General Disclaimer

One or more of the Following Statements may affect this Document

This document has been reproduced from the best copy furnished by the 

organizational source. It is being released in the interest of making available as 

much information as possible.

This document may contain data, which exceeds the sheet parameters. It was 

furnished in this condition by the organizational source and is the best copy 

available.

This document may contain tone-on-tone or color graphs, charts and/or pictures, 

which have been reproduced in black and white.

This document is paginated as submitted by the original source.

Portions of this document are not fully legible due to the historical nature of some 

of the material. However, it is the best reproduction available from the original 

submission.

Produced by the NASA Center for Aerospace Information (CASI)
n

A STUDY OF A SPACE-STATION-ASSOCIATED

LTIPLE SPACECRAFT MICHELSON

SPATIAL INTERFEROMETER

NASA Contract NASW-3756

rPPGC rah

Final Report

For the Period 24 March 1983 through 30 September 1983

Principal Investigator

Robert V. Stachnik

September 1983

" Prepared for

National Aeronautics and Space Administration

Washington, D.C. 20546

Smithsonian Institution

Astrophysical Observatory

Cambridge, Massachusetts 02138

(h ASA —CH —1754joj A SIU- _. t+ . -'i - - - L -

SPACE — STATION — ASSUCIAIED :,ULTIPLE SI:ACECEAFT

MICHELSON SPATIAL INTERFERCMETER Final

Report, 24 mar. - 30 Sep. 1983 (Smithsonian Unclas

Astro ph y sical I ,• , 1;1 r

C, G3/1 ^ ?c;u$

-rho Smithsonian Astrophysical Observatory

is a member of the

Ha rvard-Smithsonian Center for Astrophysics_

The NASA Technical Officer for this contract is Dr. Kenneth J, Frost, NASA-Goddard

Space Flight Center, Greenbelt, MD 20771.

e

r ,
A STUDY OF A SPACE-STATION-ASSOCIATED

MULTIPLE SPACECRAFT MICHELSON

SPATIAL INTERFEROMETER

NASA Contract NASW-3766

Final Report

For the Period 24 March 1983 through 30 September 1983

Principal Investigator

Robert V. Stachnik

September 1983

Prepared for

National Aeronautics and Space Administration

Washington, D.C. 20548

Smithsonian Institution

Astrophysical Observatory

Cambridge, Massachusetts 02138

The Smithsonian Astrophysical Observatory

Is a member of the

Harvard-Smithsonian Center for Astrophysics

The NASA Technical Officer for this contract is Dr. Kenneth J. Frost, NASA-Goddard

Space Flight Center, Greenbelt, MD 20771.

^^ 
TABLE OF CONTENTS

Abstract Ii

i. introduction 1

' I1. Optical Design 4

III. Operational Characteristics and Orbital Motion 9

IV. Position Sensing, Thruster Requirements 11

and Vibration Control

V. Phase Measurement and Image Reconstruction 15

VI. Signal-to-Nolse Estimation for Sample Sources 17

Vii. Astrophysical Applications 21

Vi l 1. Micro-SAMSI 24

IX. Future Studies and Laboratory Experiments 25

Acknowledgements 26

References 27

Figure Captions 29

-1-
ABSTRACT

One approach to Michelson spatial Interferometry at optical wavelengths Inwolves use of an array of spacecraft in which two widely-separated telescopes collect light from a star and direct It to a third, centrally-located, device which combines the beams in order to detect and measure Interference fringes. Earlier atudies by us have suggested that such a system (SAMSI - Spacecraft Array for

Michelson Spatial Interferometry) could be a powerful Instrument for collecting extremely high resolution (10'5

 aresecond) spatial structure Information, In the form of

power spectra, for astronomical sources. High fuel consumption, however, appeared

to restrict the observations to measurements along one-dimensional cuts through

the source. Furthermore, recovery of Information required for the reconstruction 1

actual Images at this extraordinary resolution appeared impossible.

This report describes a modified version of SAMSI which makes use of the fuel

resupply capability of a Space Station. The combination of this fuel resupply capab?lity with our recent discovery of a method of obtaining Image Fourier transform

phase information, necessary for full Image reconstruction, permits SAMSI to be

used to synthesize Images equivalent to those produced by huge apertures In

space. We here discuss synthesis of apertures In the 100 to 500 meter range.

Reconstruction can be performed to a visual magnitude of at least 8 for a 100A

passband In 9 hours. Data are simultaneously collected for Image generation from

0.1 micron to 18 microns. In the one-dimensional mode, measurements can be made

every 90 minutes (including acquisition and repointing time) for objects as faint as

19th magnitude In the visible.

i

-ii-
I. INTRODUCTION

High optical resolution for astrophysical sources can be achieved through the

use of either extremely large aperture telescopes or long baseline spatial interferometers.

In Michelson Interferometry, light from a star, or other source, is collected by

two widely-spaced telescopes and reflected to an interferoirster located n' nn

equi-optical-pathlength position. Contrast of the interference fringes as a function

of separation of the light collectors Is related In a simple way to the angular size

and light distribution of the source. Figure 1 is a schematic diagram of a Michelson

Interferometer which Includes a dispersive element allowing simultaneous multicolor

observations. One possible approach to interferometry in space requires very large

structures whose purpose la to establish the basic optical geometry and provide a

platform for the principal optical components.

An alternative to a monolithic space interferometer is a device consisting of

separate spacecraft. in such a system, two or more telescopes, aboard indept ndent spacecraft, would simultaneously observe the same source and transmit the

light they collect to a third spacecraft housing an on-board interferometer. Measurement of fringe contrast as a function of separation of the light-collecting spacecraft gives the angular size of the source. Placing the principal optical components

aboard separate spacecraft, rather than connecting them by means of a very large

structure, permits flexibility in selecting extremely long baselines, in this case, as

much as 10 kilometers. It also implies a lower mass for the entire system and

guarantees the complete vibrational isolation crucial to a large-scale Interferometry

experiment.

The key technical problems are associated with precision ranging, pointing, and

positioning of the spacecraft, since It Is necessary to equalize the optical paths

through the two light collecting telescopes to within microns. Many of the required

-1-
J

C^Z

l-O

CnHQ

O

cr

= UW

O 

_

J

J JUO

U)W

^ 

as

J w

C•7

Eno

aNW

U GUa0

U 

m .

_zw

cr

I`

WQ

Wz

cr

U) UUw

Fig. 1
.2-

sensing capabilities are already highly developed (Stachnik et a/. 1983). Fine

spacecraft angular position determination would be performed using an existing Innovative photon-counting, star/spacecraft tracker (Papailollos and Mertz 1982).

Ranging would be done uein;l either laser Interferometey or Iidar. Pathlength squallzatlon would be done with a combination of "layered" thrusters capable of precision

force application over a wide force range and a moving-mirror optical delay line In

one of the two Interferometer arms. Piezo-microstraln sensor/actuator pairs would

provide active vibration sensing and control, and a very effective method of postphoton-detection range drift compensation exists for correction of residual errors.

