid bombardment thrusters (these operate at a somewhat higher voltage or specific impulse than glass-coated grid thrusters) have undergone extensive life-testing. The 5-cm mercury-bombardment thruster has been life-tested twice at LeRC. The first test was a 1000-hr preliminary life test of a two-grid thruster with electrostatic beam deflection. ^ The second test was a 2000-hr test of a two-grid thruster with grid translation to provide beam deflection. ^0 Both tests ended successfully with a thruster life of 10,000 to 20,000 hr predicted for the 2000-hr life test. The 15-cm SERT-II mercury-bombardment thruster has been life-tested at several facilities. Thruster tests have exceeded 5000 hr; however, excessive erosion near the neutralizer was noted. Proper design and placement of the neutralizer should alleviate this problem in future thrusters. In earlier studies of mercurybombardment thruster lifetime, grid erosion rates were estimated.^1 Lifetimes of 10,000 hr for a 20-cm mercurybombardment thruster were estimated on the basis of observed grid erosion rates. Sputter yield (i.e., number of metal atoms eroded per incident ion) is an increasing function of ion energy or the potential difference between the accelerator grid and neutralizer. Therefore, increased grid life can be expected

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AUXILIARY-PROPULSION SYSTEMS

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for the 5-cm mercury-"bombardment thrust er since the negative accelerator voltage (-250 V) is nearer spacecraft potential than the SERT-II negative accelerator voltage (-1750 V). Accelerator grid erosion is also a function of current density; however, the SERT-II and the 5-cm thruster are of comparable current density. The two-grid 5-cm mercurybombardment thruster with proper neutralizer geometry should provide 10,000 hr of life.

Plasma-bridge neutralizers have been tested in conjunction with thruster life tests. One mercury-bombardment thruster test was in excess of 55000 hr. An endurance test of a SERT-II32 design neutralizer has been conducted. The plasmabridge neutralizer operated for 12,000 hr, emitting 250 mA to a collector plate. Colloid Systems Description of the Concept. The colloid thruster is defined as a device that electrostatically accelerates multiatom or multimolecular charged particles. Solid and liquid heavy particles have been studied; however, liquid colloids have proved most successful. A conceptual diagram of a colloid thruster is presented in Fig. 26. Liquid propellant (glycerol to which a small quantity of sodium iodide is added) is stored in a reservoir, which can be a rigid tank, a coiled feed tube, or a bellows tank. If the propellant is pneumatically fed as in the system presented in Fig. 26, then a propellant isolation valve is required. If a bellows is used, then the propellant NEEDLE (5 TO 15 kV; EXTRACTOR (-0.5 TO -2.0 kV) can be fed mechanically, and, thereSOLENOID VALVE fore, an isolation (OPTIONAL) valve is required

FOCUS ELECTRODE (SLIGHT NEGATIVE V ) ACCELERATOR ELECTRODE (SPACECRAFT POTENTIAL]

Fig. 26

Conceptual diagram of a colloid thruster.

only to protect the propellant from an atmosphere prior to launch. The propellant is distributed in a manifold to needles, linear slits, or annular slits (Fig. 27). A potential is maintained between the needle and the extractor electrode. The intense electric

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L. B. HOLCOMB AND D. H. LEE

field in the (b) vicinity of the thin needle rim exerts a force upon the fluid surface, producing PLATINUMelectrohydrodynamic IRIDIUM LIPS spraying of charged propellant droplets. These multimolecular droplets are charged positive when the PLATINUM TUBE capillary is main0.036cm (0.014 in. O . D . ) 0.010cm (0.004 in. I . D . ) tained at a positive (c) potential and negative when the volt-OUTER EXTRACTOR EXTRACTOR—| age potential is reversed. A focusing 0.218cm (0.086in.)electrode is re(0.200 in.) 0.008 en quired to contain (0.003 i the charged droplets within the acceler0.005cm ator gap. The (0.002 ir accelerator electrode is maintained at spacecraft potential. In recent thrusters, the focus Fig. 27 Geometry of a) standard and accelerator needles, b) linear slits, electrodes have been and c) annular slits. removed. Again, as in the ion thrusters, a neutralizer is required; however, the lower charge/mass ratio for colloid systems as compared to ion thruster systems dictates substantially lower neutralizer requirements for a colloid thruster of comparable thrust. State-of-the-Art Equipment. A Ij-.^-mN (l-mlbf) thrust colloid system is presently under development at TRW Systems Group under sponsorship of the Air Force Rocket Propulsion Lab.33 ,3^ The thruster is comprised of twelve, 36-needle modules. Each module will deliver approximately 355 M-N (80 jjLlbf) of thrust at a nominal lA ,710-N-s/kg (1500 lbf-s/ Ibm) specific impulse. This system is designed for northsouth stationkeeping of a synchronous satellite. A program goal is a 10,000-hr thruster lifetime. The if.^-mN (l-mlbf) thrust colloid system is depicted in Fig. 28.

The sodium iodide-doped glycerol is stored in a bellows tank, with propellant flow rate mechanically controlled. The

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AUXILIARY-PROPULSION SYSTEMS

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feed'system mass is 3.2 kg (7.1 Ibm) with a maximum propellant load of

11.3 kg (25 Ibm).

This allows for 10,000 hr of thrusting at 1^,710-N-s/kg (1500 Ibf-s/lbm) specific impulse and a 95% mass expulsion efficiency. The mechanical flow controller operates on a feedback loop from the beam current. Average power requirement for the controller is 1 w Fig. 28 TRW 4.^5-mN (l-mTbf) colloid with a maximum preADP thruster. dicted peak power required of 2 w. The power conditioner must provide high voltages for the needles and extractor (up to 12.3 kV) along with power for the thruster heater, neutralizer, and propellant flow rate controller. The mass of the power conditioning subsystem is 3.6 kg (7-9 Ibm) with an 81J5 efficiency.35 It will provide 70 w of conditioned power. Thruster mass is 1.4 kg (3.0 Ibm) with 2.3 kg (5.4 Ibm) of structural mass. Total fixed system mass is 10.7 kg (23.5 Ibm) with up to 11 kg (25 Ibm) or propellant capacity.

Typical operating data for a 36-needle module is presented in Table 8. The first line corresponds to the design data used in estimated 4.4-mN (l-mlbf) thruster system performance. The following lines are test data obtained on several types of modules. The annular slit performance was taken from TRW Systems Group36 and EOS37 data. The final line in Table 8 is preliminary performance data of the 4.4-mN (l-mlbf) colloid thruster system. 38 Heater and mass flow rate controller power of k.O w is required for synchronous orbit of north-south stationkeeping thrusters. Neutralizer power is estimated at 5-0 w, providing 3.8 mA of negative current for 10,000 hr. Life test data are not yet available for the 4.4-mN (l-mlbf) thruster array. A life test of a three needle

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Table 8 Colloid thruster performance data

Thruster type Typical colloid thrusters

Thrust , Geometry Hlbf)

36 needles (TRW) 36 needles (TRW)

36 needles (TRW)

6 annular slits (TRW)

Mass flow rate, ng/s

Ref.

3.91

11+ ,710 (1500)

2k. 9

Estimate

3.80

114,691 (11+98)

25.6

36

306.0

3.30

Il+,3l8 (11460)

21.U

36

3.80

114,837 (1513)

23.3

36

5.51+

15,200 (1550)

35.0

36

(120.0)

Power/ thrust , w/mN (w/mlbf)

Engine efficiency , at

"

10. 51+ (k6. 8)

70. 0

9. 95

70. 0

10. 79 (1+7. 9) 11. 01 (1+8. 9) 10. 141 (U6. 2 )

66. 0

( l 4 U . 2)

^?Yer .

conditioning efficiency,

Total pover

'

%

tr1 bd •

67. 0

gt^

72. 0

0

H >

6 annular slits (TRW)

512.0 (115.0)

5.52

114,906 (1520)

36.0

36

10. 81 (1*8. 0 )

70. 0

U U

6 annular slits (TRW)

1489.0 (110. 0 )

5.22

Il4,5ll4

35. Q

36

10. 70 (1+7. 5)

69. 0

W

89.0

0.87

13,ll4l

6.8

37

9. 80 (1+3. 5)

73. 0

'tk! H

1*4,710 (1500)

298.8

Estimate

C

56. o

15 ,090 (151+0)

309.5

38

c

63. 3C

1 annular3slit (EOS) TRW k. 5-mN (1-mlbf ) ADP colloid

Specific impulse , N-s/kg (Ibf-s/lbm)

371.0 (83.5) 383.0 (86.0)

(68.9) 3145.0 (77.6) 5314.0

36 needles (TRW)

Thruster power , w

( footnote)*

3

(footnote)b

at 0°C.

(20.0) 14,1450 (1000)

56.00

14,673 (1050)

56.00

(1U80) (13140)

C

12. 6

(56)

12. C

(53)

^12 modules of 36 needles each with the performance of the first line of this table. c lncludes 5 w for neutralizer, k w for thruster heater and mass flow rate controller.

c

81.0

69.1

81.0

66.6

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AUXILIARY-PROPULSION SYSTEMS

233

thruster has been performed. Results from over 17^0 hr of microthruster operation are presented in Refs. 39 and lj-0. Thruster operation -was good through 1300 hr of log time (time thruster was in vacuum) with a slight decrease in propellant mass flow rate. For a period of TO hr, a 0.3- to 0.5-fJ-A extractor current persisted. The current leakage was finally corrected by operating the thruster with a closed valve and a high operating temperature. During the log hour periods 1^8^ to 23^8, a constant decrease in mass flow rate was noticed and thruster temperature was increased in attempts to improve the mass flow rates. Decreased mass flow rate was later explained by a leak in the pressurant gas tank. Another life test of a colloid microthruster is reported in Ref. ^1. A three-needle array was tested for 1300 hr when a power failure permitted the propellant to form a droplet between the needle and the extractor. A second test of the array lasted 2^00 hr with a similar power failure. At test termination, needle tips were in excellent condition; however, prevention of glycerol flow to the extractor and tar deposits due to electron bombardment of glycerol are necessary for long-life colloid thrusters.

Tests of 36-needle modules have exceeded 1000 hr with little or no degradation in thruster performance.3o Current leakage between the needle and extractor was noticed several times during the thruster life test. The vacuum pressure was increased, propellant feed pressure was lowered, and the beam was vectored in an effort to "clean up" the drainage currents. This was an effective solution; however, it is not the procedure that would be used in space. At 930 hr the test was terminated when failure of vacuum equipment led to severe erosion of the thruster module. A six-annular-slit array was tested for 500 hr. Tar formation was present and eventually led to termination of the test. New designs are being studied in an effort to eliminate long-term tar deposits in annular slits. Neutralizers were tested in conjunction with the 1000-hr 36-needle tests reported in Ref. 36. The neutralizer functioned properly for the first 180 hr of the test when the chamber pressure was increased in an effort to "clean up" the thruster array. However, the neutralizer then burned up due to the presence of atmosphere. Pulsed Plasma Systems

Description of the Concept. The pulsed plasma thruster is a device that uses a "burst" of electrical energy to

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L. B. HOLCOMB AND D. H. LEE

produce, accelerate, and eject a plasma wave. Pulsed plasma accelerators have electrodes in direct contact with the gas, thus requiring a zone of conducting plasma to complete the circuit.

Pulsed plasma devices are usually classified by their discharge chamber geometry. The parallel rail accelerator (Fig. 29a) is the simplest pulsed plasma device. Two other pulsed plasma geometries are the "T tube" (Fig. 29b) and the "coaxial gun" (Fig. 29c). The typical a) pulsed plasma thruster is presented in Fig. 30. The propellant, a solid, liquid, or gas, is placed in the discharge gap. Solid propellant is maintained in the gap region by mechanical springs, Fig. 29 Various Pulsed plasma liquid is fed by accelerators: a) parallelcapillary action, rail accelerator, b) T tube, and gas is placed in and c) coaxial gun.^2 the chamber with a gas injection valve. A capacitor or inductor is required for energy storage. The anode-cathode configuration will vary as those presented in Fig. 29; however, the rail-type geometry . is presented here. No neutralizer is required for the pulsed plasma thruster. IGNITER PLUG-

FUEL RETAINING SHOULDER

Fig. 30

Conceptual diagram of the pulsed plasma thruster.

Flight Experience . A solidpropellant pulsed plasma propulsion system having four thrusters was launched aboard the MIT Lincoln Lab. LES-6 satellite in September 1968.^3 After more than 2 yr in space, the pulsed plasma thrusters continue to function well. In a simulated

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AUXILIAEY-PROPULSION SYSTEMS

235

life test at vacuum facilities, intermittency or failure to discharge capacitor energy was observed. The intermittency would not correct itself and the unit stopped firing completely. After more than 3 million pulses on each of the four flight thrusters, intermittency has been observed in three of the four thruster units. Complete failure (100$ intermittency) has occurred to only one thruster unit. More recent designs have corrected this intermittency problem. The propulsion system has functionally kept the LES-6 satellite within the +2° longitude design band.44

State-of-the-Art Equipment. Republic Aviation Division of Fairchild Killer Corp. has developed a solid-propellant pulsed plasma thruster for the MIT Lincoln Lab., LES-6.^3,45 This thruster is capable of providing a 27-M.N-s (6- (jilbf-s) impulse bit per discharge of the 2-fjiF capacitor. The flight system is charged to 1360 V, which is equivalent to a 1.85joule discharge. A solid block of Teflon is used as propellant. The thruster system is complete with feed system, power conditioning, and telemetry circuits. This system does not have thrust vectoring capability. The LES-6 thruster is presented in Fig. 31. A schematic diagram of the solid-propellant LES-6 pulsed plasma thruster is presented in Fig. 30. A small capacitor discharge (1/8 joule) across the spark igniter provides sufficient plasma to initiate a discharge between the anode and cathode. The main discharge evaporates, ionizes, and accelerates Teflon propellant. A spring is employed to feed the Teflon to the thruster. A retaining shoulder maintains the propellant between the anode and cathode. Open failure of valved liquid or gaseous systems can lead to forces or moments applied to the spacecraft. Solid Teflon propellant remains passive when not used, or in the event of a failure. The LES-6 thruster configuration employs one 2-jj,F capacitor and two separate anode, cathode, spark igniter, and propellant subsystems. The two complete subsystems provide redundancy with little addition to total mass. Each thruster system, including case and Teflon propellant, has a mass of 1.4 kg (3.0 Ibm). Each thruster includes 0.1 kg (0.3 Ibm) of Teflon, which provides 280-N-s (64 Ibf-s) total impulse. A power conditioner was developed which provides 1.32 w of power to the main and igniter capacitors of any of four thruster units. The conditioner has a mass of 0.9 kg (1.9 Ibm) and has an efficiency of

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L. B. HOLCOMB AND D. H. LEE

The thruster can operate at several discharge energies and different pulse rates. The nominal discharge energy of 1.85 joules provides 25- to 27- (JLN-S (5-T- to 6.0- p.lbf-s) impulse "bit per discharge. ^3 ,1*5 Measurements in space suggest that 2U- ^N-s (5.35fjtlbf-s) impulse bit per discharge is achieved for this thruster.^6 Typical performance data are presented in Table 9. Radio frequency interference noise has been observed on the LES-6 thrusters. Measurements of this noise are discussed in Ref. Vf . This noise has not affected normal operation of the LES-6 satellite.

Fig. 31

Fairchild-Hiller LES-6 pulsed plasma thruster (photo courtesy of Fairchild-Hiller).

Prototypes of the LES-6 thruster designed in 1968 were tested extensively in the laboratory.^3 Ground testing was performed at 1 to h pps. On several tests the thruster achieved more than 8 million discharges without failure of the capacitor or discharge initiator. Intermittency. of discharge to occur is common with the flight thruster after 1 to 2 million discharges. An intermittent thruster does not draw power, and thus the capacitor can be discharged later. However, when the intermittency becomes more frequent the thruster may never recover. During tests of the LES-6 pulsed plasma thruster at Lincoln Lab., one thruster became intermittent after 1 million pulses and did not recover.^8

Purchased from American Institute of Aeronautics and Astronautics Ta~ble 9 Plasma thruster performance data

Thruster type

Fair child Killer LES-6 thruster

Thrust , Thruster |j.N power, N ,ui ; (iii"

16.9a (3.8)

w

Power/ Specific impuls e , thrust w/mN N-s/kg (lbf-s/lbm) (w/mlbf)

l.^lb

30UO (310)

8k (371)

Specific thrust (J.N-S/J ((ilbf-s/j)

13 .3 (3 -0)

effic±

Power , ... . conditioning ,>,,..

^ '°

efficiency,

1.8

56.0

, _. Total power,

m

%

2.52 M1 tr

M

is I

C

17.8 (k. 0)

b

l.Hl

30^0 (310)

80 (353)

Ik .2 (3 .2)

1.9

56.0

2.52 o

a CO

o d

17.3 (3.9)

1.^

d

1863 (190)

81 (361)

13 .3 (3 .0)

1.2

U8.0

2.91 CO

1 (-3

Reference 1*3 at 0.667 pps. 0.1 w for telemetry.

°Reference k5 at 0.667 pps. d

T. Williams, NASA Goddard, at 0.

pps ; 1,000,000 pulses. ro

CO

co

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L. B. HOLCOMB AND D. H. LEE

Table 10

Thruster system

Stored gas

General characteristics of thruster systems Favorable characteristics

Unfavorable characteristics

Inexpensive.

Low specific impulse.

Repeatable impulse bit.

Long-term leakage, for long missions.

Flight experience.

High-pressure tankage required. Excessive mass for high total impulse missions.

Vaporizing ammonia/ electrothermal

Catalytic monopropellant hydrazine

Low-pressure storage. Repeatable impulse bit.

Relatively low reliability of present feed system designs.

Flight experience.

Thruster heater power required.

Medium specific impulse.

Leakage for long missions.

Relatively high reliability of feed system. Medium specific impulse.

Flight experience. Low leakage of propellant for long-term storage.

Electrothermal monopropellant hydrazine

Relatively expensive development.

Medium specific impulse.

Poor pulse response at low thrust.

Relatively inexpensive. Repeatable impulse bit. Flight experience.

Ion

Poor repeatability at very low impulse bit with cold catalyst bed. "Limited" catalyst bed life.

Relatively high reliability of feed system. Low leakage of propellant for long-term storage.

Plenum monopropellant hydrazine

Moderately expensive.

High specific impulse. Flight experience.

Relatively low reliability of nonpassive feed system.

Medium-to-low specific impulse.

High voltages and large power requirement. Complex system. Exhaust neutralizer required. Expensive.

High specific impulse.

Some power required.

Low power-to-thrust ratio.

High voltages.

Low propellant vapor pressure.

Complex system.

Exhaust neutralizer required. Expensive.

Pulsed plasma

High specific impulse.

Large power requirement.

Simple system.

Limited thruster life.

Flight experience. Relatively inexpensive.

RF noise. Potential feed system problems. Redundancy of systems requires addition of propellant.

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AUXILIARY-PROPULSION SYSTEMS

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IV. Qualitative System Characteristics A qualitative tabulation of thruster characteristics is

presented in Table 10. This table is not meant to serve as a selection criterion for candidate systems; however, it may

serve to narrow the selection to a few candidate thrusters.

The selection of an auxiliary-propulsion system for a particular mission requires detailed knowledge of candidate propulsion systems, their performance, mass, reliability, and cost, as well as an accurate description of mission propulsion requirements. In addition, auxiliary-propulsion system/ spacecraft interactions must be studied before a specific system can be selected. Finally a system selection technique is required to compare thruster characteristics in light of mission assumptions.258

This paper has not discussed all the data required to

perform system selections; however, it is hoped that this

paper has presented the satellite manager with an adequate overview of the propulsion systems available for communications satellites to assist him in his final selection. References Isley, W. C. and Duck, K. I., "Propulsion Requirements for Communications Satellites," AIAA Paper 72-51^, Washington, D.C., 1972, in this volume. 2

Free, B. and Huson, G., "Selected Comparisons among Propulsion Paper 72-517>

Systems for Communications Satellites," AIAA Washington, D. C., 1972, in this volume, 3

Herron, B. G., Garth, D. R., Finke, R. C., and Shumaker, H.A., "Power Processing Systems for Ion Thrusters," AIAA Paper 72-518, Washington, D. C., 1972, in this volume.

Hall, D. F. and Lyon, W. C., "Low Thrust Propulsion System Effects on Communication Satellites," AIAA Paper 72-5195

Washington, D. C., 1972, in this volume. 5

Grant, A. F., Jr. and Lee, D. H., "Evolution of the Small

Rocket Engine," AIAA Paper 67-982, Anaheim, Calif., 1967.

Lee, D. H., Levitt, B. B. , Tripp, C. N., and Young, L. D., Design Considerations in the Application of Small Thrusters to

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240

L. B. HOLCOMB MD D. H. LEE

Unmanned Spacecraft, TRW Space Technology La~b. , Redondo Beach, C a l i f . , 1964. 7

Harney, E. D.,ed.-> Low-Thrust Space Propulsion - 1967 , Office of the Director of Defense Research and Engineering, Washington, D. C., Dec. 1967. o

Holcomb, L. B., "Satellite Auxiliary - Propulsion Selection Techniques," TR 32-1505, Nov. 1970, Jet Propulsion Lab., Pasadena, Calif. Q

Shav, R.? et al., "Performance of a Propellant Feed System for the ATS Ammonia Fuel Resistojet," AIAA Paper 69-296, Williamsburg, Va., 1969. Yoshida, R. Y., et al., "Life Test Summary and High Vacuum Test of 10MU3 Resistojets," Journal of Spacecraft and Rockets ,

Vol. 8, No. 1*, April 1971, pp. klk-kl6.

Krieve, W. F., "Attitude Control and Stationkeeping Subsystem Program," AFAPL-TR-68-14, March 1968, Aero Propulsion Lab., Wright-Patterson Air Force Base, Ohio. 12 Murch, C. K., et al., "Electrothermal Thruster Performance with Biowaste Propellants," AIAA Paper 70-ll6l, Stanford, Calif., 1970.

Pugmire, T. K.,et al., "Applied Resistojet Technology," Journal of Spacecraft and Rockets, Vol. 8, No. 1, Jan. 19715

pp. 63-68. Ik

Grant, A. F.3 Jr., "Catalysts for the Thermal Decomposition of Hydrazine When Used as a Monopropellant or as a Gas Generant," Publication 15, Feb. 1953, Jet Propulsion Lab., Pasadena, Calif. Voge, H. H., et al., Shell Development Co., Emeryville, Calif., unpublished report for NASA, April-Dec., 1964. Pugmire, T. K., et al., "Electrothermal Hydrazine Engine Performance," AIAA Paper 71-760, Salt Lake City, Utah, 1971.

17Brill, Y. et al., "Propulsion for Geostationary Space3 craft," 1971 JANNAF Combined Propulsion Meeting, Las Vegas , Nev., Nov. 1-5, 1971.

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241

1o

Murch, C. K.5 et al., "Electrothermal Hydrazine Th.ruster Development ," AIAA Paper 72-1*51 > Washington, D. C., 1972.

19Worlock, R. , Davis , J. J. , Jones, E. , Ramirez, P., and Wood, 0., "An Advanced Contact Ion Microthruster System," Journal of Spacecraft and .Rockets, Vol. 6, No. k, April 1969 5 pp. U2^-^29. 20 Hunter, R. E. , Bartlett, R. 0., Worlock, R. , and James, E.' L. , "Cesium Contact Ion Microthruster Experiment Aboard Applications Technology Satellite (ATS)-IV," Journal of Spacecraft and Rockets, Vol. 6, No. 9, Sept. 1969, pp. 968-970. 21 James, E. L. and Goldner, S. J., "Ion Engine Systems Testing," AFAPL-TR-69-112, Feb. 1970, Aero Propulsion Lab., Wright-Patterson Air Force Base, Ohio. 22 Reader, P. D., Nakanishi, S. , Lathem, W. C., and Banks, B. A., "A Sub-Millipound Mercury Electron-Bombardment Thruster," Journal of Spacecraft and Rockets , Vol. 7 5 No. 11, Nov. 1970, pp. 1287-1292.

