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SPACECRAFTOPTICALTRANSCEIVERPACKAGE

FINAL REPORT

19 AUGUST,1985

JPL CONTRACT NO. 957061

S. G. LAMBERT, ET AL

(NASA-C_-179958) DES]GN AND ANALYSIS STUDY

OF A SPACECRAFT OETICAI _BANECEIVER PACKAGE

2inal Beport LMcDonnell-Douglas Astronau£ics

Co.) 286 p CSCL 22B

G3/18

This work was performed for the Jet Propulsion Laboratory.

California Institute of Technology, sponsored by the

National Aeronautics and Space Administration.

Reference herein to any specific commercial product.

process, or service by trade name. trademark,

manufacturer, or otherwise, does not constitute or imply its

endorsement by the United States Government, McDonnell

Douglas Corporation, or the Jet Propulsion Laboratory,

California Institute of Technology.

N87' 134 77

Unclas

43629

MCDONNELL DOUGLAS ASTRONAUTICS COMPANY'ST. LOUIS DIVISION

Box 5t6, Saint Louis, Missouri 63166 (314) 232.0232

CLASSIFIED DETERMrNATION' PENDING

PROTECT AS THOUGH CLASSIFIED

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MCDONNELL DOUGL_

COIPORat_N

TABLE OF CONTENTS

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ABSTRACT ........................ ................................. i

NEW TECHNOLOGY ................................................... II

l.O INTRODUCTION .................................................... l

2.0 SYSTEMS OVERVIEN ................................................ 6

2.1 Requirements ............................................... 6

2.2 EORS Characteristics ....................................... 6

2.3 Major Design Assumptions ................................... 6

2.4 System Design Rationale .................................... lO

3.0 DYNAMIC ENVIRONMENTS ....... ..................................... 12

3.1 Orbital Dynamics ......................... .................. 12

3.2 Polnt-Ahead And Doppler.. .................................. 12

3.3 Meteoroid Environment ...................................... 17

3.4 Gravitational, Magnetic, Electrical and Thermal

Radiation .................................................. 23

3.5 Vibration Environment ...................................... 23

3.6 Acoustic Environment ....................................... 23

3.7 Pyrotechnic Shock .......................................... 23

3.8 Saturnian Radiation Environments ........................... 28

3.9 Optical Background Radiation ................... _ ........... 28

3.9.1 Earth Background At OPTRANSPAC ...................... 34

3.9.2 Solar Background At OPTRANSPAC ...................... 34

3.9.3 Saturn Background At EORS ........................... 34

3.9.4 Solar Background At EORS ............................ 41

4.0 FUNCTIONAL ORGANIZATION ......................................... 45

4.1 Baseline Selection Rationale ...................... _ ........ 45

4.2 System Configuration ....................................... 45

4.3 Optical System Description ................................. 57

5.0 COMPONENT SELECTION RATIONALE ................................... 62

5.1 Laser Sources .............................................. 62

- 5.1.1 Downllnk Laser .......'........... .................... 62

S.l.2 Upllnk Laser ........................................ 62

5.2 Telescope Selection ........................................ 66
TABLE OF CONTENTS

PAGE

9.0

I0.0

II .0

5.3 Optical Mechanism ......... :................................ 66

5.4 Modulation Format Selection ................................ 68

5.5 Detector Preamplifier Selection ............................ 6B

5.6 Earth Tracker Detector Selection ........................... 72

5.7 Coding Selection ........................................... 76

6.0 LINK ANALYSIS ............. _...... ........ ,,,_,..,:,.,..,.,...,... 79

6.1 Acquisition and Tracking ................................... 79

6.1.1 Spacecraft Characteristics .......................... 79

6.1.2 Acquisition ......................................... 79

6.1.3 Tracking ............. 83

6.1.4 Earth Tracking Analysis ............................. 88

6.2 Communlcaton Links ......................................... 91

6.2.1 Transmit Galn/Polntlng Loss Relationship ............ 91

6.2.2 Downllnk Margin ..................................... 95

6.2.3 Upllnk Margin ....................................... 95

7.0 ELECTRONICS DESIGN .............................................. I04

7.I Control Electronics Assembly ............................... I04

7.2 Communlcalon Electronics Assembly .......................... 104

7.2.1 Receive Function .................................... I04

7.2.2 Transmit Function ................................... I04

7.2.3 Doppler Considerations .............................. I08

7.3 Power Conditioning Unlt .................................... I08

7.4 Redundancy Implementation .................................. lOB

7.5 Electronics Packaging ...................................... ll2

8.0 OPTICAL DESIGN..-................................................ ll5

8.1 Telescope Design ........................................... ll5

8.2 Imaging Optics Design ...................................... I15

8.3 Optical Transmission ....................................... 120

8.4 Wavefront Error and Image Quality .......................... 122

MECHANICAL DESIGN ............................................... 124

WEIGHT AND POWER ................................................ 135

GROWTH AND UNCERTAINTY .......................................... 140
TABLEOFCONTENTS

PAGE

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12.0

13.0

APPENDIX A -

APPENDIX B -

APPENDIX C -

APPENDIX D -

APPENDIX E -

APPENDIX F -

FOLLOW-ON HARDWARE DESIGN ...... ................................. 142

CONCLUSIONS ..................................................... 143

Environment Specification .......................... A-l

Component Specifications ........................... B-l

Design Notes ....................................... C-l

Interface Document ................................. D-l

New Technology Disclosure .......................... E-l

Acronyms ........................................... F-1

REFERENCES ...................................................... R-l

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ABSTRACT

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This study produced a detailed system level design of an Optical Transceiver

Package (OPTRANSPAC) for a deep space vehicle whose mission Is outer planet

exploration. In addition to the terminal design, this study provides

estimates of the dynamic environments to be encountered by the transceiver

throughout its mission llfe. Optical communication llnk analysis, optical

thin lens design, electronic functional design and mechanical layout and

packaging are employed in the terminal design. Results of the study describe

an Optical Transceiver Package capable of commun_catlng to an Earth Orbiting

Relay Station at a distance of lO Astronomical Units (AU) and data rates up to

lO0 KBPS. The transceiver is also capable of receiving l KBPS of command data

from the Earth Relay. The physical dimensions of the terminal are contained

within a 3.5' x 1.5' x 2.0' envelope and the transceiver weight and power are

estimated at 52.2 Kg (If5 pounds) and 57 watts, respectively.

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NEW TECHNOLOGY

The design of the OPTRANSPAC has required the development of the following New

Technology. The development of the Offset Dichrolc Optical Boreslght

Alignment Device provides for continuous alignment of the transmit and receive

paths (refer to Appendix E-l for details). The Offset Dichrolc Optical

Boreslght Alignment Device was conceived by W. L. Casey and first reported in

the monthly review for OPTRANSPAC (3PL Contract #957061) dated 15 February

IgBS. It was also reported in the OPTRANSPAC Midterm Review on 2-3 April

1985, Pages 41, 47, 48 and 65 and is included in this report on Pages 47, 48,

49, 57, 58, 60, 72, 75, 76, ll6, ll7, liB, 120, 126, 127, C-31 through C-Sl,

D-5, D-6 and D-7.

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l.O INTRODUCTION

The objective of this study was to produce a detailed design of an Optical

Iransceiver Package (OPTRANSPAC) for a deep space vehicle whose mission is

outer planet exploration. The optical communication system must be capable of

transmitting up to IOOKBPS of data to a relay station located in low earth

orbit. The transceiver design must provide accurate estimates of the

OPTRANSPAC size, weight and power as well as a definition of the overall operational characteristics. Operationally, the OPTRANSPAC must be capable of communlcatlng at the desired data rate from a distance of Saturn and beyond.

Figure l-l illustrates the operational OPTRANSPAC scenario.

In addition to the terminal design, the study was to provide estimates of the

dynamic environments (see Figure I-2) to be encountered by the transceiver

package throughout its mission llfe.

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Results of the study indicate a transceiver design capable of performance meeting the specified requirements. The physical envelope of the OPTRANSPAC is

illustrated in Figure I-3. The transceiver conslsts of an eleven Inch aperture

telescope fixed mounted to an optical baseplate. Imaging optics provide relay

paths to and from the telescope, laser and detectors. Three separate electronic boxes provide for power regulation, transceiver control and communication functions. The detailed weight and power=estlmates, outlined in Figure

I-4, indicate a transceiver weight of llS.0 pounds and power consumption of 57

watts.

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2.1 REQUIREMENTS - The specified system level requirements for the OPTRANSPAC

are outlined in Figure 2-I. The OPTRANSPAC. Is required to communicate wlth an

Earth-Orbltlng Relay Station (EORS) at a maximum distance of lO Astronomical

Units. Downllnklng of data from the OPTRANSPAC to the EORS at discrete rates

of IOKBPS, 30KBPS and IOOKBPS at a bit error probability of lO-3 is required.

The OPTRANSPAC must also be capable of receiving command data on an optical upllnk from the EORS at a IKBPS data rate (lO-3 BER).

Choice of useful wavelengths were set by the statement of work requirements.

The downllnk wavelength could not exceed l.O micron while the upllnk wavelength

was limited to under 2.0 microns. Outages, attributed to all factors excluding

earth blockage, were required to be less than I0%. This resulted in an availab_llty requirement of 90 percent and accounts for outages when the sun is in

close proximity to the OPTRANSPAC or EORS optical boreslghts.

Maximum Weight and Power Requirements for the OPTRANSPAC were set at 50 kilograms and 50 watts, respectively. Technology to be employed in the OPTRANSPAC

design was restricted to that "reasonable by 1988". Where the term "reasonable

by 198B" means laboratory demonstrated feasible, but not necessarily space

qualified by 1988.

2.2 EORS CHARACTERISTICS - While not specifically part of the OPTRANSPAC

design, the EORS characteristics directly affect the terminal design. The

major characteristics that contribute to OPTRANSPAC design considerations are

outlined in Figure 2-2. The EORS transmits a lO watt average power, 5 microradlan laser beam at the wavelength less than 2 microns. The EORS platform

stability and other dynamic instabilities were assumed controlled such that

open loop pointing of the transmit beam at the OPTRANSPAC is accomplished wlth

zero pointing loss. As a receiver, the EORS consist of a lO meter effective

clear aperture with a l mlcroradlan fleld-of-vlew. The receiver consists of a

photomultlpller based direct detection system. The quantum efficiency of the

detector at a nominal wavelength of 532nm was assumed to be 30 percent.

2.3 MAJOR DESIGN ASSUMPTIONS - The major assumptions made in the design of the

DPTRANSPAC terminal are outlined in Figure 2-3. The assumptions specify certain design aspects of the OPTRANSPAC that should be carefully, noted. Any

change in these assumptions may affect the completed design of the transceiver

terminal.

Spacecraft attitude control to within ±2 mllllradlans is assumed. This criteria provides for accurate pointing of the OPTRANSPAC telescope vlewfleld towards the Earth. Without this accurate knowledge, much larger telescope viewfields would have to be employed or a system design which included a glmballed

telescope would have to be implemented. In the former this would require a

large and bulky optical system causing added system weight and complexity. In

the latter decreased pointing accuracy (due to glmbal Jitter effects) would require a larger beam divergence resulting in an increase in system power in

addition to weight and complexity increases associated with including the

glmbal. Thus, this assumption provides for an OPTRANSPAC design which minimlzes weight, power and complexity of the transceiver terminal.
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In the event of attitude control disruption (outside of the ±2 m1111radlan

knowledge) and associated break-lock of the optical links, it ls assumed emergency command data is available to reestablish the ±2 mtlltradlan knowledge

and restore the pointing of the OPTRANSPAC boreslght towards the Earth. The

OPTRANSPAC system relies on this emergency control to reestablish communication

after a disturbance to the spacecraft attitude control accuracy.

The electrical interface between the spacecraft and the OPTRANSPAC is assumed

serial digital for all data transfer to and from each side of the Interface.

Data, including the spacecraft sensor inputs, spacecraft telemetry, and spacecraft attitude reference system data will be transferred to the transceiver

along these serial data ports. The redundancy philosophy Is assumed to put the

burden on the source to handle all redundancy switching. That is, signal flow

from the source (be it the transceiver or the spacecraft) will determine which

side of the load (prime or redundant) wtll be passed the signal. This technique is employed in current Lasercom systems.

All command data sent via the optical upllnk Is assumed passed to the spacecraft command decoder for processing. Commands pertinent to OPTRANSPAC operatton are decoded by the spacecraft and returned to the transceiver vla the

serial interface. Thls simplifies the command data format and allows for command decoding at only one location.

A final assumption deals with system acquisition time. It was assumed that

transceiver acquisition tlme was not a critical system parameter and was not

critical to spacecraft functions. For thls reason it was not considered a

major system driver to the transceiver design. Later, analysis indicated that

this was true and average acquisition times were calculated to be less than

three seconds.

2.4 SYSTEM DESIGN RATIONALE - For purposes of thls transceiver design study,

the following rationale outlined In Figure 2-4 was chosen. The environmental

factors that are prime system design drivers were identified. These included

optical backgrounds, orbital dynamics, polnt-ahead angles, Doppler, vibration

and range to name a few. Mission specific constraints such as the lack of a

continuous track beacon from EORS were also identified.

Candidate system designs were selected for consideration based upon the environmental effects, mission constraints and system requlrements. Preliminary

llnk analysis was performed to verify the validity of the designs. Initial

size, weight, power and complexity comparisons were then made. One or more

designs were then chosen and the link analysis was refined and components defined (lasers, detectors, telescopes, etc). Finally, a baseline configuration

was selected for its performance, size, weight, power, reliability and design

simplicity. Detailed mechanical, optical and electrical component design was

then performed to yield the final OPTRANSPAC terminal design.

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3.0 DYNAMIC ENVIRONMENTS

The dynamic environments assumed to be encountered by the OPTRANSPAC on its

mission are analyzed and defined herein. A complete dynamic environment specification is included in Appendix A.

3.1 ORBITAL DYNAMICS - A generic transfer orbit was generated for a deep space

probe from Earth to Saturn. The orbit pictured in Figure 3-I is a Hohmann

transfer, which is the minimum energy orblt for transfer. Transit time is IB2?

days (details of the orbit are given in Figure 3-I and Appendix A). This orbit

was used to calculate the relative velocities between the Earth Orbiting Relay

Station and the OPTRANSPAC spacecraft for use in the Doppler and polnt-ahead

calculations. Figure 3-2 illustrates the range versus time between the earth

and the OPTRANSPAC spacecraft using the Hohmann transfer orbit. The range Increases rapidly at first as the spacecraft is on the closest part of the orbit

to the Sun and the Earth is also swinging away from the spacecraft's orbit.

The Earth's orbital motion is the most obvious feature of the range, with osclllatlons of 2 AU superimposed upon the distance of the spacecraft from the

Sun. After one year, the range to the spacecraft is almost 5 AU, or nearly

half the distance to Saturn.

3.2 POINT-AHEAD AND DOPPLER - The Doppler shift of the light transmitted between the OPTRANSPAC and the Earth Orbiting Relay Station and the polnt-ahead

angle necessary for correct illumination were determined by the relative velocity between the two platforms. The four contributors to thls velocity when

the spacecraft is encountering Saturn are l) Earth's orbital rate around the

Sun; 2) Saturn's orbital rate around the Sun; 3) The EORS orbital rate around

the Earth; and 4) The spacecraft's flyby or orbital rate around Saturn. A flyby was assumed because it results in higher relative velocities than does a

Saturn orbit. A flyby with a perlapse of twice the radius of Saturn and an

EORS altitude of 300 nautical miles were assumed. The contributing velocities

of each of these in the directions defined are given in Figure 3-3.