In operation, the three spacecraft would deploy from the space station then

move, under their own power, a suitable distance from the Immediate vicinity of other apace station-associated experiments. The two light collectors, each having an

aperture of perhaps a meter, would point to the source to be studied and the

Interferometer-bearing central station would translate to the computed equl-opticalpathlength position. Small active mirrors in each light-coliecting satellite would then

direct narrow pencils of starlight to the central station. Aboard the central station,

back-to-back telescopes would relay the beams to an on-board Interferometer. Exfisting sensing technology allows positioning of the central station to within hundreds of microns of the true equi•path-position. At this point, the moving-mirror delay line in one optical path steps through the range of possible correct positions

until fringes are sensed. Once fringes are detected, the optical delay line continues to track them as the separaton between light collectors Is changed. The variation of fringe contrast with separation (the visibility curve) Is the basic Information

we require to obtain one-dimensional information In the form of a spatial power

spectrum, Information on the light distribution of the source. The angular resolution

Is set by the maximum separation of the light collectors. If this distance Is 10 km,

the resolution is 1 6" r3 arcascond.
-3-

Full Image reconstruction also appears possible. in this mode, the spacecraft

would be moved In a spiral pattern, as depicted In Figure 2, to permit sampling the

entire range of spacial frequencies which would otherwise be accessible only to a

single monolithic light collecting device of extraordinary size. We have, under this

contract, found a way of using simultaneous observations over a very wide

wavelength range, to estimate the image Fourier transform phase as well as ampi;-

tude (the latter being determined from the fringe contrast). These data are all

that are required for Image reconstructlon. Available array sensors used in detecting the fringe patterns would allow recovery of high-resolution Images over a

wavelength Interval extending from 0.1 to 18 micro^is, In narrow spectral

passbands. Because full Image reconstruction is time-consuming (9 hours, for complete sampling over a 100-meter -spiral) and most appropriate to brighter objects

(8th magnitude), fainter sources (19th magnitude) would be observed by making

one dimensional fringe contrast measurements, to obtain angular diameters. This

dual mode of operation is an Important characteristic of the system.

The spacecraft would be under continuous active position control and the

overall observing efficiency of the interferometer would depend on positioning the

spacecraft as rapidly as is consistent with the limitations of the thrusters and

sensing systems. At present, it appears that the combination of thrust and efflciency requirements in the propulsion system mandate principal reliance on liquid

hydrogen/liquid oxygen thursters, for which fuel use, because of the 100% duty

cycles, would be comparatively high despite their high specific Impulse. It is in this

connection that the association with the space station becomes vital. The space

station would serve as a refueling facility for the otherwise autonomous system of

spacecraft. At present, a European group (Laboyrie et ah 1982; Stachnik and Labeyris 1984) is exploring a very similar system. Wtihout the availability of a refueling facility, however, they are forced to rely on a solar sail design for which the

forces and, consequently, observing efficiency, may be nearly folir orders of magnir

Cl. ^<a^

Q C7 ap

^o

zr_a0J

•

-

r 
.4-

tude smaller.

Continuous, rapid motion of the spacecraft, made possible by periodic refueling,

would render a multiple apacecrait Interferometer an extremely powerful and versathe scientific Instrument. It would be capable of obtaining either large numbers of

one-dimensional measurements or a lesser number of two-dimensional Images. The

resolution would be a function of observing requirements since Imaging baselines of

hundreds of meters or ona-dlmenskmal baselines of tens of kikuneters will not always be needed. For special objects, unusually large or small baselines may be

employed.

The resolution achievable with the 100 m Imaging baseline we en.nhasize here

is at least 10-3 ercascond in the visible. This Is three orders of magnitude greater

than is typicrAl of routine earth-based measurements and a factor of 40 greater

than is expected from Space Telescope. In the alternate, one-dimensional mode,

10-km baselines are possible and, due to the existance of advantageous orbital

configurations, the corresponding 16'5 aresecond resolution observations can be

made at a rate of 18 sources per day.

In what follows, we describe a design based on an earlier low-thrust, nonrefuelable system for which actual Imaging did not appear possible. This earlier device Is

described in Stachnik at a/. (1983). Parts of Sections II and VII of this report are

largely taken from that source.

11. OPTICAL DESIGN

In order for the Interferometer In the central station to detect and measure

fringes, the central station must be near enough to un equhoptical-path position

that a :!-lay line in one of the optical paths can make both paths equal to within a

few wavelengths. An optical delay line is a device which (in the simplest case)

use* a moving mirror to adjust optical path length. In general, the focus of equl-eoptical-path positions is a surface defined by the condition that the difference

between the distance to light collector number 1 and the distance to light collector

number 2 is equal at every point. This surface is a hyperbala of revolution about

the axis Joining the two light collectors.

rigure 3 is a schematic diagram of the central station. It consists of two

back-to-back telescop es acting as receivers for starlight redirected by the collectIng telescopes. The back-to-back central station telescopes are also used co

transmit and receive laser light, which bounces off retroreflectors on the light collecting telescopes, for use in position sensing. The data and position sensing

beams pass through the same optical systems and may be separated either spectrally or temporally.

Initial acquisition of the fringes Involves coarse positioning of the central station with a subsequent fringe search over the region of uncertainty. A combination

of distance measurement between spacecraft and angle measurement would allow

positioning of the central station to within much less than t 0.6 mm. In the following paragraphs, we discuss factors bearing primarily on initial fringe acquisition.

In order to achieve t 0.6mm pathlength equality, for Instance, over a kilometer

0.1 aresecond tracking accuracy is required, as Is the ability to measure distances

to within fractions of a millimeter. Lidar distance measurement to better than 3

0.6mm has been demonstrated (Quersola 1979) and the existence of an Integral

dispersive element in the data collection system largely eliminates the data

besm/locator beam cross-talk problem. Although It Is angle tracking that sets the

limits on determination of optical path length, far higher distance measurement accuracy could be achip ,4od using Interferometry rather than War. Furthdrmore, use

of a multicolor laser Interferometer permits absolute distance determination. A devIce claiming one part In 109 accuracy Is described by Bien at al.(1981). 0.02μ accuracy for modulo one-meter absolute distance measurement has been demonstrat-
FQ

crJ

c^Tl

^ 

IWN

:d

ftJOUQO

aD , Ja

fOZ— FQ-UWZ

f'-

CCQJ

U)

i

w

VrW

^ a

it Q Q

J ^"

Co O

W ir

Q cn

c vw

w W^ ^ a Zwa

U.wmO

w N

U J

Q 

Z

W

J _

0. J O

Vxw

1 ^ ^ ^ c w

O Z

Z W WZ

_ZW

QN

W 

X cr0

~ W

crV

H ^N

m =

J ^

W V_

0 Q

—j J

W =

U ^V

W U

a_

1i Z N H

O cr

V 

F 

O

w v^ Z

rN

CL n h

U O w -N

a a N

c O w

U N Z

Uf- 0

Q O oC

LL

VwNW

w2N

Fig. 3

) ) 

I ^j
.aad using a 8-color laser interferometer (8ourdst and Orszag 1979). The modulo 1-

meter ambiguity Is oQslly resolved with either Iidar or radar. Coarse and emergency

location would also be done with radar.