23 King, H. J. and Schnelker, D. E., "Thrust Vectoring Systems," Journal of Spacecraft and Rockets, Vol. 8, No. 5,

May 1971, PP. 552-55^. 2k

Collett, C. R., King, H. J., and Schnelker, D. E., "Vectoring of the Beam From Ion Bombardment Thrusters," AIAA Paper 71-691, Salt Lake City, Utah, 1971.

25

Garth, D., private communication, 1972, Hughes Research Lab., Malibu, Calif. 26 Hyman, J., private communication, 1972, Hughes Research Lab., Malibu, Calif. 27 'Banks, B. A. and Bechtel, R. T. , 1000-Hour Endurance Test of a Glass-Coated Accelerator Grid on a 15-Centimeter-Diameter

Kaufman Thruster, TN D-5891, July 1970, NASA. pO

Bechtel, R. T. , Banks, B. A., and Reynolds , T. W. , "Effect of Facility Backsputtered Material on Performance of GlassCoated Accelerator Grids for Kaufman Thrusters ," AIAA

Paper 71-156, New York, 1971.

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L. B. HOLCOMB AND D. H. LEE

29 ^Lathem, W. C., "1000-Hour Test of a Dual Grid, Electrostatic

Beam Deflection Accelerator on a 5 Centimeter Diameter Kaufman Thruster," TM X-67-907, Aug. 1971, NASA.

Lathem, W. C . , "Grid-Translation Beam Deflection Systems for 5-cm and 30-cm Diameter Kaufman Thrusters," AIAA Paper 72-^85, Washington, D. C . , 1972.

Masek, T. D. and Pavlik, E. V. , "Thrust System Technology for Solar Electric Propulsion," Journal of Spacecraft and

Rockets, Vol. 6, No. 5, May 1969, pp. 557-56^. 32

Ravlin, V. K. and Kerslake, ¥. R., "SERT II: Durability of the'Hollow Cathode and Future Applications of Hollo¥ Cathodes," Journal of Spacecraft and Rockets, Vol. 7, No. 1, Jan. 1970, pp. 1^-20.

Jackson, F. A., "Colloid Advanced Development Program," Proceedings of the AFOSR Sixth Symposium on Advanced Propulsion Concepts, Air Force Office of Scientific Research, May 1971.

3^Jackson, F. A., et al., "Colloid Advanced Development Program," Interim Rept. No. 1, AFRPL-TR-72-10, Feb. 1972, Air

Force Rocket Propulsion Lab., Edwards Air Force Base, Calif. Farber, B. F. and Chester, M. S. , "Power Conditioning and Control System for a One Millipound Colloid Thruster," IEEE Publication 71 C15-AES, Pasadena, Calif. April 19-20, 1971Shelton, H. , Lear, W. C., Kidd, P. W., Huberman, M. N., Farber, B. F., and Krieve, W. F., "Charged Droplet Electrostatic Thruster Systems," AFAPL-TR-70-31, June 1970, Aero Propulsion Lab., Wright-Patterson Air Force Base, Ohio. 07 'Perel, J. , Mahoney, J. F., and Yahiku, A. Y. , "Analytical Study of Colloid Annular Thrusters," Journal of Spacecraft and Rockets, Vol. 8, No. 7, July 1971, pp. 702-709. ^Q

Zafran, S., private communication, March 1972, TRW Systems, Redondo Beach, Calif. ^Zafran, S., Beynpn, J. C., Shelton, H., and Kemp, R. F., "Colloid Microthruster Experiment," AFAPL-TR-70-55, Aug. 1970, Aero Propulsion Lab., Wright-Patterson Air Force Base, Ohio.

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Zafran, S. and Beynon, J. C. , "Colloid Microthruster System Life Test," Journal of Spacecraft and Rockets , Vol. 8, No. 2, Feb. 19T15 PP. 1^0-1^6. "^urson, W. C., Life Testing of a Colloid Thruster Source," AFAPL-TR-69-8, May 1969, Aero Propulsion Lab. , WrightPatterson Air Force Base, Ohio.

1*2 Jahn, R. G., Physics of Electric Propulsion, McGraw-Hill, New York, 1968. Guman, W. J. and Nathanson, D. M. , "Pulsed Plasma Microthruster Propulsion System for Synchronous Orbit Satellite," Journal of Spacecraft and Rockets, Vol. 7, No. 4, April 1970, pp. U09-U15. ki, Braga-Illa, A. A., "The Future of Self-Contained Control of Synchronous Orbits," AIAA Progress in Astronautics and .Aeronautics: Communication Satellites for the 70's ~ Technology, Vol. 25, edited by N. E. Feldman and C. M. Kelly, The MIT Press, Cambridge, Mass., 1971, pp. 211-227.

Vondra, R. J., Thomassen, K., and Solbes , A., "Analysis of Solid Teflon Pulsed Plasma Thruster," Journal of Spacecraft and Rockets, Vol. 7, No. 12, Dec. 1970, pp. 1^02-1^06. MacLellan, D. C., private communication, Dec. 1961, MIT, Lincoln Lab. Sicotte, R. L., "RFI Measurements of UHF on a Pulsed Plasma Thruster," Journal of Spacecraft and Rockets, Vol. 7 3 No. 3, March 1970, pp. 337-338. MacLellan, D. C., private communication, Dec. 19693 MIT, Lincoln Lab.

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SELECTED COMPARISONS AMONG PROPULSION SYSTEMS FOR COMMUNICATIONS SATELLITES

Bernard Free* and George Huson COMSAT Laboratories, Clarksburg, Md. Abstract A number of propulsion modes are compared to assist in optimization of spacecraft design. The process requires the definition and quantification (as far as possible) of the comparison criteria, including weight, thrust, power requirements, reliability, development status, etc. The importance of the several criteria varies with the type of mission. Missions considered range through attitude and longitude control, orbit inclination control, station change and orbit raising—in approximate order of complexity. Propulsion options are exemplified by cold gas, hydrazine, pulsed plasma, and electrostatic colloid and ion thrusters. Introduction This paper summarizes the important tradeoffs between two or more propulsion systems for each type of propulsion task appropriate to synchronous communications satellites. The propulsion tasks are: a) orbit raising from parking to synchronous orbit, b) initial acquisition, c) longitudinal repositioning, d) north-south stationkeeping, and e) eastwest stationkeeping and attitude control. Each of these propulsion tasks has unique characteristics which determine the suitability of the propulsion systems. No one propulsion system will be ideal for all the propulsion tasks listed. The propulsion systems discussed in this paper are at least in the Presented as Paper 72-517 at the AIAA 4th Communications Satellite Systems Conference, Washington, D.C., April 24-26. 1972. *Member, Technical Staff. +Propulsion Section Head.

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advanced developmental stage, and some have been flight tested. The selections were made by way of example, but naturally some attempt was made to match the propulsion system with the task. Orbit Raising from Parking to Synchronous Orbit The traditional method of going from parking to synchronous orbit is by means of high-thrust chemical propulsion, with some combination of upper transfer stages and apogee motor. The complete maneuver, from Earth launch to synchronous orbit, is usually completed in a few days. An alternate method of raising the satellite from parking to synchronous orbit is by the use of low-thrust electric propulsion, which has the capability of delivering much greater useful mass to synchronous orbit. However, with electric propulsion the thrust is limited by the power available, and the spirallingout process requires several months for completion.

The most significant tradeoff, therefore, is between maneuver time and useful synchronous mass. By way of example, consider an Atlas Centaur launch vehicle and a 2-kw (end-of-life) synchronous communications satellite. With all-chemical propulsion, the useful mass which can be delivered to synchronous orbit, not including the spent apogee propulsion system, is 659 to 810 kg. Two electric propulsion systems, the hydrogen resistance jet and the electron bombardment ion thruster, (The data given in this paper for ion thrusters are appropriate for both cesium and mercury propellant.) have been selected for comparison. The method of calculation and detailed propulsion system characteristics are given in Ref. 1 and summarized in Table 1. Table 1 Summary of electric propulsion system characteristics Resistance jet

Specific impulse, sec Total efficiency Power/ thrust, kw/N Specific mass, kg/kw Solar cell array (end-of-life) Power conditioning Contingency Thruster

840 0.88 5 26 23 ... 2 1

Ion thruster 2000-3000 0.63 14 32 23 4 2 3

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The propulsion task is the transfer of the satellite from a 650-km circular orbit, inclined at 28.5° with respect to the equator, to a geostationary orbit. The total impulse, expressed as velocity increment, is about 5700 m/sec. The starting mass in parking orbit is assumed to be 4400 kg (Atlas Centaur launch). Useful synchronous mass includes the mass of 2-kw (end-of-life) power but does not include any portion of the orbit-raising propulsion system or excess power used solely for orbit raising. The tradeoff is illustrated in Fig. 1, where it is evident that electric propulsion, in return for considerable power and maneuver time, yields large gains in synchronous orbit mass. The economic implications of Fig. 1 have been studied in a cursory way and are summarized in Table 2 for an all-chemical injection, a four-month resistance jet maneuver, and a six-month ion thruster maneuver. Initial Acquisition This propulsion task consists of a series of maneuvers which orient the satellite properly, establish required spin

2

.

3

4

5

6

O R B I T R A I S I N G M A N E U V E R T I M E , MONTHS

Fig. 1 Chemical and electric upper stages on an Atlas Centaur launch vehicle«

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Table 2 Major cost factors of geostationary _____________injection by several propulsion methods_____ Type of upper stage propulsion All Resistance Ion thruster chemical jet(4 mo.) (6 mo.) Launch vehicle Satellite costb Upper Stage Cost of excess solar cells @ $700/w Interest on investment Total cost of geostationary injection Useful geostationary mass Cost per unit mass

$15 $13 $ 0.5C

$15 $26 $ 0.5

$15 $46 $ 1

.. . .. .

$ 4.4 $ 0.9

$ 9.5 $ 1.5

$46 .8 1500 kg $ 0.031

$73 .0 2500 kg $ 0.029

$28 .5 700 kg $ 0.041

All costs are expressed in millions of dollars, "Assumed roughly proportional to the useful synchronous mass. G Apogee. rates, and correct the injection error. A typical example of propulsion requirements is less than 100 N sec (0.15 m/sec) to establish orientation and spin, and 30 m/sec velocity increment to correct the injection error. Neither of these tasks warrants a separate propulsion system; hence their peculiar requirements must be fulfilled by one of the propulsion systems already on board.

Kaplan has shown that the proper orientation and spin can be established in reasonable maneuver time (< 2 days) with several types of electric propulsion and a judicious use of battery and solar cell power. The most critical maneuver is rapid Earth acquisition, which may require a thrust level of

several hundred yN and 100-200 w of power for less than 1 hr, which is easily within the capacity of the batteries. Injection error depends on the method used to reach nearsynchronous conditions. If electric propulsion orbit raising is employed, there is no injection error. The assumed value of 30 m/sec is conservative for a conventional launch and apogee motor system. The injection error may be removed by either chemical or electric propulsion. The tradeoffs among maneuver time, power, and propellant mass are summarized in Table 3. The electric thrusters referred to in Table 3 will be described in more detail in a later section. For the present case, it is sufficient to state that an electric

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Table 3 Characteristics of maneuver to remove injection error Hydrazine monopropellant Thrust, N (Ib)

Maneuver time Propellent mass, kg Power, w

10(2.5) few hours 10 10

Electric thruster designed for: Station North- south changing stationkeeping 0.07(0.02) 3 days 1-2 1000

O.OOS(.OOl) 50 days 0.5-1 100-200

station-changing thruster normally will use most of the power

already onboard. If this is less than the 1000 w shown in Table 3, then the maneuver will take longer (e.g., 500 w 6 days). In any event, injection error is removed with power and electric thrusters already onboard. Thus the disadvantage of using hydrazine is the mass of propellant (10 kg); the sole disadvantage of using electric propulsion is the maneuver time.

Longitudinal Station Changing During the operating lifetime of a satellite, growth or shift in communications traffic occasionally calls for relocation of the satellite. For a 3-satellite global system, this relocation, or repositioning, is usually across 120° of

longitude. This maneuver can be accomplished, using a single thruster aligned with the roll axis, with either chemical or electric propulsion. In the case of electric propulsion, the thrust level should be high enough to minimize maneuver time by making use of all available satellite power not being used for traffic during this maneuver. The tradeoffs between hydrazine monopropellant and a large mercury ion thruster have been examined previously for this maneuver, and only a summary of the main points is given here. Details on the assumptions and method of calculation are given in Ref. 4. Thrust Program It is assumed that the thrust vector, for both highthrust (chemical) and low-thrust (electric) propulsion, is tangent to the orbit. The thrust program for the hydrazine system consists of two high-thrust pulses; one injects the satellite into an elliptical orbit with a slightly changed period, and the second recircularizes the orbit after the required longitudinal drift has occurred.

The thrust program

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for electric propulsion consists of a continuous tangential thrust over the entire maneuver with the thrust direction reversed at the half-way mark.

Summary of Tradeoffs For chemical propulsion with hydrazine, electric power is required to perform only minor functions and can be ignored. The main relationship of interest with hydrazine is the tradeoff between maneuver time and propellant consumption. For electric propulsion, the propellant consumption is -less important, and the pertinent tradeoff is between power and

maneuver time. Table 4 is a summary of pertinent data for repositioning with hydrazine in terms of maneuver time and

propellant mass. Table 5 is a similar summary for electric propulsion, with additional values included for power and thrust level. The data are shown graphically in Figs. 2

and 3.

Table 4

Longitude repositioning of a 700-kg satellite with hydrazine monopropellant

u m/sec

deg/day

4>

Time to drift -120° days

Propellant mass kg

5 10 20 30 40 50 100 150

0.88 1.76 3.52 5.28 7.04 8.80 17.6 26.4

136 68 34 23 17 14 7 4.5

1.5 3.1 6.2 9 12 16 31 .46

.

Table 5 Longitude repositioning of a 700-kg satellite with low-thrust electric propulsion u total m/sec 20 40 60 80 100 200 400

Time to drift 120° days

68.6 34.3 23.0 17.2 13.8 7.0 3.6

Mass of mercury propellant

kg

Power w

Thrust mN

0.6 1.2 1.8 2.4 3 6 12

70 210 380 600 910 3,600 14,000

2 10 21 36 57 223 875

•

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PROPULSION SYSTEMS

R E P O S I T I O N I N G M A N E U V E R TIME D A Y S

Fig. 2 Propellant consumption for 120° longitude change with high and low thrust*

S A T E L L I T E MASS: 7 0 0 k g

o

a.

Z O

4

700-kg, 2-kW S A T E L L I T E , END OF L I F E

700 kg, 400-W SATE LLITE, END OF LIFE

I

10

20

30

40

50

R E P O S I T I O N I N G M A N E U V E R TIME, D A Y S

Fig. 3 Electric power for 120° longitude change with electric propulsion* Obviously, rapid maneuver times involve penalties for both forms of propulsion. For chemical propulsion, maneuver times progressively less than about 20 days result in increasingly heavy consumption of onboard propellant. Despite this penalty, the main benefit of chemical propulsion for

60

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B. FREE AND G. HUSON

this maneuver is that emergency maneuvers can be performed rapidly. For electric propulsion, the propellant mass requirements are modest, even for rapid maneuver times, but the maneuver time is restricted by the onboard power. Lower limits on the maneuver time are indicated in Fig. 3.

North-South Stationkeeping

Because of the gravitational disturbances of the sun and moon, a satellite orbit will not remain exactly in the equatorial plane. At synchronous altitude, orbit inclination with respect to the equator accumulates at an average rate of about 0.8°/yr if allowed to go uncorrected. Hydrazine monopropellant is at present the accepted way of removing unwanted inclination, and for relatively short satellite lifetimes there are no major drawbacks associated with its use. However, the required correction is equivalent to a velocity increment of about 42 m/sec for each year in orbit, and the use of hydrazine for long duration north-south Stationkeeping results in an undesirably high propellant budget. One possible way to reduce the propulsion system mass is through the use of electric propulsion. The chief advantage of electric propulsion over chemical propulsion for north-south Stationkeeping is that higher specific impulse (propellant exhaust velocity) is attainable. The specific impulse of electric thrusters (1000 to 5000 sec) is five or more times that of hydrazine monopropellant (220 sec), which makes it possible to reduce the propellant mass by 80% or more. This weight advantage is partially neutralized by the higher fixed mass of the electric propulsion system hardware, and significant levels of electric power are required for electric propulsion. This section summarizes the tradeoffs among hydrazine and several electric propulsion systems for north-south Stationkeeping of a 700-kg synchronous communications satellite. The important comparisons involve propulsion system power and mass requirements.

Thrust Program

The thrust program for north-south Stationkeeping with high and low thrust are illustrated (exaggerated inclination) in Fig. 4. When a high-thrust propulsion device, such as hydrazine monopropellant, is used for north-south stationkeeping, the usual practice is to allow the orbit inclination to increase to some prespecified limit, for example, 0.1°. When this limit is reached, one or more short, high-thrust pulses are activated at one of the points where the inclined

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EQUATORIAL PLANE

Fig. 4

Thrust programs for north-south stationkeeping.

orbit crosses the equatorial plane (nodes). The direction of effective thrust is normal to the orbit plane, pointing north if the descending node is chosen or south on the ascending node. Enough total thrust is applied to remove the accumulated -inclination, or even to produce an equal inclination in the opposite direction.

When an electric thruster is used for north-south stationkeeping, it is desirable to -minimize both the size of the device and the power required, by employing low thrust over a long period of time. This can be accomplished by two long periods of thrust centered on the nodes as illustrated in Fig. 4. It is not desirable to apply the thrust too far away from the nodes, because thrust is ineffective (for removing inclination) near the antinodes. For this paper, it is assumed that the electric thruster is on every day for two periods of about 6 hr each, except during eclipse. The thrust programs illustrated in Fig. 4 lead to the firing requirements given in Table 6. The severe restart and endurance requirements shown in Table 6 are typical for electrical propulsion

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Table 6 Firing requirements for north-south stationkeeping with hydrazine and ________electric propulsion__________________ Assumed limit of drift

Hydrazine Teflon pulsed plasma Colloid thruster Ion thruster

Interval between firings

0.1° naa 24 hr naa 24 hr naa 24 hr

No. of firings per year

On-time/ yr, hr

1 1800 1800 1800

10-20 7 wk 300b 300 300

Daily firing can easily limit drift to below the detectable limits by any known method. For a pulsed plasma, each firing is a train of 10 ,000-100,000 pulses. tasks. Endurance and restart capability approaching the 'levels shown has been demonstrated in the laboratory, but long-endurance space tests are still needed. Propulsion System Weight Budget

The elements of the propulsion system mass can be grouped into those which are fixed (i.e., the mass remains constant with increasing time in orbit), such as thrusters, valves, propellant feed lines, etc., and those which increase as a function of time in orbit, such as propellant and tankage. In this study, propellant mass is calculated from propulsion requirements, but the other component masses are derived from published hardware figures. The mass of propellant required can be calculated from

M = M (1 - e~U/V) P o

•

where M is the total in-orbit mass (700 kg), u is the total velocity increment (42 m/sec times years in orbit) and v is the exhaust velocity (specific impulse times 9.8 m/sec ). Tankage and feed system mass is calculated by using a single value for published data on flight-designed hardware, and by applying the formula r.-r

_

r.-r

/ nx

/\n

\

/ ^

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PROPULSION SYSTEMS

-

255

where W is the tank weight, and M is the propellant mass. The subscript 1 refers to published data, the subscript 2 implies a calculated value. The component masses for exemplary propulsion systems are given in Table 7. Hardware data for the hydrazine and colloid systems were derived from Refs.5 and 6, and the hardware data for ion thruster systems were derived from Refs. 7 and 8. Fig. 5 illustrates the growth of the propulsion system mass as the in-orbit lifetime increases. . Attitude Control and East-West Stationkeeping

East-west Stationkeeping requirements are almost trivial in terms of total impulse (> 2 m/sec per yr), and thrust levels exceeding 100 yN (25 ylb) are more than enough to correct worst-case drift. For any given station, the drift rate can be determined accurately as a function of time, and corrections can be made by either periodic or semicontinuous thrust. For these reasons, east-west Stationkeeping does not pose any severe propulsion problems and is not expected to be a critical element in the selection of a propulsion system.

Table 7 Elements of propulsion system mass for candidate north-south Stationkeeping thrusters Propulsion system

Component

Component mass, kg 1 yr 5 yr 10 yr

Propellant Tankage and feed system Thrusters (4)

13.5 3.2 1 17.7

64.4 9 1 74.4

122 15 1 138

Colloid thruster Propellant (I = 1500 sec) Tankage and feed system Thrusters (4) Power conditioning and controls

2 2.2 6

10 6.3 6

2010 6

3.6 13.8

3.6 25.9

3.6 39.6

0.8 0.3 6

4.2 0.5 6.

8.4 0.8 6

8 15.1

8 18.7

8 23.2

Hydrazine (I = 220 sec)

Ion thruster Propellant (I = 3500 sec) Tankage and feed system Thrusters (4) Power conditioning and controls

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B. FREE AND G. HUSON

S A T E L L I T E MASS:

HYDRAZINE

700kg

4

6

Y E A R S IN O R B I T

Fig. 5 Propulsion system mass vs years of north-south stationkeeping*

On the other hand, attitude control is probably the most complex and controversial of all of the propulsion tasks for communications satellites. In very simple terms, there are two general methods of maintaining orientation: periodic

unloading of angular momentum devices (spinning satellite, momentum wheels), and limit cycle operation (i.e., the use of all-thruster control to neutralize or overpower the torques generated by the synchronous environment. The possibilities involved in combinations of these two techniques are far too complex for adequate coverage here, and we will use a single situation to exemplify the strong points of several propulsion

systems.

The example chosen is a 700-kg satellite designed for a 10-yr lifetime and equipped with momentum wheels on each axis. These wheels are unloaded when necessary by one of three candidate propulsion systems: cold nitrogen gas, monopropellant hydrazine, or teflon pulsed plasma. The propellant budget in each case is designed for operation with the momentum wheels functioning so that the onboard propellant is sufficient to generate about 38,000 N sec of impulse for attitude control, about 13,000 N sec for east-west stationkeeping, and 310,000 N sec for north-south stationkeeping.

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Table 8 Propulsion system characteristics for attitude control functions

Specific impulse, sec Mass of propulsion system, kg Propellant and tankage Thrust ers (12) Momentum wheel system

cold nitrogen

Hydrazine

Teflon- pulsed plasmaa

70

220

330

143 120 3 20

43 20 3 20

50 13 17 20

These values are estimated for the thruster type used on the LES-6 satellite^ but exhibit suitable increases in the component masses for the present propulsion task. Under these conditions, the propulsion system mass chargeable to the attitude control function is given in Table 8. With the momentum wheels functioning, the weight advantage of hydrazine and pulsed-1-plasma systems over the cold gas system is obvious. It is also apparent that, in the case shown, the final mass of the electric thrusters is more than enough to cancel any weight advantage resulting from their higher specific impulse.