Figure 3-4 shows the maximum contributions to Doppler shift and polnt-ahead

angles from the velocities noted in Figure 3-3. For polnt-ahead, all four contributors may have their total velocity perpendicular to the llne-of-slght and,

thus, apply to polnt-ahead. However, the maximum angle contribution from the

EORS is limited by the extremes of where it Can be in its orbit around the

Earth. As the OPTRANSPAC range increases, the effective angle traversed by the

Relay Station back and forth around the Earth is inversely proportional to the

range. The maximum point-ahead angle which the OPTRANSPAC must use is 440

mlcroradlans. Because the contribution of Saturn's velocity is only in one direction, the maximum polnt-ahead angle in the opposite direction is 312 microradlans.

The Doppler shift is caused by the instantaneous velocity along the llne of

sight. The contribution of the EORS was therefore, not limited as it was for

polnt-ahead. Saturn's velocity, however, only contributes through the sine of

the minimum angle of itself with respect to the llne-of-sight. Defined by the

minimum distance from the Earth to Saturn and the maximum crossrange of the

Earth's position. The Earth and spacecraft flyby contributing velocities are

the same as for the polnt-ahead case. The four contributions were summed to

get a maximum Doppler shlft at I064 nanometers of 2.3 Angstroms and at 532

nanometers of 1.15 Angstroms.

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Figure 3-5 through 3-8 analyze Doppler and point-ahead. Prior to the Saturn

encounter, only the Earth, the EORS, and the spacecraft transfer orbit motion

apply. Figure 3-5 shows the velocities of the Earth and spacecraft in the directions of the axes of the transfer orbit. The spacecraft has an orlginal veloclty of about lOKm/second more than the Earth's In the minor axis direction

and a ve]octty the same as the Earth's tn the major axls direction. As the

spacecraft travels- farther from the Sun, its velocity decreases as potentlal

energy increases. The Earthis velocity follows the expected slnewave pattern

as it orbits the Sun.

The velocities of Figure 3-5 were rotated into the llne-of-slght direction between the Earth and the spacecraft. The relattve velocities between the Earth

and spacecraft In this coordinate system, parallel to and perpendicular to the

llne of sight are plotted versus flight time (See Figure 3-6). These, when

combined with the Earth-Orblting Relay Station maximum velocity contribution,

noted previously, determine the necessary point-ahead angles and expected doppler shift during the Earth to Saturn transfer.

Figure 3-7 shows the necessary point-ahead angles for the OPTRANSPAC during the

Hohmann transfer between the Earth and Saturn due to the relattve velocities

shown in Figure 3-6. Adding the point-ahead angle needed to compensate for the

EORS motion (51 mtcroradtans maximum) creates a total point-ahead angle of 322

mlcroradians during the transfer as compared to 440 mtcroradians during the

Saturn encounter.

The maximum Doppler shift expected during the Hohmann transfer without the

Relay Station contribution Is plotted versus flight time (See Figure 3-8).

Adding in the maximum velocity contribution of the Relay Station yields an expected Doppler shift of approximately 1.7 Angstroms (x = 1064 nanometers)

during the Hohmann transfer from _arth to Saturn. This value ts less than the

maximum expected during the Saturn encounter.

Figures 3-5 through 3-8 illustrate the effects of Doppler and point-ahead durtng the Hohmann transfer are less than those during the Saturn encounter.

Thus, the OPTRANSPAC design must accommodate the maximum Doppler and pointahead effects defined by the Saturn encounter and In doTng so will also accommodate any Doppler or point-ahead contributions during transfer.

3.3 METEOROID ENVIRONMENT "_ Figure 3-9 tabulates the meteoroid environment

design requirement for travel to Saturn. This data was derived from the

Galileo Orbiter environmental design requirements, GLL-3-240 Rev. A, 3PL, 1982.

The integral fluence numbers from the Galileo environment were increased by the

ratio of relative transit times to Saturn instead of Jupiter. This ratio calculation assumes 90% of the Galileo environment fluence was attributed to the

asteroid belt and did not change. The mean relative speed was increased by

one-half the velocity change difference from transfer to Saturn instead of

3uplter.

Contact with the meteoroids occurs when the spacecraft co111des with the slower

moving particles. Upon encountering the mlcrometeorolds the spacecraft's velocity vector is assumed normal to the meteoroids' velocity vectors. Nominal

spacecraft velocity (approxlmately 15 km/sec) is orders of magnitude greater

than the velocities of the mlcrometeorltes. Vector algebra will show the contribution from the particles' velocities are negligible when compared to the

17
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18
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22
spacecraft velocity. Thus, it is concluded that the collisions occurs along

the direction of the velocity vector.

t

m

z- t

The critical component In the OPTRANSPAC system when considering the micrometeoroid environment is the telescope mirror. Since the primary mirror Is not

shielded it could possibly be exposed to the incident meteoroids. However, the

OPTRANSPAC telescope boreslght Is assumed to be pointed towards the Earth (approximately directly opposite the velocity vector). Due to this reasoning,

shielding requirements for the OPTRANSPAC terminal are assumed no larger than

the nominal shielding required for the entire spacecraft.

3.4 GRAVITATIONAL, MAGNETIC, ELECTRICAL AND THERMAL RADIATION - The data given

In Figure 3-10 was collected from several sources. They are given below.

A) Galileo Orbiter Environmental Design Requirements,

GLL-3-240 Rev. A, 3PL, 1982.

B) Space Vehicle Design Criteria (Envlronment), The

Planet Saturn, NASA SP-8091, 1970.

C) Space and Planetary Environment Criteria Guidelines

for use in Space Vehicle Development, NASA TM-82501,

1982.

D) Space Shuttle System Payload Accommodations, NASA

3SC-07700, Volume XIV, Rev. H, 1983.

The maximum gravitational, magnetic, electrical and thermal radiation environments are within present Lasercom capabilities, the Earth, the shuttle and

near-Earth orbit provide the largest values of these environments to be encountered by the OPTRANSPAC. Saturn's environments are less than these because

encounter is assumed at a distance of twice the Saturnian radius from the

planet's surface. Minimum thermal radiation at lO AU may require a thermal

design that includes heaters to keep the components within specified operating

limits.

3.5 VIBRATION ENVIRONMENT - The Space Shuttle vibration environment during

launch is illustrated in Figure 3-11. This data was taken from Space Shuttle

System Payload Accommodatlons, NASA 3SC-077_0, Volume XIV, Rev. H, 1983. This

environment is within present Lasercom system capability, Current Lasercom

systems are designed and tested to withstand the Space Shuttle vibration environments.

3.6 ACOUSTIC ENVIRONMENT - Figure 3-12 illustrates the Space Shuttle acoustic

environment during launch. The data was also taken from Space Shuttle System

Payload Accommodations, NASA 3SC-07700, Volume XIV, Rev. H, lgB3. This environment Is also within present Lasercom system capability. Current Lasercom

systems are designed and tested to withstand the Space Shuttle acoustic

environment.

3.7 PYROTECHNIC SHOCK - Figure 3-13 depicts the spectrum of the expected pyrotechnic shock environment. This is the maximum spectrum during spacecraft

launch/separatlon. It wlll be attenuated by intervening spacecraft structure.

The data was taken from the Galileo Orbiter Environment Design Requirement,

GLL-3-240, Rev. A, 3PL, 19B2. Present Lasercom design capability for pyrotechnic shock is a factor of lO less at the spacecraft/Lasercom interface. If

the attenuation of the spacecraft structure (which is as yet undefined) is

23
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24
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greater than a factor of ten, the design will be within present Lasercom capability. If not, shock isolation must be provided between the spacecraft and

the OPTRANSPAC system.

3.8 SATURNIAN RAD._IIA_TIONENVIRONMENTS - The OPTRANSPAC radiation environments

associated with the charged particle and radioisotope thermoelectric generator

are given in Figure 3-14. The trapped proton and electron environments were

taken from the Pioneer II and Voyager missions. Electron and proton doses are

representative of a Jovian flyby environment. From the data It was determined

that thls environment was worst case for Saturn, and therefore, representative

radiation levels for the OPTRANSPAC system were chosen as lO0 KRADS total dose

and l x lO14 (l MeV damage equivalent) neutrons/cm 2 displacement damage.

The OPTRANSPAC can expect to see a Galactic particle environment on the order

of that shown In Figure 3-15. Thls environment interacts with the system electronics introducing soft errors In the LSI/VLSI devices. Hard errors or single

event latch-up of devices will be eliminated by device selection.

Mitigation of the single event upset environment can be handled by system clrcumvention (i.e., redundant storage of critical data, error detection/

correction, watchdog timer, reasonableness of data checks, etc.). Appropriate

device selection can reduce the probability of soft errors and aid In the mitigation of these effects.

The mission environments contributing to total dose include the interplanetary

electrons, protons and solar flare particles plus the Saturn electron and

proton Van Allen belts. Figure 3-16 illustrates that under nominal spacecraft

shielding of one gram per cm2 (Al), the total dose experienced by semiconductor electronics will be 60 KRADS (Si).

The major contributing environments to displacement damage in semiconductor

electronics are the protons trapped In Saturn's Van Allen belt and the interplanetary cosmic ray protons. Minor contributions are made by the nuclear

power source and mission electron environments. Figure 3-17 indicates under

nominal rience 2.5spacecraft x lO13 neutrons/cm shleldlnq_ (l(l MeVgram/cm neutron 2 (Al)) equivalent the electronics damage in silicon will )expe- .

The total dose level of lO0 KRADS as determined from the electron/proton data

is currently within the capability of Lasercom systems. Development data and

hardness assurance results from Lasercom hardware programs on electronic,

electro-optlc and optical piece parts support this conclusion. The required

radiation design margin for the system can be accommodated wlth a hardness assurance test program during the production phase of the program.

The displacement damage fluence Is currently higher than Lasercom design levels, however, within the capability of selected electronics. Care In device

selection and appropriate development testing may be necessary.

3.9 OPTICAL BACKGROUND RADIATION - Five specific background noise source scenarios were defined for the OPTRANSPAC design. Figure 3-18 outlines the flve

scenarios. On the upllnk to the OPTRANSPAC from the EORS, background radiation

from the Earth as well as off-axls scattering of sunlight onto the detector

must be considered. The earth tracker must also deal with off-axls

2B
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29
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33
scattering of light onto the array detector. On the downllnk, the EORS receiver must concern itself wlth the radiation from Saturn and also off-axls

scattering of solar radiation.

3.g.1 EARTH BACKGROUND AT OPTRANSPAC - The calculation of the Earth background

at the OPTRANSPAC follows from the followlng analysis. The energy incident

upon the Earth in a particular wavelength band was reduced by the albedo

factor. The albedo factor of the Earth was taken to be approximately 30%.

[Ref. l]. The Earth was then assumed to re-radlate thls energy into a 2_

steradlan half sphere. The energy flux at the OPTRANSPAC is then simply a

function of the range from the OPTRANSPAC to the Earth. Figure 3-19 shows the

range relationship on Earth background flux at the OPTRANSPAC for three dlstinct wavelengths of interest. Note, although the Earth moves through phases

as viewed from the OPTRANSPAC, the radiant background energy was calculated on

a "full Earth" assumption.

3.9.2 SOLAR BACKGROUND AT OPTRANSPAC - The amount of solar background incident

upon the OPTRANSPAC detectors Is a functlon of the Sun angle geometry, the offaxls solar rejection function and the wavelength of interest. For the

OPTRANSPAC, the Sun angle geometry Is sketched In Figure 3-20. The 90% availability requirement set forth In the statement of work defines the Sun angle

requirement. As is seen in Figure 3-20, the Earth orbit has two places in each

half orbit where the OPTRANSPAC to Sun Angle, _ , is identical. Simple geometric algebra yields the expression of _ as a function of llnk distance, D.

Figure 3-21 shows the minimum Sun angle is given as approximately one degree

off of boreslght.

To determine the amount of solar energy that reaches the detectors, an off-axis

stray light model was defined for the optical system. The off-axls stray light

rejection model is pictured In Figure 3-22 and was taken from previous analysis

performed for a similar telescope design on another Lasercom project. The reJection function, S(_ ), was obtained from a complex modeling of the stray

light rejection properties of anti-reflectlve coatings, aperture, stops, telescope cleanliness levels, and baffle designs for specific telescopes. The assumption made here, wlth high confidence, was extrapolation to similar telescope designs wlth differing aperture sizes and vlewflelds was valid. The

optical background power on the detector was readily obtained from this model.

The effective background radiance at the OPTRANSPAC due to the off-axls solar

scattering is plotted versus link distance. The worst case angle (i.e. smallest off-axis angle) occurs at the maximum range from the Earth (smallest solar

intensity). The two phenomenon, range loss and scatter function, tend to offset each other. As can be seen by the curves in Figure 3-23, the solar scatterlng background flux can be modeled as a constant over all range. To this

end, the background radiant flux was modeled as the constant levels illustrated

in Figure 3-23 for each of the three wavelengths.

3.9.3 SATURN BACKGROUND AT EDRS - The background flux at the EORS platform due

to Saturn was taken from recorded Earth orbit measurements of Saturn's Irradlance [Ref. 2]. Since the EORS characteristics specify a one mlcroradlan viewfield and Saturn subtends approximately 95 Microradlans when viewed from the

Earth, the Saturn background flux must be expressed in terms of a radiance

function (i.e. only a fraction of Saturn, about one ninetieth is viewed by the

EORS). Figure 3-24 tabulates the effective Saturn background available at the

34
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EORS. The radiance function was used directly wlth the field-of-view, aperture

size, optical ftlter bandwidth and optics transmission, to obtain the background optical power on the EORS detector.

3.9.4 SOLAR BACKGROUND AT EORS - Once again the amount of solar background

radiation Is a function of the Sun angle geometry and off-axls solar rejection. The EORS Sun angle geometry (See Figure 3-25) was calculated similar to

the OPTRANSPAC geometry (See Section 3.9.2) except the 90% availability reQuirement encompasses only the "backside" EORS position. The angle y was

fixed at 18 degrees and the off-axls angle, _, was given by the range between

the EORS and the OPTRANSPAC. The off-axls scattering angle versus range for

the EORS receiver boreslght is illustrated in Figure 3-26.

The scatter function, S(0), defined in Figure 3-22, was used in conjunction

wlth the solar radiance measured in the Earth's atmosphere to define the background radiance. Figure 3-2? shows the effective background radiance as a

function of the llnk distance for the EORS. The background radiant flux is

larger at close range and becomes less at long ranges (i.e., the off-axls angle

increases). The llnk margin close in was sufficiently large to offset the Increase In background power. Thus, the links were designed to accommodate maximum range background.

41
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t

4.0 FUNCTIONAL ORGANIZATION

Various system configurations were envisioned as possible solutions to the

OPTRANSPAC design. From these, one configuration was selected as the baseline

design. A detailed system block diagram and functional description was also

generated for the baseline design. A detailed optical layout was then generated based on the configuration selected.

4.1 BASELINE SELECTION RATIONALE - Eight possible system configurations were

consldered for the OPTRANSPAC terminal design. Flve major functions were Identlfled as trade space. The design rationale was to minimize system complexity

and weight. Continuous or static alignment was necessary to maintain transmit/

receive path alignment during OPTRANSPAC lifetime. Polnt-ahead compensation

was necessary to correct for llne-of-slght variations due to platform velocity

aberrations. Earth tracking was considered as necessary to provide track capability when no beacon is present (i.e., the EORS disappears behind the Earth).

Beacon track existed as a trade parameter_or option. Beacon communications was

required by the statement of work.

Figure 4-I shows the baseline configuration selection rationale trade matrix.

Figures 4-2 through 4-9 i]lustrate the optical systems associated with each of

the eight systems configurations. The baseline selection process quickly ellminated options B,C,E,F, and G since no beacon communication was possible.

Option H was ruled out because the wide vlewflelds required severely taxed the

quadrant detector Earth tracker option. Options A and D became the only viable

solutions out of the original eight. Option A was chosen over option D because

it offered a lighter weight and less complex (fewer mechanisms) solution to the

OPTRANSPAC design.