With highly accurate distance determination possible, the principal requirements

for fringe position prediction are precise spacecraft tracking and large angle measurement. 0.04 aresecond (2mm at 10km) tracking In space applications was reportad 10 years ago in Air Force Program 40M (Elson 1973). If the back-to-back wwntral station telescopes have 20 cm diameters and operate at the diffraction limit In

+.;to visible, 0.1 arcesc tracking (8mm at 10 km) requires positlo ►t sensing to onefifth of the aiffractlon limit. SOT, the 1-meter Solar Orbiting Telescope, is designed

for tracking to an accuracy 0.2 times Itr, diffraction limit on a far more difficult object, the tiun (Dunn 1980). Space telescope Is designed for tracking to 0.16 times

its diffraction limit (Jones 1979). Spacecraft pointing would be done with the usual

combination of control moment gyros and momentum dumping by means of thrusters.

The attidue of the three spacecraft with respect to an ext6rnal coordinate system

would be obtained from measurements of light collector positions against the grid of

background stars made at the central station.

Once located, the fringes must be tracked. Analyses of tracking and measurement errors arising for photon-noisy fringes are found In Kibblewhlte (1979) and In

more detail by Hardy and Wallner (1978). An analysis of signal-to-noise in MichelV son Interferometry also appears in Tango and Twiss (1980), and In Greenaway

L (1970).

Figure 4 shows one of the light collector satellites. This design could, In princlpie, be replaced by a large flat, or, for cases where the source is far off the normal to the line along which the three spacecraft are displaced, a pair of fiats.

Substitution of large fiats, however, seems to offer no advantage of cost or simplicity of design or fabrication, and requires larger telescopes on the central station
TERS

THRUSTERS

i

iSER )STARLIGHT

SER

LIGHT COLLECTOR 1

F ig. 4
-7-

to accept the unconverged beams.

The light collectors must be diffraction-Ilmited optical systems and require full

three-axes stabilization to, perhaps, a fifth of a diffraction resolution element.

(The diffraction limit is 0.1 arcaec for a 1-meter aperture.) An acquisition and tracking system is required to control the small flat directing light from the collecting

telescope to the central station, with fine tracking based on signals transmitted

from error sensors on the central station.

A sample optical layout for the central station Including a scheme for pointing

and range error signal generation, appears In Figure 5. In this system, lidar, rather

than interferometry Is used for ranging. A legend describes the basic optical components. In the diagram, quad cells are four-element position sensing devices,

which generate corrective error signals If the detected signal is not balanced

among all four sensors. Nutating mirrors or other systems may prove more sensitive. Eitsments marked "active mirrors" are tip-tilt devices except for the delay

line mirror which has linear motion only. The dichroics transmit laser light and reElect light from the source to be studied. Coarse pointing may be done with small

motions of the two central station telescopes. Fine pointing is done with smaller

active mirrors.

A pulsed laser in each of the two tracking and pointing systems periodically

transmits pulses outward, through the active optical system, in the direction of the

light collector it Is to track. We shall consider the case of system number 1, trackIng light collector no. 1.

Part of the outgoing pulse is picked off (here by the back surface of the dichroic) and relayed to the ranging pulse sensor, D1. The pulse is returned to the

central station by a retro reflector mounted on (possibly at the center of) the active fiat on light collecting satellite no. 1. The pulse returns to detector D1 after

passing through the delay line, through the dichroic, and reflecting off one beam

E.
ui

^__ W I I I ---, J

WI I p (

U. I N

a I ( p

f

m I V I I J !V

Z I I p J

I• I I

INm

_^

r 

Z

O0 I H a m IsZ I IYau II ^^ ^

I~ I G I O_ O • u

^x l g

4 I J la m

f/f

IZ

^ I u

f O

> I^ W

Fig. 5

J W

IQ

F-00a

U_QJNUWi

cQU

2

Z M

^ n JJ U M

<

Ji
-8-

splitter then passing through another. The difference between the distances obtained from systems 1 and 2 is compared with that calculated for the expected location of the fringes and the central station Is moved as appropriate. Fine adjustment is made by means of the delay line. The component of the return pulse reflected off the second beamsplitter goes to a quad cell or a similar device to generate pointing error information.

The sensing beams are coincident with the data beams for which we wish to

obtain Interference. The data beams are reflected off the front surfaces of the dF

chroics and into an Interferometer. The Interferometer in Figure b Is shown

schematically. In fact, a modification of Michelson's original method of interfering

the beams would be used. The beams would be brought together at a narrow angle

after passing through a dispersive element permitting simultaneous observation of

the resulting fringes over a wide range of wavelengths. Figure 6 shows the fringe

patterns falling on a set of 4 array detectors covering a spectral range between

0.1 and 18.0 microns.

An extremely important sensor characteristic, now demonstrated at visible

wavelengths (Papallolios and Mertz 1982) is image d3tection In a photon cataloguIng mode. Since certain photon-counting array sensors record detected photons as

an x, y, t catalog, t being the time of detection, it is possible in the case of a regularly changing pathlength (fringe drift) error, first to detect the photons, then

record them as times and addresses, perform drift compensation in software, and flpally, use the new positions to generate a drift-corrected fringe pattern and measure a contrast. If the absolute drift rate Is unknown, different values are tried

with the goal of optimizing contrast. Furthermore, observed changes in velocity

between integration Intervals can be used to model accelerations affecting contrast measurements during an Interval.

. 
irZ a

 cr

Ena

ooa

w ir

o

a w

w z a co

N

0

J,

O'

o^

0v

J

Fig. 6

J
-9-

III. OPERATIONAL CHARACTERISTICS AND ORBITAL MOTION

In operation, the three spacecraft comprising SAMSI would deploy from Space

Station powered by thrusters capable of producing a net thrust of 0.6 newton in

any direction. This thruster size was chosen to guarantee station-keeping for

1000 Kg. spacecraft In a low earth orbit at all orientations, and for separations of

several hundred meters (after Columbo et a/., 1978). An estimate of 1000 kg per

1 meter telescope/spacecraft appears safe since cube-law scaling from the 10600

kg mass of the, 2.4-meter Space Telescope, suggests a mass of 770 kg. The mass

of IRAS, with a 80 cm aperture, is 812 kg. Of this, 608 kg Is associated with the

cryogenics (Trace and Rosing, 1983).

The spacecraft would move an appropriate distance, perhaps a few kilometers,

from Space Station, and the two light collectors would point at the source to be

studied. Crude relative positioning of the major optical elements would be followed

by fringe search and acquisition. This process would be done with the spacecraft

at a small separation to assure high Initial fringe contrast. Concern about moving

large spacecraft in such close proximity to one another is legitimate; for the system to work at all, however, station-keepir;g must be maintained to within microns.

Unless this requirement is met, the device is useless. With a properly designed experiment, one can insure that failure of one or all onboard systems will not lead to

collision.

Either one or two dimensional measurements can now be made. For comparetively bright sources (mode 1), full image reconstruction can be achieved if the

light collecting satellites are flown in a spiral of constantly Increasing diameter so

that measurements of fringe contrast and position as functions of color can be obtained over a grid of positions. The fringe contrast and position Information are

used to obtain Image Fourier transform amplitudes and phases. The Inverse

transform yields an Image of the source with a resolutlo:,, in radians, of order a/D
-10-

where A Is the wavelength of the light and D Is the maximum distance between

light collecting spacecraft. )1/D is also the resolution of a single monolithic aperture of diameter equal to the maximum spacecraft separation.