For the second part of the comparison, we assume that the momentum wheel system fails at some point in the satellite lifetime, and that attitude control goes into a limit cycle mode of operation. In this case, propellant consumption for limit cycling only (i.e., excluding disturbance torques) is proportional to the impulse bit size squared, which is illustrated in Fig. 6 for the candidate propulsion systems. The minor effect of specific impulse is illustrated for the case of the teflon pulsed plasma. (If a 1000-sec ISp pulsed plasma device were used, the masses given in Table 8 would not be applicable.) It is apparent from Fig. 6 that the rates of propellant use generated by limit cycle operation in the case of a teflon

pulsed-plasma device and a nitrogen gas system are small and would not adversely affect the propulsion system lifetime. In the case of hydrazine, the rate of propellant use depends greatly on the impulse bit capability. An electrothermal hydrazine thruster-LO operating with a 7-mN-sec impulse bit would not adversely affect the propulsion system lifetime in limit cycle operation. However, in the case of a catalytic'

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B. FREE AND G. HUSON

HYDRAZINE

w

10 CATALYTIC

g

ELECTROTHERMAL

10'2

TEFLON-PULSED PLASMA

IMPULSE BIT, mN-sec

Fig. 6

Minimum propellant use rate for attitude control by limit cycle operation*

thruster operating with a 100-mN sec impulse bit, the rate of propellant use for this operation alone would be more than twice that for all other propulsion tasks combined (including north-south stationkeeping), and the remaining propulsion system lifetime would be reduced by a factor of three.

Over-All Comments

The cursory tre'atment given here is a far cry from a complete propulsion system study and is not intended to generate firm conclusions. Also, this survey has treated the several propulsion tasks as if they were completely separate, whereas in reality the entire onboard propulsion task must be addressed as a single entity. Nevertheless, some very obvious generalizations can be made. First, in cases in which substantial and rapid emergency maneuvers are mandatory, highthrust chemical propulsion is the only choice. Second, in cases in which maneuvers requiring a large total impulse must be performed but time is available, some form of electric propulsion should eventually replace chemical propulsion because electric propulsion will result in an overall propulsion system mass reduction. Third, maneuvers which require small or moderate total impulse can be performed efficiently by almost any propulsion method. Finally, thrusters capable of operating with small impulse bits should be used for many, if not all, limit cycle operations.

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References

Free, B., "Satellite Raising to Synchronous Orbit," Journal of the British Interplanetary Society, Vol. 23, 1970, pp. 669690. 2

Free, B., "Economic Tradeoff Studies for Electric Propulsion Missions on Communications Satellites," AIAA Paper 71-683, Salt Lake City, Utah, 1971.

Kaplan, M., "All-Electric Thruster Control of a Geostationary Communications Satellite Which Employs Narrow-Beam Antennas," AIAA Paper 72-436, Bethesda, Md., 1972. 4 Free, B., "Chemical and Electrical Propulsion Tradeoffs for Communications Satellites," CQMSAT Technical Review, Vol. 2, No. 1, Spring 1972, pp. 123-145. COMSAT Labs Staff Members, In-house calculations, Propulsion Section, 1971, Clarksburg, Md.

Jackson, F., et al., "Colloid Advanced Development Program,"

Interim Final Rep. 1, AFRPL-TR-72-10, TRW Systems Group for Air Force Rocket Propulsion Laboratory, AF Systems Command, Edwards Air Force Base, Calif., February 1972.

Finke, R.C., private communication, 1971, Propulsion Components Section, NASA Lewis Research Center, Cleveland, Ohio. Q

James, E., et al., "A One Millipound Cesium Ion Thruster System," AIAA Paper 70-1149, Stanford, Calif., 1970. g

Guman, W. and Nathanson, D., "Pulsed Plasma Microthruster System for Synchronous Orbit Satellite," Journal of Spacecraft and Rockets, Vol. 7, No. 4, April 1970, pp. 409-415.

Murch, C. and Hunter, C., "Electrothermal Hydrazine Thruster

Development," AIAA Paper 72-451, Bethesda, Md., 1972.

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POWER PROCESSING SYSTEMS FOR ION THRUSTERS B. G. Herron* Hughes Research Laboratories, Malibu, California D. R. Garth^ Hughes Aircraft Company, Los Angeles, California R. C. Finke? and H. A. Shumaker^ NASA Lewis Research Center, Cleveland, Ohio . Abstract The proposed use of ion thrusters to fulfill various communication satellite propulsion functions, such as eastwest and north-south stationkeeping, attitude control, station relocation, and orbit raising, naturally leads to the requirement for lightweight, efficient, and reliable thruster power processing systems. Collectively, the propulsion requirements dictate a wide range of thruster power levels and operational lifetimes, which must be matched by the power processing. This paper describes the present development status of such power processing systems, presents system design alternatives, and projects expected near future power system performance.

Presented -as Paper 72-518 at the AIAA Vth Communication Satellite Systems Conference, Washington, B.C., April 2^-26, 1972. •^Head, System Technology Section, Ion Device Physics Department. iSenior Project Engineer, Power Systems Department, Space and Communications Group. ?Head, Propulsion Component Section, Ion Physics Branch, Electromagnetic Propulsion Division. §Head, Power System Technology Section, Power Electronics Branch, Spacecraft Technology Division.

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Introduction The future of global communication systems will require the development of medium- and high-power satellites having long operational lifetimes. These satellites probably will be 3-axis stabilized spacecraft requiring N-S and E-W stationkeeping and employing large flexible solar arrays. Spacecraft launch-mass considerations may preclude the use of gas jets or other low-specific-impulse control devices, since the percentage of the total spacecraft weight attributed to propellant becomes an unacceptably large part of the total.

Within the family of electric propulsion devices that have been brought to various levels of readiness for applications ,^ the electron-bombardment ion thruster appears particularly well suited to fulfill all of the propulsion control needs that can be potentially listed for a communication satellite. In addition, the larger thrusters appear appli3 In each thruster application, conditioned and controlled power must be made available for thruster operation. This means that the solar array must be sized to provide additional power in excess of that for the prime communication needs and that allowances in spacecraft design must be made for the inclusion of the thruster power processing equipment. To date, considerably more development work for ion thruster power conditioning has been done for thruster sizes more useful for prime propulsion.^>5 In the near future, however, complete auxiliary thruster systems will be carried through flight qualification. A 1 mlb Cs bombardment thruster system will be flown on ATS-F^7 and a 5-cm Hg bombardment thruster with electrostatic deflection grids allowing±10° of 2-axis beam deflection has been selected as an experiment to be flown on the Canadian Communication Technology Satellite.°>9 Thruster Load Requirements and Characteristics A .block diagram which presents the complement of power supplies required to power the individual thruster elements is shown in Fig. 1 for a multikilowatt, hollow-cathode, electronbombardment ion thruster. Lower power auxiliary thrusters (for N-S and E-S stationkeeping and attitude control) have essentially the same configuration with the exception that

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Fig. 1 Power sapply/thruster block diagram. present thruster designs derive both beam and cathode vapor flows from a single vaporizer and require additional supplies to implement thrust vector deflection. The major portion of the thruster power is associated with the screen (beam) and discharge (arc) supplies. Typical output voltages of the screen supply will range between 1 and •2 kv although considerably higher voltages have been flown,, for example, on the Space Electrical Rocket Test (SERT) II. Ion beam current levels may range from as low as 25 to 35 mA for a 5-cm auxiliary thruster and up to 2 amp for a 30-cm orbit-raising thruster. The discharge supply is referenced at the screen potential and must have a current rating, of 6 to 10 times the screen supply for prime thrusters and generally greater than 10 times for auxiliary thrusters. Typical output voltages for this supply lie between 35 and ^5 v.

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The accelerator supply output voltage requirements generally will fall in the range of -0.8 kv and -2.0 kv, and under normal thruster operating conditions will be loaded to less than 1$ of the beam current. However, under abnormal conditions, arcs between the screen and accelerator electrodes can produce the equivalent of a short between the screen and accelerator supplies; hence both supplies must be capable of withstanding high-peak transient currents without failure. The remaining heater, vaporizer, and keeper supplies, except for the requirement of output transformer high-voltage isolation when referenced at the screen potential, generally can be designed and implemented in a straightforward manner.

For the thruster tc be started and controlled at a desired operating point, sequential and dynamic closed-loop control must be provided as an integral part of the power processing system. As indicated in Fig. 1 the control loops can be implemented internally to the system using sensed power supply voltages and currents without recourse to external sensing equipment or measuring devices. It appears that the thruster process can be satisfactorily controlled with no more than three moderately low-gain control loops when individual supplies are internally regulated for either output voltage or current. The problems of implementing the required supplies and controls in the form of a space-qualified power processing system continues to lessen as a result of the following three factors: (l) rapid improvements in electronic power conditioning devices; (2) continued improvements in the thruster characteristics when viewed as a process to be controlled and a load to be powered; and, perhaps most importantly, (3) accumulated experience in designing, fabricating, and integrating thruster/power conditioning systems.

Total System Considerations To fully assess the impact of introducing a prime or auxiliary propulsion system into a satellite spacecraft configuration, a number of spacecraft and propulsion parameters and their interrelationships must be considered. The weight of the thruster hardware and the propellant are factors that are dictated by the type of thrusters employed, the specific impulse and mode of operation chosen, and the total duration of thruster service during a given mission. These factors are determined by requirements established ,during mission analysis studies and by the characteristics of implemented hardware

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from thrust design activities. Seldom are power system trade possibilities considered in making the foregoing judgments and selections.

The thruster power system activities are initiated upon receiving various inputs from the mission analysis and thruster areas which usually include the total thruster system power alloted and its distribution at the thruster load; size, weight, and reliability specifications; the desirable characteristics or mode of regulation for each supply; the thruster electrical parameters to be controlled which indirectly establish the number and kind of dynamic control loops; and the thruster power-time use profile which describes startup and shutdown requirements as well as whether eclipse operation at full or reduced power is called for. In addition, from the spacecraft area comes specifications of bus voltage magnitude (whether supplied directly from the solar array or battery subsystem) , the range of voltage variation and the peak transient loading allowed, EMI compatibility requirements, and general specifications of the thermal and vibrational environments.

No matter how the power conditioning problem is posed it always reduces to satisfying the totality of specifications and requirements while the emphasis is placed on the major and related items of weight, efficiency, and reliability. The total weight of the electrical power system attributed to the inclusion of electric propulsion is a function of the thruster power conditioning specific weight (which is in turn a function of total power level, efficiency, and reliability), the specific weight of the solar array, the efficiency of the power conditioning, and the incremental specific weight of any special structures required for mounting electronics or for thermal control of power conditioning losses.

A simple representation of the above functional relationship describing the power system specific weight associated with the thruster system is given by

where #pg = total specific weight associated with thruster power system, ape = specific weight of thruster power conditioning, a $A = specific weight of solar array, a gg = specific weight of special structures specifically added to the spacecraft for the thruster power system, P = power to thruster (s-), r\= efficiency of power conditioning including line losses, and R = power conditioner. reliability.

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SHUMAKER

Equation (l) in slightly modified form (including

thruster and propellant weight considerations) has "been used very successfully in deriving optimum high-power prime propulsion thruster systems for both interplanetary and satellite orbit-raising missions. Its application in guiding the design of lower-power systems typical of auxiliary propulsion presently is limited since scaling laws describing the functional relationships and their parametric variations become more difficult to derive accurately as the power level is decreased; yet the number of power supplies and the control requirements remain essentially constant. In addition, at the lower-power levels the discontinuities of the functional relationships due to the discrete size and weight of available components become more pronounced. Transformer high-voltage isolation also becomes an important factor in determining size and weight of the magnetics, and the weight and bias power requirements of the control electronics can be the same for a 100-w system and 1-kw.system. Required thruster power levels and mission reliability gene-rally establish the power conditioning configuration (the number of inverters and transformers, the manner in which redundancy is implemented, etc.) and determine the type of components to be used and their ratings. A prime opportunity then exists for optimizing total power system specific weight in terms of £fficiency. For fixed thruster power P and system reliability R, the weight optimum system (assuming a-p^ and tfgg are continuous in q) occurs for efficiency rj^-which satisfies

(8^/811) (P,T]*,R) - (*SAAi*2) + (a*ss/^) (P^*) = o

(2)

For increasing r\ (neglecting discontinuities due to component availability) the first term monotonically increases while the latter two monotonically decrease, and an optimum rj * exists and can be determined if sufficient scaling (variational) information is available. For presently achievable power conditioning weightefficiency penalty variations, and for solar panel specific weights between 30 to 33 Ib/kw, values of r| between 90 to 93% appear near optimal for power requirements 'of 1 kw and above. It is not expected that the optimum value of TJ will change significantly in the future. The reason for this is that while lower values of cn^ can be expected to occur in the near future (tending to move the optimum r\ to a lower value), expected improved power conditioning components will lead to lower incremental weight penalties at high efficiencies

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(tending to move the optimum r\ to a higher value). The net result should have a stabilizing effect with respect to the optimum r\ even though a-pg could be reduced significantly. In general, the optimum r\ for the lower-power thruster systems falls at lower values of n than for the high-power systems. Although, as previously mentioned, the same functional expressions are valid regardless of power level, the control power for sequential and dynamic control is a larger percentage of the thruster -input power for low-power systems; hence the value of r\ for which QtfpQ/drj rapidly increases is less than for higher-power systems.

Power Processing for Auxiliary Propulsion Thrusters The design of power processors for auxiliary ion thruster s represents a major challenge because of the long mission times and the inherent problem of maintaining low specific weight and high efficiency for controllable power supplies with low- output power capabilities. To achieve the goals of long-term operation in an electrical environment imposed by the thruster on the power conditioner, including high-level noise and arcing, the power conditioner design must assure considerable derating of components. This is only possible with discrete components, since integrated circuit operation is characterized at specified voltages and will not operate properly at derated supply voltages. Use of integrated circuits is highly feasible, however, if precautions are taken to limit transients induced in wiring caused by thruster arcing. Although usually not specified by manufacturers except in rare cases, integrated circuits can withstand fairly high line transients, usually on the order of two times their normal rating. Their use at specified levels is thus an implied derating of up to 50$. The duration of the transients is considered of importance in specifying component reliability but has not yet been resolved on other than a pass- fail basis. The implied derating is therefore of little value since transients are superimposed on the operating levels and in severe cases come close to the two-times rating. Discrete components, on the other hand, have well-defined safe limits. Derating of 50$ on all parameters is a typical practice for thruster power conditioning, whereas derating of only 30$ is normally used on other spacecraft equipment. Redundancy is another critical parameter in determining reliability, but it also infringes heavily on weight.

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Redundancy can Toe carried to the point of a total duplication of all circuitry. The opposite extreme is no redundancy at all, which may meet the over-all mission reliability criteria. Obviously the lightest system is the completely nonredundant system. Partial duplication has been used in some kilowatt systems where the semiconductors of a power inverter, "but not the heavier magnetic component, are replaced via relay transfer switching. In the case of the small thruster power conditioners little or no redundancy has been employed because replacement of an inverter or a portion might constitute up to 20 C - > 197Z; this vo lume. Herron, B. G. , Garth, D. R. , Finke, R. C. , and Shumaker, H. A. , "Power Processing Systems for Ion Thrusters, " AIAA Paper 72-518, Washington, D. C. , 1972; this volume. Hall, D. F. , Newnam, B. E. , and Womack, J. R. , "Electrostatic Rocket Exhaust Effects on Solar-Electric Spacecraft Subsystems, " Journal of Spacecraft ana Rockets, Vol. 7, No. 3, March 1970, pp. 305-312.

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Reynolds, T. W. , "Mathematical Representation of Current Density Profiles from Ion Thrusters, M AIAA Paper 71-693, Salt Lake City, Utah, 1971. Hall, D. F. , "Electrostatic Propulsion Beam Divergence Effects on Spacecraft Surfaces, " Final Rept. , Vol. II, JPL Contract 952350, Jan. 1973, TRW Systems, Redondo Beach, Calif. 8

Staggs, J. F. , Gula, W. P., and Kerslake, W. R. , "The Distribution of Neutral Atoms and Charge-Exchange Ions Downstream of an Ion Thruster, " Journal of Spacecraft and Rockets, Vol. 5, No. 2, Feb. 1968, pp. 159-164. Q

Sellen, J. M. , Jr. , "Solar Electric Propulsion/Instrument Subsystems Interaction Study, " Mid-term Briefing, Contract NAS2-6940, Sept. 1972, TRW Systems, Redondo Beach,

Calif.

Lyon, W. C. , "Monopropellant Thruster Exhaust Effects upon Spacecraft, " Journal of Spacecraft and Rockets, Vol. 8, No. 7, July 1971, pp. 689-701. Dugan, J. V. , Jr. , "Upper-limit Charge Exchange Cross Sections For Mercury"1" on Molybdenum and Cesium"1" on

Aluminum," TM X-2527, March 1972, NASA.

Sellen, J. M. , Jr. , "Interaction of Spacecraft Science and Engineering Subsystems with Electric Propulsion Systems, " AIAA Paper 69-1106, Anaheim, Calif., 1969.

Goldin, D. and Sellen, J. M. , J r . , "Integration of Scientific Pay loads with Multimission Electrically Propelled Spacecraft, " AIAA Paper 70-1141, Stanford, Calif., 1970. 14 Sellen, J. M. , Jr. , "RF Emission from an Electron Bombardment Ion Thruster, " TRW Document 4360. 4. 2. 72-15, March 1972, TRW Systems Group, Redondo Beach, Calif. Pawlik, E. V. , et al. , "Solar Electric Propulsion System Evaluation, " Journal of Spacecraft and Rockets, Vol. 7, No. 8, Aug. 1970, pp. 968-976. Rulis, R. J. , "SERT II: Design Requirements for Inte-

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grating Electric Propulsion into a Spacecraft, " Journal of Spacecraft and Rockets, Vol. 8, No. 3, March 1971, pp. 209-213. 17

Reynolds, T. W. and Richley, E. A . , "Propellant Condensation on Surfaces Near an Electric Rocket Exhaust, " Journal of Spacecraft and Rockets, Vol. 6, No. 10, Oct.

1969, pp.

1155-1161.

18

Kemp, R. F., et al. , "Effects of Electrostatic Rocket Material Deposited on Solar Cells, " AIAA Paper 72-447, Bethesda, Md. , 1972. 19

Hall, D. F. , "Evaluation of Electric Propulsion Beam Divergence and Effects on Spacecraft, " 08965-60 13-RO-OO, Final Rept. , Contract NAS7-575, Sept. 1969, TRW Systems, Redondo Beach, Calif. Reynolds, T. W. and Richley, E. A. , "Contamination of Spacecraft Surfaces Downstream of a Kaufman Thruster, " TN D-7038, Jan. 1971, NASA. Staskus, J. V. and Burns, R. J. , "Deposition of Ion Thruster Effluents on SERT II Spacecraft Surfaces, " AIAA Paper 70-1128, Stanford, Calif., 1970. Hall, D. F. , "Electrostatic Propulsion Beam Divergence Effects on Spacecraft Surfaces, " Final Rept. , Vol. 1, JPL Contract 952350, Aug. 1970, TRW Systems, Redondo Beach, Calif.

23

Hall, D. F. and Kelley, L. R. , "Experimental Techniques to Determine Electrostatic Rocket Exhaust Effects on Spacecraft Surfaces, " AIAA Paper 70-1144, Stanford, Calif. , 1970. 24

Hall, D. F. and Luedke, E. E. , "Degradation of Thermal Control Coatings by Cesium Atom Beams, " Interim Report, JPL Contract 952350, June 1972, TRW Systems, Redondo

Beach, Calif. 25

Hall, D. F. and Green, H. E. , "Erosive and Chemical Effects of Energetic Mercury Ions Bombarding Spacecraft Surface Materials, " AIAA Paper 72-446, Bethesda, Md. , 1972.

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26

Kelley, L. R. , Luedke, E. E. , and Hall, D. F. , "Damage of Thermal Control Coating Properties by Energetic Mercury Ion Bombardment, " AIAA Paper 72-445, Bethesda, Md. , 1972.

Hall, D. F. , "Electrostatic Propulsion Beam Divergence Effects on Spacecraft Surfaces, " Progress Report for Nov. 19, 1971-Feb. 25, 1972, JPL Contract 952350, March 1972, TRW Systems, Redondo Beach, Calif. 28

Milder, N. L. and Sovey, J. S. , "Optical Radiation from Regions Downstream of Mercury Bombardment Thrusters, " AIAA Paper 72-441, Bethesda, Md., 1972. 29 Lyon, W. C. , "A Study of Environmental Effects Caused by Cesium from Ion Thrusters, " HIT-487, March 1971, Hittman Associates, Inc.', Columbia, Md.

Huberman, M. N. , "Measurement of the Energy Dissipated in the Electrostatic Spraying Process, " Journal of Applied Physics, Vol. 41, No. 2, Feb. 1970, pp. 578-584.

Lord Rayleigh, Philosophical Magazine and Journal of Science, Vol. 14, 1882, p. 14. Zafran S. and Beynon, J. C. , "Colloid Microthruster System Life Test, " Journal of Spacecraft and Rockets, Vol. 8, No. 2, Feb. 1971, pp. 140-146. 33

Huberman, M. N. , et al. , "Present Status of Colloid Microthruster Technology, " Journal of Spacecraft and

Rockets, Vol. 5, No. 11, Nov. 1968, pp.

1319-1324.

34

Zafran, S. , et al. , "One-Millipound Colloid Thruster System Development, " AIAA Paper 72-1153, New Orleans, La., 1972. Huberman, M. N.,et al. , "Exploratory Development of Advanced Colloid Thrusters, " AFRPL-TR-128, Nov. 1971, TRW Systems, Redondo Beach, Calif.

Huberman, M. N. , "Annular Colloid Thruster for ThreeAxis Stabilized Military Satellites, " Report No. 1, Air

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Force Contract FO46 11-72-C-OO18, Nov. 1971, TRW Systems, Redondo Beach, Calif., Jackson, F. A. , "Colloid Advanced Development Program, " Status Rept. 29, AF Contract F336 15-70-C-1694, Nov. 1972, TRW Systems, Redondo Beach, Calif. 38

Physical Properties of Glycerine and Its Solutions, Glycerin Producers Association, New York. 39 Jackson, F. A. , "Colloid Advanced Development Program, " Status Rept. 27, AF Contract F33615-70-C-1694, Sept. 1972, TRW Systems, Redondo Beach, Calif.

40

Zafran S., et al. , "Colloid Microthruster Experiment, " AFAPL-TR-70-55, Aug. 1970, TRW Systems, Redondo Beach, Calif. 41

Jackson, F. A. , "Colloid Advanced Development Program, " Status Rept. 25, AF Contract F336 15-70-C-1694, July 1972, TRW Systems, Redondo Beach, Calif. 42

Jackson, F. A. , "Colloid Advanced Development Program, " Status Rept. 24, AF Contract F33615-70-C-1694, June 1972, TRW Systems, Redondo Beach, Calif.

43

Thomassen, K. I. and Vondra, R. J. , "Exhaust Velocity Studies of a Solid Teflon Pulsed Plasma Thruster, " Journal of Spacecraft and Rockets, Vol. 9, No. 1, Jan. 1972, pp. 61-64. 44

Vondra, R. , personal communication, March 30, 1970, MIT Lincoln Labs. , Cambridge, Mass.

45

Guman, W. J. and Nathanson, D. M. , "Pulsed Plasma Microthruster Propulsion System for Synchronous Orbit Satellite, " Journal of Spacecraft and Rockets, Vol. 7, No. 4, April 1970, pp. 409-415. 46

Lyon, W. C. , "Thruster Exhaust Effects Upon Spacecraft, " TM-X-65427, X-460-70-401, Oct. 1970, NASA.

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47

Guman, W. J. , personal communication, March 31 and April 7, 1970, Fairchild-Hiller Corp. , Farmingdale, N. Y. 48

Chirivella, J. E. , Moynihan, P. I. , and Simon, W. , "Small Rocket Exhaust Plume Data, " JPL Quarterly Technical Review, Vol. 2, No. 2, July 1972, pp. 90-99.

49

Hill, J. A. F. and Draper, J. S. , "Analytical Approximation for the Flow from a Nozzle into a Vacuum, " Journal of Spacecraft and Rockets, Vol. 3, No. 10, Oct. 1966, pp. 1552-1554. Chirivella, J. E., personal communication, Jan. 1973, Jet Propulsion Laboratory, Pasadena, Calif. To be published as Chirivella, J. E. and Simon, W. , "Molecular Flux Measurements in the Backflow Region of a Nozzle Plume, " 7th JANNAF Plume Technology Meeting, Huntsville, Ala., April 1973.