4.2 SYSTEM CONFIGURATION - The top level OPTRANSPAC terminal configuration is

shown in Figure 4-I0. The OPTRANSPAC system at the box level consists of an

Electronics Assembly and Electro-Optlcs Assembly. The Electronics Assembly

consist of a Power Conditioner Unit, a Communications Electronics Assembly and

a Control Electronics Assembly. The Electro-Optlcs Assembly contains a telescope, imaging optics, downllnk laser, Earth tracker head assembly, and beacon

communication detector. Figure 4-11 illustrates the OPTRANSPAC signal flow

from, and to, each of the subassemblies. The functional description of each

subassembly is listed below:

Fixed Mounted Telescope:

- The uncertainty area of the platform is within its vlewfleld.

- Allows diffraction limited transmit operation.

- Collects beacon and Earth radiation and relays it to the imaging optics.

Imaging Optics:

- Provides for dlchrolc separation of transmit and receive paths.

- Allows for co-allgnment of transmit and receive paths.

- Transfers optical energy from laser _6 telescope and from telescope to

approprlate detectors.

- Contains tracking mirrors for vernier tracking function and vlewfleld

scan.

Downlink Laser:

- Generates 532 nm laser radiation in a pulsed format encoded with

information from the communication electronics.

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Earth Tracker HeadAssembly:

- An array detector used for point-ahead and Earth tracking functions.

Beacon Communication Detector:

- An avalanche photodlode used for beacon communication

Communication Electronics:

- Provides for pulse position modulation conversion of input data from

the spacecraft.

- Encodes Input data from spacecraft Into a rate 7/8 Reed-Solomon code

word.

- Decodes rate 7/8 Reed-Solomon data from beacon link.

- Oemodulates the pulse position modulated signal in a "greatest-of"

format.

Control Electronics:

- Provides for Earth track capability.

- Controls all OPTRANSPAC modes of operation.

- Controls alignment function.

- Interfaces with spacecraft attitude reference system and Earth tracker

error signals to provide TMBS track commands.

- Calculates point-ahead angle offsets to be used by the Earth tracker.

- Communicates with spacecraft control processor for time of day and

orbital information.

Power Conditioner Unit:

- Provides all prime power conditioning for the various OPTRANSPAC

components.

- Provides telemetry multlplexlng of monitor signals for telemetry link.

- Controls all OPTRANSPAC redundancy switching

4.3 OPTICAL SYSTEM DESCRIPTION - The system level layout of the optical train

Is given In Figure 4-12. Descriptions of the optical elements employed In the

OPTRANSPAC system are given in Figure 4-13. The optical design consists of a

diffraction limited eleven Inch modified Cassegrain Telescope coupled Into an

imaging optics assembly. The telescope maintains an image quallty of at least

x/20 RMS at the transmit wavelength of 532 nanometers over a ± 2.5 milllradlan transmit vlewfleld.

The imaging optics contains beam steering mirrors for vernier tracking and optical relay elements that transfer the transmit laser energy to the telescope

as well as transfer the received energy to the appropriate detectors. Dichrolc

filters were used to differentiate between the transmit and receive paths which

allowed use of common path optics. Redundant paths around active elements were

used to eliminate single point failures. Redundant switching of active elements, such as lasers and detectors was accomplished through use of solenoid

driven "pop" mirrors. The Earth track optics path offset dlchrolc boreslght

alignment device allowed for Earth track, continuous alignment, and electrical

polnt-ahead functions to be implemented in one optics path design.

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5.0 COMPONENT SELECTION RATIONALE

The rationale behind selection for each major OPTRANSPAC component Is outlined

below. Also, rationale for selection of modulation type and coding is given.

Speclflcatlons for major components (laser, telescope, optics, detector/

preamplifier and Earth tracker) are detailed in Appendices B-l through B-5.

5.1 LASER SOURCES - Laser source selection for the OPTRANSPAC encompassed comparisons of a variety of laser sources. Three types of lasers were considered; gas, solld-state, and semiconductor. Figure 5-I outlines the laser

sources considered as candidates for the OPTRANSPAC terminal. The figure Illustrates the lasers' operational wavelengths and defines their typical operational characteristics. The selection of the downllnk and upllnk laser sources

was made simple by the reduced trade space set forth by the system requirements. It was determined through prellmlnary llnk analysis that the Gallium

Arsenlde sources could not provide enough power to sufficlentiy operate at a

range of lO AU. The Helium Neon source provided low rellabillty in meeting the

lifetime needed. The Carbon Dioxide laser did not fall within the wavelength

requirements. The only viable sources found to close th_ OPTRANSPAC links were

the Nd:YAG and frequency doubled Nd:YAG sources. The doubled Nd:YAG laser was

chosen as the downllnk source as it met the wavelength requirement and the PMT

receiver provided an adequate quantum efficiency at 532 nanometers. The Nd:YAG

source was then selected as the upllnk source. In addltlon, high quantum efflclency APD's enabled good upllnk operation at the Nd:YAG frequency.

5.1.1 DOWNLINK LASER - The downllnk laser on-board the OPTRANSPAC terminal was

specified as a frequency doubled Nd:YAG (532 nm) source in a diode pumped slab

configuration. The diode pumped slab deslgn enables high efficiency laser operatlon effectively reducing the required prime pump drive power. A functional

specification for the downllnk laser is given In Appendix B-l.

The downllnk laser was specified to operate at three distinct pulse rates corresponding to the three distinct data rates. The nominal pulse rates of operation are 14.2B5 KPPS, 4.285 KPPS and 1.429 KPPS. The laser pulsewidth was specified at lO nanoseconds (FWHM) and the average output power was specified to

be 400 millIwatts. An isometric view of an external cavity frequency doubled

diode pumped slab laser is illustrated in Figure 5-2. For high efficiency, the

laser is pumped by laser diode bar arrays matched in wavelength to the absorption band of the Nd:YAG material. The frequency doubling can either be internal or external to the laser cavity. Figure 5-3 outlines the specific design

considerations associated with each device. A specific choice of cavity design

was left as a design choice further into actual hardware implementation.

5.1.2 UPLINK LASER - While not a part of the OPTRANSPAC design, the characterlstlcs of the upllnk laser are governed by the OPTRANSPAC command receiver

implementation. To that end, the upllnk laser shall transmit lO nsec FWHM

pulses at a nominal rate of 142.86 PPS. The transmitted wavelength shall be

1064 ± 0.5 nm. l'he energy per pulse, referenced to the nominal wavelength,

shall be 70 millljoules and the Interpulse period shall be 70 msec + 25.6

wsecond.

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65

FIGURE 5-3
5.2 TELESCOPE SELECTION - To meet the specified requirements outlined in Appendix B-2, a mirror based reflective telescope was chosen over a refractive

system. An ll" aperture refractor would likely require several lens elements

to meet the image quality requirements over the vlewfleld and, and a result,

would be heavy. A refractive design would also be too prone to thermal gradients, which adversely affect transmitted wavefront by the large temperature

dependence of the glass refractive indices

Figure 5-4 illustrates the wide variety of reflective telescopes available.

The image quality requirements over the vlewfleld limit the choice of telescopes. The Newtonlan form was unacceptable over this field due to the coma

inherent In the base parabolold mirror. The Gregorian form was too long and

bulky. The Maksoutoff arrangement tended to be either too long (with a thin

corrector plate) or too heavy (the corrector plate thickness increases as telescope length is reduced to maintain the required correction). The Cassegraln

design form, in one of Its many variants, offered the best approach to meeting

the diffraction limited performance requirements over the vlewfleld with a compact, light weight, stable package.

The "Classical Cassegraln" consists of a parabolold primary mirror with a

hyperbolold secondary mirror. This arrangement can cover a larger vlewfleld

than the single parabolold mirror with substantially better imaging performance. It will, however, not be able to cover the full _2.5 milllradlan field

with the required near diffraction limited performance.

Several variants to the "Classical Cassegraln" offer improved image quality

over the field at a cost In complexity. One variant Is the Ritchey Chretlen

which consist of aspherlc primary and secondary mirrors whose zonal curvatures

are slightly weakened with respect to the parabolold/hyperbolold mirrors of the

Classical Cassegraln. Another variant Is the Schmldt Cassegraln which incorporates a full aperture aspherlzed reflective plate to correct the resulting

spherical aberration. This plate covers the full aperture and is prone to

thermal gradients perturbing the wavefront quality. A final variant Is to Incorporate small refractive correcting elements in the converging beam after the

secondary. This allows additional aberration correction. It was thls final

variant that offered the best approach to meeting the image quality over the

full telescope vlewfleld.

5.3 OPTICAL MECHANISMS - The use of optical mechanisms should be minimized to

provide a design with the fewest possible moving parts. However, gimbals, beam

steerers, and optical mechanisms (select mirrors, shutters, adjustable, field

stops, etc.) are usually needed (in some combination) to accommodate dynamic

system fluctuations and provide for redundancy switching.

lhe OPTRANSPAC system design accommodates pointing and tracking with a single

set of beam steerers. Gimbal pointing was determined as not being necessary.

Many types of beam steering mechanisms were considered; piezoelectric deflectors, torque motor types and acousto-optic deflectors. One and two axis deflector mechanisms were considered. Piezoelectric steerers require high

voltage drivers for the crystal. They also exhibit an inherent mechanical

weakness of the crystal to the mechanical deflection inputs at the structural

resonant frequency. Acousto-optlcal deflectors, although widely used commercially, exhibit optical transmissions of about 70% for a single axis and about

50% for both axes. It was thls loss in optlcal signal power that made AO deflectors undesirable for the OPTRANSPAC design.

66
TELESCOPE CONFIGURATIONS

,fAPERTURE STOP

COLLIMATED BEAM

PARABOLOID

REFLECTOR

I

I

AXIS OF PARABO_.OID--'_I

I

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MIRROR

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I

CONVEX SECONDARY

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l

CONCAVE PRIMARY

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ASPHERIC

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OBSCU_ _/

SECONDARY POINT.

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PRIMARY

SECONDARY _PRIMARY MIRROR

MIRROR FOCUS

/

FOCAL POINT

(b) GREGORIAN

IW

CONVEX SECONDARY CONCAVE PRIMARY

FOCAL POINT

(d) CASSEGRAIN WITH REFRACTIVE ELEMENTS

II

SPHERICAL

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(f) MAKSOUTOFF

FOCAL

POI NT

67

FIGURE 5-4
The OPTRANSPAC design employs two single axis torque motor beam steerers in an

X-Y configuration. The torque motor devices were chosen because of their long

development heritage and low risk design. Space qualified torque motor beam

steerers already exist and are being integrated into a spaceborne optical communlcatton terminal.

Other mechanisms required In the OPTRANSPAC destgn Include a commandable focus

drive to ensure diffraction limited beam quality over the design life and redundancy select flip mirrors that allow selection of redundant optical paths or

electro-optic devices (detectors, laser). The use of solenoid and motor driven

select mirrors for redundancy switching Is a proven method for redundancy management. Redundant dr%yes were associated with each mechanism to ensure rellable switching.

5.4 MODULATION FORMAT S_LECTION - Laser sources cannot be modulated In all

possible modulation formats. Because of thls fact, the choice of a particular

laser affects the selection of an applicable modulatlon type. The selection of

the Nd:YAG lasers and the high peak powers needed to communicate from deep

space, yleld a short pulse, high peak power laser source.

The most efflclent pulsed modulation format to use wlth a narrow pulse laser Is

an M-ary pulse position format. M-ary Pulse Position Modulation (PPM) with

"greatest-of" detection provides the most sensitive receiver in terms of blts

of information per photons transmitted. The PPM format employed in the

OPTRANSPAC system Is illustrated in Figure 5-5. Electronic speeds as well as

hardware maturity drove the design to a 256-ary (8 bits/pulse) pulse position

format. Compatibility wlth eight bit Reed-Solomon encoders and decoders made 8

bit PPM an attractive solution to the OPTRANSPAC design.

5.5 DETECTOR-PREAMPLIFIER SELECTION - The command upllnk communication detector selection involved comparisons of five applicable optical detectors. The

detector types chosen for comparison are listed along wlth their important

characteristics in Figure 5-6. Front end detector gain, quantum efficiency,

internal noise mechanisms (i.e. dark current, ionization coefficient) and rellability are listed for each detector. The selection criteria for the upllnk

communication detector are listed in Figure 5-7. Moderate to high front end

gain is necessary to partly offset the effects of the preamplifier noise contributions; hlgh quantum efficiency Is necessary to provide high efficiency

conversion of optical to electrical energy; low internal noise yields good

receiver sensitivity; and high reliability Is necessary for the long life required. As is seen (Figure 5-7), the silicon APD provides the best fit to the

selected criteria. The low-k device with dimpled design provides good quantum

efficiency and low noise operation. An avalanche galn on the order of 200

helps to reduce the effective preamplifier noise contributions and the proven

reliability of the silicon APD offers the desired lifetime operation.

Low noise transimpedance preamplifiers were interfaced wlth the command upllnk

receiver avalanche photodlode to provide good overall system sensitivity. The

noise current effects of the preamplifier were reduced by the front-end avalanche gain of the APD. The increased gain, however, increases the internal

APD noise as well as the received background noise. An optimum APD gain was

found when the contributions of the preamplifier noise and detector noise (includes background) were equal. Thus, low noise preamplifiers are important in

achieving the greatest possible receiver sensitivity.

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For the OPTRANSPAC system, the selection of an appropriate preamplifier with

low noise occurred from the extrapolation of present measured device values to

the system technology time frame. The preamplifier noise current is bandwidth

dependent and Figure 5-8 shows this relationship for a few low noise devices

manufactured by MERET, Inc. Extrapolation to the 50MHz preamplifier bandwidth

required to pass the received lO nsec pulsewldth resulted In a selection of a

preamplifier with a noise equivalent current of approximately 1.4 plcoamps per

root Hertz. Figure 5-g illustrates the effect on upllnk margin sensitivity

resulting from an increase or decrease In the preamplifier noise current. The

OPTRANSPAC detector and preamplifier design parameters are specified in

Appendix B-4.

5.6 EARTH TRACKER DETECTOR SELECTION - In order to provide track capability by

the OPTRANSPAC such that a narrow divergence beam can be transmitted back to

the Earth, a method of beacon tracking had to be devised. Most narrow divergence optical communication systems cooperatively track each others transmissions. However, the OPTRANSPAC scenario required communication with an EORS

that disappeared every 45 minutes. With round trip transmit times to Saturn of

approximately 2.75 hours, tracking the EORS command beacon was not feasible.

A solution to this problem involved tracking the Earth wlth a quadrant or an

array detector. Quadrant detectors dld not provide the sensitivity (due to _nsufficient background rejection) necessary to track the Earth under the specified availability requirements. An array tracker implementation provided the

best background rejection available due to the small pIxel angular vlewfields.

Random readout and access techniques offered multiple target tracking. This

multiple target tracking enabled the elimination of the polnt-ahead beam steerer group and made the selection of option A (See Section 4.1) possible.

The selection of an array tracking detector enabled the same detector to perform polnt-ahead, boreslght alignment and Earth tracking functions. The pointahead function of the OPTRANSPAC was designed to be an electrical rather than a

mechanical (i.e. beam steerer) compensation. An area of the array detector,

not used for Earth tracking, was used to determine the boreslght of the transmlt signal (See F_gure 5-1O). The artificial polnt-ahead boreslght was commanded by the processor based on a known polnt-ahead angle and with respect to

the transmit boreslght. Alignment of the transmit and receive paths can also

be achieved in thls manner.

The requirements of the OPTRANSPAC on the Earth tracker (track, polnt-ahead,

and alignment) made the choice of a Charge Injection Device (CID) as the Earth

tracker, a good one. Random access (single pIxel addressing) and nondestructive readout capability led to the choice of the CID over other array detectors

such as Charge Coupled Devices (CCD). CID's, where readout does not involve

destruction of the stored charge, have increased sensitivity over CCD's.