We have calculated that a 0.5 newton thruster permits a pair of 1.0-meter opta • Ical collectors with Individual masses of 1000 kg to completely fill the area of a

100-meter equivalent aperture In 9 hours. We note that so finely-sampled a grid of

measurements will not, in general, be necessary. The observing strategy Is subject

to significant optimization as a function of source. The visible wavelength resolution associated with a 100-meter aperture is 1 milliaresecond, approximately 40

times that of Space Telescope. Five hundred meter spirals are possible for the

thruster size and fuel use we have specified.

Because the predominant thrust requirement is provision of a centripetal force

for the light collectors circling the central station, the use of tethers is an obvious

possibility. Due to problems of interforence with light beams and transmission of vibrations we have not considered their use here but combining tethers with thrustars would yield a significant fuel use advantage.

An alternate mode of operation (mode 2) is more appropriate to the case of

faint sources, for which poor signal-to-noise in the fringes dictates slower spacecraft motions and longer integration times. In the case of these objects, onedimensional measurements with very slow spacecraft drift rates represent the most

I ' efficient use of observing time. Very long one-dimensional baselines are a possibillI, ty at any light level above M my = 19, and for faint sources the signal-to-noise requirements are greatly relaxed by concentrating only on accurate measurement of

the separation corresponding to the first fringe null, all that Is needed to perform a

diameter determination. The observing strategy described in Stachnik of al. (1983)

suggests that selection of the proper orbital configuration permits very long baseline measurements with extremely small expenditures of fuel.

---- 
- 11 -

IV. POSITION SENSING, THRUSTER REQUIREMENTS AND VIBRATION CONTROL

It is clear that the chief technical problem to be solved is that of assuring

sub-micron pathlength equality at the fringe sensor surface as the major optical

elements move In a complex pattern through a differential gravitational field. Very

high level precision thruster control Is required and onboard vibration must be

minimized. Residual contributions to path-length inequality must be sensed and compensated.

The most direct method of s ,)using path error depends upon measurement of

the fringes themselves. Introduction of a wie wave differential path error will

result in displacement of the fringe pattern by a fringe width. Problems arise when

noise masks the fringe modulation used to generate the error signal. In the ideal

case, the noise Is due purely to quantum fluctuations and becomes a serious problem when the source brightness is low, when the fringe amplitude to be measured

Is small or when uncompensated vibrations or accelerations necessitate extremely

short "exposure times."

We shall see in Sections V and VI that for full reconstruction of complex Image

structure It is necessary to measure fringes of very low contrast; however, the

spectral cross-dispersal technique assures that, at some wavelengths, relatively

high contrast fringes will always be available for error signal generation. At our

maximum Fourier plane sampling rate, set by the available thrust of 0.5 nt, we suggest In Section VI a sampling time at each spacecraft separation and position angle

of 3 seconds, and Indicate that this will result In a photon count adequate for

phase recovery for 8th magnitude stars. This limit is chosen to Insure 1%, or

better, measurement of fringe position over the Integration time, the accuracy we

believe necessary for sources of Interest.

Is it possible to Insure a high degree of pathlength stability over the 3-second

Integration period? Several levels of sensing and correction are available to help
- 12-

satisfy this condition. The pathlength correction systems Include the thrusters

themselves, the optical delay line (which may Incorporate both fine and coarse

correction loops), active plazo vibration damping of key structures, and postdetection fringe drift compensation of the sort described at the end of Section 11.

Sources of error signals, other than the fringes themselves, Include precision accelerometers, piezo structural vibration sensors, piezo thrust sensors and modeling

of thruster effects and residual vibrations.

Thruster uniformity requirements vary with frequency range. For the case of rF

gid bodies at least, the large inertia and small thrust of the engines make the

spacecraft surprisingly insensitive to large thrust fluctuations at high frequencies.

Forces required to produce 0.02μ displacements (0.04 of a fringe at a wavelength

of 0.5 μ) are given below as a function of frequency.

Thruster Force Error Required

to Produce a 0.024L Displacement Error

(1000•kg Rigid Spacecraft)

Integration Force Error

Time (sec) (nt)

0.01 4 x 10"

0.1 4x 103

1.0 4 x 10'r3

10.0 4 x 10-7

If correction updates are available every 3 seconds, we require 4μ.nt thrust

uniformity of our 0.5-nt thruster. Thrusters In this range exist at present.

Typical bi-propellent thrusters have 2-6X thrust uniformity over a wide range of

frequencies (Sutherland and Maes, 1088). An accepted technique for obtaining

higher uniformity Is to use a "layered" system consisting of a large thruster which

produces the basic Impulse and a smaller thruster which, when provided with the
-13-

appropriate error signal, compensates, at the several percent level, for fluctuations

In the larger device. Three to five layers would provide the uniformity required

here. The lowest thrust "layer" could be provided by variable reflectivity microshutters changing both reflectivity and reflectivity angles of the skins of the different spacecraft (Labeyrie, et al. 1982, C.E.H., 1983).

Each thruster layer Is mounted through a plezo strain gauge measuring the

force provided by that particular motor. The resulting measurement force for each

motor provides control signals for the next lower level of thruster. The entire propulsion package is mounted through a strain gauge monitoring overall force uniformity. At the lowest frequencies, the applied force measurements might be supplemented by signals from precision accelerometers of the sort used in the Discos

(Johns Hopins, 1974), Cactus (Bernard et al., 1982), and Super-Cactus (Bernard et

at., 1982) experiments. Super-Cactus Is a particularly interesting device since it Is

capable of measuring accelerations as small as 10-11 m/sec,. Super-Cactus is, however, non-linear in response, having Its full sensitivity only at low accelerations.

Devices having a similar sensitivity at accelerations as high as 10"6 m/3ec 2 (Bernard et al., 1982) are projected for the near future. This approaches typical SAM•

SI accelerations of 6 x 10 4m/sec2. Accelerometers measure only nongravitational effects, however, and derivation of positions from accelerometer data

also require modeled differential gravitational potential Information derived from

ultra-precise gravity gradiometer maps. In general, we believe that the best and

most practical method of relative position determination will be fringe measurements.

Continuous operation of the thrusters even at a modest thrust level, suggests

the use of a high specific impulse propulsion system. A Siquid hydrogen-liquid oxygen (LH2-LOX) combination has a specific Impulse of 480 sec (Forward, 1983), and

a 87 kg. propellant mass would assure continuous 0.6 nt operation for a week. As-
.14-

suming that only a top-level thruster, having a #3% thrust uniformity, uses LH2-LOX,

and that kowerlevei thrusters use much more controllable but less efficient cold

gas technology, the maximum additional propellant would be 26 kg: per week.

Use of cryogenic propellants has the disadvantage of requiring large dewars.