Price, T. W. and Evans, D. D. , "The Status of Monopropellant Hydrazine Technology, " NASA CR-92742, Feb. 1968, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, Calif.

Sutherland, G. S. , et al. , "Monopropellant Hydrazine Reaction Control Systems--A Five Year Status Report, " Aviation and Space: Progress and Prospects; Proceedings of the Annual Aviation and Space Conference, Beverly Hills, Calif. , 1968, ASME, 1968. "Spacecraft Attitude Control Gas Systems Analysis, "

CR-86661, April 1967, NASA. 54

Esenwein, F. T. and Walker, S. C. , "Effects of Hydrazine Exhaust Plumes and Propellant Spills on Selected Spacecraft Materials, " paper presented at the Hydrazine Monopropellant Technology Symposium, Nov. 28-30, 1967, published in CPIA Publication 160, Dec. 1967. 55

"Martinkovic, P. , personal communication to E. N. Borson, June 1968, U. S. Air Force Rocket Propulsion Lab. , Edwards Air Force Base, Calif.

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Borson, E. N* , "Rocket Plumes as Contamination Sources Sources, " paper ^presented at the Symposium for Optical Contamination in Space, Aug. 13-15, 1969, Optical Society of America, Aspen, Colo.

Carlson, R. A . , Blumenthal, J. L. , and Grassi, R. J. ,

"Space Environment Operation of Experimental Hydrazine Reactors, " R e p t . 4715.3. 68-37, July 1968, TRW Systems, Redondo Beach, Calif. C o

Brill, Y. C. , Stechman, R. C. , and Reis, R. J. , "Effect of Hydrazine Rocket Fuel on Spacecraft Materials, " paper presented at The Institute of Environmental Sciences, 14th Annual Technical Meeting, St. Louis, Mo., 1968. 597

Kesten, A. S. , "Analytical Study of Catalytic Reactors for Hydrazine Decomposition, " CR-89791, May 1967,

CR-92988, Jan. 1968, NASA.

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RESULTS FROM TESTS OF A LARGE LIGHTWEIGHT SOLAR ARRAY UNIT K. L. Hanson* General Electric Company, Philadelphia, Pa. Abstract Results from environmental and performance tests of a 250-sqft, 79. 3-lb rollup array unit intended to establish technology readiness of the design concepts are presented. Conventional techniques were used for the vibration, acoustic, pyrotechnic shock, and stowed thermal-vacuum tests. The deployed dynamics and thermal-vacuum tests involved unusual techniques. A gravity cancellation deployment aid was utilized for the thermal-vacuum and performance deployment. The deployed dynamic test results agreed with predictions for out-ofplane motion but not for in-plane motion. A blanket wrap tension of 0. 226 Ib/in. was found sufficient to provide stability in stowed vibrations where the peak amplification factor was 4. 3. The dominant test problem for large, light structures is the accommodation of gravity forces without interfering with test results. Presented as Paper 72-569 at the AIAA 4th Communications Satellite Systems Conference, Washington, D. C. , April 24-26, 1972. This work was performed for the Jet Propulsion Laboratory, California Institute of Technology, sponsored by NASA Contract NAS 7100. The writer acknowledges the contributions of the many individuals at General Electric Co. and the Jet Propulsion Laboratory that contributed to the test program described in this paper. Individuals who made major contributions include: W. Hasbach and R. Ross of the Jet Propulsion Laboratory and N. Shepard Jr., P. Perez, C. Stahle, R. Tuft, R. Sipple, Carl Petroff,and S. Kaplanofthe General Electric Co. In particular R. Ross derived Eqs. (1) and (2) and developed elemental mass and stiffness matrices for the solar array blankets. * Manager, Space Power Programs, Space Division.

307

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I. Introduction The testing of a new design configuration and/or the simulation of new environments and operating conditions can be a technical challenge equal to or exceeding the technology required to design and analyze it. This paper is based on the results from the testing of a large area lightweight rollup solar array. The testing was the final step in a program intended to achieve a maturity of the technology of a 250-sq-ft rollup solar array unit producing 30 w of power/lb of weight that would make the design concepts acceptable for new spacecraft designs. The program results have been documented in detail. The purpose of the paper is to discuss the significant problems, their solutions, and the application of the experience gained to future programs.

Solar arrays are the power source for current and projected long life spacecraft. Conventional solar arrays are an interconnected assembly of silicon solar cells mounted on a rigid substrate and have an extensive history of successful applications. During the past several years, there has been active interest in developing lightweight, large area, low stowed volume solar arrays for future applications. Advanced communication satellites are one application for these arrays and provide one of the justifications for the development of this technology. Summaries of these developments have been provided by Abbott9 and Lockheed. 10 One approach to reduce the weight and stowed volume of solar arrays is to reduce the thickness of both the silicon solar cells and the cover glass and mount the cells on a flexible substrate, typically a sheet of kapton with a thickness of 0. 002 in. The interconnected solar cells on the substrate, termed a blanket in this paper, are stowed for launch by rolling the blanket on a rigid drum or folding it on equally spaced parallel lines and packaging it in a housing. For operation, the system is deployed by a mechanical system which also provides the main structure for the deployed system. Typically the mechanical system consists of deployable booms, end members, and a means for tensioning the deployed blanket. The geometry of the deployed blanket is maintained by tension. Usually two sides edges of the blankets are free.

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To minimize weight the structural members are small and the blanket tension low. As a structure the system is flexible and has low natural frequencies of oscillation compared to conventional rigid substrate solar arrays. The incorporation of this type structure on a spacecraft introduces a number of additional design considerations. Some of these were not recognized in the past and resulted in problems as discussed by Bouvier and Likens.

The successful launch of the flexible solar array on the U.S. Air Force SESP-71-2 Thorad Agena Flight on October 17, 1971, and its subsequent successful operation resolved the questions as to

whether or not a flexible substrate solar array can be launched and deployed. However, additional knowledge and data are required before the application of the flexible substrate solar arrays can be considered routine. The technical questions which require answers for each mission include the following: 1) Will the design survive the environments ? 2) What will be the performance of the design? 3) Will the design interact with other spacecraft systems ? In developing the design of a rollup array for a specific application, the following questions arise: 1) What is the relation between the environment levels, the design configuration, and the stress levels ? 2) What methods of analysis are applicable to the design and what is their accuracy? 3) How can the results with previous designs be applied to the design configuration of interest? Testing is one method of resolving these questions and, in some cases, the only approach short of operational experience. Results to date provide many answers and a good basis for proceed ing with the utilization of flexible substrate solar array technology.

H. Program Summary The configuration of the solar array unit is shown in Fig. 1 and pertinent design parameters are given in Table 1. A detailed design description is given.3 The design was a result of an initial feasibility study followed by the detailed design and analysis of the system. During the feasibility study an engineering demonstration model was built and key development tests were run to provide data for the final design.

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a) Stowed

b) Fully deployed

CENTER SUPPORT

STORAGE DRUM

LEADING EDGE MEMBER

A R R A Y BLANKET

c) Configuration (coordinate system shown)

Fig. 1 Prototype test model.

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Table 1 Summary of design characteristics Solar cell blanket assemblies Weight breakdown Total Center support Lending edge member and brackets Boom actuator Outboard end supports Drum assembly Blanket assemblies Blanket tension

2 Ib/side

Flight system

Test unit

79.3 Ib 1.33 i. 07 11.73 4.10 15.60 45.48

82.5 Ib 1.33 1.07

11.73 4.10 17.60 46. 98

Natural frequency requirement above 0.04 Hz (zero gravity)

Flight system

Test unit

5 ,176 2 cm x 2 cm cells 2 2 cells in series 1 4 modules 2 19px 22s 1 2 19px 20s Electrical performance

11 active sol ar cell modules 8 19 p x 20 s 2 18 px 20 s 1 12 px 20 s

131. dummy modules s 'ith glass platelets

2466 w at 102 v (at 55°C)

19 x 20

4 filter dummy modules

2 1 x 20 2 4 x 20

The unit is made up of two storage drums mounted on a center support structure. Each drum has a bearing system, a slip ring assembly for the transfer of power and signals, and a Negator spring motor that provides a constant tension in the solar array blanket. A deployable boom is mounted on the center support and attached to a leading edge member. The solar array blanket consists of an interconnected assembly of cells mounted on a flexible substrate to form a solar array blanket. A blanket is rolled onto each drum, with the outboard edge attached to the leading edge member. The system is deployed by extending the boom. The deployed boom and the leading edge member comprise the primary structure. Outboard end supports are provided in the launch configuration and are pyrotechnically released before deployment. Table 2 Summary of environmental tests Pyrotechnic shock

Deployed

-130°C +1400C Thermal shock between -130°C and 140°C

Requires long dwell time to Selection of control points.

-130°C +140°C Thermal shock between -130°C and 1-10°C

Large differences in thermal mass c ture differences in system. Selectic

Low temp (-13U°C) High temp (140° C)

Long dwell times to produce u n i f o r n Selection of control points. Deployr requires support fixture.

150 dB over-all spectrum specified Vibration Sinusoidal

5 10 225 555

-

10 Hx. 225 Hz 550 Hz 2000 Hz

9 0 - 7 0 0 Hz 20 - 90 Hz 700 - 2000 Hz Mechanical shock

O.'JO in DA 4 . 6 g's (O-P) 0.00176 in. DA 27 g's (O-P)

Three widely space supjxn-t points.

1 G 2 /Hz Increasing at 6 dB/octave Decreasing at 6 dB/octave

Three widely separated support points.

250 G. 0.5 m sec terminal sawtooth

Large mass of test unit imposes ex on vibration equipment.________

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Table 3 Summary of performance tests Test

Technical comments

Deployment/Retraction

External support required for 1 G operation. Need to minimize tracking restraints.

Electrical Performance

Large area involved. Solar simulation desired.

Chip and Crack Inspection

Large area and large number of units involved.

Deployed Thermal Bending

Realistic test conditions involve large area illumination, gravity effects, and thermal environment.

Deployed Dynamics

State of the art testing problem involving low frequency regime, aerodynamic effects, gravity effects, and measurement of blanket motion.

In-Plane Structural Characteristics

Blanket tension forces should exceed gravity forces. Solve problem discovered in deployed dynamics tests.

Wrap Tension for Stability in Vibration

Solve problem discovered in stowed vibration tests.

The environmental and performance test program requirements were spelled out in considerable detail in SS 501407 12 and were intended to verify the design and analysis techniques. The environmental test program requirements are summarized in Table 2. The results are discussed in Sec. III.

The performance tests were intended to verify the system performance before and after environments or to verify analyses, performance prediction, or techniques. These tests along with key technical factors are tabulated in Table 3 and discussed in Sec. IV. The last two tests listed in Table 3 were added to the program as a result of knowledge acquired in the deployed dynamics test and the stowed vibration test. in. Environmental Test Results

The tests are discussed in the order which they were run. Pyrotechnic Shock

The array system contains two separation nuts (each armed with two active squibs) on the outboard end supports and the tests involved simultaneous firing of both separation nuts. Data were obtained from 24 piezoelectric shock accelerometers mounted at selected locations in the system. Three runs were made. The highest accelerometer reading was 6000 g on the outboard caging arm about 4 in. from the pyrotechnic nut assembly. The shock level on the mounting frame a similar distance from the pyrotechnic nut was approximately the same. The shock levels were much lower at greater distances from the shock source. The out-

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ATS-A SEPARATION TEST. SPACECRAFT RESPONSE (REF 1)

313

RA 230 - F I X T U R E

MARINER SEPARATION TEST. SPACECRAFT RESPONSE (REF 2)

IXTURE SIDE CENTER SUPPORT

REF 1 GSFC T&E MEMO REPT NO. 671-18DTO OCT 9. 1967 R£F2GEPIR6230-SM£-190S M. KAPLAN DTD MAY 2. 1968

Fig.

2 Comparison of various spacecraft shock spectra with the test specification and realized shock spectra.

board ends of the drums had peak shock levels of 1365 g and the peak shocks on the outer wrap of the blankets were only 125 g. Shock spectra were determined at 19 locations. Spectra for two locations are shown in Fig. 2, Figure 2 also shows the shock spectra induced in spacecraft structures by the firing of spacecraft separation band pyrotechnic release devices and the spectrum for a 250-g, 0. 5-msec terminal sawtooth shock pulse. Comparison of these spectra leads to the conclusion that the shock loads from the array pyrotechnic nut assemblies produce spectra equivalent to those induced by spacecraft separation band pyrotechnics and higher than those produced by the terminal sawtooth shock pulse. The aforementioned, coupled with the fact that the solar array was not damaged by the pyrotechnic tests, provides guidance to the planning of tests for future designs, Thermal-Vacuum Tests

The primary objective for this test was to achieve the specified temperature levels with realistic heating and cooling rates. Solar simulation with the intensity required to .produce 140°C was judged

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to be not essential and no serious effort was made to obtain access to the one known existing facility that could provide this environment. It was considered of little interest to impose test conditions on the blanket areas covered with dummy glass so the heat sources were concentrated in the areas occupied by the active solar cell modules (with the system deployed), the boom and leading edge member (deployed), and the drums, center support, and actuator. Quartzline infrared lamp assemblies were used for thermal input and the system was installed in a 32 X 54-ft vacuum chamber with cryogenic walls to simulate cold black space. Temperature was measured at 109 locations with thermocouples. Deployment at both high and low temperatures was required. Therefore, the unit used for performance deployment tests was installed in the test chamber. The major results of the thermalvacuum tests were as expected, and predicted heating and cooling rates were confirmed. Data are shown in Figs. 3 and 4.

B. STEM ACTUATOR HOUSING (T/C NO. 62) OUTBOARD END SUPPORT (T/C NO. 22)

STORAGE DRUM SHELL (T/C NO. 10 ~ Y THERMISTOR ON E6 MODULE (FROM THI

I

I

I

ELAPSED TIME (HOURS)

Fig. 3 Stowed transient temperature history.

In the stowed state the solar array blanket layers insulate the inner wraps and the drums from the external environment. The leading edge member and the leader on the blankets are directly exposed to the environment. Thus it is difficult to obtain a uniform system temperature at either high or low levels. Very long soak times are required to establish thermal equilibrium.

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40

60

80

100

120

140

160

180

200

220

240

260

280

300

320

380

315

400

ELAPSED TIME (MINUTES)

Fig. 4 Deployed transient temperature history. A greenhouse phenomena was observed in stowed testing (see E6 Module Temperature in Fig. 3). Radiant energy in the visible spectrum penetrated the wraps of the stowed blankets, but could not be radiated to the cold walls because the test blankets were opaque to the long wavelengths radiated by the blanket wraps. Consequently, the interior temperatures were higher than the surface temperatures. This would occur to a much lesser degree when the blankets are covered with solar cells because they are opaque in the visible portions of the spectrum.

The thermal characteristics of the deployed system were similar to those of the stowed system, except that the entire blanket is

exposed to the environment rather than just the blanket leader. The rate of change of temperature varies widely throughout the system. This fact was recognized in planning and conducting the tests by providing heat input over the dummy cell blanket area to prevent exceeding the low-temperature specification while the drum system was cooling. Acoustic Noise

The acoustics environment test was run without incident in an acoustic test chamber. With over-all input levels of 150 db or

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greater, accelerometers mounted on the outer wrap of the blanket measured rms g T s ranging from 3. 6 to 15. 6. Vibration-Sinusoidal & Random

The sinusoidal vibration test had the dual objective of determining the frequencies, mode shapes, and damping ratios of the resonant modes below 100 Hz and of subjecting the system to qualification level sinusoidal environments. The test setup utilized a single MB220 exciter. The stowed rollup array was mounted on a box beam vibration fixture and hydrostatic slip bearings were utilized as required. The setup for one axis of vibration is shown in Fig. 5. The instrumentation consisted of 78 accelerometers and 37 strain gages. Testing was accomplished without major problems. There were no major structural resonances below 100 Hz; therefore, no modes were investigated. No blanket modes with a classical response (peak quadrature response with zero inphase response) were discovered. Of great interest are the low amplification factors throughout the system. Selected data are tabulated in Table 4. The low amplification factors result in low dynamic loads and low stress levels. During the tests, it was noted that the blankets became slack at the outer wrap as the vibration caused the blankets to tighten on the drums. A stable condition was reached and was the subject of an investigation discussed in Sec. IV.

Fig. 5 Y-axis test setup. Table 4 Amplification factors for sinusoidal excitation Frequency

Amp factor

Frequency

2.4 3.4 3.9

125 120 300

1.6 2.3 2.0

60

2.2 4.3

130 350

1.7 2. 1

40 50

2.2 1.9

No data 150

1.9

(Hz) Actuator

25 GO 00

Leading-edge member (midspan) Drum

25

(Hz)

Amp factor

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System response in random vibration tests was low and no damage resulted. It was concluded this environment can be tolerated by rollup solar arrays.

IV. Performance Test Results Deployment/Retraction Deployments and retractions of the system were to be performed in both ambient and vacuum conditions to develop confidence in the tracking of the blankets, particularly during retraction, and in the deployment mechanism. Gravity forces were the dominant problem as the lightweight structural system could not tolerate the force of gravity for either upward, horizontal, or downward deployment. For example, approximately 33% of full vertical deployment could be accomplished without external support.

Vertically upward deployment was selected for this program, and a deployment aid was designed that provided a support force equal to the weight of the deployed portion of the blanket and boom. This aid consisted of two chains unrolling from a drum to provide the desired linearly varying force. The effectiveness of this approach, shown in Fig. 6, is proven by the data in Fig. 7. When using the aid, the deployment boom operated against the loads it would experience in zero g,viz., the blanket tensions created by the Negator springs.

Fig. 6 Upward deployment aid (installation in 32- x 54-ft space simulator).

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318

80

100

120 140

160 180

200 220

240 ?60

280

300

320

340 360

380 400

420

ARRAY DEPLOYMENT (INCHES - M6614 INSTALLATION!

Fig. 7 Upward deployment aid force - deflection curve.

This deployment aid provided some restraint at the boom tip, which aids tracking of the blankets on the drums. Since this condition would not occur in zero g operation, misalignment was deliberately introduced as shown in Fig. 8. The blankets successfully tracked until the misalignment, 6 , reached 10 in. Note that gravity forces increase the severity of the misalignment because the blankets tend to hang vertically from the leading edge member. Since blanket tension forces tend to reduce in-plane misalignments in zero g conditions, it was concluded that tracking should not be a problem in actual operation. Electrical Performance

Uniform illumination which is required to accurately measure solar array performance, would be difficult to achieve over the entire array area. The active modules were individually tested as they were not electrically interconnected. The individual module area in this design was relatively easy to illuminate to the required conditions.

Chip and Crack Inspection Chip and crack inspection of the system requires examination of a large number of units. This operation is not a technical prob-

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lem, but merely involves an effort proportional to the area of the system. Deployed Thermal Bending

Knowledge of the thermal deflection of the deployed system is needed to assess both mechanical and electrical performance. However, determination of the deployed thermal bending of the system is difficult. At this time, it appears that accurate prediction of the thermal bending of the deployment boom, in this case a BI-STEM device, ^ is beyond the state-of-the-art. Estimates can be made by using techniques developed for gravity gradient booms. Since the booms are the key structural element, prediction of system performance which also involves the effects of blanket tension is even more difficult. Since the thermal ** "Y bending forces are small relative to gravity forces a system thermal bending Fig. 8 Schematic of track- test appeared unfeasible. ing demonstration test setup.

The approach selected for this program was to conduct tests in a thermal bending test facility at NASA Goddard Space Flight Center on a 10 ft-boom specimen. Test results were extrapolated to the 33. 5-ft length of the deployed boom and the additional deflection caused by blanket tension was superposed to predict total displace-

ment. The predicted deflection based on test data was less than predictions based on the gravity gradient boom techniques, which were within specification.

The foregoing approach was intended to be conservative but did not include several effects which may be important. As already stated, the thermal bending characteristics of the BI-STEM are not well understood. The test data6 indicates that the boom tip moves transverse to the sun-boom line (in-plane) about the same distance it t A BI-STEM boom consists of two overlapping U C M sections. Each o section has an included angle of about 340 .

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moves along the line (out-of-plane). For inplane tip displacements beam column effects increase the tip displacement. For out-ofplane tip displacements, the blanket tension forces both inhibit and increase tip deflection. Thus, the accurate prediction of tip deflection is complex and the accuracy of the above approach is not known. Deployed Dynamics

Knowledge of the dynamic characteristics of the deployed system is needed because the flexible structure may interact with the spacecraft attitude control system. One of the major program goals was to develop analytical models for the system and verify them with test data. Before testing, the dynamics of the deployed system were predicted for out-of-plane symmetric (bending), out-of-plane antisymmetric (torsion), and in-plane motion. The analysis included the effects of gravity for a vertically downward deployment, the selected test condition. Thus, predicted resonant frequencies, mode shapes, and response to base excitation as a function of frequency were available for test planning and test operation.

The test setup and operations involved advances in the state-ofthe-art of low-frequency tests and have been described by Shoulberg and Tuft. The setup is shown in Fig. 9. Tests were run in a vacuum chamber at a pressure level of about 1 torr. This approach was verified by the 20 to 1 reduction in damping ratio and the 50% increase in first natural frequency in going from testing at 14.7 psi to 1 torr. The array was deployed downward and excited at the drum end to best represent the boundary conditions which would exist in an application. With this arrangement, the leading edge of the system was free and the blanket tension at the leading edge is minimum. The drums were free to rotate individually because each was fitted with an auxiliary torsion spring adjusted to support the weight of the blanket. Displacements were measured with six Optron Model 800 optical trackers mounted on a frame and focused on targets on the blankets. These devices which had no physical contact with the test unit could be scanned along the blanket to provide displacement data at 60 discrete locations. Resonant frequencies were determined by observing

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the response of a selected target to a slow sine sweep (constant amplitude). The response was separated into its colinear and quadrature components. Test runs were made to stimulate the three motion types: bending, torsion, and inplane bending. The principal test results are given in Tables 5-7.

Fig. 9 Deployed dynamics test setup.

Table 5

The test results for the two types of out-ofplane motion agreed well with the analytical results with some modes having excellent agreement. However, the natural frequency for in-plane motion was much higher than predicted. The reasons will be subsequently discussed.

Summary of resonant frequencies from, deployed dynamic testsa

A. Out-of-plane

Frequency (Hz)

Mode number

Symmetric excitation 1

Ambient Vacuum B.

Measured 2

3

4

Predicted 5

6 *

*

0.16 0.251

0.55 0.632

0.781

0.248

0.55

0.94

0.12 0.174

0.50 0.65

0.99 0.74

0.232

* 0.58

0.96

1.00 1.015

-

_

* 0.38

Out-of-plane Antisymmetric excitation Ambient Vacuum

C.

In-plane excitation Ambient Vacuum

*

Asterisk indicates that no analysis was made which included aerodynamic effects.

*

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322

Table 6

Modal frequency (Hz) 0.251

Damping coefficients Table 7 Damping coefficients foi for out-of-plane modes out-of-plane modes - symmetric excitation* antisymmetric excitation Damping Coefficients Method 1 Method 2 Method 3 £a £ Input level f b t in D.A. 0.025 0.0042 0.0055 0.003 0.003 0.0053 0.050 0. 0055

0.632 0.781

3- Notes:

Mode frequency (Hz)

fa 0.178 0.178

0.10 0.20

0.0275 0.0226

1) Method 1 determined from decay of motion, a is decay from response to step input and b is decay from dwells at a natural frequency. 2) Method 2 is determined from frequencies of peak in-phase response to a slow sinusoidal sweep. 3) Method 3 is determined from comparison of response with analytical responses.