The Earth tracker design was based on developed CID technology. The CID consist of an array of photosensitive picture elements (plxels) which integrate

incident signal charge, store it, and enable readout at the individual plxel

site. The OPTRANSPAC Earth tracker was designed around a ST-256-CID developed

by General Electric. The physical size of the array was 5.12mm x 5.12mm and

encompasses 256 x 256 pixel. The l mrad by l mrad OPTRANSPAC field of view was

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mapped into a 200 x 200 plxel area providing excellent background rejection due

to small (5 mlcroradlan) plxel vlewflelds. A specification for the Earth

tracker is given In Appendix B-5.

5.7 CODING SELECTION - Both the upllnk and downllnk were selected as candidates for channel coding. Coding of these links provides for a reduction in

aperture size and laser power at a small cost in weight and complexity. Using

a block code over the pulse position modulation frame provides the largest

degree of error correction [Ref 3]. The selection of the Reed-Solomon block

code for the OPTRANSPAC links was based primarily on the maturity of the technology, particularly the 8-blt Reed-Solomon encoders and decoders.

lhe spacecraft or command data is encoded into B-blt Reed-Solomon channel

symbols by a rate 7/8 Reed-Solomon encoder. These B-blt channel symbols are

further encoded into PPM symbols by appropriate pulse timing into one of 256

slots. At the receiver the PPM symbols are decoded in a "greatest-of" fashion

selecting the slot with the greatest pulse amplitude out of the 256 possible

slots. The PPM symbols are then converted back to channel symbols for the

Reed-Solomon decoder, which in turn converts the channel symbols to source bits.

Figure 5-11 shows the relationship between the uncoded PPM bit error probabillty and the Reed-Solomon encoded bit error probab_llty It can be seen that to achieve a final bit error probability of lO-3, the PP,7 channel must support a bit error probability of approximately 0.02 for a rate 7/8 Reed-Solomon

code. Further coding gain is possible with higher order codes (i.e., rate 3/4,

rate I/2) but the increase in laser pulse, rate required to accommodate the

higher order codes reduces the available peak laser power per pulse. This effectively offsets a portion of the coding gain and from Figure 5-12 it is

readily seen by the upllnk example that the rate 7/8 code provides the largest

available coding gain.

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6.0 LINK ANALYSIS

The analysis that supports the validity of the OPTRANSPAC system operation is

presented below. Because the two are interrelated, the system links from an

acquisition and tracking standpoint as well as from a communication standpoint

were analyzed.

6.1 ACQUISITION AND TRACKING - System concerns when considering the acquisition and tracking performance of the OPTRANSPAC include spacecraft dynamic

characteristics, spacecraft uncertainty, tracking detector accuracy, boreslght

uncertainty and beam steerer noise to name a few. The acquisition and tracking

system must compensate for some of these concerns and operate through others.

6.1.1 SPACECRAFT CHARACTERISTICS - The following spacecraft characteristics

were assumed in order to estimate the spacecraft attitude dynamics for

OPTRANSPAC inputs [Ref. 4]. A spacecraft whose mission is close scrutiny, deep

space, has a mass of between lO00 and 3000 Kg. By choosing the smaller mass, a

more conservative estimate of attitude motion was assumed. It was assumed a

three axis stablllzed attitude control system was needed to perform precision

pointing and rapid retargetlng tasks. The spacecraft was also assumed to employ an articulated science platform with a mass of about 80 Kg for tracking a

target during close scrutiny. The respective control bandwldths of the spacecraft and the platform were assumed l and O.l Hertz. Control accuracy of the

spacecraft was specified as 2 mllllradlans and the control stability of the

platform was assumed to be 550 mlcroradlans. During target tracking the

science platform slews at 1.2 mllllradlans per second and slews at a maximum

rate of 35 mllllradlans/second during retargetlng. Figure 6-1 outlines the

spacecraft assumptions used.

6.1.2 ACQUISITION - The assumptions made during OPTRANSPAC acquisition of the

Earth are outlined in Figure 6-2. The figure lists the driving assumptions for

acquisition scanning and acquisition time. The spacecraft attitude control

system, with a 2 milllradlan accuracy, places the Earth inside the telescope

fleld-of-vlew. A one mllllradlan square vlewfleld detector provides angle

error updates at a 200/second rate. The false alarm rate was assumed to be

once in two seconds and the computer switching time is O.l second. The beam

steerer bandwidth required for acquisition was calculated to be 20Hz.

In order to cover the entire uncertainty area of the telescope vlewfleld, the

detector array was assumed scanned in the pattern shown in Figure 6-3. This

type of pattern, a raster scan, was chosen over a conical scan because of ease

of implementation and no need for the most rapid average acquisition time posslble. In order to assure no gaps are left in the scan coverage, an overlap of

around 30 percent is usually allowed. For the telescope fleld-of-vlew and detector array size noted in Figure 6-2, a pattern of six swaths was determined

necessary. This yielded an overlap factor of about 29 percent. Two alternatives existed for the sequence in which swaths were searched. In the first,

the scan covered the inside two swaths first, as they were most likely to contain the target, and then proceeded outward to the next most likely swaths. A

disadvantage of this general sequence in some laser communication systems was

the separation between sequential swaths, which allowed more time for attitude

drift between adjacent swaths. This was true of this application, where spacecraft attitude drift rates, on the order of 350 mlcroradlans per second, could

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cause the target to move from one swath to the adjacent one before the detector

array arrives at the original target position. In the other alternative for

the search sequence of the swaths, the first swath was on one edge of the uncertainty region, and the next sequential swath was the adjacent one. This sequence did not search the highest probability areas first, so the average acquisitlon time was higher than the first alternative. However, since the adjacent swaths were searched sequentially, the attltude drift rate will not carry

the target out of the swath before the detector array arrives at that posltlon. Since acquisition time was not pressing In this application, this type

of pattern and sequence was chosen. Given the above scan description, the acquisition time was calculated using the rationale outlined in Figure 6-4.

6.1.3 TRACKING - Tracking error contributors are separated into two categorles, random and systematic. Random errors (or Jitter terms) are assumed to be

higher frequency and uncorrelated, so a root sum square gives the total expected error from these sources. Systematic errors (or bias terms) are assumed

to be low frequency, and may be correlated, so the maximums may all occur simultaneously over a long period of time. The systematic errors must therefore

be summed together. The total error budget for the tracking system Is the sum

of the 3 sigma random errors (RSS'ed together) and the systematic errors.

Figure 6-5 indicates the tracking error contributions to the transmit beam

pointing. While this represents a budget for each particular entry, the

numbers were determined considering device limitations. The error magnitude

determinations are evident from the comments, except for the last two. Spacecraft dynamics and spacecraft attitude are further explained below.

Three sources of spacecraft dynamic tracking error were considered (See Figure

6-6): Spacecraft maneuvers, with an attitude control bandwidth of l Hz; spacecraft vibration from control moment gyros or other machinery; and science platform slewing at maximum rate for retargetlng. For a detector array update rate

of 200/second, a tracking loop bandwidth of at least 20 Hz can be supported.

Spacecraft maneuvers wlth a l Hz loop will therefore cause no significant

errors. Vibration is expected to cause the science platform stability accuracy

error quoted earlier. Since the science platform has a bandwidth of O.l Hz,

and the OPTRANSPAC has a bandwidth of 20 Hz, the expected OPTRANSPAC error due

to vibration was calculated to be less than 0.02 microradlans. The largest

contributor to tracking error is expected to be the science platform retarget

slewing. The platform will Jump to maximum rate as fast as possible to minimize retargetlng time. This rate wlll be transmitted to the spacecraft by the

ratio of the platform and spacecraft Inertlas, which are roughly proportional

to the square of their masses. A step in attitude rate of that size, into a 20

Hz type II tracking loop wlth the beam steerers, causes a tracking error of 0.3

mlcroradlans. This value was inserted on the tracking budget chart.

Figure 6-7 depicts a three dimensional layout of the polnt-ahead angle necessary between the OPTRANSPAC and the target (Earth Relay Station). The tracking

llne Is the llne between and perpendicular to the reference plane of the

OPTRANSPAC and the reference plane of the target. The point ahead angle is the

angle away from this tracking llne at which light has to be sent in order to

compensate for relative velocities between the two platforms. This angle

should lle within the plane defined by the tracking llne and the velocity

vector. The creation of this angle by the OPTRANSPAC may not be perfect,

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because the spacecraft may have an attitude error around the axis of the tracktng line of up to 2 mrad. The exploded vlew In Figure 6-7 shows the effect of

thls attitude error projected by the point-ahead angle. The tracking error due

to this effect is nearly 900 nanoradtans. Thls error was too large to allow In

the tracking budget.

The error can be reduced by compensating for the rotational error of the spacecraft around the tracking llne wtth the point ahead, if the attitude of the

spacecraft is known. Assuming that the attitude knowledge of the spacecraft ls

0.1 times the control accuracy, the point-ahead induced pointing error can be

reduced to less than 0.1 mlcroradian. If the spacecraft attitude knowledge Is

worse than the above assumption, tt would still be possible to reduce the

tracking error contribution to about the same value by implementing a star

tracking function around the tracking line wlth the detector array, and essentially determining the OPTRANSPAC's inertial attitude Independently of the

spacecraft.

6.1.4 EARTH TRACKING ANALYSIS - Figure 6-8 shows the relationship between

Earth tracker angle error noise and aperture slze at various ranges. The

tracking budget amount is also indicated. The Earth tracker error becomes a

function of the distance from the Earth, the optical resolution and the centroldlng algorithm.

The range between the OPTRANSPAC and the Earth determlnes the amount of signal

energy incident upon the array detector. The solar energy reflected from the

Earth was assumed to be Lambertlan distributed in intensity. The factor associated with the partial Earth phase was assumed to be 5%. That is, only 5% of

the energy reflected by a full Earth Is reflected by the partial Earth. The

range also plays an important role In determining the spot slze on the detector. The physical angular subtense of the Earth varies from 85 mlcroradlans at

l AU to 8.5 mlcroradlans at lO AU. The more Earth tracker pixels wlth 5 microradlan fields of view required to receive the incident energy, the more noise

incurred in the angle error measurements (i.e. the more signal required to

maintain an acceptable signal to noise ratio).

The actual spot slze on the detector array is a function of the optical resolution of the system. Figure 6-9 shows the relationship between the blur circle

on the detector versus range for various receive path optical resolutions.

Analysis determined a comfortable 30 mlcroradlan resolution would be sufficient

for the receive tracking path part of the optical system.

The centroldlng algorithm is the final measure of the angle error accuracy.

The algorithm used must be able to accurately centrold the energy received even

under Earth partial phase conditions. Knowledge of orbital position is imperative In the Earth tracker design, The actual centrolding algorithms used are

beyond the scope of this report, rather the analysis was based on star tracker

designs employing CID technology.

It is widely accepted [Ref. 5] that for a uniform circular spot, the slope at

the origin is a function of the spot angular slze and is given by

SF = 2.55 / Ospot

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the detector noise terms specified in Appendix B-5. The received signal Is

composed of a complex function involving non-destructive readouts and additional averaging t4mes. These techniques are documented tn [Ref. 6].

6.2 COMMUNICATION LINKS -Wlth component values identified In Section 5.0 and

background levels determined in Section 3.9, the final sizing of the OPTRANSPAC

links involves definition and trades of laser powers and aperture sizes. The

design ratlonale Involving ilnk margin analysis ls outlined in Figure 6-10.

The environment and pointing control as well as communication llnk parameters

must be traded to obtain an overall acceptable ltnk design with confidence. By

the tlme the llnks are finalized, many iterations have occurred Involvlng the

control and communications aspects. Only the flnal links are given. The intermediate iterations are not shown, but their importance In optimizing the

design should be noted.

6.2.1 TRANSMIT GAIN_LPOINTIflG LOSS RELATIONSHIP - The transmit galn intensity

of a Gaussian fed Cassegraln telescope falls off tn magnitude as a function of

off-axls angle. Zero pointing error results tn maximum on-axis gain. However,

perfect pointing ls unachlevable wtth most real platforms and thts results in a

loss (from maximum on-axis gain) associated with the magnitude of the pointing

error. This loss Is referred to as pointing loss. Figure 6-11 illustrates the

resulting transmit/gain pointing loss relationship for various diffraction

limited transmit apertures and pointing errors. Note that large diffraction

limited apertures cannot be supported by large pointing errors. Thus, minimum

pointing errors become a deslgn driver.

The tracking budget developed tn Section 6.1.3 indicates a maximum tracking

error' of 1 mlcroradlan. From the curves tn Figure 6-11, It Is evident the

maximum diffraction limited aperture size is approximately sixteen inches. The

minimum aperture slze wtll be set by one of two criteria; the transmlt gain

required on the downllnk or the receiver galn required for the upllnk. From a

system point of view, the aperture stze was determined by optimizing OPTRANSPAC

weight and power. Figure 6-12 shows the OPTRANSPAC telescope weight versus

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prime power delta (from 400 mW average transmit power) needed to close the

downllnk with 3 dB of margin. The system design point is indicated. Weight

and power indicate a baseline design of ll5 pounds and 57 watts. The minimum

system aperture size was limited to eleven inches by the upllnk command beacon

margin. As can be seen, an increase of 12 pounds (14" aperture) would yield a

decrease in required prime power, but this amount is minimal when high efficiency lasers are used. The baseline design point appears on the bend in the

curve and any increase or decrease in aperture size causes either a large

weight increase for small power gain or large power need for small weight reduction. Weight was assumed most critical in meeting the overall OPTRANSPAC

requirement, and thus, an eleven inch OPTRANSPAC aperture was chosen as basellne.

6.2.2 DOWNLINK MARGIN - The OPTRANSPAC downllnk consists of a frequency

doubled Nd:YAG laser modulated in a PPM format at one of three distinct rates

(lO0 KBPS, 30 KBPS, and lO KBPS). The transmit source radiates 400 mlllIwatts

of 532 nanometer average power into a 3.4 mlcroradlan beam. The maximum range

between the two platforms is lO AU. The EORS detector is a photomultlpller

with a 30 percent quantum efficiency at 532 nanometers. The receiver vlewfleld

is one mlcroradlan and the worst case background is the off-axls scattering of

solar energy into the telescope. The sensitivity of the receiver was calculated to be o.g nanowatts, peak (at lO-3 BER, RS coded). The llnk is given

in Figure 6-13 with a design margin of 3dB.

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The downllnk margin versus aperture size relationship is given in Figure 6-14

for various ranges and data rates. The increase in margin at the lower data

rates can be used to decrease the system bit error qrobabillty. Figure 6-15

indicates that bit error probabilities less than lO-_ can be achieved at the

lower data rates with llnk margins in excess of 3dB.

Extrapolation of the downllnk data rate to 300 KBPS requires modification of

the transceiver design. If the same 8 bit per pulse PPM format is used, laser

power or aperture size or both would need to be increased to accommodate this

change. Increased prime power, added weight and possibly a new laser design

may be required. If the PPM format is changed (more bits per pulse), a new

Reed-Solomon encoder/decoder deslgnand a new PPM electronics design would have

to be required. Added weight and prime power may also be required. Overall, a

modification to the present design by increasing the data rate will ripple

through the entire system design. "

6.2.3 UPLINK MARGIN - The upllnk command beacon consists of a Nd:YAG Laser

modulated in a PPM format at IKBPS. The transmit source radiates lO watts of

average power into a 5 uradlan beam. The maximum range between the two termInals is lO AU. The OPTRANSPAC receive aperture is II inches in diameter and

has a 22% obscuratlon. The vlewfleld is l mllllradlan and is driven by the

Earth tracker and alignment requirements.