In the case of IRAS, 76 kg. of liquid helium required a 432-kg. dewar (Irace and

Rosing, 1983). IRAS, however, had a hold-time of 12 months; we require a holdtime of a week. Furthermore, the more easily held liquid oxygen would presumably

provide a Jacket for the liquid hydrogen. The resulting dewar design could be considerably lighter and less sophisticated. Cryogenic material will be required, as

well, for measurement of signals In the thermal Infrared and the same cryogenic

reservoir feeding the propulsion system can supply, as well, an LH 2 LOX fuel cell

supplementing the estimated 600 watts available from solar cells covering the bodies of the light collectors. No additional power may be needed, since 600 watts

Is all the power required by IRAS.

The foregoing suggests that reasonable amounts of fuel can be carried by the

SAMSI spacecraft for a week's observing. This time may be extended significantly

If use of tethers, now under study, proves practical. When fuel levels become

unacceptably low the spacecraft would return to Space Station for refueling.

Experiments reported by Forward (1980, 1981a, 1981b) describe considerable

success In sensing extremely minute vibrations In structures by means of plezo

strain gauges of the type used In gravity wave detectors. The appropriate signals

are then sent to other piezo devices, used as drivers rather than sensors, In order

to achieve active mechanical damning of the structure. The result has been damping of vibrations In optical systems of as much as three orders of magnitude over a

range from ten to 6000 Hz. Forward (1980) states that measurements corresponding to differential displacements of 10-6 microns, one tenth the diameter of an

atom, have been made and that In one experiment, active damping of a membrane
-16

mirror support was so good that, witn the damping loop closed, reflections from the

mirror were identical whether or not vibrations were being excited In the sucport.

This promising technology can be extanded to frequencies below 1 Hz (Forward,

1983).

V. PHASE MEASUREMENT AND IMAGE RECONSTRUCTION

The problem of full Image recovery using Michelson Interferometry has been

less extensively studied than has the simpler case of determining source angular

diameters. This Is true because the latter problem requires only measurement of

the light collector separation at which fringe contrast drops to zero (the contrast

rises again for larger separations). Full image reconstruction requires both fringe

contrast and phase measurements (to be described shortly) at many position angles and separations at far larger baselines, for which the contrast drops to

perhaps a few tenths of a percent.

Ground-based measurements of such extremely low contrast features are complicated by the fact that the atmosphere imparts severe, and randomly changing,

distortions to the wavefronts transmitted to the central station by the light collecting elements. An early approach to minimizing this problem was to pick off a small

portion of wavefront of characteristic size, r o, for which

1 cm < ro < 10 cm.

Over ro the distortion is small and the wavefront may be regarded as essentially

plane. The rapid and random pathlength changes remain, however, and the same

size of ro means that few photons are available for interference. Ingenious alternative schemes have been devised permitting use of larger portions of the wavefront but they must still contend with unpredictable pathlength fluctuations and, in

general, pay the price of measuring not the fringe contrast directly but, rather, the

square of this quantity. This Is not a serious shortcoming when one wishes only to
- 1elocate a first null, but, when the measured modulation is very low, the signal-tonoise may be too poor for even extremely bright sources.

Comparison with the circumstancer surrounding space-based Michelson Is strikIng. The wavefronts are as plane as available optical technology can make them.

Vibrations are readily modeled and can, In any case, be directly measured by highly

sensitive piezo sensors. There Is no atmospheric absorption of light outside the

visible range and there is no need for amplitude-square measurement techniques.

Furthermore, in space, in the absence of random wavef rant deformations, a

technique proposed for use on the ground by Koechiln (1977), to use distortions In

dispersed fringes In the measurement of Fourier transform phase, becomes for more

powerful. The problem in measuring very small displacement of a fringe from the

geometric equi-path position, corresponding to zero phase, Is In determining the

reference position. Koechlln pointed out that the technique of Labeyrie (1980) by

which fringes are dispersed along their length to permit simultaneous measurement

of contrast over a range of wavelengths contained a possible solution. Since resoiution varies as X/D, the resolution at a given light collector separation Is a linear

function of wavelength. It is, In fact, possible for an object to be unresolved at

one wavelength and well resolved at another. Since a completely unresolved point

has no structure contributing to fringe displacement, the zero phase reference provided by a star observed at one wavelength may be used to measure phase displacement at other wavelengths. "Wiggles" as a function of wavelength (shown In

F Figure 8) become markers of phase as a function of wavelength. This sort of differential procedure becomes still more powerful when a very large range of

wavelengths can be simultaneously measured. !:rsy detectors exist over a spectral range extending from 0.1 to 18 p., and there is no reason, In principle, why this

entire region cannot be observed. Furthermore, this implies a resolution ratio of

180:1 at a single baseline.
- 17-

In fact, of course, complications exist. Detector sensitivity curves, source

spectral dixtributions and changing source morphology as a function of wavelength

make both measure , .ient and Interpretation of fringe wiggles difficult Very high

contrast Fringes P 15%) .xlst only for the case of an unreso:ved scu co. Thus,

such very high contrast fringes can always be taken to mark a zero phase referonce. Obviously, even If they occur In disadvantageous spectral emission or detector response regions, very high contrast fringes can be easier to measure t;•-n low

contrast fringes.

If no high contrast fringes can be measured at adequate signal-to-noise anywhere in the spectral range, the problem becomes somewhat more difficult in that

selection of the best approaches to phase recovery may depend upon having some

knowledge of the object. In the very prelimianry experiments we have done,

descriL,ed in Section VI, model dependency seems very weak. For stellar-type

sources, local maxima of the fringe amplitudes occur at locations where the phase

takes on values near either 0 or t !r. Information on detailed source structure, in

the form of fringe "wiggles" occurs away from these maxima. It Is pos.:'ble, that

reference phases may always be relatively easy to find, with Zero displacement occuring at specifiable amplitude maxima.

We note that a .curate recovery of the phases Is far more Important in successfully Inverting the Fourler transform to obtain an Image then is -tccurate messurement of the amplitudes (Andrews and Hunt 1977). We undertake discussion of

the problems of measuring both in the next section.

VI. SIGNAL-TO-NOISE ESTIMATION FOR SAMPLE SOURCES

The "magnitude limit" of SAMSI will depend on details of the system technical

characteristics, numerous specifics of the observing protocol (Including spectral

bandwidth and integration time) and nature of tl source. Furthermore, as we have
.isalready indicated, full image recovery (mode 1) has far more stringent signal-tonoise requirements than does simple diameter determination. The limiting magnitude

will therefore differ substantially. We have not yet attempted a full signal-to-noise

analysis. What follows is an extremely simple examination of the general character

of the fringe behavior combined with reasonable assumptions about system characteristics in an effort to arrive at preliminary estimates of visible wavelength magnitude limits for both observing modes.

Figure T shows several model stars for which we investigated the general

behavior of the associated fringe patterns. All are characterized by limb darkening

coefficients of 0.5 and between 15 and 20 circular surface features ranging in

size between 6% and 35% of the stellar diameter. Feature contrasts vary in the

range 10% to 407E One star, C, has a random, 107E modulation granulation pattern.

The stars are 32 resolution elements in diameter.