Damping coefficients ( £ ) Method 1

0. 0069

f

b

0.0060

0.0025

0. 0054

0.0050

0 037 0 043

0.01 0.05 0.10 0.25

0.010 0.010 0.011 0.012

0.01 0.05 0.10 0.25

0.015 0.015 0.018 0.019

0.65

0.74

Method 2 Input level t in D. A.

As anticipated, very slow sweep rates were required to determine the natural frequencies. Figure 10 shows the sweep used to define the first bending mode frequency. Seventy-one minutes were used to sweep the frequency range from 0. 23 to 0.26 Hz. Long test times are required both for frequency sweeps and for mode shape measurements at the low natural frequencies that are characteristic of this class of mechanical system. Note that gravity forces were significant in these tests. The approach used was to include the effects of gravity in the analysis and compare the analytical and test results. If agreement was achieved the analytical model had good prospects of being valid for zero g conditions when gravity effects were eliminated.

In-Plane Structural Characteristics The deployed dynamic test results for in-plane motion did not agree with analytical predictions. Review of the test analysis and test results indicated the stiffness of the system had been underestimated. Blanket tension effects were suspected to be the cause and an experimental program was initiated. The Engineering Demonstration Model (EDM) from the feasibility study 1 was selected as the test unit. The EDM is half the width of the 250-sq-ft unit and is fitted with bare Kapton blankets. The mechanical arrangement is similar,, except that the unit is not equipped with slip rings. The EDM was selected because blanket

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OUT OF PLANE SYMMETRIC EXCITATION INPUT • 0.05 IN (DA)

»— PREDICTED RESONANT FREQUENCY

2.67 DISPLACEMENT

f

IN PHASE RESPONSE

IDA)

2.67

1 V 1

2.67 OISPUCEMCNT IN IDA)

k r^\

V

2.67

A] |f

SWEEP (UNCf 0.23 TOOL 26 Hz

i 0.2

OUT OF PHASE RESPONSE

L 71 J SWEEP RATE |~MIN 1 17). 100 SECONDS/ t DECADE t 0.25 0.3 FREQUENCY (Hz)

Fig. 10 Narrow band frequency sweep.

Fig. 11 Over-all view of test setup.

DEPLOYED LENGTH. L ' 8 FT. BLANKET TENSION, T - 2 LB. PER SIDE

DEFLECTION SCALE:

Fig. 12 System force deflection characteristic (in-plane displacement)

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tension effects can be made significantly larger than gravity effects, due to the lighter blankets and the reduced width. The full-scale aspect ratio can be achieved at half the deployed length where the boom can sustain higher column loads thereby allowing higher blanket tensions. The test set-up is shown in Fig. 11. The force deflection characteristics of the system were determined by measurements at the boom tip. Blanket tension was found to contribute stiffness to the system in a nonlinear fashion. A typical set of data is given (Fig. 12) in which the small inner loop shows a small deflection cycle and the outer loop a large deflection cycle. The large deflection force characteristic can be represented by the relation 3

L (1 + 4a)

"

£T________ U , IW 9

15 (1 + 4a) L

*6

T

(

*

where F = transverse force (pounds) a

=

EI/KRL

W

=

blanket width (ft)

El

=

2 section modules of BI-STEM (Ib-ft )

K

=

root stiffness of BI-STEM (ft-lb/rad)

L

=

deployed length (ft)

T

=

blanket tension (Ib)

6

=

tip deflection (ft)

In the region near the undeflected position (denoted Region 1) the stiffness is greater. In this region the blanket tension is being redistributed on the drums. This stiffness was determined empirically from the test data as

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where k

=

region 1 stiffness (Ib/ft)

k

=

region 2 stiffness (Ib/ft)

T

=

blanket tension (Ib/ft)

L

=

deployed length (Ib/ft)

C

=

empirical constant

Ci

C had an average value of 696 for the test conditions of this program. The results of this investigation show that the stiffness for inplane deflections is increased about 30% in zero g by blanket tension and provide an analytical representation of the observed characteristics for use in dynamics analysis. The in-plane frequencies will be 15-20% higher than the out-of-plane modes because of the increased stiffness. High tension in the deployed dynamics tests, caused by gravity forces had a more significant effect. Wrap Tension for Stability in Vibration A parametric investigation was made to determine the value of blanket wrap tension required to stabilize the blankets on the drums during stowed vibration. This value was found to be 10.4 lb/side or 0.226 Ib/in. of blanket width. This tension can be provided by an auxiliary tensioning device used when the blanket is being stowed before launch (or a stowed vibration test).

V. Summary and Conclusions

The test program described achieved the objective of providing data for a specific configuration of a large area, lightweight deployable solar array. In addition, test techniques applicable to other configurations were developed and demonstrated. Unusual test techniques were necessary because of the large size of the system and its lightweight structure. State-of-the-art advances were made in low-frequency dynamics testing and in measuring the displacements of structural members with electro-optical instruments.

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K. L. HANSON

The environmental tests demonstrated that the design was capable of withstanding a range of environments that should include all launch vehicles applicable to this equipment. Data on the dynamic response of rolled up solar array blankets were obtained. The amplification factors were low with respect to all forms of excitation: pyrotechnic shock, acoustic, and mechanical vibration. Large size and lightweight structure are characteristic of a category of spacecraft equipment of particular interest in communication satellites. The dominant test problem is to accommodate the gravity forces in a manner that does not interfere with the tests because minimum weight structures in this category are not selfsupporting in Earth T s gravity field. In this program this was a factor in deployment tests, deployed dynamics, and deployed thermal vacuum. Special considerations were necessary in each case.

Aerodynamic forces, significant because of the large area of these systems, can be eliminated by testing in a vacuum.

References "Final Report - Feasibility Study: 30 Watts Per Pound Rollup Solar Array, n Rept. 68SD4301, June 21, 1968, General Electric Co.,

Philadelphia, Pa.

2

"Quarterly Report No. 1 - Rollup Subsolar Array, t f Rept. GE-SSO69SD4282, June 12, 1969, General Electric Co., Philadelphia, Pa.

2

4

"Quarterly Report No. 2 - Rollup Subsolar Array, " Rept. GE-SSO69SD4351, Dec. 15, 1969, General Electric Co., Philadelphia, Pa. "Quarterly Report No. 3 - Rollup Subsolar Array, " Rept. GE-SSO69SD4373, Dec. 15, 1969, General Electric Co., Philadelphia, Pa. "Quarterly Report No. 4 - Rollup Subsolar Array, " Rept. GE-SSO70SD4225, March 13, 1970, General Electric Co., Philadelphia, Pa.

"Final Report - Rollup Subsolar Array, Volume 1 - Program Summary," Rept. GE-SSO-70SD4286, Feb. 1, 1971, General Electric Co., Philadelphia, Pa.

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327

"Final Report - Rollup Subsolar Array", Volume H - Detailed Test Results," Rept. GE-SSO-70SD4286, Feb. 1, 1971, General Electric Co. , Philadelphia, Pa. "Final Report - Rollup Subsolar Array, Additional Tasks, " Rept. GE-SSO-71SD4239, May 20, 1971, General Electric Co. , Philadelphia, Pa.

9

Abbott, D.D., "Lightweight Large Area Solar Arrays, " Proceedings of the Fourth Intersociety Energy Conversion Engineering Conference, Paper 699095, Washington, B.C., Sept. 1969. "Evaluation of Space Station Solar Array Technology and Recommended Advanced Development Programs," Rept. LMSC-A981486, Contract NAS9-11-309, Dec. 1970, Lockheed Missiles & Space Co., Sunnyvale, Calif. Bovier, H.K. and Likins, P. J. , "Attitude Control of Nonrigid Spacecraft," Astronautics & Aeronautics, Vol. 9, No. 5, May 1971, pp. 64-71.

"Detail Specification for 30 Watts Per Pound Rollup Cell Array, " JPL Specification SS 501407 Rev. E . , Oct. 22, 1969, Jet Propulsion Lab., Pasadena, Calif.

12

13

Shoulberg, R.H. and Tuft, R.H., "Equipment Considerations for Ultra-Low Frequency Model Tests," Proceedings of the 42nd Annual Shock and Vibration Symposium, sponsored by the Naval Research Laboratory Shock and Vibration Information Center, Key West, Florida, Nov. 1971.

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POLARIZATION ISOLATION CHARACTERISTICS OF A DUAL-BEAM REFLECTOR ANTENNA J. W. Duncan,

*

+ ± S. J. Hamada, and W. C. Wong*

TRW Systems, Redondo Beach, Calif. Abstract The isolation between two orthogonally polarized beams radiated by a single paraboloid has been determined analytically and experimentally. "Frequency reuse" in satellite communications systems requires adequate («30 db) isolation between orthogonal beams. Measurements with a 30-in.-diam paraboloid at X-band show that 30- to 50-db isolation can be realized over the half-power beamwidth when the beams are linearly polarized. The isolation between coincident elliptically (or circularly) polarized beams depends upon the axial ratio of the antenna, but 30-db isolation is feasible. Polarization coupling to an Earth receiving antenna (polarization tracking) is treated in detail. I.

Introduction

Recently considerable attention has been focused on the technique known as "frequency reuse" which offers the possibility of significantly increasing the data capacity of a

Presented as Paper 72-531 at the AIAA 4th Communications Satellite Systems Conference, Washington, D.C., April 24-26, 1972. This work was performed by TRW Systems under the sponsorship of the International Telecommunications Satellite Consortium (INTELSAT). Any views expressed are not necessarily those of INTELSAT. The authors are pleased to acknowledge the technical support and interest of D. DiFonzo and R. W. Kreutel of Comsat Laboratories, Clarksburg, Kd. ^Manager, Electromagnetic Technology Staff. +Member Technical Staff. ^Staff Engineer.

331

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332

J. W. DUNCAN, S. J. HAMADA, AND W. C. WONG

satellite communications link. Frequency reuse is the technique whereby two or more separate data channels within the same frequency band are transmitted (or received) on separate antenna beams. A case of particular importance is represented by two antenna beams radiated by the same reflector with the beams polarized orthogonally with respect to one another. Sufficient beam isolation must be realized between the orthogonally polarized beams. Under operating conditions, the two beams may be coincident or, alternatively, the beams may be partially overlapping (adjacent) with a specified angular separation between the beam axes. This paper presents the results of an extensive analytical and experimental investigation of the polarization isolation characteristics of a dual-beam reflector antenna.

The analytical study considers the cross polarization properties of a paraboloidal reflector that is illuminated by two different types of primary feeds. One feed is a conical corrugated horn. The second feed is an open-ended circular waveguide. The experimental study utilizes a 30-in.-diam paraboloid (F/D = 0.375) as the test antenna. The half-power beamwidth of the radiation pattern is approximately 3.5°. Measurements are carried out in the frequency range 7.4 to 8.4 GHz. Three different feed systems are used in the reflector to obtain the dual orthogonally polarized beams. In the first case, the two beams are coincident and linearly polarized. In the second case, the beams are displaced (adjacent) and linearly polarized. In the third case, the beams are coincident with opposite senses of circular polarization. Measurements of the isolation between the two orthogonal feed ports (beams) are presented as a function of the coordinate location (6,(j>) relative to the reflector axis. From measurements of the vector field of the reflector, it is possible to calculate the polarization coupling of the

dual-beam reflector to an Earth receiving antenna located at an arbitrary position (9,(j>). In general, the radiation is elliptically polarized. Assuming that the polarization ellipse of the Earth antenna is adjusted for a null response to one of the beams, the coupling of the Earth antenna to the second beam is derived. Alternatively, assuming that the

polarization ellipse of the Earth antenna is adjusted for maximum coupling to one of the beams, the resultant coupling (isolation) to the second beam is derived. Polarization coupling data are presented for the reflector antenna that radiates two coincident on-axis beams.

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POLARIZATION ISOLATION CHARACTERISTICS

II.

333

Cross-Polarization Characteristics of a Paraboloidal Reflector

If a paraboloidal reflector is excited by a linear current element, the secondary field contains a cross-polarized component that is generated by the curved surface of the reflector.2-4 in addition to this mechanism, cross-polarized radiation is produced by feed and strut (feed support) scattering in a conventional front-fed paraboloid. Furthermore, the primary feed itself is necessarily imperfect and radiates a certain amount of cross-polarized field. All of these factors contribute to the total cross-polarization properties of the reflector antenna and, therefore, determine the amount of beam (polarization) isolation that can be achieved between coincident or adjacent beams that are produced by dual-polarized feeds in a reflector. When the two beams are coincident, beam isolation can be realized only by means of polarization discrimination. When two separate (orthogonally polarized) feeds are used to produce two displaced beams, isolation can be accomplished by means of both polarization and amplitude discrimination. The purpose of the analytical study is to evaluate the cross-polarization properties of the reflector for specific primary feeds and, thereby, to determine a feed that is satisfactory for the dual-beam application.

Consider the reflector coordinate system defined in Fig. 1. Two orthogonally polarized feed elements EI and E2 are located at the focal point of the paraboloid. Ej_ is x directed, and £2 is y directed. The origin of the coordinate system is at the paraboloid focus, and point P(r) in the far zone radiation field has rectangular coordinates (x, y, z) and conventional spherical coordinates (r, 6, ). A second set of spherical coordinates (r, 6, $*) with corresponding unit vectors ar, SQ', a^, is used to define the principal and crosspolarized vector field components. Assume for a moment that the primary feed is simply a short electric dipole polarized in the x direction, which has unit amplitude and a length dx. The current J induced on the reflector surface by the dipole field is3 J = ^——.———— Ar l

*•**».

*——

.

sin

V

I

""

z La x

9

(-cos6 + 2 sin (j> cos /

/-S

I

O

- a (sin 2(j> cos

. Therefore, if the primary feed exhibits identical patterns in the principal planes, and if the feed phase center is unique, the feed illumination will cancel the reflector induced cross-pol, and the aperture field will be linear. The dual-mode conical horn excited in the TEn and

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POLARIZATION ISOLATION CHARACTERISTICS

335

modes nearly satisfies the required characteristics.6 However, the dual-mode feed is a narrowband device and therefore is unsatisfactory for our application. The open-ended circular waveguide and the conical horn excited in the dominant TE;Q mode exhibit radiation patterns that are nearly of the form of Eq. (2). However, because f]_(0) = f2(6) is not satisfied identically, in both cases the radiator provides only partial cross-pol compensation. A small square waveguide excited in the TE^Q mode is substantially less satisfactory than the circular waveguide. The conical corrugated horn,' which has excellent pattern symmetry and phase center characteristics, is expected to yield better cross-pol compensation than either the circular guide or the conventional conical horn. To establish the definitions of principal and crosspolarized field components for a dual polarized antenna, again consider Fig. 1 with E^ = axE]_ defined as the x port feed element and £2 = ^y^2 defined as the y port feed element. Each feed element produces a beam in the direction of the z axis and, in the far zone of the reflector, the field resulting from each port can be expressed (and measured) in terms of the vector field components Eg* and EJ*. Thus

6* 'Q



The designation of Eg* or Er as a principal-polarization or cross-polarization component depends upon the feed (or source) orientation. The z axis is the axis of symmetry of the paraboloid. For high gain antennas, the region of interest is a small solid angle centered on the z axis. It can be shown-1that near the boresight axis of the reflector (0«1 rad) Eg' is principal-pol and E$~is cross-pol for the x directed feed, whereas Ep'is principal-pol and-Eg"is cross-pol for the y directed feed. In fact, in the principal planes of the antenna (the xz plane and the yz plane of Fig. 1), these definitions are exact for all values of 9. It also follows that near the z axis unit vector IQ'w -ax and unit vector aV w -ay. III. Analytical Study

Two particular horn type radiators are considered as primary feeds for the paraboloidal reflector. The first is a conical corrugated horn, which has a radiating aperture 1.86 in. in diameter. Because of the corrugations, the maximum diameter of the horn is about 2.8 in. The second radiator is a l.O-in.-diam circular waveguide excited in the TE11 mode, that is, an open-ended circular waveguide. The radiation

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336

J. W. DUNCAN, S. J. HAMADA, AND W. C. WONG

patterns of the paraboloidal reflector are calculated by means of a vector diffraction computer program called "MEASURE." The computer program determines the vector field of the reflector by integrating the current distribution J that is induced on the reflector surface by the primary feed illumination. The surface current Js derives from the incident vector field by means of the physical optics equation Js = (2/c)nx (kxEp), where £ = 120ir is free-space wave impedance, k is a unit vector in the direction of wave propagation, and Ep is the electric field intensity of the primary feed in the direction k. Input data to program MEASURE are the measured amplitude and phase of the vector field components Eg' and E^ of the primary feed pattern. Measurements of the feed pattern are made in an anechoic chamber, and the measured data are digitized on punched cards for input to the computer analysis. Referring to Fig. 1, when the feed pattern is measured in the chamber, the feed horn aperture is located in the xy plane with the electric field being x directed and with the horn radiating in the positive z direction. Thus, ag-defines the principal-polarization component and a^the cross-polarization component of the feed pattern. Program MEASURE takes the measured primary feed pattern and, in effect, rotates the feed horn orientation by 180° so that the boresight axis of the horn is in the negative z direction. Thus, the complex vector field incident on the reflector is completely specified, and ag' and ay define the principal and cross-polarized vector field components of the reflector radiation pattern. The data file defines the complex vector feed pattern at 225 coordinate locations encompassing the reflector surface. Program MEASURE uses linear interpolation to define the field at coordinate locations intermediate to the 225 data points when performing the integration. MEASURE provides for numerous options in the antenna configuration. The reflector surface can be arbitrarily distorted. The aperture can be elliptical as well as circular. The primary feed may be defocused by lateral and axial feed displacements. The feed may be rotated or tilted (angular offset) , and several identical feeds can be included in a multiple feed configuration. The integration of the vector surface currents yields the secondary radiation pattern within the constraints of physical optics. The pattern can be calculated in the Fresnel region as well as the Fraunhofer region of the antenna. The only mechanism that is unaccounted for in the analysis is the effect of feed and strut scattering on the secondary pattern.

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POLARIZATION ISOLATION CHARACTERISTICS

337

An extensive analysis was made to determine the crosspolarization characteristics of reflectors. The study considered the effect of reflector focal length (F/D ratio) , the effect of feed defocusing (axial and lateral displacements), and, of course, compared the relative performance obtained using the two particular primary feeds. In the interest of

brevity, only a small portion of the results from the analytical study will be presented.

Computed secondary patterns for the case where the reflector is illuminated by the conical corrugated horn are shown in Fig. 2. The feed pattern data were measured at 7.9 GHz. The reflector diameter is D = 20A, where A is the wavelength and the ratio F/D = 0.375. Referring to Fig. 1, the feed is polarized in the x direction as indicated by vector E]_. We are not considering a dual-polarized feed; therefore, £ 2 = 0 . The xz plane denotes the E plane, and the yz plane denotes the H plane of the reflector antenna. The principal-pol (E>) pattern and the cross-pol (Ey) pattern are calculated in the cj) = 45° plane as defined in Fig. 1. The patterns are normalized with respect to the on-axis (6 = 0) gain for the principal-pol component. The interesting result is the equivalent of a Condon lobe in the 45° plane, which reaches a maximum of -30.8 db at 6 = 3.4°. Over the half-power beamwidth of the principal-pol pattern, the cross-pol amplitude is in the range -32.5 to -34.8 db. Neglecting the effect of feed and strut scattering, Fig. 2 indicates that the polarization isolation between two coincident on-axis beams should be on the order of -32.5 db. When the reflector is illuminated by the open-ended circular waveguide, the patterns of Fig. 3 are obtained. The feed pattern data were measured at 7.9 GHz, As in the case of the corrugated horn, the feed is polarized in the x direction, and we show the principal-pol and cross-pol patterns in the $ = 45° plane. The on-axis cross-pol amplitude is -33 db compared to -34.8 db that resulted with the corrugated horn feed under the same conditions. Thus it can be concluded that the corrugated horn provides slightly better X-pol compensation than the openended circular waveguide feed. To assess the cross-pol canceling properties of the two feeds, the secondary patterns resulting for a modified Hertz dipole feed have been calculated. The dipole feed is located at the paraboloid focus and is polarized in the x direction. The dipole feed pattern E (6,) is the function

Ep(9,cj>) = a~ cos cos 0 + sin cj> cos 6

(4)

•IN

P

O

H»

rt

O P

•

HH P CT O CD H-

O X) CD

CO

hi CD

CD hj

PJ rt rt CD

CO

rt £ O o

o.

O ON 3 ^

CO

h{ i-{ CD

O h- CD C CD O h-1 P *O O

P ^e- rt PJ rt CD *~s & rt • CD CD rt Pa O PJ

h-1 CD rt

CD

W rt

•

• a.

,£* CD H' CD N >n

CO >n

(D M

hh (D O

VD O

CD -vl CO CO

-Prt Ln O OHs~* O hh rt pj < H N-Pl—» PJ1 53* —'

O

H' O

a. h{

§CO

hi p CD H* PJ hh PJ O rt rt (D hi

rt p^ CD H' CD hi

O CD -e- rt {x| p co

C *

hh bd

CO

H- C

P

CD P.

CD

PJ hi P PJ P.* PJ

CD

P* rt

H P hh

SI

I3

CD Td (JQ O PJ CO M ft O CD C O P . O " P' O

rt O PJ rt rt

co c

CD h{ P

PJ

h{

O h{ O

rt p

PJ rt

*o

| *

O

P. P O H« O O hi PJ CD pJ M rt O1 X h{ rt hi h-» P. O h- I C

•TCJ O hh ^

•

OQ TJ CD O >d

(D O

CD rt CD rt P- CD

o o n P

OQ

O rt

O

rt hh1 P- H

H» CD

K>

= 45° OPRINCIPAL-POL AX-POL D/\ = 20 F/D = 0.375

-40

0

2

4

6

POLAR ANGLE 8(DEGREES)

Fig. 4

Reflector radiation pattern with Hertz dipole feed.

scans in near coincidence with the principal-pol pattern and the peak amplitude of the X-pol pattern remains nearly constant. For example, the peak X-pol amplitude increased 0.6 db when the beam was scanned 13° from boresight, that is, about three times the half-power beamwidth. Thus, off-axis coincident beams obtained using a dual-polarized displaced feed should provide polarization isolation of about the same order as an on-axis feed. The l-in.-diam (2X/3 at 7.9 GHz) circular waveguide feed is approximately 1/3 the diameter of the corrugated horn. Placing one feed on-axis at the focus and a second feed imme-

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POLARIZATION ISOLATION CHARACTERISTICS

341

diately adjacent to provide a second displaced beam, the minimum beam separation is about 4°, using circular waveguide feeds, and is about 13° using two corrugated horns. An important aspect of the experimental investigation was to measure the polarization isolation between two closely spaced beams. Because of this fact, the circular waveguide feed was chosen for the experimental investigation rather than the corrugated horn. The reflector is a 30-in.-diam paraboloid (20A at 7.9 GHz) with focal length F = 0.375D. Feed and strut scattering are expected to introduce X-pol radiation so that the resulting beam isolation for on-axis coincident beams should be something less than -33 db.

IV.

Test Antenna Descriptions

Three basic antenna configurations are considered in the experimental investigation. In the first case, the antenna radiates two coincident on-axis beams that are linearly polarized. In the second case, the antenna radiates two separate linearly polarized (adjacent) beams. In the third case, the antenna radiates two coincident on-axis beams that are polarized with opposite senses of circular polarization. A photograph of the reflector/feed configuration for radiating two coincident linearly polarized beams is presented in Fig. 5. The feed is supported in front of the reflector by four symmetrically placed struts. Two semirigid coaxial cables are fastened along two adjacent struts and are connected to the orthogonal ports of the dual-polarized feed, which is shown in Fig. 6. The feed consists of a rectangular waveguide orthomode transducer (OMT), coaxial probe terminations of the waveguide OMT, a tapered transition (circular), and the l-in.-diam circular waveguide radiator. It will be seen in results to follow that scattering from the asymmetrical feed structure affected the magnitude and symmetry of the Fig. 5 Reflector/feed for two coincident linearly beam isolation data. polarized beams. This particular feed

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342

J. W. DUNCAN, S. J. HAMADA, AND W. C. WONG system was used in the experimental investigation because the OMT was an available off-the-shelf component.