The beacon communication detector is an avalanche photodlode which, under the

worst case background conditions (i.e., solar light scattered into the aperture) and associated preamplifier noise, provldes a sensitivity of 2.6 nanowatts, peak (at lO-3 BER, RS coded). Figure 6-16 shows the OPTRANSPAC upllnk

sized for 3dB design margin. Figures 6-17 and 6-18 illustrate the effect of

receive aperture size on upllnk margin and the effect of llnk range on upllnk

margin. The OPTRANSPAC is limited in aperture size to II inches, minimum, by

g5
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96

FIGURE 6-13
=_

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_= 9P 'NI9_VW

97

FIGURE 6-14
DOWNLINK BIT ERROR RATE

VERSUS SYSTEM MARGIN

10--1

10-2

10-3

10_4

---. 10-5

_ 10_6

10--7

10-8

I

I

1 I

2

100 Kbpz 30 Kbl_

I

I i _ I !

6 8 10 12

DOWNLINK MARGIN (dB)

SPECI FI ED

BIT ERROR RATE

I EFFECTIVE DATA

RATES

BIT ERROR RATES

IN EXCESS OF 10-7

ARE POSSIBLE AT

THE TWO DISTINCT

LOWER DATA RATES

SPECI FLED, WITH

3 dB OF MARGIN

I I I

14 16 18

FIGURE 6-15

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100
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FIGURE 6-18
the command uplink. The maximum range offered to maintain the 3dB design

margin is 10 AU. Command link operation out to approximately 14 AU may be

possible with no design margin.

The specified command llnk bit error probability may require a decrease from

lO-3 to lO-6 under certain command upllnk conditions. Figure 6-19 shows

the relationship between error probability and upllnk margin for the OPTRANSPAC

system. At all ranges less than 9.3 AU, this increase In bit error probability

Is offset by added margin. However, at ranges greater than 9.3 AU, the margin

for a lO-B command upllnk is less than 3dB, becoming 2.4 dB at lO AU. Thus,

the maximum cost of maintaining a lO-6 command upllnk is approximately O.6dB

at IKBPS.

102
UPLINK BIT ERROR RATE

VERSUS SYSTEM MARGIN

=

0

mm_

m0

mmz

m

.J

rE

10-1

10-2

10-3

10 -4

10-5 -

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SYSTEM UPLINK MARGIN (dB)

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FROM 10-3 IS APPROXIMATELY 0.6 dB

1 Kbps EFFECTIVE DATA RATE

FIGURE 6-19

103
T.O ELECTRONICS DESIGN

The electronics as configured consist of three assemblies which provide control, communication and power conditioning functions. A digital computer,

Input/output circuitry, and drive circuits within the Control Electronics Assembly provide the acquisition and tracking functions. The communication functions of codlng/decodlng and modulatlng/demodulatlng are provided by circuitry

w1thln the Communication Electronics Assembly. The Power Conditioner Unlt converts the spacecraft power to the required secondary voltage levels, provides

redundancy switching, mechanism and heater control, as well as command and

telemetry interfaces.

7.1 CONTROL ELECTRONICS ASSEMBLY - The Control Electronics functional block

diagram Is depicted in Figure 7-I. Command signals to position the torque

motor beam steerers (TMBS's) to direct the incoming and transmitted optical

signals is the principle output of the Control Electronics dlgltal processor.

These commands are based on orbital ephemeris data, spacecraft position sensor

data and Earth tracker angle error calculations. These outputs are provided at

a 200 Hz rate to support the 20 Hz TMBS loop bandwidth and become the drivers

for the operations rate of the control processor. Command signals for posltlonlng the command focus mechanism are initiated by ground command. Mode control of the OPTRANSPAC operation is provided by the digital processor or via

ground commands. Figure 7-2 describes in general terms the processor operations in each operating mode. Further description of the control electronics

assembly including TMBS position commands and mode logic are given In Appendix

C-l; a design note on the OPTRANSPAC electronics.

7.2 COMMUNICATION ELECTRONICS ASSEMBLY - The functional block diagram of the

Communication Electronics Assembly is depicted in Figure 7-3. Both receive and

transmit functions are required. The receive function consists of condltlonIng, decoding and formatting of the data pulses from the communlcatlon detector. The transmit function involves data formatting, encoding and generation

of the required laser modulator drive signal. Further details of the Communlcation Electronics can be found in the design note in Appendix C-l.

7.2.1 RECEIVE FUNCTION - The output of the communication detector preamplifier

consists of the electrical signal amplified and applied to a PPM decoder that

operates on a "greatest-of" principle to select the slot containing the signal

pulse. The output of the decoder is an NRZ data stream containing rate 7/8

Reed-Solomon encoded data at I142.B6 symbols per second. This data Is applied

to a Reed-Solomon decoder which produces a lO00 bit per second output. Finally, the lO00 bit per second data Is formatted as spacecraft command data. The

output of the formatter is applied to a command decoder located within the

Power Conditioning Unit.

7.2.2 TRANSMIT FUNCTION - Transmit data consists of a fixed rate serial telemetry data stream From a telemetry formatter located In the Power Conditioner

Unit or a serial data stream from the spacecraft sensor. Spacecraft sensor

data rates can be lO,O00 bits per second, 30,000 bits per second or lO0,O00

bits per second. This data is assumed to contain overhead including synchronlzatlon information. When spacecraft sensor data Is available for transmission,

the telemetry data is not transmitted. The output of the data selector is rate

7/8 Reed-Solomon encoded and output to a PPM encoder. The output data from the

Reed-Solomon encoder is applied to a PPM encoder that produces output symbol

104
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PROC.

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NO

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I OPERATIONS I

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FIGURE 7-2

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107
information at 1428.57, 4285.71 or 14285.71 symbols per second depending upon

the sensor data rate. Symbols are encoded by outputting a pulse in one of 256

data slots following a fixed time Interval that is referenced to the previous

fixed time interval using an Internal clock. The output of the PPM encoder is

input to modulator drive circuitry. The circuitry converts the logic level input to signal levels required to operate the laser modulator located within the

laser assembly.

7.2.3 DOPPLER CONSIDERATIONS - Variations in timing due to Doppler has been

analyzed. Maximum velocities encountered are defined In Section 3.2. The temporal shift is dependent upon the instantaneous pulse rate. Because Pulse

Position Modulation involves a variable Interpulse period, a minimum and maximum temporal shift is associated with each specific data rate. For the specified slot widths at the receiver of lO0 nanoseconds, the clock center frequency

is lO MHz. The Doppler shift associated with the specified data rates is calculated to be 2400 Hz. This equates to 240 parts per million which can be

readily covered by a voltage controlled crystal oscillator driven by ephemeris

data.

7.3 POWER CONDITIONING UNIT - Primary power from the spacecraft°s electrical

system is conditioned to provide secondary power for the OPTRANSPAC equipment.

The power conditioning unit pictured functionally in Figure 7-4, provides the

conversion and regulator circuitry required to produce the redundant secondary

voltage outputs. Command and telemetry circuitry are also located within the

power conditioning unit. Serial command data from the spacecraft's command decoder or the Communications Electronics Assembly is decoded. The output from

this circuitry is provided to discrete drivers which control the OPTRANSPAC redundancy switching mechanisms, the active focus mechanism, or power control

relays. Serial magnitude commands are relayed to the Control Electronics Assembly via bl-dlrectlonal serial interface circuitry located within the PCU.

This serial interface receives telemetry and heater control data from the Control Electronics Assembly. The telemetry from the Control Electronics Assembly

as well as internal analog and discrete data is multiplexed to form a serial

data output. This data is transmitted to the EORS via the Communication Electronics Assembly. Circuitry required to condition internal analog and discrete

telemetry parameters is provided within the PCU. OPTRANSPAC heater drive circuitry is also located internal to this assembly. Further description of the

Power Conditioning Unit is found in Appendix C-l.

7.4 REDUNDANCY IMPLEMENTATION - Figure 7-5 shows the OPTRANSPAC redundancy

philosophy. All OPTRANSPAC electronics and mechanisms are redundant with the

exception of the command focus mechanism. Cross-strapplng for selected functional circuit blocks is provided to minimize the probability of system failure.

The elements of the control function and the cross-straps provided are described in Figure 7-6. All elements of the function are cross-strapped to other

redundant elements with the exception of the beam steerers. Experience has

proven that cross-strapplng of the drivers and beam steerers does not provide

significant enhancement of system reliability since the reliability of the

drivers and beam steerers is high compared to other elements. In addition,

this makes the design of the beam steerer drivers much simpler.

108
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The elements of the communication function and the cross-strapplng provided are

detailed In Figure 7-7. Cross-straps are not provided between the communication avalanche photodlode and the preamplifier due to the low signal levels and

noise sensitivity at this point. Similarly, design conslderatlons are the

principle reason for not cross-strapplng the Communication Electronics to the

laser assembly.

Figure 7-8 details the elements and cross-strapplng of the power conditioning

function. Implementation of these cross-straps ts accomplished by use of relays controlled via EORS generated commands. Fuses are provided at the input

to each converter to prevent load or converter failure propagation to the redundant converter.

7.5 EiECTRONICS PACKAGING - The OPTRANSPAC electronics are housed in packaged

assemblies that are designed for minimal size, weight and power and provide

thermal radiation through coverplates. A detailed breakout of the size, weight

and power Is given in the Electronics Design Note in Appendix C-l.

The Control Electronics design is based on 14 ceramic and 2 polylmide circuit

boards. Digital circuits In leadless chip carrier packages will be mounted on

the ceramic circuit boards. The polyimide boards will be populated with analog

integrated circuits and discrete components. D1menslons of all circuit boards

are approximately 4.6" by 3.8" Both prime and redundant circuitry will be Implemented on the 16 circuit boards.

A total of 12 circuit boards are estimated for the redundant communication

electronics. A pair of polylmlde boards will be used for the analog integrated

circuits and discrete components required. All other circuit boards will be

populated with leadless chip carrier digital integrated circuit packages.

The circuitry required for the converters, regulators and relays will be implemented using a module concept. Each module will consist of two (prime and redundant) polylmlde circuit boards housed in a frame. The dimensions of each

module will be approximately 6" by 6" by 0.3". A total of 5 modules is required.

112
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114
8.0 OPTICAL DESIGN

The optical system consists of two basic subassemblies: A telescope and an

imaging optics assembly (IOA). The system performance requirements of each

subassembly are outlined in Appendlces B-2 and B-3. an optical system design

was developed from the functional requirements through the thin lens stage.

The details of the optical analysis can be found In the optical system design

note - Appendix C-2. An optical isometric showing the lens layouts is given in

Figure 8-I. An orthographic of the optical system Is pictured in Figure 8-2.

In the design of the telescope and imaging optics for a spaceborne system, the

size and number of optical elements used are kept to a minimum providing necessary functions within the required system performance levels. System optical

transmission Is usually maximized while system size, weight and procurement

costs are minimized. Diffraction limited optical systems such as OPTRANSPAC

require the use of multiple element lens components with moderately slow optical speeds to meet image quality requirements. Field of view and spectral

range requirements further add to the design complexity. These requirements

exact design constraints on optics transmission, optical system envelope and

optical system weight.

8.1 TELESCOPE DESIGN - The telescope selected for the OPTRANSPAC system Is a

modified Cassegraln with refractive correctlng elements in the converging beam

after the secondary. The selection rationale is outlined in section 5.2. The

telescope design is shown in Figure 8-3. Geometric ray tracing of this design

both on-axis and #t full field confirms that this design fulfills the image

quality requirements over the ±2.5 milllradlan vlewfleld. Diffraction analysis

indicates this telescope's wavefront quality Is approximately x/30 RMS at a

wavelength of 532 nanometers. The telescope aperture diameter is eleven inches

(297.4 mm) with a central obscuratlon of twenty-two percent. This section of

the optical system is designed to be afocal and maps the 0.5 Inch diameter

axial bundle of the imaging optics paths to the eleven inch primary mirror

diameter.

The Cassegraln telescope optics are designed to produce an f/lO output cone.

This allows for a component package whose optical speed is still s_low enough to

ensure good image quality and _readily manufacturable Components. Further telescope ahalysts can be found in Appendix C-2.

8.2 IMAGING OPTICS DESIGN - For clarification the _tmagtng optics paths are

subdivided into four areas: the common path (prime and redundant), the transmtt path, the receive cornmuntca_:tons path, and the receive _;racktng path.

In the common path and throughout the imaging optics it was desirable to use

only a few basic lens groupings In a repetitive way to simplify the design and

procurement of the optics. Such repetitive lens groupings must relay the telescope pupll through the imaging optfcs with minimal vtgnetttng. This requires

that sections of collimated space used for locating beam steerers or beamsplitters be rather short to minimize the need for large (> 50 mm diameter)

optics. Further details and ray traces of the common path are given in

Appendix C-2.

The optics in the transmit path are handling monochromatic (532 nanometer)

light which is of small divergence and remains on-axls (all steering of the

115

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118
CASSEGRAIN WITH REFRACTIVE CORRECTING ELEMENTS

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FIGURE 8-3
transmit beam is accomplished In the common optics path). The prime and redundant lasers are selected by a "pop mirror" mechanism. The output laser light

passes through an afocal relay. One of the components of thls relay can be

shlfted axially to introduce specific amounts of focus error to partially compensate for wavefront errors introduced by component shifts. These component

alignment shifts can arise from the forces during launch, from stress relief of

structural parts over time, or from long term temperature changes during the

course of the mission. (In a diffraction limited system even such minor

motions can have a serious effect on performance). Upon exiting the focus

mechanism, the beam enters an afocal beam expander to increase the beam In dlameter to 12.7 mm for entry into the common optics path. Specific details of

the transmit optics path are found In Appendix C-2.

The receive communication path consists of all elements from the recelve/transmit dlchrolc beamsplltter through to the communlcatlons detectors. The received I064 nanometer light from the EORS is received by the telescope and relayed through the common path optics. The received energy is reflected at the

recelve/transmlt beamsplltter and relayed to the communlcatlons/tracklng dlchrolc beamsplitter where the I064 nm energy is reflected and filtered by a 25

Angstrom bandpass filter and focussed onto a communication detector (avalanche

photodlode). The avalanche photodlodes are operated in a slightly defocussed

position to keep the spot size large enough that the variation In detector responslvlty over its surface has a negligible effect on communication llnk

margin. Further design of the communication optics path details are given in

Appendix C-2.

To accommodate the tracking and polnt-ahead functions on the same array detector an offset is introduced between the broadband reflected sunlight from the

Earth and the 532 nanometer transmit spot retroreflected from the transmit

beam. A dlchrolc beamsplltter separates the broadband and 532 nm light Into

two separate channels. A wedge prism deviates the 532 nm light prior to enterIng the tracking detector lens. The broadband light enters the tracking detector lens undevlated yielding the angular offset required. The tracking detector lens axis Is decentered with respect to the input receiver axis to accommodate the two spatially separated Earth track and polnt-ahead track channels.

Details of the receive track channel are given in Appendix C-2.

8.3 OPTICAL TRANSMISSION - A summary of overall optics path transmission is

given in Figure 8-4. The transmission estimates for beginning of llfe and end

of llfe are listed. The individual component performance estimates given are

readily manufacturable specification values and are based on the measured performance of optical components built and used In Lasercom systems over the past

eight years. The component substrates and coating materials are restricted to

materials with demonstrated radiation resistance. Transmissions are based on

optical surface counts for each path. Specific transmission values for lenses,

mirrors, filters and dlchrolcs are multiplied to yield the values in Figure

8-4. Monochromatic performance numbers for the transmit and receive communication paths are based on measured values from previous systems. The coatings

are optimized for the monochromatic paths.

The antlreflectlon coating performance will be substantially poorer over the

broadband visible and near infrared band of the reflected sunlight. The same

will be true over the blue-green end of the spectrum for the hlgh reflectance

coatings on telescope primary and secondary minors. The overall transmission

120
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FIGURE 8-4

121

|

=

II

=
for the tracking receive path will be substantially less because of this.