Figures 8a, 9a, 1 Oa, and 11 a show the associated fringe (Fourier transform)

amplitudes and 8b, 9b, 1 Ob, and 11 b show fringe phases along one dimensional

cuts. Most ground-based long-be.^ line interferometry concerns itself with measurement of the location of the first null and the secondary maximum. Full image reconstruction at a resolution corresponding to 32 elements across the stellar disk requires measurement of the entire curve shown in these figures. The horizontal

"spatial frequency" scale could equally well be labeled "light collector separation"

or "baseline." For a circular source, the first null occurs at a baseline of 1.22 )./a,

where a Is the wavelength and a, the source angular diameter in radians. Thus, our

four sample Images depict the 5000A wavelength resolution one might achieve with

a 100-meter baseline for a 33-milllaresecond diameter star.

In addition to amplitude cuts, we have also generated phase cuts. Each ampli- 

tude null is associated with a w phase change. For a completely centro-symmetric

object, these changbs are discontinuous. For more complex stellar surfaces, how-
Or?•, f ir.' ;' ,^

OF POUR

F ig. 7
19-

ever, the phase assumes values near either zero or * w only when the amplitude Is

at a maximum and otherwise takes on intermediate values. For the 18384 point

two-dimenskmal Fourier transforms associated with the four Images in Figure 7, histograms of the phase show that over half are within * 0.03 radian of zero and 10%

are within * 0.03 Radian of * w. For a fringe phase resolution of * 0.10 radian,

26% of the measured phases are seen to have values different from zero or * ir.

It is these values that carry information on the non-centro-symmetric image structure. As we have already pointed out, accurate phase recovery appears (Andrews

and Hunt, 1977) to be of greater importance than accurate determination of the

amplitudes. Clearly, the foregoing does no more than to suggest the level to which

measurements must be made for reconstruction of surface features to be possible.

It is of Interest to estimate the magnitude of a star for which phases can be determined to within * 0.10 radian.

After Hardy and Waliner (1979) for the case of direct -not quadratic-fringe contrast measurement, and with contrast reduction due to photon statistics alone, the

RMS fringe amplitude and phase measurement errors, v,l, 

A 

and Q4 are given by

f

s,r A = T

f

Q#

where N is the number of detected photons In a measurements interval and y A is

T the Image Fourier transform amplitude. a,? 

A 

measures the fringe contrast error relative to the contrast at 7A = 1 and s 4 measures the phase error at a given fringe

contrast, 'yA, in radians.

Inspection of the Fourier transform amplitude curves suggests that fringe contrasts In the range 0.006 to 0.001 must be measured in order to obtain useful Information. At a fringe contrast of 0.003, 106 photons are required for * 0.1 radian
a

Q

I--

Q

J

Lcr—

W

O

D

Z

at

w

bo

ao

m

Q

.)

U u-

~O

Q^

U)

O^

co Z

uO

D

_W

crH

0

0 0 to 0 Lo 0 to

O —— N N M

1 1 1 I 1 I

3anil-ldVYV DOI 03Zl'lb'WaOR

^pi,
~ 

<{~ r-t- LL 

uC/) 

-

~ 

.... 

~ 

v 

o 

w Z 

0 

(DO: 

"j<{ 

~ 

:::> 

0: -

W 

0: 

.... 

C. 

C 

( 

( 

o 

I 

rt) 

I 

N 

o 

I 

o 

~ 

) 

~ 

~ 

.J 

~ 

~ 

o 

o 

3S'\1Hd 

I -

o 

I 

1 

N 

o 

I 

~ 

~ 

rt) 

o 

1 

I 

o -----

. 

V 

I 

(f) 

~ 

.... -

« 

...J 

u.. 

0: 

W 

0 

:::> 

~ 

Z 

(.) 

>-

t 

00 

.J:l 

btl 

. 

~ -
w

UO

^cr

CID g

U)

O 

CO Z

LL

O

D

w_

H

cr

O O o In O in O in O

O N N M M

1 1 1 i ^

30nllldVYV 001 03zI'lb'waON

w

O

w

Z

V

t

LL

w

a

Q

H

Q

J
O

OD

FEw

0

00

00000

0

M

N

-O^

N

M

3SVHd

J 

cr W0DWZVr1

w

aaa
U

1

cn

aH

Q

J

LW

O

D

W

w

r,

0

O ZQ

m(n

W O

U ^

U

W

O

D

cr

W_

w

H

cr

O In O O ^? O

N M O6 ^ ^ N ^ ^

3onil -IdWd 90-1 03ZI-lVVY80N

^^l
d WZV1

o

crW

w

aQ

F-QJ

1

^tO

O

CO

W Q

F0

ILOD

tYW I I I v ^?

O

O

M

N

-

O-I

N

M

I^

3SVHd
w

d

D

w

Z

V

a

Q

H

Q

J

LL w

.^p

m

w 

N

u.

CD Z

Oa

UO

_w

cr

H

cr

0 to O O Ir. O

O O N N M M

I I I 1 I

3anll -ldWd JO". 03ZI"IdINa0l",J'

oil
O Fm m

^ Q

W Z

LL

O

IY

W_

it

W

Z

U

}

1

r, W r,

LAcr

w

.^p

a

Q

H

Q

J

O O O O O O O v O

M N — O — N M

I ^ 1 I

3SVHd

, J
.20-

phaae measurement.

Figure 12 shows number of detected photons per 100A spectral band for stars

belonging to ei' -- spectral types. The assumed spectral response Is that of an S20 photocathode. The eight stars have the same V-magnitude (note that the

curves cross at A - 0.640L). Other assumptions Include:

- 2 nne-meter telescopes

-mv=6

- Integration time = 3.0 sec.

- Optical efficiency = 0.2

The 3-3ec integration Interval allows time for 11000 samples in Fourier space

during a 9-hour observing period. For the case of the 100-meter spiral, typical light

collector velocities are 0. 16 m/sec, Implying spacecraft motion In a 3-3econd Interval of approximately half an aperture width. The restriction of a 3-3ec maximum Integration time Is set by the 0.4μ-nt/Integration time thrust uniformity-controllability

requirement already discussed.

It is clear that, for all spectral types, many more than 106 photons are present

In each 100A spectral band for the specifications used In Figure 12. At this same

100A bandpass, stars 3 magnitudes fainter could be measured over a spectral

range of at least several thousand angstroms. This implies imaging In tens of spectral bands simultaneously. A magnitude limit of 8 Includes all stars resolvable at

baselines of several hundreds of meters. For sufficiently bright stars, it is possible

to narrow the bandpass to permit acquisition of, In effect, "spectroastrograma"-

single spectral line stellar images.

For the case of simple angular diameter determination (observing mode 2), we

need only measure fringe contrast to within 10% in order to Insure detection of the

secondary maximum. Thus, since,

` N o,^ 0.1 030

A
6.0

0.0 1--

0.35

FOR

2 ONE METER TELESCOPES

INTEGRATION TIME = 3.0 sec

OPTICAL EFFICIENCY = 0.2

BANDPASS = 100 x

S-20 PHOTOCATHODE

mv=5

mo

M5

0.40 0.45 0.50 0.55 0.67 0.65 0.70

WAVELENGTH (,.)

0

O

X 5.0

a

a

0

z

m 4.0

_ZL

O

O

W 3.0

a.

z

O

F-

= 2.0

a

U-O

w

m 1.0

D

z

0

B

A

F

0.75

Fig. 12
-21 .

300 photons per measurement are required.