A photograph of the feed used to radiate two separate linearly polarized beams is shown in Fig. 7. The on-axis beam

Fig. 6 Primary feed for dual-orthogonal linear polarization.

Fig. 7 Primary feed for two displaced linearly polarized beams.

results from exciting only one port of the two-port feed. The displaced beam is produced by a second l-in.-diam circular waveguide that can be fixed in three lateral positions. As a result, the off-axis beam can be located at angles of 5.5°, 8.4°, and 12.6° from the boresight axis. Referring to Fig. 1, the off-axis feed is displaced in the +x direction; thus the beam is squinted in the xz plane in the direction defined by = IT, ? > 90°. The on-axis beam is polarized in the y direction; the off-axis beam is polarized in the x direction.

The primary feed structure used to radiate coincident onaxis beams that are polarized with opposite senses of circular polarization (CP) is shown in Fig. 8. An X-band circular waveguide polarizer, which has a 1.2-in.-diam was adapted to the OMT to provide circular polarization. The x oriented port produces right-hand circular polarization; the y oriented port

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343

POLARIZATION ISOLATION CHARACTERISTICS

Fig. 8 Primary feed for

coincident circularly

polarized beams.

produces left-hand circular polarization. The CP feed, which does not represent an optimum engineering design, exhibited an axial ratio of about 3.0 db on-axis. As will be seen subsequently, the axial ratio of the dualpolarized CP feed is of prime importance in determining the polarization isolation between coincident CP beams.

V. Polarization Isolation Measurements It was pointed out in Sec. II that Eg* is principal-pol and E$ is cross-pol for the x port beam, whereas E$ is principalpol and Eg is cross-pol for^ the y port beam. When measurements are performed on the antenna pattern range, the x axis of the coordinate system (Fig. 1) is oriented horizontally. The z axis remains normal to the paraboloid aperture. The axis of the azimuth rotation of the test antenna is perpendicular to the plane containing the x axis and radius vector r, which is directed toward the transmit source at position P. Thus, the xr plane is the horizontal plane on the pattern range. Azimuth rotation of the test antenna corresponds to variation of angle (T while holding angle If constant (cf pattern cuts). Different If pattern cuts are obtained by tilting the paraboloid (the z axis) relative to the horizontal (the xf plane). Because the xr plane is horizontal, the transmit source at P is polarized horizontally (vector a^) to measure Eg' and is polarized vertically (vector a^) to measure Ey. In this manner, the principal and cross-polarized vector field components are measured in the 0",) denote the gain of the antenna in the direction (0,) of the Earth antenna and let Gr denote the peak gain of the receiving antenna. The power Pr received by the Earth antenna can be expressed in the form

Pr = Pt(A/47rR)2 Gt(0,4>) GrF

(8)

where P is the transmitted power, R is the distance between the satellite and Earth antennas, X is the wavelength, and F is the "polarization coupling factor." F takes into account the polarization of the receiving antenna relative to the incident wave. When the polarization ellipse of the receiving antenna is identical to the polarization ellipse of the incident wave, F = 1, denoting total coupling to the incident wave. The polarization of the receiving antenna also can be orthogonal to the incident wave, in which case F = 0, denoting zero coupling to the incident wave. Coupling factor F is given by°

(l+r?)(l+r?) ± 4rlr9 + (1-r?)(1-rJ) cos 2^

where r^ and TJ^ are the axial ratio and angle T of the incident wave polarization ellipse, r2 and i^ are ^ne axial ratio and angle T of the receiving antenna, and ip = (T2~T^) is the angle between the major axes of the two polarization ellipses. The sign of the term "t4r^r2 ^s cnosen + or - as the polarization ellipses have the same or opposite sense, respectively. Equations (8) and (9) apply to the fields resulting from both the x port and the y port of the dual-beam antenna. As a result, two polarizaiton coupling factors can be defined which correspond to the simultaneous coupling of the receiving antenna to the x port and y port beams. Two particular polarization conditions for the Earth receiving antenna define the unique coupling factors F0 and F]_. First, assume that the polarization of the receiving antenna is adjusted to be orthogonal to the x port beam; then F0 is defined as the resulting coupling factor for the y port beam. Thus, Fo denotes the polarization coupling to the y port beam under the condition of a null response (F = 0) to the x port beam. Second, assume that the polarization of the receiving antenna is adjusted for maximum coupling (F = 1) to the x port beam; then F^ is the resulting factor for the y port beam. Thus, FI describes the isolation of the receiving anten-

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POLARIZATION ISOLATION CHARACTERISTICS

349

na to the y port beam under the condition of maximum coupling to the x port beam. If the x port, y port designations are interchanged, the result is that the coupling factors FQ and Fj_ are unchanged . In other words, if the receiving antenna is adjusted for a null response to one beam, Fo is the resulting coupling factor for the second beam.. Conversely, if the receiving antenna is adjusted for maximum (unity) coupling to one beam, F-^ is the resulting coupling factor for the second beam.

It is convenient to use a parameter S to define the sense of an elliptically polarized wave, thus:

S = +1, denotes clockwise rotation CW (wave approaching) S = -1, denotes counterclockwise rotation CCW (wave approaching) In a given direction (0,), the field from either port (beam) is elliptically polarized with polarization ellipse parameters r, T, and S defining the axial ratio, the angle of major axis orientation, and the polarization sense, respectively. If we use subscripts x and y to denote the x port and y port fields, the coupling factors F0 and F^ are given by the following expressions^ :

(l+r2)(l+r2)-4(S .S )r r -(1-r2) (l-r2)cos 2(r -T )

2 2 (l+rx2)(l+r2)+4(Sx .S )r x r +(l-r x ) (l-r )cos 2(T x -T )

y

y

y

y

y

The complex vector field corresponding to each port of the antenna is required to calculate coupling factors Fo and F-L. If we assume a symmetrical antenna structure, measurements can be performed when only the x port of the antenna is excited. Because of rotational symmetry, a matrix transformation can be applied to reconstruct the field that would result for the orthogonal y port excitation. The ellipse parameters for the x port and y port fields then can be derived from the magnitude and phase of the Eg% Ey components.

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350

J. W. DUNCAN, S. J. HAMADA, AND W. C. WONG

The x port field of the antenna was measured over a conical angle centered on the paraboloid axis. Measurements were performed at 7.4, 7.9, and 8.4 GHz. The procedure outlined above was then used to determine F0 and F^ as a function of coordinate position (0,). Typical results are given in Table 1 for the frequency 7 . 9 GHz . Table 1 Polarization coupling factors (7.9 GHz)

F

+ (deg)

(deg)

(db)

F l (db)

0.0 0.0 0.0 0.0 0.0 0.0 0.0

0.0 1.0 2.0 3.0 4.0 5.0 6.0

-1.811E-03 -2.313E-03 -3.133E-03 -7.259E-03 -1.150E-02 -1.186E-02 -3.497E-02

-3.380E+01 -3.274E+01 -3.142E+01 -2.777E+01 -2.578E+01 -2.564E+01 -2.096E+01

45.0 45.0 45.0 45.0 45.0 45.0 45.0

0.0 1.0 2.0 3.0 4.0 5.0 6.0

-1.811E-03 -1.713E-03 -2.281E-03 -4.051E-03 -1.991E-01 -1.028E+00 -2.621E4-00

-3.380E+01 -3.404E+01 -3.280E+01 -3.030E+01 -1.349E+01 -6.762E+00 -3.438E+00

90.0 90.0 90.0 90.0 90.0 90.0 90.0

0.0 1.0 2.0 3.0 4.0 5.0 6.0

-1.811E-03 -1.062E-03 -5.688E-04 -3.793E-04 -1.682E-03 -5.771E-03 -2.640E-02

-3.380E+01 -3.612E+01 -3.883E+01 -4.059E+01 -3.412E+01 -2.877E+01 -2.218E+01

e

0

The polar angle 6 has the range 0 < 8 < 6° in 1° increments. Recall from Fig. 3 that the principal-pol pattern level is -30 db relative to peak gain at 6 = 6°. We show data for the cases where = 0, 45°, and 90°. The coupling factors are expressed in db. The second column of the coupling factor indicates the exponent of 10 associated with the magnitude given in the first column. For example, at the position = 45°, 6 = 4°, we find that F0 = -0.199 db and Fx = -13.49 db. These data indicate that, if the receiving antenna is adjusted for a null response to one beam, only 0.2-db polarization loss results to the second beam. Alternatively, if the antenna is adjusted for unity, coupling to one beam, the coupling loss (isolation) to the second beam is -13.5 db.

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POLARIZATION ISOLATION CHARACTERISTICS

351

A study of the results obtained from the frequencies of 7.4, 7.9, and 8.4 GHz indicates that essentially perfect beam

isolation can be realized if the polarization parameters of the In fact, polarization tracking is effective even in the sidelobe region of the dual-beam reflector. Earth antenna can be adjusted to the required values.

VII.

Coincident Circularly Polarized Beams

The reflector/feed configuration produced coincident onaxis beams with opposite senses of elliptical polarization (rather than circular) because the axial ratio of the feed was on the order of 3 db. The range transmit source radiated a right-hand elliptically polarized beam (0.5-db axial ratio) at a frequency of 7.9 GHz. Data from the CP isolation measurements [as defined in Eq. (7)] are presented in Fig. 13. The port-to-port isolation is about 21 db over the half-power beamwidth. These low values of beam isolation are consistent with the axial ratio of the feed (the feed axial ratio results from the polarizer/OMT and not the circular waveguide radiator). Beam isolation can be substantially improved (increased) by reducing the axial ratio of the feed; that is, the feed characteristics effectively determine beam isolation for coincident on-^-axis beams assuming, of course, that the incident wave has a small axial ratio.

92°

91°

90°

Freq = 7.9 GHz Tx - RHCP

Mon'z. Port (P.HCP Vert. Port (LHCP) Date: 9/23/71

90°

- 3 dB CONTOUR

Fig. 13

Polarization isolation for coincident

elliptically polarized beams.

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352

J. W. DUNCAN, S, J. HAMADA, AND W. C. WONG

The port-to-port isolation is determined by the coupling of the incident wave to the oppositely polarized ports, and this depends on the ellipse parameters associated with each port and the incident wave. To discuss the effect of the feed axial ratio, we make the simplifying approximation that beam isolation is equal to the polarization coupling factor for the orthogonal (opposite sense to the incident wave) port. If we assume an elliptically polarized wave incident on the dualbeam reflector, the orthogonal port coupling factor is-*-

F = ——— ———— ±————

——— ±———— £-

( 1 2 )

where r-j_ is the axial ratio of the incident wave, and r2 is the axial ratio of the orthogonal port beam. The plus sign in Eq. (12) yields the minimum coupling loss (coincident ellipse axes) and the minus sign yields the maximum coupling loss (orthogonal ellipse axes). If we assume 0.5-db axial ratio (r]_ = 1.059) for the incident wave, the maximum and minimum beam isolation values as derived from Eq. (11) are shown as a function of ^2 in Fig. 14. Notice that, if ^2 = 0.5 db, the orthogonal port isolation can be between 25 db and infinity (F = 0) depending on the relative orientation of the ellipse axes. If the major axes of the two ellipses coincide, the isolation is less than 30 db for all values of r2 except r£ = 0 db. The test data of Fig. 13 indicate that the axial ratio of the test antenna was on the order of 2 db . If the incident wave is circularly polarized, the orientation of the orthogonal port ellipse to the incident wave is immaterial, and the polarization coupling factor has only the single value F=101og10[2(l+r2)/(l-r2)] db

( 1 3 )

A graph of Eq. (13) is shown in Fig. 15. Thus, if the incident wave is circularly polarized, beam isolation will exceed 30 db when the orthogonal port axial ratio is less than 0.55 db.

The importance of the CP feed design is illustrated by considering the effect of polarizer phase error on the feed axial ratio and beam isolation. The polarizer is designed to provide a 90° phase differential between the orthogonal field components of the OMT. If 6 (degrees) denotes the phase deviation of the components from the required 90°, the resulting feed axial ratio is

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POLARIZATION ISOLATION CHARACTERISTICS

353

INCIDENT WAVE

AXIAL RATIO = 0.5 dB

RECEIVING ANTENNA AXIAL RATIO (dB)

Fig. 14

Orthogonal port isolation vs. antenna axial ratio. Axial ratio = 0.1516 db

Equation (14) shows that a 3.3° polarizer phase a 0.5-db axial ratio and, from Fig. 15, 30.8-db the condition of an incident circularly polarized larger phase error will cause the isolation to be 30.8 db.

(14) error produces isolation under wave. A less than

The experimental measurements and analysis of Sec. VI can be extended to derive the isolation characteristics for coincident circularly polarized beams. Introduction of the time phase factor e-j^/2 into the complex y port field and then summation of the x port and y port fields yields a circularly polarized field (on-axis) by definition. This approach allows the effects of feed axial ratio and phase error to be separated out from the reflector X-pol characteristics. In other words,

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354

J. W. DUNCAN, S. J. HAMADA, AND W. C. WONG

0.2

0.3

0.4

0.5

0.6

RECEIVING ANTENNA AXIAL RATIO (dB)

Fig. 15

Orthogonal port isolation to a circularly polarized wave.

the analysis assumes a perfect feed. The axial ratio of the orthogonal port is calculated at a given location (§^,T), and the maximum and minimum isolation values corresponding to that axial ratio are extracted from Fig. 14. The analysis assumes an incident wave with 0.5-db axial ratio. The results of this analysis are shown on the coordinate grid in Fig. 16. The calculated axial ratio and corresponding isolation values are entered at each coordinate intersection point. These data show that 30-db isolation is feasible for coincident on-axis beams provided that a first-rate circularly polarized feed is utilized. The data are not symmetric about the polar axis for the reasons discussed in Sec. V.

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POLARIZATION ISOLATION CHARACTERISTICS

355

Axial Maximum Ratio (dB) Isolation (dB) Minimum Isolation (dB)

Qo

89U

90U

91U

92°

92U

91U

1.87

1.16 28.5 20.2

1.08 29.8

88U

0.75 37.1

22.8

0.91 1.8

1.29

0

30.8

3.5

30.8

12.3

0.59 45.0

22.8

.34.5

0.64 42.0 23.6

0.51 5 0.01 0.9831.3 21.2

0.26r.O

27.1

23.5

0.91 32.8

0.75 37.1

26.0

24.0

0.0

0.67 40.7

0.36 41.2

0.88 33.2

0.42 44.9

1.19 28.2

21.8

Fig. 16

1.385.8 9.2

1.14 29.0 20.4

Orthogonal port coupling factors assuming a perfect CP feed.

VIII.

Summary and Conclusions

The antenna considered in this investigation is a paraboloidal reflector that radiates two coincident on-axis beams or, alternatively, two separate displaced beams. The orthogonal beams can be linearly or circularly polarized. The results show that 30-db isolation between orthogonal ports can be realized over the half-power beamwidth of the beams. Under many conditions the isolation is much greater than 30 db.

Cross-polarized radiation from the reflector/feed system determines the realizable polarization isolation. When used

Purchased from American Institute of Aeronautics and Astronautics

356

J. W. DUNCAN, S. J. HAMADA, AND W, C. WONG

as primary feeds, the conical corrugated horn and the openended circular waveguide radiate fields that reduce the crosspolarized currents generated by the reflector surface. _The differences in port-to-port isolation that resulted for a0* source orientation and a^ source orientation emphasize the importance of complete antenna and feed symmetry. Antenna symmetry is particularly important when the coincident beams are circulatly polarized. Analysis shows that coincident circularly (or, more correctly, elliptically) polarized beams can yield 30-db isolation over the half-power beamwidth provided that the axial ratio of the elliptically polarized antenna is sufficiently small (< 0.5 db). The exact amount of isolation depends upon the axial ratio of the incident wave. The CP feed must be designed with great care and be rotationally symmetric to insure good CP performance. Over the half-power beamwidth of coincident beams, orthogonal linear and orthogonal circular polarization are equally effective methods of achieving polarization isolation. The calculation of polarization coupling to an Earth antenna using a dual-polarized linear feed shows that essentially perfect polarization isolation can be achieved if the polarization parameters of the Earth antenna can be adjusted to the required values. References

•4)uncan, J. W., "Antenna Beam Isolation Investigation," Kept. 17565, Dec. 1971, TRW Systems Group, Redondo Beach, Calif. 2 Condon, E. U., "Theory of Radiation from Paraboloidal Reflectors," Westinghouse Research Rept. SR-105, Sept. 1941, Pittsburgh, Pa.

o

J

Jones, E.M.T., "Paraboloid Reflector and Hyperboloid Lens Antennas," IRE Transactions on Antennas and Propagation, Vol. AP-2, July 1954, pp. 119-127. ^Kerdemelidis, V., "A Study of Cross Polarization Effects in Paraboloidal Antennas," Tech. Rept. No. 36, May 1966, California Institute of Technology, Antenna Lab., Pasadena, Calif. Rusch, W.V.T. and Potter, P. D., Analysis of Reflector Antennas , Academic Press, 1970, pp. 29-30. Potter, P. D., "A New Horn Antenna with Suppressed Sidelobes and Equal Beamwidths," Tech. Rept. No. 32-354, Feb. 1963, Jet Propulsion Lab., California Institute of Technology, Pasadena, Calif.

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POLARIZATION ISOLATION CHARACTERISTICS

357

'Clarricoats, P.J.B. and Saha, P. K., "Propagation and Radiation Behavior of Corrugated Feeds; Part 2, Corrugated Conical Horn Feed," Proceedings of the ZEE, Vol. 118, Sept. 1971, pp. 1177-1186. ^Rumsey, V. H., Deschamps, G. A., Kales, M. L., and Bohnert, J. I., "Techniques for Handling Elliptically Polarized Waves with Special Reference to Antennas,11 Proceedings of the IRE, May 1951, pp. 533-552.

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MEASUREMENTS OF EARTH STATION ANTENNAS G/T RATIO BY RADIO STARS AND SATELLITES Henry J. Kochevar* Westinghouse Electric Corporation, Baltimore, Md.

Abstract A new technique has been developed to measure accurately the G/T ratio of a small aperture antenna using geostationary satellites and the well-established radio star method. This technique is based on the use of two antennas. An accurate measurement of the G/T ratio of a large aperture antenna (reference antenna) is first obtained by using a radio star as a source of known power. Then a small aperture antenna is used in conjunction with the large antenna; two C/N ratios are thus obtained in terms of the signals radiated by a satellite. After normalizing the two C/N ratios to the system noise temperature of the large antenna, the G/T ratio, and hence the gain of the small antenna, is determined.

Presented as Paper 72-528 at the AIAA 4th Communications Satellite Systems Conference, Washington, D.C. , April 24-26, 1972. The test measurements for the technique described herein were performed under contract for the NASA Applications Technology Satellites (ATS) program, Goddard Space Flight Center, Greenbelt, Md. I would like to thank in particular Mr. E. Metzger, Operations Chief of the ATS program, for his cooperation and aid in providing the ATS Data Acquisition Facilities for this test. Appreciation is also acknowledged to Mr. J. Miller at NASA/GSFC for his helpful criticism in reviewing the manuscript. *Senior Engineer. 359

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360

H. J. KOCHEVAR

Nomenclature

Ae (C/N) r Cr Cs f G k n Pn PS

Pst r Rr Rs T TS T

st

oc AT X 0f 0n

00 04

= GX2/47r effective area of the antenna = 20. 5 db carrier-to-noise ratio of reference antenna = atmospheric attenuation for the reference antenna = atmospheric attenuation for the small aperture antenna = measuring frequency of reference antenna system in GHz = gain of Earth station antenna = Boltzman's Constant = 1. 38 x 10~23 Joules/°K = number of years from Jan. 1, 1968 to June 13, 1969 = 1.45yr = noise power of Earth station corresponding to system noise temperature T = 00Ae = 00 (GX2)/4*" total flux power received at the ground station receiver from the satellite = noise power received from radio star = signal-plus-noise-to-noise ratio (measured) = (P s t+P n )/P n = 37,292 km range of reference antenna to satellite = 3 8 , 8 8 7 km range of small aperture antenna to satellite = system noise temperature of the Earth station = system noise temperature of reference antenna when not pointed to the satellite, but at the same elevation angle = system noise temperature of reference antenna when not pointed to the radio star Cassiopeia A, but at the same elevation angle = elevation angle of corresponding antenna = correction factor for normalizing the G/T value to a reference system noise temperature = wavelength of measuring frequency = 0. 0718 m (4.178 GHz) = flux density at the measuring frequency in w/m 2 /Hz = flux density (noise power density) radiated from the radio star Cassiopeia A at the measuring frequency and time corrected from Jan. 1, 1968 in w/m 2 /Hz = flux power density received from the satellite in w/m2 = established flux density at 4 GHz and at the date of Jan. 1, 1968 - 1065 x 10"26 w/m 2 /Hz (Table 1)

Subscripts f n ns o or os r s st

= frequency = number of years = normalized small antenna = received power = received power by reference antenna = received power by small antenna = reference - small antenna and satellite = star

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361

G/T MEASUREMENTS

I. Introduction

The reference antenna G/T determination by the radio star method is discussed and then followed by the computation of the G/T value. The small aperture antenna G/T is next discussed and followed by a presentation of the relation of the G/T ratios between the small and the reference antennas. The final step in this procedure is the computation of the resultant G/T value of the small aperture antenna. Significant characteristics of both antennas are presented in the antenna characteristics tables shown in the text.

A derivation of the expression which determines the G/T by the radio star method appears in Appendix A. Appendix B contains a detailed derivation of the equation which relates the G/T ratios between the small and the reference antennas. II. Reference Antenna G/T by the Radio Star Method

The establishment of the reference antenna G/T ratio by the radio star method was performed with an Earth station Cassegrainian type of antenna having a diameter of 85 ft and located at the ATS Ground Station, Rosman, N.C. Positioning of the antenna to follow a desired pattern is accomplished with its MXM and !I Y M axis system that is driven by a programmed tape with a computer. From a source of three radio stars (Cassiopeia A, Cygnus A, and Taurus A), the star Cassiopeia A has the largest power density,6 as shown in Table ll at the 4-GHz frequency. Therefore, it was selected for the G/T measurement. The radiated power density from the other two radio stars is too low in power level for accurate G/T measurement of the 85-ft reference antenna. Table 1 Radio Star

Cassiopeia A 1065xlO~ 26 W/M 2 /Hz

Flux density at 4 GHz 04 Flux density at Other frequency

Flux density of radio stars

0~

^-0.75

Established value for Jan. 1, 1968.

Taurus A a

717xlO" 26 W/M2/Hz

'*$ )-°- 25

Cygnus A 495xlO~ 26 W/M 2 /Hz

Mi)"' 19

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362

H. J. KOCHEVAR

III. Reference Antenna G/T The first step in determining the G/T value for the reference antenna was to measure the signal-plus-noise power of the received radiation from the star Cassiopeia A. With the antenna position fixed relative to Earth and pointing to the star, the antenna beam was allowed to move approximately 10° away from the star's position in space by Earth's rotation. At this cold sky position the system noise temperature (T) was measured where a minimum radiated sky noise power was received. An average noise temperature of 68 °K was measured at this antenna position. At least six measurements of the above two parameters were taken to insure repeatable results. An average signal -plus-noise-to-noise (Pst + Pn)/pii P°wer ratio of 3.48 db was established from the above two measured parameters with a maximum variation of 0. 13 db in the values measured.