Detailed transmission estimates on an element basis Is glven In Appendix C-2.

8.4 WAVEFRONT ERROR AND IMAGE QUALITY - In order to meet the image quality

requirements at lO AU, a near diffraction limited wavefront needs to be radiated from the transmitter. Both the telescope and imaging optics are required

to maintain X/20 RMS wavefront quality at 532 nanometers. Thls equates to a

total system performance of x/14 RMS. Analysis indicates surface quality of

X/f04 RMS (X/20 - X/30 peak to valley) per element is required. This Is

within the state of the art for today's optical fabrication technology but Is

considerable more labor intensive than for the more commonly encountered X/4

or X/8 peak to valley surfaces. Other sources of wavefront error including

glass Inhomogenelty, thermal gradients, mounting stresses, mlsallgnments and

residual uncorrected design aberrations are not included In the above design.

Further discussion of transmit wavefront error is given In Appendix C-2.

The image requirements are less stringent for the receive path than for the

transmit path. Receive image quality is driven by the tracking requirements.

Figure B-5 illustrates the variation in the size of the Earth image on the

tracking detector as a function of range. Diffraction limited image quality Is

not necessary to meet the required tracking uncertainty. An optics resolution

in the 15-25 mlcroradlan range is acceptable.

122
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FIGURE 8-5
9.0 MECHANICAL DESIGN

The OPTRANSPAC terminal design is the result of various trades which culminated

In the final baseline mechanical design. Maximum use of existing space qualified optical communication components and packaging techniques were employed in

the final design. Weight and power guidelines were imposed to obtain the optimum conflguratlon. The envelope slze was not considered a large design driver

because a specific host platform was not known.

The OPTRANSPAC terminal is pictured isometrically In Figure 9-I. An orthographic vlew Is given in Figure 9-2. A single baseplate concept was employed

which provides interface mounting surfaces between the host platform, the optical assembly and the electronics. The plate is thermally isolated from the

spacecraft to maintain and adiabatic interface. Three lap pads are provided to

allow a strain free attachment to the spacecraft. The plate Is composed of two

Beryllium facesheets with an aluminum honeycomb core. The strength to weight

ratio is optimized with thls baseplate concept and the Beryllium facesheets

provide a good optical bench. For the telescope mount, a truss structure has

been designed to provide a strain free, three polnt mount to the baseplate.

The cover has been designed for minimum weight and size and does not obstruct

the laser's view of space for a radiator to reject heat.

The optical elements are all supported by the single baseplate. The optics

path includes two fllp mirrors which provide a totally redundant optics path.

An optical isometric Is pictured in Figure 9-3. The baseplate supports optical

elements from both faces, and openings in the baseplate are provided to relay

the optical signals from the top side to the bottom slde of the optical baseplate.

The electronics have been divided into three boxes; the Power Conditioning

Unit, the Control Electronics Assembly, and the Communications Electronics Assembly. The Power Conditioning Unit was located at the bottom of the baseplate

with the laser assembly so It can reject heat Into space. Figures 9-4 and 9-5

depict isometric and orthographic views of the PCU. The PCU has been deslgned

using the module concept where appropriate. These modules are attached to a

thermally conducting top housing plate with the components mounted in such a

manner as to place the hottest closest to the plate.

The Communication Electronics Assembly (see Figures 9-6 and 9-7) and the Control Electronics Assembly (see Figures 9-8 and 9-9) are structured using the

doubled sided card concept. Thls allows tight packaging into reduced volumes.

The boxes provide slide mounts wlth end connectors for each of the electronics

cards. The heat is dissipated through the top face plate surface away from the

optical baseplate.

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FIGURE 9-9
lO.O WEIGHT AND POWER

The OPTRANSPAC final weight and power summary is presented In Figure lO-l.

These results of llS.0 pounds and 5?.0 Watts are broken down by the major components defined by the functional block diagram. The weight allocations were

optimized within the constraints of the functional requirements, materials,

environments and cost. Whenever possible these allocations were also modified

to reflect 1988 technology to further ensure the achievement of a minimum

weight design.

The power allocations for the OPTRANSPAC system were established by reviewing

each piece of equipment to determine It's power requirements. This data was

compared to similar hardware program equipment, and/or data taken from breadboard circuits, in order to provide substantiation. This analysis also allowed

for a 3 percent llne loss and conversion efflclencies of 68 percent. An additional lO percent was allocated for heater power. The electrical power design

provides the user with both flexibility and growth.

The OPTRANSPAC component weight summary is presented In Figures I0-2 and I0-3.

All of the weights presented are nominal weights. Nominal weight is the most

probable weight and it is composed of a base weight plus contingency. Contingency is defined as that weight which must be added to a base weight to obtain

a nominal weight. Contingency includes such things as an allowance for gage

tolerances in sheet metal or machined parts, weight from correlation factors,

or an assigned weight based on the type of subsystem and Its current development state. In general the base weights for this study were derived using one

of the following methods: analytical models, empirical equations, detailed

component analysis, or weights of similar existing components with allowances

for modifications.

All of these methods of weight analysis have been employed In the electrooptlcs

assembly. The basic telescope weight was derived from an analytical model.

Even though this approach gives proper trends for the critical system parameters, the model fails to account for the entire system weight. Therefore, by

using the model on existing hardware, a plot of estimated versus actual weights

was obtained. The correlating factor which resulted from this plot accounts

for the added contingency. The structural weights in the imaging optics and

laser assemblies were derived by using a schematic. A weight was determined

for each component on the schematic. These component Weights we-re based on

similar on-golng laser communication hardware program components and on vendor

supplied data. The weight of all the other components of these assemblies such

as wire and fasteners were based On empirical equations. The detector and

Earth tracker assemblies were prlmariiy based on vendor supplied data.

In the electronics area the weight is primarily driven by the design and slze

of the circuit boards. The control electronics design features 14 ceramic

boards populated with leadless chip carrier packages containing the digital

circuits, and 2 polylmlde boards populated with analog circuits and discrete

components. The communication electronics is a similar design with lO ceramic

and 2 polylmlde boards. The size of each of these boards is approximately

4.6 inches by 3.8 inches. The weight of these boards was derived from actual

weights of similar existing boards. This resulting pounds per square inch was

applied to the area of the OPTRANSPAC boards. All other weights in these

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138
assemblies were derived from modifications to existing hardware. The Power

Conditioner unit weight was based on vendor supplied data. The circuitry required for the converter, regulators, etc. is in module form. Each module consists of two polyimide circuit boards housed in a frame. Five modules are required for the OPTRANSPAC design.

In the structures area, the weight of the support structure was based on the

material gages necessary to provide strength and structural stiffness, while

the miscellaneous structural weight was based on empirical equations. The wire

and connector weight was obtained from existing wire bundles modified to meet

the requirements of the OPTRANSPAC design.

139
ll.O GROWTH AND UNCERTAINTY

Areas In the OPTRANSPAC design that may require further examination beyond the

scope of thls study are identified. At the level of system maturity and analysis performed under thls study, detailed thermal design, structural dynamic

design, electronics design, optical design and host interface design have not

been performed. Detail design considerations associated with each of these may

slightly perturb the OPTRANSPAC design In terms of size, weight or power.

Figure ll-l outlines the OPTRANSPAC design considerations where a variation

from the baseline in size, weight or power may occur.

Due to the harsh Saturnian radiation environments identified In Section 3.8,

additional shielding may be required beyond the nominal shielding provided by

the baseline design. The Earth tracker and the command upllnk detector may require localized shielding to reduce radiation induced false detections. Radiators and/or heaters may be required to maintain the OPTRANSPAC components within the specified operating and non-operatlng temperature limits. Structural

analysis using launch loads and acoustic environments may indicate the need for

strengthening and prudent material selection. Further detailed optical analysis (ray tracing, telescope baffling design, and element definition) will

determine the exact number of elements required.

Because of the detail design that already exists and because of the few design

considerations still unknown, a growth and uncertainty factor of only I0-15%

for both weight and power is recommended. The margin for weight and power

growth then becomes ll.5 to 17.25 pounds and 5.7 to B.55 watts above the ll5

pounds and 57 watts determined for the baseline design.

140
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141

FIGURE 11-1
12.0 FDLLOW-ON HARDWARE DESIGN

To achieve the 1988 technology base required for the OPTRANSPAC design, funding

of key component technologies is required. Although some of these technologies

are advancing through other funding centers, design criteria specific to the

OPTRANSPAC may not be advanced to the desired state-of-the-art level by lgB8 if

development is not pursued with the OPTRANSPAC terminal In mind. The specific

technologies that require special attention are: l) The downllnk laser; 2)

the Earth tracker; and 3) the offset dlchrolc optical boreslght alignment device in the Earth tracker optics.

The OPTRANSPAC technology advancements required In the area of the downllnk

laser include the development of a high efficiency (> I0% optical output power

to pump source input power) frequency doubled (532 NM) Nd:YAG laser and the development of a compatible cavity dumped modulator capable of operation at pulse

repetition rates from l KBPS to 25 KBPS. These technologies are vital to the

OPTRANSPAC system design.

The advancements required in the area of the array tracker involve Earth track

capability along with continuous tracking of the transmit laser to provide

alignment and electrical polnt-ahead capability. Centrold algorithms that provide the required accuracies and resolutions at the desired output rates must

be determined and made compatible with the array tracker design. The capabillty of Earth track is imperative to the OPTRANSPAC design and the development of electrical polnt-ahead will provide a reduced weight optical deslgn by

eliminating additional redundant point-ahead beam steering mirrors.

The final technology area needing development involves the novel offset alignment device that allows a con_on boreslght to be tracked at two positions on

the array tracker. Fabrication and testing of this design will allow for verlflcatlon of the design and will define the optical tolerances required for such

a design.

Beyond tracking and developing pertinent hardware designs for the OPTRANSPAC,

breadboard design of the terminal to demonstrate system concepts appears as the

next logical step. Because a host platform is still unknown, the breadboard

design will demonstrate the operational feasibility and allow for design modifications where appropriate.

lhe final hardware follow-on involves the design and fabrlcatlon of an

OPTRANSPAC terminal for a specific JPL spacecraft employed in outer planet exploration.

142
i 13.0 CONCLUSION

The net effect of this study has been to prove the feasibility of operating a

laser communication llnk from a deep space probe to an Earth Orbiting Relay

Receiver. A system level terminal design of an optical transceiver package has

been performed. System analysis supported by communication llnk analysis indicates an OPTRANSPAC system design that )is capable of communication from Saturn at data rates up to lO0 KBPS (lO-3 BER . Reception of command data at l KBPS

Is also possible. First order optical, electrical and mechanical analysis and

design has been performed. A system mechanical envelope based on the optical

system layout and electrical component packaging has been developed. System

weight and power estimates of If5 pounds of 57 watts have been established.

All system level requirements, outlined in Section 2.1, with the exception of

the weight and power estimates, were met. Although not within the specified

limits required, the weight and power estimates establish the best fit terminal

design based on data from existing laser communication components and system

designs.

143
THIS PAGE LEFT INTENTIONALLY BLANK
APPENDIXA

!

!

I

!

OPTICAL TRANSCEIVER PACKAGE (0PTRANSPAC)

DYNAMIC ENVIRONMENTS

SPECIFICATION

28 JUNE 1985

A-I
1.0 SCOPE

i.I Definition

This document defines the dynamic environments that the Optical

Transceiver Package (OPTRANSPAC) is expected to encounter during

its mission.

1.2 OPTRANSPAC Definition

0PTRANSPAC is an optical communications terminal to be mounted to a

deep space vehicle whose mission is outer planet exploration. The

transceiver will communicate with an Earth-orbiting relay station

(EORS) from distances of Saturn and beyond, at downlink data rates

up to i00 Kbps, and an uplink rate of I Kbps. The transceiver

package must be reliable enough to ensure full operation over the

i0 to 20 year life of such a mission.

2.0 PURPOSE

This document shall be used as a reference to develop adequate

design specifications to ensure proper OPTRANSPAC operation in the

presence of the dynamic environments specified herein.

A-2
3.0 OPTRANSPAC DYNAMIC ENVIRONMENTS

3.1 Orbit Analysis

3.1.1

3.1.2

3.1.3

3.1.4

3.1.4.1

Earth to Saturn Transfer Orbit - A generic Hohmann (minimum energy)

transfer orbit for a deep space probe from Earth to Saturn is shown

in Figure 3.1-1. The seml-major axis of this ellptical orbit is

5.67 astronomical units (AU) and the eccentricity is 0.826.

Transit time is 1827 days. The true anomaly at launch is 6.9

degrees and at Saturn encounter is 172 degrees. This orbit was

used to calculate the relative velocities between the Earth and the

OPTRANSPAC spacecraft for Doppler and point ahead calculations.

Earth to Spacecraft Range - Figure 3.1-2 is a plot of range versus

time between the Earth and the OPTRANSPAC spacecraft for the

transfer orbit. The 2 AU oscillations are due to Earth orbital

motion.

Earth and Spacecraft Velocities - Figure 3.1-3 shows the velocities

of the Earth and spacecraft in the directions of the axes of the

transfer orbit. The spacecraft has an original velocity of about

i0 km/sec more than the Earth's in the minor axis direction and the

same as Earth's in the major axis direction. The relative

velocities between the Earth and spacecraft in axes parallel and

perpendicular to the instantaneous line-of-slght are plotted versus

time in Figure 3.1-4.

Point Ahead Angle and Doppler Shift

Point Ahead and Doppler Contributors - The Doppler shift of light

transmitted between the OPTRANSPAC and the Earth relay satellite

and the point ahead angle necessary for correct illumination are

determined by the relative velocity between the two platforms. The

four contributors to this velocity when the spacecraft is
encountering Saturn are the Earth's and Saturn's orbital rates

about the sun, the relay satelllte's orbital rate around the Earth,

and the spacecraft's flyby or orbital rate around Saturn. The

maximumvelocity contribution of each is given in Figure 3.1-5. Up

to the time of Saturn encounter, only the Earth, the relay

satellite, and the spacecraft transfer orbit motion apply.

3.1.4.2 Point Ahead An_le - Figure 3.1-6 shows part of the necessary point

ahead angle for the OPTRANSPAC during the Hohmann transfer between

Earth and Saturn due to the relative velocities in Figure 3.1-4.

The other part of the necessary point ahead angle is that needed to

compensate for the relay satellite motion. The maximum angle

contribution from the Earth relay station is limited by the

extremes of where it can he in its orbit about the Earth. That

contribution of up to 51 microradlans, which is reduced by the

inverse of range after a range of 1.8 au, creates a total point

ahead angle of 322 microradians maximum for the Hohmann transfer.

i

I

During the spacecraft's encounter with Saturn, Saturn's velocity

also contributes to point ahead. Since all four contributors may

have their total velocity perpendicular to the llne-of-sight, the

maximum velocity of each contribute to point ahead. The maximum

contribution to point ahead of each velocity is given below.

Point Ahead Maximum Contributions (_ tad)

Earth

Relay Satellite

Saturn

Spacecraft Flyby

199

51, range 1.8 AU

92.7/R, range 1.8 AU

64

167 .....

440, range = 9.5 AU

A-4
Since the contribution of Saturn's velocity is only in one

direction, the maximum polntahead in the other axis direction is

only 312 microradians.

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3.2

3.2.1

Doppler Shift - The Doppler shift expected during the Hohmann

transfer without the relay satellite contribution is plotted versus

flight time in Figure 3.1-7. Adding In the maximum contribution of

the relay satellite gives a maximum expected Doppler shift of 1.67

angstroms during the transfer from Earth to Saturn.

During the Saturn encounter, Saturn's velocity contributes through

the sine of the minimum angle of it with respect to the

llne-of-sight, defined by the minimum distance from Earth to Saturn

and the maximum cross-range of the Earth's position. The Earth and

spacecraft flyby contributing velocities are the same as the point

ahead case. The maximum Doppler contributions are shown below.