Figure 12 shows that at least 1.5 x 10 8 photons are available from a 6th magnitude source under the specified observing conditions. Increasing both the

bandpass and integration time by an order of magnitude is permissible (for one di-

;;; mensionai measurements, the propulsion system, the principal source of unmodeled

fringe wander, can largely be shut down In favor of using differential orbital effects

to move the spacecraft), bringing the photon flux to 1.5 x 10 8, 14 r+agnitudes

brighter than is needed for diameter determination only. Given the stated condltions, then, we have a magnitude limit of 19.

Diameter determination alone, for very largo baselines (> 10 km), has olready

been explored by us with the aid of NASA "Center of Excellence" and "Innovative

Research Program" funding (Stachnik et d. 1983). All of the findings of that earlier

study apply to the system described here. The key result of that work was that,

for one-dimensional spatial measurements, extremely favorable orbital configurations

exist, for which "automatic" baseline scanning is obtained with only very limited expenditure of fuel. E for the full 10 km baseline, a now object can be studied

every 90 minutes. All of the ,`unctions of the original design, with the addition of

full imaging, can be fulfilled with the present system.

Compared to full Image reconstruction with its more limited baseline and less

favorable magnitude limit, this second observing mode has at least as great, or

greater, a scientific Importance.

VII. ASTROPHYSICAL APPLICATIONS

The astrophysical applications of SAMSI differ considerably depending on which

of the two observing modes—full Image reconstruction or one-dimensional power

spectrum measurements--Is being discussed. The high information density and ease

of Interpretation characterizing the former is offset by Its less favorable magnitude

.Y t

7
-22-

limit, longer data collection time and more limited resolution.

In the full Imaging mode, expected resolutions are as follows:

Resolutfon in Microareseconds

Baseline (m) Wavelength (m)

0.1 0.6 6.0 20.0

100

600

0.2 1.0 10.0 40.0

0.04 0.2 2.0 0.6

+ The magnitude limit is at least 8.

Perhaps a half-dozen supergiants have angular diameters sufficiently large to

permit visible wavelength resolution of 30 elements with a baseline of 100 meters.

Surface structure on several hundred stars could, however, be monitored at 32-

element resolution, at a baseline of 600 meters. This need require only a slightly

longer observing time and a modest increase in full usage. Such imaging, especially

In narrow spectral bands and with time-sequences, could contribute significantly to

our appreciation of tl a solar-stellar connection. At a 600 meter baseline, betelgeuse would be resolved at over 160 resolution elements.

In general, angularly large stars are cool and above the main sequence. With

the 600-meter baseline, we estimate that 30-element resolution is possible for main

sequence stars as early as late-F. Lesser resolution is, of course, possible for still

earlier stars. Resolution is, in principle, possible for later-type stars but problems

could arise because the (approximately) 10-hour measrirement time becomes a significant fraction of the stellar rotation period.

Comparatively detailed studies of additional astrophysical applications of very

high angular resolution measurements app. ar In a doctoral thesis (Vatoux 1979)

and, In English translation, in a manuscript by Praderie (1979), In Davis (1979), and

in Field et al. (1981).
.23-

Full image reconstruction will be possible for stellar surfaces, close, and

E perhaps interacting, binary stars and circumstellar gas and dust emission features.

In general, this observing mode will be most appropriate to problems of stellar physics and dynamics, for which light levels are high. A more complete evaluation of

the scientific implications of full Imaging at these resolutions over a very wide

range of wavelengths will be carried out in a future study.

The second observing mode, for which very high resolution, one-dimensional

measurements, are posible, has been studied in more detail.

The resolution achievable with an (expandable) 10-km one-dimensional baseline

Is at least 10,5 aresecond. This is five orders of magnitude greater than is typical

of routine earth-based measurements and several orders greater than may be expected from other near-term space or ground-based imaging techniques. Meaningful

extrapolations regarding the likely significance of such ultra high-resolution observations have dubious value but it is possible to make some useful statements.

10-6 aresecond resolution is adequate to resolve an earth-sized object at 10

pc, a white dwarf at 40 pc, and the Sun at 1000 pc. Supergiants could be

resolved in nearby galaxies. With good signal-to-noise ratio, stellar limb darkening

can be measured and some inferences drawn about the presence of surface structure for stars far smaller than those which can be imaged. It would be possible to

measure directly angular diameters for stars of 1j spectral and luminosity classes.

Simultaneous multi-color observations are a a natural consequence of our interferometer design. Where direct distance measurements are available, reliable

linear diameters can be obtained, and one is left with the prospect of a powerful

geometric distance indicator calibrated in a very straightforward manner against

nearby stars. The number of specific astrophysical problems for which such information would be useful is very large.

Within the galaxy, Interesting measurements would result from studies of such
-24-

prosaic sources as rotationally-flattened stars and contact binaries and such exotic

objects as SS 433, (perhaps at a larger baseline) the Crab optical pulsar, and expanding nova shells at a very early stage in the explosion. In all cases, simultaneous collection of Infrared and ultraviolet data would permit a far deeper understanding of the underlying physical processes in these systems.

A list of extragalactic objects for Investigation would certainly Include quasars

and galactic nuclei as well as, at somewhat larger baselines, individual sources

within nearby galaxies. The expanding shell due to a supernova explosion in a

nearby galaxy would be observable within hours of the beginning of an outburst.

Extending the galactic measurements, there is the prospect of developing

geometric distance indicators to tie intragalactic distances to local group extragalactic distance (using stars) and using geometric measures of quasars or galactic

nuclei to check or buttress conjectures about the large scale structure of the

universe.

VIII. MICRO-SAMSI

Studies have now been undertaken for two versions of SAMSI. The original

large-scale design had as its goal acquisition of large numbers of one dimensional

spatial power spectrum measurements for different objects, at very high (10-5

aresec) resolution, with a high observing efficiency (18 objects/day). The current

Space-Station associated version, known Internally, for a time, as "mini-SAMSI",

would fulfill all of the functions of the original SAMSI but would also make use of

the resupply capability of Space Station to replenish micro-rocket propellants needed to move the SAMSI spacecraft through the more complex maneuvers required

for full Image reconstruction.

In order both to assess the technology requirements of the larger systems and

to undertake some extremely interesting scientific experiments In the relatively
-25-

rm, it would be useful to consider a ath! smaller device, micro-SAMSI. Micro-

SAMSI would be a Shuttle-deployed package incorporating as many design simpliflcations as possible while still retaining the essential feature of SAMSI -- a multiple

spacecraft approach with potential for very high angular resolution measurements.

The design simplifications would Include elimination of the following capabilities

of the version of SAMSI described In fl- 4 report: observation over a wide

wavelength range, measurement of extremely low contrast fringes, Fourier

transform phase recovery, in-apac nr"ueP, ig and high-speed, high-efficiency twodimensional image Fourier transform :sampi`ng. Micro-SAMSI might rely heavily on

tethers or booms for spacecraft motion, hr*ve a simpler fine pointing system and be

designed primarily for measurenei. of high contrast fringe amplitudes at a small

number of position angles on bright sources. This would still permit simple angular

diameter measurements, 1,-.cluding limb darkening, for hundreds, possibly thousands,

of stars.

We hope to pursue a study of micro-SAMSI as part of a phased program leading

to eventual development of a full scale imaging device.