The G/T ratio of the reference antenna is next derived from the (Pgt + Pn)/Pn value obtained and the established radiated power density from the star Cassiopeia A. Calculation of the G/T ratio is then determined from the expression recommended by CCIR^-» 2 and derived in Appendix A:

G/T = [87rk(r-l)]/X20n

(1)

The flux density or noise power radiated from Cassiopeia A at the frequency of 4. 178 GHz was derived by using the established flux density value of 1065 x 10~26 w/m2/Hz shown in Table 1 and the following expression: -75

(2)

Substitution of the value of 04 and a measuring frequency of 4. 178 GHz into Eq.(2) results in ./

-2fi4 1 7 8 ~ ^

0f - 1065 x 10

(~^-J

= 1030 x 10

o£ Zb

o

w/ni /Hz at 4.178 GHz

and date 1/1/68. This flux density also can be obtained from the graph in Fig. 1 but to a smaller degree of accuracy. Figure 1 and Eq. (2) show the relationship of the radiated flux density (Cassiopeia A) with frequency and indicate the flux density decreases with an increase in frequency. The flux density from Cassiopeia A also decreases with time. Therefore, before it can be applied to Eq. (1), it must be corrected for the time elapsed from Jan. 1, 1968, of the established flux density 0p to the time of the G/T measurement for the reference

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G/T MEASUREMENTS

363

antenna. This new flux density value 0n is determined from 0n = 0 f (0.989) n

(3)

Substitution of the values of the foregoing two parameters into Eq. (3) results in a flux density that is corrected for frequency and time 0 - 1030 x 10"26 (0. 989)1' 45 - 1014 x 10~26 w/m2/Hz

for date 6/13/69. This value of time-corrected flux density also can be determined, but less accurately, from the graph in Fig. 2. The G/T ratio for the reference antenna now can be calculated by applying the known parameters into Eq. (1) and becomes 8.x 1.38x10 -

00

23-1)

=




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Purchased from American Institute of Aeronautics and Astronautics

DESIGN OF THE INTELSAT IV TRANSPONDER

385

Fig. 14 Input Multiplex assembly on completed communication compartment. position of the output TWT's and power supplies on a completed communications compartment.

Output Multiplexers

The output of each HLTWT is fed to one of four output multiplexer assemblies through a harmonic filter and a network of switches as shown in Fig. 13. Four output multiplexers are included on each spacecraft to accomplish the routing to the two global and two spot-beam transmit antennas. Odd channels can be connected to one of the global horns or to the West spotbeam transmit antenna, whereas even channels go to the second global or to the East spot-beam antenna. Channels 1-8 can be switched on command to either a global or a spotrbeam antenna, whereas channels 9-12 can be used only for global transmission. Each global output multiplexer assembly contains six filters which are interconnected by a waveguide manifold to reduce losses. The spot beam multiplexers require only four filters each. Channelizing into odd and even sections simplifies the filtering in the output section by providing 40-MHz spacing between adjacent channels on each manifold. Each filter and the interconnecting manifold is constructed of Invar, and then silver plated. The filters are

Purchased from American Institute of Aeronautics and Astronautics

S. B. BENNETT AND I.

386

DOSTIS

f c + 18MHz

fc-18MHz

fc

f c +18MHz

Fig. 15 Input multiplexer performance.

six-section Tchebycheff filters which are tuned for approximately a 40-MHz equiripple bandwidth. This tuning is wider than the 38-MHz equiripple bandwidth used on the input filtering to produce only small in-band contributions to amplitude and delay variations. Response of a typical channel in the output section is presented in Figs. 18a and 18b. Figure 19 HLTWT

25MHz

OUTPUT POWER VS FREQUENCY

16 MHz

16MHz

Fig. 16a HLTWT performance.

25MHz

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DESIGN OF THE INTELSAT IV TRANSPONDER

387

HLTWT A 0 V S DRIVE TYPICAL RESULT SPEC

- ———J 2

Q

20

SAT-6 db

SAT

SAT-3 db RELATIVE POWER INPUT

0

HLTWT TYPICAL P0 VS P | N

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SAT. P 0 = 7.9 dbw

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POWER INPUT RELATIVE (db)

Figs. 16b and c HLTWT performance. shows the output multiplexer assemblies installed on the communications compartment.

Purchased from American Institute of Aeronautics and Astronautics

388

S. B. BENNETT AND I. DOSTIS

Fig. 17 HLTWT and EPC layout on despun compartment. OUTPUT FILTER INSERTION LOSS AND GAIN SLOPE VS FREQUENCY

--1.0 db

"_,

--1.2db

g co - - 1 . 4 d b

O CO

LLJ Q_ O

--1.6 db

0.03 db/MHz

CO

0.02 db/MHz 0.01 db/MHz

-18

-9 0 +9 FREQUENCY RELATIVE TO BAND CENTER (MHz)

+18

OUTPUT FILTER DELAY VS FREQ.

-18

-9 0 " +9 FREQUENCY RELATIVE TO BAND CENTER (MHz)

+18

Figs. 18a and b Output multiplexer performance.

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DESIGN OF THE INTELSAT IV TRANSPONDER

389

Fig. 19 Output multiplexer on despun compartment. SPOT e.i.r.p.

•36-

©---«y--*' SPECIFIEDMINIMUM

•35©- —-0 '34-33-

GLOBAL e.i.r.p.

•24-

.-0-—O--® SPECIFIED 1

2

3

4

1 I NIMUM

•22-

5

6 7 CHANNEL

8

9

10

11

12

Fig. 20 Spot and global beam e.i.r.p. ••70

O MEASURED X PREDICTED

J0--" X

X

•54

SPECIFIED 1

2

3

4

5

6 7 CHANNEL

MINIMUM 8

9

Fig. 21 Repeater G / T .

10

11

12

Purchased from American Institute of Aeronautics and Astronautics

390

S. B. BENNETT AND I. DOSTIS

80

©

©

0

0 ©

5

© 0 82-

6 7 CHANNEL

8

9

10

11

12

Fig. 22 Intelligible crosstalk performance. TYPICAL FREQUENCY RESPONSE VS, INPUT BACKOFF ABSOLUTE SCALE REMOVED -0 db

SAT

-0.0 db -0.1 db -0.2 db -0.3 db -0.4 db

-0.5 db J24.0db BACKOFF f - 14.4 MHz f + 14.4

J f c - 1 6 . 2 MHz

-0.6 db

f c + 16.2 MHz j

f,- 18MHz

fn + 18 MHz

Fig. 23 Complete transponder frequency response. ©RECEIVER "A" B RECEIVER "B"

©

El

©

©

.-16 ©

-17 © .-18 MINIMUM

SPECIFIED .-19

6 7 CHANNEL

8

9

10

11

12

Fig. 24 Typical illumination for saturation (0 db attenuator)

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DESIGN OF THE INTELSAT IV TRANSPONDER

391

System Performance

The major over-all performance parameters for the INTELSAT IV communications system include e.i.r.p., receive G/T, IXTR, frequency response, and system gain. Typical measured values for each of these parameters are compared to the applicable specification in Figs. 20-24. The results presented indicate that adequate margins exist for all specified parameters. The presented values for global e.i.r.p., receiver G/T, system gain and IXTR were measured in an anechoic chamber. The spot-beam e.i.r.p., presented is derived from a combination of unit power and far-field antenna measurements. The frequency response data presented were measured during the integration of the communications system and show the fre*• quency response for a large range of input drives.

A preliminary comparison of in-orbit values measured for global and spot-beam e.i.r.p., receive G/T, system gain, and frequency response to those presented indicated favorable agreement. References 1

COMSAT Technical Review, Vol. 2, No. 2, Fall, 1972 (entire issue).

2werth, A. M., "SPADE: APCM FDMA Demand Assignment System for Satellite Communications," presented at the INTELSAT/IEE International Conference on Digital Satellite Communications, London, England, Nov. 1969. 3Schmidt, W. G., et al., "MAT-1: INTELSAT1s Experimental 700Channel TDMA/DA System," presented at the INTELSAT/IEE International Conference on Digital Satellite Communications, London, England, Nov. 1969.

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A TIME DIVISION MULTIPLE-ACCESS SYSTEM FOR INTELSAT D. J. Withers* and A. K. Jefferis^ Telecommunications Development Department, British Post Office, London, United Kingdom Abstract The main advantages to be gained by using time division multiple access (TDMA) in the INTELSAT system, in place of frequency division multiple access, are higher traffic capacity per satellite, increased flexibility, better circuit quality, and more efficient integration of satellite facilities into a woi*ld communication network that is making increasing use of digital services and digital transmission techniques. This paper reviews the principal features likely in a TDMA system now being defined for consideration for INTELSAT. Among these are automatic acquisition of synchronization, a modular interface with the terrestrial network, common frame timing between transponders and "frequency hopping" of transmit and receive chains to reduce Earth station equipment requirements. The system is intended for use with INTELSAT IV satellites, but one prime objective will be flexibility to accommodate to subsequent types of satellite. TDMA might carry the major part of INTELSAT!s traffic in the 1980s. Presented as Paper 72-538 at the AIAA Uth Communications Satellite Systems Conference, Washington D.C.,April 2h-26,1972. This paper is based upon work performed under the sponsorship of the International Telecommunications Satellite Consortium (INTELSAT). Any views expressed are not necessarily those of INTELSAT. The system concepts presented arise largely from work done by the members of the ICSC/T Working Group on TDMA and in COMSAT Laboratories on behalf of INTELSAT. Acknowledgment is made to the Senior Director of Development of the Post Office for permission to publish this paper. Staff Engineer, Space Communications Systems Branch. Assistant Staff Engineer, Space Communications Systems Branch. 393

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394

WITHERS AND JEFFERIS

I. Introduction

Multiple access has been used in the INTELSAT system for an but its first satellite, "Early Bird"; that is, several or many Earth stations use a satellite simultaneously for their mutual traffic. Frequency division multiple access (FOMA) has been used so far, with each Earth station being assigned one or more frequency slots for its transmissions within the bandwidth amplified by the satellite transponders. Multichannel frequency division multiplex (FDM) basebands are used for telephony and similar services to frequency modulate the radiofrequency carriers which are transmitted in these frequency slots. Another system, called SPADE, using Phase Shift Keyed (PSK) carriers, each carrying a single digitized telephone channel, is about to be introduced in the Atlantic region; this system is particularly attractive for telephony on lowtraffic routes. However, a radically different multiple access technique, in which Earth stations take their turn in time sequence to transmit their traffic into a high capacity satellite highway, promises to be better than either FDM/FM/ FDMA or SPADE for most traffic. This new technique is called time division multiple access (TDMA).

In a typical TDMA system each Earth station is allocated one time slot per time frame (usually 125> jisec or a multiple thereof). During its time slot, an Earth station transmits a carrier burst, phase shift keyed in digital form in time division multiplex by signals from each of the channels it has to transmit. All the other Earth stations in the group receive this burst, and each extracts and demodulates the channels addressed to it. The process is then repeated, with each of the other Earth stations transmitting a burst in turn in its time slot until the frame has been completed. Then the first Earth station transmits a second burst, and so on. INTELSAT IV satellites are coming into service. Use of these satellites would give TDMA the following major advantages : (a) Given a TDMA system operating at the maximum symbol rate permitted by a usable transponder bandwidth of 36 MHz, only one carrier would be present in the satellite TWT at any instant. The TWT therefore could be operated at saturation output with maximum efficiency and without intermodulation. This must be compared with the noisy, backed-off condition arising with FDMA. A U-phase PSK TDMA system could transmit over 900 channels through a global beam transponder, compared with about 500 channels by FDM/FM/FDMA. Similarly for a

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TIME DIVISION MULTIPLE-ACCESS SYSTEM

395

spot beam transponder, a TDMA system using 8-phase PSK would provide perhaps 1UOO channels, compared with about 1000 for FDMA. Thus TDMA would provide more channels than FDMA for multiple access.

(b) The number of channels allocated to each Earth station in a TDMA system can be made exactly equal to requirements, so there need be no idle capacity. Furthermore, the allocation can be readily rearranged by reprogramming, if traffic requirements change. On the other hand, FDM/FM/FDMA would be unmanageably complicated if every carrier were equipped for the exact number of channels needed and bandwidth was apportioned accordingly. Consequently, all FDM basebands conform to one of a dozen or so standard capacities, and channels that are surplus to the requirements of the Earth station to which they are allocated cannot be used by any other Earth station. Furthermore, with FDM/FM/FDMA it is a major replanning exercise to redistribute capacity between the various stations from time to time when the traffic forecasts are revised. Thus, TDMA avoids a major management problem and permits the full theoretical extent of its enhanced channel capacity to be utilized, whereas FDMA locks 10# or 20% of its capacity away in basebands where it is not wanted. (c) Telephone circuit quality should be better when digital transmission is used. In particular, the noise and cross-talk levels will be lower. (d) Given suitable interface equipment, digital data traffic can be transmitted much more efficiently in a PCM system than in an FDM/analog system. Thus, for example, a l±8-kbit/sec data channel, with additional bits for protection against errors, can be exchanged for a single telephone channel in a PCM system, but displaces twelve telephone channels in an FDM system.

In addition to these benefits, realizable with INTELSAT IV satellites, TDMA offers two important long-term advantages: (e) The national long-distance telephone networks of the world are being converted from analog transmission to digital. It will be a slow process, and it has scarcely started yet, but significant progress will have been made in 10 years1 time, and digital operation may be general by 1990. At that stage, digital transmission over the satellite system will substantially reduce the cost of Earth station equipment.

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396

WITHERS AND JEFFERIS

(f) Future satellite systems for international telephony services probably will employ "frequency re-use." High gain antennas on the satellite will serve small areas of Earth's surface, each of which will have independent access to the whole frequency spectrum covered by the satellite. In this way traffic capacity can be made large and orbit utilization can be made efficient. These independent links between the satellite and the various service areas must be interconnected at the satellite, and one attractive way of doing this, using a time-switched interconnection matrix, would be feasible only if TDMA is used.

Along with these advantages, there are of course disadvantages, including the cost of re-equipping Earth stations and retraining staff. INTELSAT has set up a working group to develop an optimum specification for TDMA equipment, evaluate its use in the INTELSAT system, and plan the introduction of the system if it is found to be desirable. None of these three tasks has been completed yet, but much progress has been made in the first two. This paper includes an account, as seen by the authors, of current thinking in the group as to the probable configuration of a TDMA system, which might be given a field trial in prototype form in about 3 years1 time, and might be brought into service in the second half of the 1970s if the final appraisal is favorable.

Much work on the technical basis of TDMA already has been done. A low-speed system (MATE) was developed and demonstrated by COMSAT for INTELSAT as early as 1966, and at least 3 high-speed systems have since been developed to, the point of testing and demonstration over satellite links. ^ However, all these systems must be regarded as experimental, designed to explore solutions to such fundamental problems as modem design and burst formation, synchronization, and identification. In no case was effort concentrated on meeting the traffic and economic needs of a practical TDMA system working in a practical satellite system. Perhaps this point can be made more clearly by referring to some of the problems that the INTELSAT working group is having to study: (a) The TDMA system would probably go into service first with INTELSAT IV satellites. However, new satellites of greater traffic capacity will be needed soon, and yet another generation of satellites may come into service long before the TDMA equipment is 10 years old. These new satellites may be very different from INTELSAT IV. In particular, the transponder bandwidth probably will be greater, the effective power output may be greater - or less, and there may be on-board

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TIME DIVISION MULTIPLE-ACCESS SYSTEM

397

beam switching. The optimum TDMA system characteristics for these future satellites certainly will be different from those of INTELSAT IV. It is essential that the TDMA equipment should be adaptable, so that it can be reconfigured to work with future satellite systems at little cost.

(b) Future generations of satellites will use high-gain satellite antennas for most if not all of their capacity. This complicates system synchronization. (c) The system should be able, in principle, to carry efficiently any kind of communication service within its information capacity range. Most of the traffic will, no doubt, be voice channels for the international telephone network and here the problem is made more difficult by the failure, to date, of the CCITT to standardize the companding law and multiplexing characteristics to be used for international digital telephony. Facilities may be needed for telephony transmission economy techniques such as demand assignment or digital speech interpolation (DSI). However, the system also will have to carry services such as telegraphy, telex, data, and videophone. Finally, it must be capable of interfacing with both digital and analog terrestrial networks. II, System Description

Before proceeding with the description of the system it may be helpful to mention the salient features of the INTELSAT IV satellites with which the TDMA system will be working initially. INTELSAT IV has 12 transponders, each of 36-MHz usable bandwidth. The nominal transmitter power per transponder is 6w under single-carrier-saturated conditions. Four

of the transponders are connected permanently to global coverage antennas, while each of the other 8 can be switched to a global antenna or to one of the two U*5°-wide spot-beam antennas. The EIRP at the edge of the global beam is 22 dbw minimum and at the edge of the spot beam, 33*7 dbw minimum. The carrier-to-noise ratio in 36 MHz achieved at a standard INTELSAT Earth station (of G/T ratio UO? db:) when the satellite is at 5° elevation angle, that is, the worst case, is typically 19 db for a global beam transmission and about 30 db for a spot beam transmission. The basic TDMA parameters and objectives applicable to systems using an INTELSAT IV global transponder will be as follows:

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398

• One Frame - 7^0 pS • Burst from Station A

Station B

Station C

Station D

Station E

4 —————— -j—— —c———— — ——

Reference Burst (Preamble Only)

Traffic

I Bits » / Guard Tijne

...001100110... J1?

.

.I

Bits

Carrier Phase and Bit Timing Acquisition

20 Bits

.'

8

2 Bits

1

1

Bits 4

A

Unique Word (Non-Reference)

2 Bits A

I

1*8

Bits

I

/ Telephony / Voice / Signalling / Order Wire Station Telegraph Identification Order Wire

Fig. 1 Frame and burst format for prototype TDMA system.

to A and B

i

PGM Encoded 60 ch FDM

liOO samples (approx.) 9 Bits/Sample

to A

to B

60 ch Voice Encoded PCM 60ch PCM (Part 60/30 ch DSI Equipped) 1.. ....... ..30 1. .................... ..60 1..........2ii 30 Time Slots 6 Samples/Slot 8 Bits/Sample

60 Time Slots 6 Samples/Slot 8 Bits/Sample

2h Time Slots 6 Samples/Slot 8 Bits/Sample

Fig. 2 Possible assembly of traffic burst, from Station C.

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tion.

399

(a) MODULATION METHOD: U-phase PSK with coherent detec-

(b) BIT RATE: In the region of 60 to 65 Mbit/sea (precise value still to be determined). (c) FRAME PERIOD: ?£0 Jisec; e.g., one TDMA frame will contain 6 PCM samples per telephone channel at the standard 8-kHz channel sampling rate. (d) MARGIN (for degradations not dependent on bit rate): not less than 5>'5 d^ (e) MARGINS (for degradations dependent on bit rate, such as intersymbol interference and adjacent transponder interference): to be determined. (f) BIT ERROR RATE: Not to exceed 1 in 10 , measured over a period of 1 min, for more than 0*3% of the worst month of the year. o

(g) FRAME LOSS RATE; Not to exceed 1 in 10 when the bit error rate is 1 in 10 . Provision will be made in the system for possible conversion to 8-phase PSK for use in spot beam transponders of INTELSAT IV or subsequent satellites.

General Description of the System Figure 1 shows the format of one frame of the TDMA system. The reference burst and the traffic bursts from the various stations are arranged to arrive at the satellite in a prearranged sequence during the ?50-)isec frame period. A nominal guard time between bursts of 12 bits duration, say 0*18 psec, has been found to be adequate to prevent overlapping in the various tests that have been made with experimental TDMA systems. All Earth stations synchronize their burst timing relative to the reference burst, which has three parts. The first part is a U8-bit sequence to allow all Earth stations to acquire carrier phase and bit timing, the second part consists of a "unique word," and the short third part contains auxiliary signals which provide superframe timing for the whole system*-7 The unique word must be chosen with great care to ensure that a correlation detection process can identify the precise time of occurrence of the word with a high degree of

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WITHERS AND JEFFERIS

confidence even in the presence of some bit errors. Any suit*ably equipped Earth station can provide the reference burst* Use of a reference burst that does not carry traffic makes it a simple matter for a standby reference station to take over from the nominated reference station in the event of an Earth station failure, since other stations do not need to change their burst timing.

A traffic burst consists of a preamble followed by a stream of bits carrying traffic. The preamble contains a carrier phase and bit timing acquisition section, followed by a unique word, like the reference burst. However, these are followed by a transmitting station identification code and bits providing for a common channel telephony signaling system (if required) and voice and telegraph service channels for maintenance and operational communication between Earth stations. The same unique word is used for all traffic bursts, but a different word gives a positive identification for the reference burst. The traffic part of the burst would be assembled according to the destination of the traffic* The length of burst would be just sufficient to meet the traffic requirements of the station and normally would be changed only to meet traffic growth or other long-term variations. No short-term variation to match the "instantaneous" traffic demand is envisaged. Figure 2 shows how several modules of traffic capacity (described later) might be assembled to form the traffic part of the burst. Figure 3 is a simplified block diagram of a TDMA Earth station terminal. A brief outline of its method of operation will be given, followed by a more detailed description of certain aspects. The signals received from the terrestrial network, which may be in analog form (e.g., individual voice channels or FDM assemblies of voice channels) or in digital form (e.g., terrestrial PGM or data traffic), must first be converted into the digital form required for transmission over the TDMA system. This is carried out in the terrestrial interface modules, three different types of which are shown in Fig. 3The multiplexer unit calls the traffic bits out of the interface module compression buffers and assembles the traffic section of the burst at a time determined by the burst synchronization unit. These traffic signals are then scrambled by a pseudo-random code to avoid emitting a radio emission spectrum with strong spectral lines at particular frequencies, which

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401

TIME DIVISION MULTIPLE-ACCESS SYSTEM Terrestrial Interface Units

Buffers

SG Coder

Voice ^|

60/30 ch DSI

Voicev I

60 ch PCM

rii

——

Voice

!

60 ch PCM

it:

60 ch

---

PCM

--

~^" Order ->- Wire Unit

Signal Unit

)>Ut

Entry Modem

:. z i. ~ n ~ ~i^~- — *— __

1 l_ De-multiplexer

Voice

i

,

IF

1

/x

SG Decoder

60/30 ch DSI

se Modula tor

s —————————————— |

L. f— t

Voice i

"H Differentia Encoder

I-P

(Equipped for 2Uch)

60 ch FDM

—— —7 HPA Gating Control

X

) Voice

—

Preamble Generator

—— >--

Control Processor 1

60 ch FDM

r

I ,— _- —— —— —— —— —— —— —* Un-Cc nverter 1 ; Switc h Control

Station Identification Unit v /

1

V

Preamble Detector

-

Burst Synch . Unit

^

Auto Entry Unit

A i

— ^- -

Generator

—— ^—— Differential Decoder

—

h Phase Demod.

DownConverter Switch Control

JF In

— — — MAIN CONTROL LINES • SIGNAL PATHS

Fig. 3 Block diagram of TDMA terminal.

might cause harmful interference to other systems. The scrambled signals are added to the preamble and the complete sig-

nal is differentially encoded to allow resolution of the fourfold ambiguity in carrier phase that is inherent in the carrier extraction circuits of U-phase coherent demodulators. The combined bit stream is then fed to a U-phase PSK modulator. The modulator produces a burst output at an intermediate frequency, probably at UHF, and the IF signal is fed to the Earth station radio transmission equipment. In the receive side, a number of the functions are simply the reverse of the corresponding functions in the transmit side. The main difference lies in the need for unique word detection on which the burst synchronization and the demultiplexer timing is based. The functions of some of the more interesting items shown in Fig. 3 are covered by the detailed descriptions that follow.