Doppler Maximum Contributions (_)(I _ 1.064_ m)

Earth 1.06

Relay Satellite .27

Saturn .08 sin (arc tan (2/8.5))factor

Spacecraft Flyby .89

2.30

Solid Particles - Meteoroids

Nominal Environment Meteoroids - The OPTRANSPAC should expect to

encounter the meteoroid fluences and particle characteristics

tabulated in Figure 3.2-1. for a deep space mission which includes

crossing Saturn's E-rlng. Column two of the figure provides the

total number of impacts on each square meter of exposed spacecraft

surface for meteoroids having mass greater than the value specified

in column one. The meteoroids are omnidirectional having no

preferred direction.

A-5
3.3 Masnetic Field

The OPTRANSPAC will be subjected to the magnetic fields defined

below.

3.4

Earth (non-operating)

interplanetary

Saturn

Shuttle (non-operating)

Gravitational Field

5 x 104 nT @ surface

25 nT

8 x 103 nT @ 2 Rs

32 mT

The OPTRANSPAC will encounter the following gravitational field

differential accelerations.

Earth (non-operating)

Sun

Saturn

3.5 Thermal Radiation

3.0 x 10 -6 m/sec/m @ surface

7.9 x 10 -14 m/sec/m @ 1 AU

4.3 x 10 -8 m/sec/m @ 2 Rs

3.6

Exposed portions of the OPTRANSPAC should expect to encounter the

thermal radiation levels specified below.

Earth Reflected

Solar

Saturn Reflected

57.3 mWcm -2 @ RE

163.0 mWcm -2 @ i AU

i.i mWcm -2 @ Rs

Electrical Field

The maximum expected Saturn induced electric field is

116 v/m @ 2 Rs.

A-8
3.7 Launch Vibrations

3.8

3.9

The expected Space Shuttle vibration environment is shown in Figure

3.7-1. The OPTRANSPAC shall be designed to operate after exposure

to this environment.

Launch Acoustic

The OPTRANSPAC shall be designed to operate after exposure to the

expected Space Shuttle acoustic environment shown in Figure 3.8-1.

pyrotechnic Shock

Figure 3.9-1 shows the maximum shock spectrum during

spacecraft/launch vehicle separation. This spectrum shall be

attenuated by intervening spacecraft structure.

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3.10

3.10.1

3.10.2

Radiation Environments

Total Dose - The mission environments contrlbuting to total dose

include the interplanetary electrons, protons, and solar flare

particles plus the Saturn electron and proton Van Allen belts. The

proton and electron-dose environments are shown as a function of

shielding in Figures 3.10-1 and 3.10-2. The total dose environment

is shown in Figure 3.10-3.

Displacement Damage - The major contributing environments to

displacement damage in semiconductor electronics are the protons

trapped in Saturn's Van Allen belt and the interplanetary cosmic

ray protons. Minor contributions are made by the nuclear power

source and the mission electron environments. The proton,

electron, and total displacement damage environments are shown as a

function of shielding in Figures 3.10-4, 3.10-5, and 3.10-6,

respectively.

A-7
3.10.3 Galactic Particles - The OPTRANSPAC can expect to see a galactic

particle environment on the same order as shown below.

Atomic Number Partlcles/cm2-sec

i (proton) 4.0

2 (alpha particle) 0.5

8 (oxygen) 3.0 x 10 -2

14 (silicon) 7.0 x 10 -3

26 (iron) 3.0 x 10 -4

3.11 Optical Background

3.11.1 Earth Background at OPTRANSPAC - The expected worst case Earth

background at OPTRANSPAC is shown in Figure 3.11-1. These data

were developed assuming "full Earth" illumination.

3.11.2

3.11.3

Solar Background at OPTRANSPAC - The effective background radiance

at the OPTRANSPAC due to the off-axls solar scattering is plotted

versus range in Figure 3.11-2.

Saturn Background at EORS - The background flux at the Earth

Orbiting Relay Satellite (EORS) platform due to Saturn is given in

Figure 3.11-3. These data are from recorded Earth orbit

measurements of Saturn's irradlance. The radiance function is used

directly with the field-of-view, aperture size, optical filter

bandwidth, and optics transmission to obtain the optical background

power on the EORS detector.

3.11.4 Solar Background at EORS - The solar background radiance at the

EORS due to off-axis scattering is shown in Figure 3.11-4.

A-8
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THIS PAGE LEFT INTENTIONALLY BLANK
i •

q,

I

Type

Configuration

Output Wavelength

Interpulse Period

Nominal Repetition Rates

Pulsewldth

Energy Per Pulse(4)

Efficiency

(Regulated Pump Power

To Optical Output)

Weight

Power

Envelope

APPENDIX B-1

DOWNLINK LASER

SPECIFICATION

Frequency Doubled Nd:YAG

Diode Pumped Slab

532 Nanometers _ 5 Angstroms

70 ± 25.6 _sec (1)

233.3 ± 25.6 gsec (2)

700.0 ! 25.6 psec (3)

14285.71 PPS (1)

4285.71 PPS (2)

1428.57 PPS (3)

10 nsec FWHM

28 _ 3oules (1)

93 _ 3oules (2)

280 p 3oules (3)

>10%

< 20 LBS

< 4 Watts m

8.2" x 9.5" x 5.0"

(1) 100 KBPS Data Rate

(2) 30 KBPS Data Rate

(3) 10 KBPS Data Rate

(4) Based On Nominal

Repetltlon Rates

I

i

I

B-1
THIS PAGE LEFT INTENTIONALLY BLANK
APPENDIX B-2

TELESCOPE

SPECIFICATION

i

• i

Type

Clear Aperture

Obscuratlon

Wavelength

Receive

Transmit

Quality

Wavefront Error

Viewfleld

Collimated Beam Size

Transmission

Magnification

Weight

Fixed Mounted

II inches (279.4 mm)

61.47 mm

400-1100 Nanometers

532 Nanometers

Diffraction Limited Transmit

at 532 Nanometers

< x/20 RMS at 532 Nanometers

+ 2.5 Milllradlans

< 0.5 inches

95% (excluding wavefront loss)

< 24:1 m

18 pounds

_t

i

B-2
THIS PAGE LEFT INTENTIONALLY BLANK
APPENDIX B-3

OPTICAL SYSTEM

SPECIFICATION

Type

Transmission

Wavefront Error

Filter Bandwidth

Ftlter Transmission

Collimated Beam Slze

Commandable Focus

Alignment Path

Weight

Common Path With Otchrolc

Comblned Transmit - Receive

> 65% Transmlt at 532 NH

> 65% Receive Com at 1064 NM

> 20% Earth Track

> 40% Allgnment at 532 NH

< X/20 RRS Transmit

25 Angstroms Centered About

]064 NR

> 70% at 1064 NM

< 0.5 inches n

Yes

Yes

< 13.5 Pounds

m

I

!

i

|

B-3
THIS PAGE LEFT INTENTIONALLY BLANK
APPENDIX B-4

DETECTOR/PREAMPLIFIER

SPECIFICATION

Detector

Type

Configuration

Quantum Efficiency

Ionization Coefficient

Avalanche Gain

Noise Equivalent Power

Bandwidth

Preamplifier

Type

Bandwidth

Noise Equivalent Current

Silicon Avalanche Photodiode

low -k, dimpled

> 40% at 1064 nm u

< .007

210

< 2.1 x iO-14 W/Hz I/2

Based on RCA 30954E

> 50 MHz

Transimpedance

> 50 MHz i

< 1.4 x lO-12 A/Hz I/2

|

B-4
I

THIS PAGE LEFT INTENTIONALLY BLANK
APPENDIX B-5

EARTH TRACKER DETECTOR

SPECIFICATION

I

l

!

Type

Sensitive Area

Pixel Size

Earth Track Active Area

Alignment Active Area

Earth Track Viewfleld

Earth Track Pixel FOV

PIxel Saturation

Dark Current

Dark Current Variations

Readout Noise Constant

Response Variation

Quantum Efficiency

Charge Injection Device (CID)

5.12ram x 5.12mm

20 microns

4.0mm x 4.0mm

4.0ram x 1.12ram

1.0 mrad x 1.0 mrad

5 urad x 5_rad

> I .5 x 106 e6.2 x lO4 e-/sec-plxel @ 23°C

.022 (RMS) for adjacent plxels

< 3.16 e-/Hz I/2

.006 (RMS) For adjacent plxels

>40% for 0.4 wm <_),<_ 1.0 wm

I

!

_i

|

B-5
4

THIS PAGE LEFT INTENTIONALLYBLANK
APPENDIX C-1

OPTRANSPAC

DESIGNNOTE

!

J

TITLE:

ELECTRONICS SYSTEM DESIGN

NUMBER 1_

REVISION

DATE 22 July 1985

TOTAL SHEETS

SUMMARY:

The baseline OPTRANSPAC Electronics are described. The electronics as

configured consist of three assemblies which provide the control, communication,

and power conditioning functions. The functional requirements of each assembly

are presented. A redundancy implementation based on attaining the highest

achievable probability of mission success is presented. The physical

characteristics of the electronic assemblies are presented.

PREPARED BY:'-_S_" _-_" _)7_.__:_,-_

H. J. MINGO " ;J

SECTION CHIEF ELECTRONICS

CHECKED BY: _- (_, __/_--_

S. G. LAMBERT

(_TRANSP_ ENGIneERING MANAGER

APPROVED BY: _-6 /-2"__

_. A. PAUTLER

OPTRANSPAC PROGRAM MANAGER

DISTRIBUTION:

J. P. Carter, S. G. Lambert, H. J. Mingo, J. A. Pautler

C-1

MCDONAII(I.I. OOUOLAII AS"rRONAUTICII COMPANY.ST. LOUIS DIVlilIOI_I
OPTRANSPAC ELECTRONICS

1.0 INTRODUCTION

The OPTRANSPAC Electronics assemblies (reference shaded blocks of Figure i)

provide the acquisition and tracking control function as well as the

communication function. A digital computer, input/output circuitry, and drive

circuits within the Control Electronics Assembly provide the acquisition and

tracking function. The communications function Is provided by circuitry within

the Communication Electronics Assembly. The Power Conditioning Unit converts

the spacecraft power to the required secondary voltage levels, provides

redundancy switching, mechanism, and heater control as well as command and

telemetry interfaces.

2.0 FUNCTIONAL REQUIREMENTS

2.1 CONTROL ELECTRONICS (Reference Figure 2)

Command signals to position the torque motor beam steerers (TMBSs) to direct

the incoming and transmitted optical signals is the principal output of the

Control Electronics Digital Processor. These outputs must be provided at a

200 Hz rate to support the 20 Hz TMB$ loop bandwidth and are the drivers for

the processor operation rate. Command signals for positioning the command

focus mechanism are initiated by ground command. Mode control of the

OPTRANSPAC operation is provided by the digital processor or via ground

commands. Figure 3 describes (in general terms) the processor operations in

each operating mode.

2.1.1 _4BS Command Operations - Command signals for TMBS position are computed

using the formula described in Figure 4.

2.1.2 Command Focus - Positioning of the command focus mechanism is ......

accomplished via ground command. The ground command is processed to generate a

discrete output which causes the focus mechanism to move until its position

sensor signal fed back to the processor is equal to the commanded position.

2.1.3 Mode Logic - The operational modes of the OPTRANSPAC are depicted in

Figure 5 and the mode logic is described in Figure 6.

2.2 COMMUNICATIONS ELECTRONICS (Reference Figure 7)

Both receive and transmit communication functions are required. The receive

function consists of conditioning, decoding, and formatting of the data pulses

from the communication detector. The transmit function is data formatting,

encoding, and generation of the required laser modulator drive signal.

2.2.1 Receive Function - The output from the Communication Detector's

preamplifier consists of an electrical signal with the following

characteristics:

Rate: 142.86 pulses/see

Coding: PPM data with 8 bits/pulse

Pulse Width: i0 nS

Amplitude: 0.i00 to 1.0 volts-peak

C-2
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0

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ts4

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C-4

I-- Z

i.- Z u.l

00_

FIGURE
OPTRANSPAC OPERATIONS LOGIC

EARTH TRACK

I

i_ !

)

I

l

CLOSE TMBS I

LOOPS

.

OUTPUT TRACK EARTH TM /

PROCESS EART_

TRACKER INPUT_

_

POSITION PROCESS INPUTS TMBS /

J

COMPUTE TMBS

COMMANDS

_YES

_YES

FIGURE 3

SHEET 1 oF 6

C-5
OPTRANSPAC OPERATIONS LOGIC

BEACON COMMUNICATION

INPUTS!

YES

NO

I PROCESS TMBS I

POSITION INPUTS

I_ i

COMPUTE TMBS

COMMANDS

DECODE COMM, BEACONI CMDS,

I SWITCH CMD I

SOURCE FROM

S/C TO BEAqON

[ INCREMENT i

TIMER i

c-B

YES

FIGURE 3

SHEET 2 OF 6
OPTRANSPACOPERATIONSLOGIC

DOWNLINKCOMMUNICATION

ACTIVATE I

LASER

I SELECTRATEDATA ]

_ROCESS EARTH I

ACKER INPUT_

i

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i

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l_

I PROCESS POS. INPUTS TMBS

ITMBS CMDS I

C-?

YES

FIGURE 3

SHEET 3 OF 6
OPTRANSPACOPERATIONSLOGIC

ALIGNMENT

I ACTIVATE LASER

PROCESS TRACK INPUTS EARTH

NO

OFF

NO

EARTH

TRACK

YES

EARTH

TRACK

YES

FIGURE 3

SHEET 4 OF 6

C-8
OPTRANSPAC OPERATIONS LOGIC

STANDBY

I PROCESS I

THERMISTOR

INPUTS

OUTPUT TM

'r

INDICATE I

IN TM

FIGURE 3

SHEET 5 OF 6

c-g
OPTRANSPAC OPERATIONS LOGIC

EARTH ACQUISITION

kRTH

CQUISIT

TRACKER TO 1

ACQ, MODE

VIEWFIELD SCAN 1

YES

_ 1N°

TIMER

NO

NO

FIGURE 3

SHEET 6 OF G

C-10
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C-i2 FIGURE 5
OPTRANSPAC MODE LOGIC

NO

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QUIESCENT MODE ]

OFF YES

NO

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PROC,

OFF

NO

YES

NO EARTH

(

YES

I OPERATIONS EARTH ACQ.

C-13

1

@ _"_G/J"ES6

SHEET 1 OF 3
OPTRANSPAC MODE LOGIC

(CONTINUED)

OFF YES

NO

YES STANDBY

NO EARTH

TRACK

NO

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OFF YES

NO

YES

ACQ

NO

®

COMM

YES

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®

C--14

NO

(

FIGURE 6

SHEET 2 OF 3
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(CONTINUED)

® ®

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OPERATIONS

OFF YES

NO

YES

EARTH

NO

NO

YES

DOWNLINK COMM I

AND

ALIGNMENT OPS

I BEACON_NO

YES

BEACON COMM I

OPERATIONS

C-15

FIGURE 6

SHEET 3 OF 3
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C-16

FIGURE 7
This signal is amplified and applied to a PPM decoder that operates on a

"greatest of" principle to select the slot containing the signal pulse. The

output of the decoder is a NRZ data stream containing rate 7/8 Reed-Solomon

encoded data at 1142.86 bits/sec. Thls data is applied to a Reed-Solomon

decoder which produces a i000 bit/sec output. Finally, the i000 bit/sec data

is formatted to be identical to the command data received vla the radio

frequency command llnk. The output of the formatter is applied to a command

decoder located within the Power Conditioner Unlt.

2.2.2 Transmit Function - Transmit data consists of a fixed rate serial

telemetry data stream from a telemetry formatter located in the Power

Conditioner Unit or a serial data stream from the Spacecraft sensor.