IX. FUTURE STUDIES AND LABORATORY EXPERIMENTS

Studies of SAMSI so far have concentrated on determining that plausible solutions exist for problems arising In all relevant technical sub-areas. Future investigations should concentrate on detailed analysis in a more limited number of key

technology areas. Four such areas stand out:

- Study of the image reconstruction problem with signal-to-noise

analysis and computer simulation.

- A laboratory study of the fringe measurement and active path

length compensation problems.

- Detailed orbital analysis for the required complex spacecraft mo-
-28-

tion in a differential gravitational field with special attention given

to use of tethers.

- Scientific assessment of experiments possible in the full imaging

mode of operation.

To date, extremely valuable technical contacts have been developed with European scientists in France, Germany and the United Kingdom working on a parallel

project under ESA auspices. We would like to explore future Involvement with a

NASA Center and, perhaps, with Industry technical representatives.

ACKNOWLEDGEMENTS

This work was done with the assistance of Keith Ashelin, of M.I.T., soon to be

with the Los Angeles Air Force Station. Sharron Hamilton, of Wellesley College, also

participated. Beryl Langer was the artist for the color illustration.
-27-

REFERENCES

Andrews, H.C., Hunt, B.R. 1972, Digital Image Restoration.

Bernard, A., Gay, M., and Quillerat, R. 1982, The Cactus Accelerometer, AGARD Dograph No. 264, 6.1.

Bien, F., Comac, M., Caulfield, H.J., and Ezekiel, S. 1980, Ap. Opt., 20, 400.

Bourdet, G.L., and Orazog, A.G. 1979, Ap. Opt., 18, 226.

Colombo, G., Arnold, D.A., Binsack, J.H. (MIT-CSR), Gay, R.H., Grossi, M.D., Lautman,

D.A., and Orringer, O. (MIT-CSR) 1978, SAO Report In Geoastronomy No. 2.

Davis, J. 1979, The Application of High Angular Resolution Interferometry to the

Study of Single Objects, Proceedings of IAU Colloquium No. 60, ed. J.

Davis and W.J. Tango, Chatterton Astronomy Dept., Univ. of Sydney, NSW,

Australia.

Dunn, R.B. 1980, Proceedings: Optical and Infrared Telescopes for the 7990's, Adelaide Hewitt, Kitt Peak National Observatory, Tucson, AZ, 168.

Elson, B.M. 1973, Av. NK. ane Sp. Tech., 98, 63.

Field, G., Katz, J., Muller, R.A., Press, W., Shapiro, I. 1981, Working Paper of the Astrophysics Working Group, Nat. Acad. Sci. study.

Forward, R.L. 1979, Ap. Phys., 61, 6601.

Forward, R.L., and Swigert, C.J. 1981, J. Spacecraft, 18, No. 1, 6.

Forward, R.L. 1981, J. Spacecraft, 18, No. 1, 11.

Forward, R.L. 1983, personal communication.

Hardy, J.W., and Wallner, E.P. 1979, Active Controlfor Michelson-Stellar Interferometers, Proceedings of IAU Colloquium No. 60, ed. J. Davis and W.J. Tango,

Chatterton Astronomy Dept., Univ. of Sydney, NSW, Australia, 10-1.

Hill, P.G., and Peterson, C.R. 1986, Mechanics and Thermodynamics of Propulsion,

371.

Irace, W. and Rosing, D. 1983, J. British Interplanetary Soc., 38, 27-33.

John Hopkins University Ap. Phys. Lab. Staff, Space Dept., and Stanford University
-28-

Staff, Guidance and Control Lab., J. Spacecraft, 11, No. 9, 637.

Jones, C.O. 1979, Opt. Eng., 18, 273.

Kibblewhite, E.J. 1979, report submitted to ESA.

Koechlin, L. 1977, Optical Telescopes of the Future, , ESO Conference, Geneva,

Switzerland, 476.

Labeyrie, A. 1980, Proceedings: Optical and Infrared Telescopes for the 1990's, ed.

Adelaide Hewitt, Kitt Peak National Observatory, Tucson, AZ, 786.

Labeyrle, A., Savarla, E., and Schumacher, G. 1982, Flute or Trio: Different Approaches to Optical Arrays In Space, IAU Symposium on VLBI, Toulouse,

France.

Papaliollos, C., Mertz, L. 1982, Instrumentation In Astronomy IV, SPIE Vo. 331, 380.

Praderle, F. 1979, Angular Interferometry from Space, unpublished report of meetInys and experiments of French astronomers at Meudon and CERGA.

Querzola, B. 1979, Ap. Opt., 18, 3036.

Stachnik, R.V., Melroy, P., McCormack, E.F., Arnold, D., and Gezari, D.Y. 1982, Instrumentation in Astronomy 5, SPIE Proceedings, 445.62.

Stachnik, R.V., and Labeyrie, A. 1984, Optical Astronomy from Clusters of Satellites,

Sky and Telescope, scheduled for publication February.

Sutherland, G.S., and Maes, M.E. 1988, J. Spacecraft and Roc., 3, No. 8, 1163.

Tango, W.J., Twlss, R.Q. 1983, Progress In Optics, 17, 241.

Vatoux, S.A. 1979, Etude Theoretlque et Experimentale dun Inte-ferometre Stellaire

Optique, a Usage Spatial, Doctoral Thesis, submitted to Universite de

Paris-Sud, Centre d'Orsay.
-29-

FIGURE CAPTIONS

Figure 1. Schematic diagram of a Michelson stellar Interferometer. Light from

widely-separated locations on the wavefront of a star are combined at the surface of a detector. If the pathlengths are equal

.. to within a few microns, fringes, whose contrast are determined

E 4 by the baseline and by the angular diameter of the source, are

observed. Spectral dispersion out of the plane of the page permits simultaneous fringe measurements over a wide spectral

range.

Figure 2. Illustration of the spiral pattern flown by SAMSI during image reconstruction, following deployment from Space Station, described In the

text.

Figure 3. Central station functional schematic diagram, showing a pair of back-toback telescopes which track the two light-collector satellites.

Light transmitted to the central station Interferes and fringe contrast is measured as a function of separation of the light collectors. Figure is from an earlier study In which Ion, rather than

chemical, thrusters were the primary propulsion devices.

Figure 4. Schematic diagram of one of the light-collector satellites, showing how a

converged and collimated beam from the source is transmitted to

the central station. The electric propulsion system, shown here,

would be replaced by LH2/LOX thrusters.

Figure b. Central station optical schematic diagram.

Figure 6. Dispersed fringes falling on area detector arrays spanning a spectral

range from 0.1 to 18.0 microns.

Figure 7. 4 stellar Images for which Fourier transform phase and amplitude characterlatics were studied. All stars have 32-element resolution, a

0.6-limb-darkening coefficient and between 16 and 20 surface
-30-

spots having a range of sizes and contrasts.

Figures 8, 9, 10 and 11, A and B. to ourlsr transform amplitude and phase cuts,

respectively, for the stellar images In Figure 7.

Figure 12. Numbers of photons per 100A passband in the spectral Interval 0.36 to

0.76 microns for stars of 8 spectral types. Other parameters are

specified on the figure.