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WITHERS AND JEFFERIS

Synchronization and Acquisition

Transmit burst synchronization. The clocks of the TDMA system are free-running and have an accuracy of the order of 1 part in 10 . With a perfectly stationary satellite the transmitted burst position therefore could drift relative to the reference burst at the rate of one symbol duration every 1$ sec. A considerably greater contribution to burst position error can arise from satellite orbit errors, especially inclination. For an orbit inclination angle of 1°, for example, and with the least favourable locations of the stations, there would be a drift in relative timing corresponding to one symbol duration every O7 sec. The method of correcting the drift is as follows. Every 300 msec each station compares the timing of its own unique word detection pulse with the reference unique word detection pulse, delayed by an amount corresponding to the required position of the burst within the frame. If the pulses are not coincident, the time difference will be measured and used to advance or retard the timing pulse that

determines the start of the transmit burst, at the rate of one symbol per frame, for the appropriate number of frames. In normal operation and with satellite orbits of readily attainable precision, the burst position error will not normally exceed 1 symbol during 300 msec. Measurement and correction more frequently than once every 300 msec could cause instability because of the round trip delay between the Earth station and the satellite. Initial acquisition and rapid re-acquisition. Initial acquisition will be by the transmission of a continuous signal, at least 20 db below normal level, probably consisting of a pseudo-random sequence. The signal will be generated in a separate modulator and detected in a correlation detector. The timing of the detected signal will be compared with the time of receipt of the reference unique word; when the correct timing between transmit and receive has been established, normal transmission can begin.

To prevent two or more stations attempting to gain access together, the reference station would define an acquisition timing cycle of about l min, and each station would be allotted a window within this cycle during which acquisition could be attempted.6 The entire acquisition process would be initiated automatically as soon as synchronization had been lost for more than say 30 sec.

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A rapid re-entry technique' is provided to minimize the

effect of brief outages on telephone calls in progress. After an interruption of less than, say, 30 sec, re-entry would take place on a basis of predicted burst timing. (Rapid reentry is not very important if the interruption exceeds 30 sec since most speakers will abandon their call if a longer

disconnection than this occurs, although it should be quite safe to use it even after 2 or 3 min of outage.) In the event of an outage the local clock, running from a protected power supply, would continue to operate and so allow the precise duration of the outage to be determined. The control processor will use this information, together with the rate of

change of burst phase immediately prior to the outage, to generate new burst timing pulses for re-entry. Before reentry actually is accomplished the processor will have to reload a number of registers in various parts of the control system that may have lost their contents during the outage.

Spot beam synchronization. The basic synchronization and acquisition methods described above rely on reception by each Earth station of its own transmitted burst after retransmission by the satellite. When spot beams are being used, this usually will not be possible and other means of synchronization are required. However, INTELSAT IV satellites always will have some global beam transponders, and so may future generations of satellites. It is proposed, therefore, that common frame timing shall be used for all the TDMA signals passing through any one satellite, controlled by a single

reference station. A reference burst would be transmitted

through one global beam transponder, and all Earth stations would use this burst and this, or another, transponder to acquire system burst timing, which would then be valid for all their bursts, whether transmitted by that transponder or any other. It is still considered that some form of confirmation will be required that a burst is in its correct position, and this can be provided by means of a feedback channel from a station that is receiving the burst.

Detection of Preambles in Received Bursts The PSK demodulator will be designed to acquire the phase of the carrier during the first part of the preamble, and to maintain it throughout the burst, with sufficient accuracy to give a low error rate at the operating carrier-to-noise ratios. Similarly the demodulator will rapidly acquire the clock phase and retain it through the burst. Consequently correct detection of incoming bits should be taking place before the start of the unique word. The time of arrival of the unique word

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WITHERS AND JEFFERIS

can be predicted to an accuracy of about - 1 symbol duration from one frame to the next (a movement of 1 symbol per frame being the maximum rate at which stations correct their burst phase). Consequently there is no point in running the risk of false detections by having the unique word detector operative all the time. Instead, an aperture generator allows the generation of unique word detection pulses .only at times within - 3 symbols of the predicted time. The probability of missing^or falsely detecting the unique word, must be less than 1 in 10 when the bit error rate is 1 in 10 . This can be achieved with the 20-bit word chosen and a 7-symbol aperture. Two correlation detectors are required, one for the reference burst and one for nonreference bursts. The aperture generated for the reference unique word is only 5 symbols wide to reduce further the probability of false detection. Five false or misdetections of the reference unique word are necessary before the reacquisition process is initiated. During acquisition the aperture generated for the reference unique word detector is fully opened to enable it to be found quickly. The aperture generator also provides timing pulses for the operation of down-converter switches if frequency hopping is being used (see "Multiple Transponder Operation" below).

The remainder of the preamble is routed to the station identification unit which informs the demultiplexer of the origin of the burst, and to the control signal unit which extracts the order wire (i.e., service channel) circuits and routes them to the order wire unit.

Terrestrial Interface Modules Some earlier TDMA systems have employed direct high-speed PGM encoding at the TDMA system bit rate. However, the INTELSAT prototype system will be designed on a modular basis employing, in general, low-speed PCM encoding and decoding. The chosen arrangement will allow a wide range of interface modules, both for PCM encoding at the Earth station and also for direct digital interfaces when the terrestrial links to Earth stations are also digital. Special modules (for example, for data traffic or demand assignment of telephony or videophone) also could be introduced when a need for them is identified. For the prototype trials, the following interface units are likely to be built: (a) A PCM codec for up to 60 telephone channels, expandable 1 channel at a time using 8-bit, A-law encoding. This

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TIME DIVISION MULTIPLE-ACCESS SYSTEM

405

will be the normal equipment for voice channel PGM encoding and decoding using an analog/digital interface at the Earth station. (b) A digital speech interpolation unit. This also may have expandable capacity but probably with a minimum of 60 input channels, interpolated to 30 time slots. (c) A supergroup codec for 60 telephone channels in FDM. This interface will be attractive for the interim period while most national telephone networks use analog transmission. Encoding standards have not been determined yet. The 11-bitsper-sample contemplated for terrestrial systems provides a standard of performance which is unnecessarily high for systems that can span intercontinental distances without recoding; 9-bits-per-sample, perhaps even less, should be quite adequate and should raise transmission efficiency significantly. (d) A 2-OU8-Mbit/sec .digital interface unit. This unit will be for the particular case of a terrestrial digital link that can be frame synchronized with the satellite TMA. system. Direct digital interface units for separately synchronized systems will be the subject of subsequent development work. Each interface module will contain a compression buffer, in which the low-speed continuous signals are converted to high-speed "sub-bursts" at the bit rate required by the TMA system, and an expansion buffer to perform the reverse operation. The output of each interface module will be two parallel channels P and Q, each operating at half the TDMA bit rate, i.e., about 30 Mbit/sec. The P and Q channels will be carried in parallel right through to the modem, which, being a li-phase modem, can conveniently use parallel binary inputs and outputs. This technique reduces the operating speed requirements of practically all the high speed equipment in the TDMA terminal. Multiplexers and Demultiplexers One of the important potential advantages of TDMA compared with present multiple-access methods is in flexibility to meet changing patterns of traffic. The main source of this flexibility in the prototype system will reside in the multiplexer and demultiplexer. All the timing functions of these units are controlled by a nonvolatile memory, the contents of which can be modified when required by the control processor. The multiplexer unit controls the formation of the transmit

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406

WITHERS AND JEFFERIS

burst by generating pulses, at the appropriate time interval^, after receipt of the start of burst pulse, which causes the preamble unit, the control signal unit, and then each of the terrestrial interface modules in turn to read out the contents of their buffers to the modulator. The demultiplexer unit performs the equivalent receive function* This involves selecting from each received signal only those parts of the burst containing traffic for the station in question and directing these to the appropriate interface module expansion buffers.

Multiple Transponder Operation INTELSAT IV has 12 transponders. If only one transponder is used for TBMA, clearly every participating station needs one set of terminal equipment. However, this would provide only about 900 telephone channels. If two transponders carried TDMA, some Earth stations might need to be equipped for both TDMA systems. Going one step further, if, let us say, 10 transponders were carrying TDMA traffic, then an Earth station wishing to communicate with many other stations could find itself having to receive from most of the 10 transponders. Alternatively, the traffic might be destinationoriented, and the station would be asked to transmit to many transponders. This would be uneconomic. The numbers of transmit and receive chains at Earth stations could be considerably reduced if a single transmit' chain or receive chain could be switched from one transponder to another during the TDMA frame. A possible method of implementing this "frequency hopping" technique is shown in Fig. h for the transmit side, and a similar arrangement would be possible for the receive side. There are two basic requirements for multiple-transponder operation by means of frequency hopping; first, the frame synchronization must be common for all the transponders involved, and second, the allocation of traffic to bursts must be such that there is least overlap between the bursts transmitted by one Earth station and least overlap between the bursts received by it. Frame synchronization is also desirable for another reason - namely, to simplify operation in spot beam transponders. The least-overlap constraint may reduce the efficiency of use of the TDMA system, but this inefficiency can be made acceptably small if traffic is transmitted not necessarily in a single burst, but in two or three, if this helps to eliminate overlap.

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407

Fig. 4 Transmit chain switching.

A study of the effect of this frequency hopping technique showed that, for a traffic matrix with 2£ stations, it could reduce the number which required more than one transmit chain from 10 to 1. Similarly, the average number of received chains per station was reduced from U*5 to 1 *3« The efficiency of use of the TDMA capacity in this case was reduced by about 1% because of the need to avoid overlap. The only modifications required to the TDMA equipment itself will be the provision for transmitting multiple bursts per frame and the generation of up and down-converter switching pulses. It would be necessary, however, to provide special facilities in the radio equipment of the Earth station.

Operation and Maintenance

Some of the functions of the control processor already have been explained. The use of a processor permits changes in the operational functions of the TDMA terminal to be achieved by means of modification of the software. This is particularly useful when changes are made to the numbers of channels on different routes. A simple change to the program can alter the length of the transmit burst and the locations of the channels selected from received bursts. Even routes to new stations can be set up with no more hardware provision than the addition or expansion of the appropriate terrestrial interface module. The control processor also will make regular checks on the operation of the various units and will switch over to a standby unit in the event of a fault. A separate off line

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WITHERS AND JEFFERIS

maintenance computer will be provided for the diagnosis of faults in any unit.

Consideration of Successor Satellites to INTELSAT IV

Although it is not yet decided what form the successor satellites to INTELSAT IV will take, it is highly likely that they will make greater use of spot beams, and it is possible that the bandwidth of their transponders will differ from those of INTELSAT IV. It is also possible that they eventually will use on-board time division switching between beams, in which case transponder bandwidths much greater than those of INTELSAT IV would be used. The prototype system will be designed so that operation with subsequent satellites should require little modification. Operation to wider bandwidth transponders would mean increasing the bit rate. The TDMA equipment specification will call for logic capable of operating at higher bit rates than initial service demands, and so an increase in bit rate would be implemented by increasing the clock rate up to the limit at which the basic logic can operate - say around 60 Mbit/sec for the prototype equipment, permitting U-phase PSK at 120 Mbit/ sec, and perhaps more for subsequent production. If still higher speeds are required, they could be achieved either by providing additional TDMA transmit and receive chains with secondary multiplexing and demultiplexing at the PSK modem input and output or, alternatively, by providing a high-speed buffer to compress the bursts into a shorter period at higher bit rates. A new PSK modulator and demodulator probably would have to be provided in any case.

As already explained, the synchronization method proposed will be suitable for use with spot beams. However, an essential feature of the method is that there is at least one global beam TDMA system for which timing information can be obtained. Hopefully the successor satellites will have some global capacity available. If on-board satellite time division switching between beams were used in a successor satellite, it is less likely that any significant amount of global beam capacity would be available, but on the other hand the existence of an on-board switch facility would permit a different and basically simpler solution to the synchronizing problem to be used. Thus, it 'could be arranged that each beam received in the satellite could be switched to a down-beam serving the same geographical area for a period in each switching cycle. Each Earth station could use this period to lock itself to a reference signal generated in the satellite

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TIME DIVISION MULTIPLE-ACCESS SYSTEM

409

and locked to the switching cycle. On-board switching also would require each Earth station to transmit several bursts, at least one for each down-beam of interest, but this multiple burst facility is to be provided in the prototype system for other reasons as indicated above (i.e., to enable a good compromise to be reached between the cost of terminal equipment and efficiency of utilization of the system). In these ways, it is thought that the TDMA equipment can be made to accommodate to future satellites at little cost. Ill. Conclusion A TDMA system could be designed, along the lines described which would be technically and operationally suitable for INTELSAT and should be flexible enough to be used with few and cheap modifications for many years to come. It has yet to be demonstrated that a changeover to such a system would be costeffective in the near future, but this is being studied actively. References Puente, J. G., Schmidt, W. G., and Werth, A. M., "MultipleAccess Techniques for Commercial Satellites," Proceedings of the IEEE, Vol. 59, February 1971, pp. 218-229. 2 Schmidt, W. G., Gabbard, 0. G., Cacciamani, E. R., Maillet, W. G., and Wu, W. W., "MAT-1, INTELSATfs Experimental 700 Channel TDMA/DA System," INTELSAT/IEE International Conference on Digital Satellite Communications, London, November 1969, pp. U28-liUO. 3

Nosaka, K., "TTT System, 50 MBPS PCM-TDMA System with TimePreassignment and TASI Features," INTELSAT/IEE International Conference on Digital Satellite Communications, London, November 1969, pp. 83-9U.

Eckhardt, G., Reidel, B., and Rupp, H., "A TDMA System Proposed for the Experimental German-French Communications Satellite fSymphonic1," INTELSAT/IEE International Conference on Digital Satellite Communications, London, November 1969, pp. 331-3U2. Schmidt, W. G., et al., "A TDMA Satellite Communications System Having Special Reference Bursts," U.S.A. Patent Application, August 11, 1971.

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WITHERS AND JEFFERIS

°Schmidt, W. G;., et al., "A TDMA Satellite Communications System with an Aperture Window for Acquisition," U«S«A» Patent

Application, August 11, 19?1.

7 Schmidt, W. G,, et al., "A TDMA Satellite Communications System with Rapid Automatic Re-entry," U.S.A. Patent Application, August 11, 1971.

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SYNCHRONIZATION OF EARTH STATIONS TO SATELLITE-SWITCHED SEQUENCES R. A. Rapuano" COMSAT Laboratories, Clarksburg, Md. N. Shimasaki Nippon Electric Co., Ltd., Kawasaki, Japan Abstract A multiple-access and multiple-spot beam satellite with cyclically switched beam interconnections controlled by a stable onboard clock permits spectrum reuse with greatly increased communications capacity. All Earth stations would transmit a sync burst at the beginning of a frame, which would be used to determine the time of occurrence of the sync window. The sync window is a connection in the switching sequence of the switch matrix, which returns any signal from a particular spot beam zone to the same zone. Synchronization is achieved by centering the sync burst in the sync window. Three options, one of which uses a PSK sync burst, the other two using FSK sync bursts, have been analyzed. Their ability to establish synchronization and their technical features are compared for use in a TDM system with 300-MHz bandwidth. The choice of a digital-controlled synchronizer which uses the FSK sync word is justified. A synchronized timer was constructed which used an FSK sync burst and a digital control Presented as Paper 72-545 at the AIAA 4th Communications Satellite Systems Conference, Washington, D.C., April 24-26, 1972. Additional work done on the unit since then is given in the last section. The authors wish to indicate their appreciation to J. G. Puente for his steadfast encouragement and valuable suggestions. Particular thanks are due W. G. Schmidt, 0. G. Gabbard, and T. R. Dobyns for their advice on the details of implementing the concepts in high-speed logic, and G. Davidson Collins for converting a paper design to a working model. ^Presently, Consulting Scientist, Data Acquisition Directorate, Raytheon Company, Wayland, Mass. ^Engineering Manager, R&D Project Office .

411

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412 loop.

R. A. RAPUANO AND N. SHIMASAKI Timing errors of about +15 nanosec were obtained under

laboratory conditions. has been observed.

Some departure from ideal performance

Introduction In a previous paper N. Shimasaki described the basic concept of the onboard distribution center of a multiple access satellite which would be used with many spot-beam antennas. To permit Earth stations in one beam to communicate with stations in another beam, the beams are cyclically interconnected in a rapid sequence by a satellite borne distribution center controlled by an onboard timing device of high stability. Called SDMA/SS-TDMA, this concept requires that each Earth station synchronize its transmissions to the satellite switching sequence, and thus use TDM/TDMA for communicating with other stations in the same spot beam or with other stations in any other beam. Synchronization would be accomplished by using a sync burst that is looped back to the transmitting station by a switch connection for an interval called the sync window.

During this short interval, once per frame, each station can receive its own transmissions. Thus, if a sync burst signal radiated by each station is timed to arrive at the satellite during this interval, it will be detected. Proper placement of this sync burst (or sync word) with respect to the sync window will establish the correct timing for each subsequent set of data bursts so that they will enter the proper data window to be routed to the appropriate receiving station. Synchronization

Before the system can be synchronized accurately, it is necessary for the sync burst to be returned to the station that sent it. Since each station can receive its own transmission during the sync window period, it is necessary first to institute a search for this interval by sending out the sync burst at various time phases of the frame period until part of the word is returned. Then error signals can be generated to shift the timing in finer steps until the sync burst is properly positioned in the sync window.

The relationships between the sync window and uplink burst transmission are shown in Fig. 1. For this study, a

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413

SYNCHRONIZATION OF EARTH STATIONS

DATA BURSTS

DATA BURSTS UP-LINK

BURST TRAIN OF SPOT BEAM =1

FROM FROM FROM FROM FROM = 11 E/S = 12 E/S -13 E/S = 11 E/S =12 E/S

FROM FROM FROM -11 E/S -12 E/S = 13 E/S

FROM REF. E/S TO -1 ZONE

TO OPERATION OF SATELLITE SWITCHING MATRIX

DATA BURSTS

SYNC BURST

TO -2 ZONE

TO -1 ZONE

SYNC WINDOW

TO =3 ZONE

DATA WINDOW — 2

DATA WINDOW -1

FROM =13 E/S

DATA WINDOW =3

(UTILIZED TO LOOP BACK DATA BURSTS TO THE SAME ZONE)

DATA BURSTS

DATA BURSTS

DOWN-LINK

BURST TRAIN OF SPOT BEAM =1

DATA BURSTS

SYNC BURST FROM

FROM

REF. E/S

-11

FROM

FROM

E/S =12 E/S =13 E/S

FROM

-31 E/S

FROM

FROM

=32 E/S =33 E/S

FROM

TO ±?1 ZONE

TO -1 ZONE

TO -1 ZONE

FROM

1

f

OFF———/

SYNC

\

ii/viNnnw /

DATA WINDO

DATA WINDOW -rr

(NOTE) SYNC BURST IS

ASSUMED AS

ALLOWABLE LAG LIMIT PHASE_ __

- = (m-1) THE LEADING

THE LAST

DATA BURST OF

DATA BURST OF

'

#1 DATA WINDOW

~1DATA WINDOW

DATA WINDOWS

THE DATA BURST OF #m DATA WINDOW

b) Fig. 1

FROM

=21 E/S =22 E/S ^23 E/S

Relationships between sync window and uplink bursts.

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R. A. RAPUANO AND N. SHIMASAKI

frame rate of 8000/sec (frame period Tp = 125 //sec) and a sync window duration of 970 nanosec was chosen, but the techniques described will work as well for a wide range of frame rates and sync window durations. For a particular sync burst to pass through the sync window and be received back, it must be transmitted earlier than the time of occurrence of the sync window by the one-way trip time To to the satellite (about 135,000 //sec). However, since the frame length is Tp = 125 //sec, proper synchronization will occur whenever the sync burst is transmitted at a time T, T = T

- NTF

(1)

(where N = an integer) earlier than the occurrence of the sync window. Therefore, to find the sync window, the time of occurrence of the sync burst need only be varied over an interval of one frame. This technique does not require precise satellite range knowledge; however, it does require that the ground clock and the space clock have the same period to a high precision and that both have very good short-term stabilities. The two clocks should be within about 10"" of each other in frequency, so that most of the signal still will pass through the sync window after coarse search has been terminated. The short-term stability for a high data rate system should be about 10" . The use of the sync window only to synchronize each

Earth station to the autonomous satellite distribution center eliminates the concept of a reference Earth station. The satellite becomes the reference station for all users and makes possible both frame synchronization among spot-beam zones for SDMA and frame synchronization among stations in a spot-beam zone for TDMA.

Sync Window Search All of the options which have been considered use essentially the same techniques for searching for the sync window. Figure 2 shows a circuit, similar to the one in the spacecraft, which generates the spacecraft switching signals, but is modified to permit search. This circuit can a) generate a sync burst at the frame rate, and b) slip its time of occurrence in increments about equal to the width of the sync burst to provide for the search function. In this circuit in the nonsearch mode, a 1.032-MHz stable source is divided by 128 with synchronous dividers. The output of the divider then sets a flip-flop through logic circuits, which inhibits one

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SYNCHRONIZATION OF EARTH STATIONS

clock pulse from entering the divider chain, and then resets the circuit. Therefore, 129 clock pulses are counted, with the last pulse corresponding to the sync burst.

To permit this circuit to perform the search function, a modification of the 129-to-2 divider chain is made, which converts it to a programmable divider. This is set to count 130 pulses for a single frame and then revert to the normal value. This simple modification uses another counter to count the number of sync burst signals.

STABLE CLOCK FROM DIVIDERS OR VCO 1.032 MHz

415 BURST GATES

TO SYNC BURST GATE GENERATORS SYNC BURST "PRESENT SYNC BURST PET.

BAND PASS FILTER

TO ERROR MEASURING CIRCUIT

Fig. 2 Block diagram of frame generator with search function.

The exact number of pulses to be counted is not critical, but the total time interval should exceed the round trip to the satellite. In the absence of a search inhibit signal from the sync burst detectors, the counter is arranged to count to 2560 and then reset itself. (The value of 2560 is derived from eight binary stages and a decade counter.) This means that, at a particular phase of the spacecraft frame time, the sync burst is transmitted for about 0.312 sec, which is appreciably longer than the round-trip time to the satellite (0.267 sec). If, by the end of this period, no output has been obtained from the sync burst detector, then the 2560:1 counter activates a synchronized flip-flop which prevents one clock pulse from entering the frame divider chain so that the sync burst pulse is slipped by about 1 //sec. This process is repeated until an output is obtained at the sync burst detectors, i.e., until the sync burst has slipped enough so that it enters the sync window and is returned to the ground station. The output of the sync burst detector then exceeds a preset threshold and acts as a search inhibit signal, which disables the flip-flop or one-shot. The frame divider chain remains in the 129-to-1 mode, and the search mode is terminated.

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416

R. A. RAPUANO AND N. SHIMASAKI

Fine Phase Control

After the search mode has ended and the sync burst, now trimmed by the sync window, has been received, it is necessary for final synchronization to center the sync word in the sync window. This process, fine phase control, is done by operating on the received signal to a) obtain an error signal from the sync burst, which has been modulated by the pulse time modulation (PTM) function of the sync window; and b) use this error signal to center the sync burst in the sync window and to hold it there in the presence of drift and jitter caused by spacecraft motion, relative clock rates, noise, and logic circuit behavior. Two general methods for achieving fine phase control are possible; one uses a high-stability primary clock, digital dividers, and logic circuits to make the phase corrections in discrete steps. The other uses an analog voltage-controlled clock and, in principle, can make corrections to any desired degree of fineness. The position of the sync burst word with respect to an ideal, i.e., rectangular, sync window can be measured accurately. For example, if a PSK sync burst with a bit rate of 100 megabauds were used, then a single bit error corresponding to 10 nanosec could be detected. In the case of an FSK sync burst, the accuracy of a single measurement depends upon the

bandwidth of the FSK detectors.

.The timing error is given as the rms sum of four timing errors, i.e., the leading edge, trailing edge, and the two transitions between the two words. The timing error of a single sample of the sync burst was calculated for two filter bandwidths, 1 MHz and 10 MHz and a carrier-to-noise ratio of +38 db in a 1-MHz band.

6r = 2/B v/S7N (Refs. 2 and 3) where B = 1 MHz,