Spacecraft sensor data rates can be i0,000 blts/sec, 30,000 bits/sec or i00,000

bits/sec. Figure 8 illustrates the transmit data frame. Overhead Is limited

to less than 3%. When spacecraft sensor data is available for transmission,

the telemetry data is not transmitted. The output of the data selector (either

telemetry or sensor data) is rate 7/8 Reed-Solomon encoded and output to a PPM

encoder. The Reed-Solomon encoder adds a coding blt to each input word

resulting in the following output rates:

Input Rate

i0,000 bits/sec

30,000 bits/sec

i00,000 bits/sec

Output Rate

11,428.57 blts/sec

34,285.71 bits/sec

114,285.71 blts/sec

The output data from the Reed-Solomon encoder Is applied to a PPM encoder that

produces output symbol information at 1428.57, 4285.71, or 14,285.71

symbols/sec depending upon the sensor data rate. Symbols are encoded by

outputting a pulse in one of 256 data slots following a fixed time interval

that is referenced to the previous fixed time interval using an internal

clock. The following are the timing requirements for the PPM encoder:

Slot Width: i00 nsec

Fixed Interval: I/fs, where fs Is the transmitted symbol rate

The output from the PPM encoder Is input to modulator driver circuitry. The

circuitry converts the logic level input to signal levels required to operate

the laser modulator located within the Laser Assembly.

2.3 POWER CONDITIONING (Reference Figure 9)

Primary power from the spacecraft's electrical system is conditioned to provide

the required secondary power for the OPTRANSPAC equipment. The Power

Conditioning Unit provides the conversion and regulator circuitry required to

produce the following secondary outputs:

C-17
REPRESENTATIVE DATA FRAME

100 KBPS DOWNLINK

t

I

34

PPM SYNC

12 WORDS

FRAME SYNC DECODER SYNC

3 WORDS I WORD

SPACECRAFT

ENCODED

DATA

(4352 BITS)

..... im

FRAME RATE 26.26/SEC

OVERHEAD PERCENTAGE ~ 2.9%

128 BITS

c-18 FIGURE 8
Z

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C- 1 9 FIGURE 9
Levei Current

+5Vdc 1.2A

+ 5 Vdc 0.3 A

+ 5 Vdc 0.2 A

+ 5 Vdc 1.3 A

+ 5 Vdc 0.6 A

+ 5 Vdc 1.0 A

+ 5 Vdc 0.5 A

+/- 15 Vdc 0.i A

+/- 15 Vdc 0.013 A

+/- 15 Vdc 0. ii A

+/- 15 Vdc 0.05 A

+/- 15 Vdc 0.01 A

+/- 15 Vdc 0.02 A

+ 28 Vdc 0.01 A

+ 28 Vdc 0.04 A

+ 400 Vdc 250 _A

User

Earth Tracker

Control Electronics

Control Electronics

Control Electronics

Control Electronics

Comm. Electronics

Power Conditioning

Earth Tracker

Control Electronics

Control Electronics & TMBS's

Comm. Electronics

Comm. Detector

Power Conditioning

Power Conditioning

Comm. Electronics

Comm. Detector

Figure I0 describes the converter implementation. Command and telemetry

conversion circuitry is also located within the Power Conditioning Unit.

Serial command data from the spacecraft's command decoder or the Communication

Electronics is decoded. _= The output from this circuitry is provfded to discrete

drivers which control OPTRANSPAC redundancy switching mechanisms, the active

focus, or power control relays. Serial magnitude commands are relayed to the

Control Electronics via bl-dlrectlonal serial interface circuitry located

within the Power Conditioning Unit. Thls serial interface receives telemetry

and heater control data from the Control Electronics. The telemetry from the

Control Electronics as well as internal analog and discrete data is multiplexed

to form a serial data output. This data is transmitted to the Earth station

via the Communication Electronics or on the spacecraft radio frequency links.

Circuitry required to condition internal analog and discrete telemetry

parameters is provided within the Power Conditioning Unit. OPTRANSPAC heater

drive circuitry is also located internal to this unit.

C-20
v

m

e

sic

BUS

A

s/c

BUS

B

PREREG, A IPREREG. B

POWER CONVERTERS

__$I @ I

%LH I

_Ll@l

% Ld® ITiL_

+5V

I +15V

+5V

+5V

+5V

+5V

ira,

+5V

F_+-_5v--=

+5V

L+ibV, -,,- Ira,,,

+15V

+15V

i+28V =

--+I5V

t +Z_V

_+15V

_v___v

(CONTINUED ON NEXT PAGE)

C-21

i:"IGuI_ I0

PAGE i OF 3

....... ; .... L ;Z -
POWER CONVERTERS

(CONTINUED)

POWER

BUS ------_

CONVERTER NUMBER

DENOTES COMMAND

X

+15V

+15V

+5V

{+lSV :--

+5V

.li,,,

+5V

i i

B

FIGURE i0

PAGE 2 OF 3

C-22
_UM±/I -4

o

3

4

5

6

7

8

9

i0

II

12

13

14

15

16

17

18

Figure 7 Continued

Power Converter Desi_natlons

A

Earth Tracker

B

A

Earth Tracker Readout Elect.

B

A l

Serial Interface and 'ARS

! B

A} CPU

B

A

TMBS and Drivers

B

A

Comm. Electronics

B

A

B

A

PCU Internal

!

S

A

Memo ry

B

C-23

FIGURE I0

PAGE 5 OF 5
3.0 _U_

3.1 CHARACTERISTICS

Redundancy implementation shall be provided to eliminate single-polnt failures

and provide the highest possible probability of success consistent with

practical weight and complexity constraints. The following "ground rules"

regarding cross-strapplng of redundant elements have been established.

A. All cross-straps at source.

B. All cross-strap selection via power command in the PCU and digital

serial magnitude commands for addressing the selected memory block or

I/O port.

C. The Communication Electronics is block redundant with cross-strapplng

at the output to PCU Command Decoder and Inputs (provided by source)

from PCU TM and S/C Sen_or.

D. The Power Conditioner Unit converters are dedicated to the load. There

is no power converter output cross-strapping.

3.2 REDUNDANCY IMPLEMENTATION

All OPTRANSPAC electronics and mechanisms are redundant with the exception of

the Command Focus Mechanism. Cross-strapplng for selected functional circuit

blocks is provided to minimize the probability of system failure.

3.2.1 Control Function - The elements of the function and the cross-straps

provided are described in Figure II. All elements of the function are

cross-strapped to other redundant elements with the exception of the TMBS's.

Cross-strapping of the drivers and TMBS's does not provide significant

enhancement of system reliability since the reliability of the drivers and

TMBS's is high (based on previous designs) compared to other elements. In

addition, a simpler design for the TMBS drivers is possible.

3.2.2 Communications Function - The elements of the communication function and

the cross-strapplng provided are detailed in Figure 12. Cross-straps are not

provided between the Communications Avalanche Photodlode and the preamplifier

due to the low signal levels and noise sensitivity at this point. Similarly,

design considerations are the principle reason for not cross-strapplng the

Communication Electronic - Laser interface.

3.2.3 Power Conditlonln_ Function - Figure 13 details the elements and

cross-strapping of the power conditioning function. Implementation of these

cross-straps is accomplished by use of relays controlled via Earth station

generated commands. Fuses (reference Figure i0) are provided at the input to

each converter to prevent load or converter failure propagation to the

redundant converter.

t

C-24
s

Z0

S,==-4

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h-- ._

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FIGURE 11

C-25
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C-26

FIGURE 12
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C-27

FIGURE 13
4.0 DESIGN IMPLEMENTATION

The OPTRANSPAC Electronics are housed in package assemblies having the

following characteristics (estimated):

Assembl 7 Power Disc. Weight Length Width Hel_ht

Control Electronics 14.5 W 12.5 ib 9.0" 6.0" 6.0"

Communication Electronics 7.7 W 9.0 ib 7.0" 6.0" 6.0"

Power Conditioning Unit 17.1 W 19.5 ib 6.0" 8.0" 7.5"

Earth Tracker 4.6 W 4.5 ib ---

Comm. Detector 0.4 W 1.0 ib ---

TMBS's 1.4 W (Included in Imaging Optics)

Laser 4.0 W (Included as Laser)

Wire and Misc. 6.7 W

4.1 CONTROL ELECTRONICS

The Control Electronics design is based on 14 ceramic and 2 polyimide circuit

boards. Digital circuits in leadless chip carrier packages will be mounted on

the ceramic circuit boards. The polyimlde boards will be populated with analog

integrated circuits and discrete components. Dimensions of all circuit boards

is approximately 4.6" x 3.8". Both prime and redundant circuitry will be

implemented on the 16 circuit boards.

4.2 C_4MUNICATION ELECTRONICS

A total of 12 circuit boards are estimated for the redundant Communication

Electronics. A pair of polylmide boards will be used for the analog integrated

circuits and discrete components required. All other circuit boards will be

populated with leadless chip carrier digital integrated circuit packages.

4.3 POWER CONDITIONING UNIT

The circuitry required for the converter, regulators, relays, and etc. will be

implemented using a module concept. Each module will consist of two (prime and

redundant) polylmlde circuit boards housed in a frame. The dimensions of each

module will be approximately 6" x 6" x 0.3". A total of 5 modules are required.

C-28
OPTRANS PAC ELECTRONICS

Control Electronics Circuit Boards

Qty*

2

2

2

2

2

2

4

16

Function

CPU and Clock

Analog I/O

Memory

Serial Interface & ARS Logic

Discrete I/O

TMBS and Command Focus Drive

Earth Tracker Readout Logic

Communication Electronics Circuit Boards

2

2

2

2

2

12

Function

Post Amp. & PPM Decode

Rate 7/8 Decode & Sunc. Def.

Data Mux., Data Format., & Clock

Modulator Drivers

Temperature Controller

Power Required

2.0W

4.4W

3.0W

1.0W

0.5W

2.1W

I.SW

ir755w

Power Required

1.0W

1.0W

I.OW

2.0W

1.5W

7.7W

*Prime and redundant, only i board powered.

C-29
OPTRANS PAC ELECTRONICS

Power Conditioning Unit Power

Converter +28 Vdc

Earth Tracker

Earth Tracker Readout

Serial Int. & ARS

CPU and Clock

I/O (analog & dlscrete)

Memory

TMBS, TMBS and

CMD. Focus Dr. 0.01

Comm. Elect. 0.04

APD

Mist/Wire

Output Current in Amps

+/-15 Vdc +5 Vdc Bias

0.i 0.32

0.3

0.2

0.4

0. 013 O. 9

0.6

0. ii

0.05 1.0

0.01 250 pA

Laser

Total Power

4.6W

1.5W

I.OW

2.0W

4.9W

3.0W

4.1W

7.7W

0.4W

6.7W

4.0W

TOTAL

Converter Losses (based on 70% eff.)

39.9 W

17.1 W

C-30
i I

i

APPENDIX C-2

OPTRANSPAC

DESIGNNOTE

TITLE:

OPTICAL SYSTEM DESIGN

NUMBER 2

REVISION A

DATE 19 AUGUST 1985

TOTAL SHEETS

SUMMARY:

The optical system performance requirements of the OPTRANSPAC system are

summarized. An optical system design is developed from functional requirements

through the thin lens design stage. Component performance requirements and

packaging/envelope requirements are discussed.

PREPARED BY: f __"¢_

T. C. WILLIS

OPTICAL SYSTEM ENI__LEER CHECKED BY: _.._'_Y_:'_;-.,-/--_

S. G_ CA_BERT "

APPROVED BY: _PTRANSPA_3-_NGINEERING_. c'7 /,_Jc-_Z_Iv -

MANAGER

C_/j.A. PAUTLER

OPTRANSPAC PROGRAM MANAGER

DISTRIBUTION:

J. P. Carter, E. C. Clarke, L. M. Gnojewski, S. G. Lambert, H. J. Mingo,

J. A. Pautler, M. Ross, J. D. Zino

C-31

It4CDOI41RiII,.I. O04,10&.Am AJTIeORiJA#JTIC, CORI41mA/tlt" .ST. &,O4.J#4B Ir)QVISION
l.O OPTICAL SYSTEM REQUIREMENTS

The basic configuration of the OPTRANSPAC optical system is defined In Figure

I. OPTRANSPAC optical system performance requirements are summarized in Table

I. The receiver aperture slze and path transmissions are governed by the

range losses out at the maximum range of lO AU. The fields of view are

selected for background rejection and to support the chosen acquisition sequence. The angular steering range is governed by the limitations on the

pointing ability of the spacecraft attitude control system. The wavelength

range Is set by the need to track the Earth In broadband visible light (reflected sunlight) and the need to monitor the frequency doubled Nd:YAG laser

transmitter.

The optical system consists of two basic subassemblies: a telescope and an

imaging optics assembly (or IOA). The optical system performance of Table l

can be flowed down to the individual subassemblies. This Is illustrated in

Table 2.

2.0 TELESCOPE SELECTION

The performance requirements of the telescope flow down from the system level

requirements for OPTRANSPAC. These telescope requirements are defined in

Table 2.

To meet these requirements a mirror based (reflective) telescope is the system

of choice. An ll" aperture refractor would require several lens elements to

meet image quality requirements and would be heavy as a result. A refractive

telescope would also be prone to thermal gradients, which would adversely

affect the transmitted wavefront by the temperature induced change of the

glass refractive indices.

The wide variety of reflective telescope design forms is illustrated in Figure

2. The imagequallty requirements over the field of view limit the cholce of

telescopes. The Newtonlan form is unacceptable over this field due to the

coma inherent in the base parabolold mirror. The Gregorian form Is too long

C-32
and bulky. The Haksoutoff arrangement tends to be either too long (with a

thin corrector plate) or too heavy. The corrector plate thickness must increase as the telescope length is reduced to maintain the required state of

correction. The Cassegraln design form, in one of Its many variants, offers

the best approach to meeting the diffraction limlted performance requirement

over the field of view with a compact, ltght weight, stable package.

I

m_j

The Classical Cassegratn consists of a parabolold primary mirror with a hyperbolold secondary mirror. This arrangement can cover a larger field of view

than the single parabolotd mirror with substantially better imaging performance. It wtll however not be able to cover the full 5 mrad field with the

required near diffraction limited performance.

There are several varlants of the Classical Cassegrain which offer improved

image quality over the fleld at the cost of an increase In design complexity.

One of these variants is the Rltchey Chretlen. This configuration consists of

aspheric primary and secondary mirrors whose zonal curvatures are sltghtly

weakened wtth respect to the parabolotd/hyperboloid mirrors of the Classical

Cassegratn to reduce coma over the field. Another variant ts the

Schmidt-Cassegraln which dispenses with the asphertc primary mirror and incorporates a full aperture aspherized refractive plate to correct the resulting

spherical aberration. This plate must cover the full aperture and is prone to

thermal gradients perturbing wavefront quality. Another variant on the

Classical Cassegraln (and on the Rttchey Chretlen) ts to incorporate small

refractive correcting elements in the converging beam after the secondary.

Thts allows additional aberration correction. It Is this final variant that

offers the best approach to meeting the rigorous image qualtty requirement of

k/20 RMS over the relatlvely large 5 mrad field of view.

In the OPTRANSPAC telescope design the primary mirror ls an f/1.75 hyperboloid

(only sltghtly changed from a paraboloid). A hyperbolold secondary mirror,

operating at a magnification of 5.7, produces a final f/# of _f/10 for the

telescope. The refractive correctors are of zero net power and affect the

C-33
final f/# very little. A field lens of =600 mm focal length relays the system aperture stop (at the primary mirror) Into the rest of the optical

system. A thick lens design of this telescope was not completed during the

study. However, a very slmllar telescope deslgned on a different program was

available for study. This telescope design, incorporating small refractive

correcting elements, ls shown In Figure 3. A detailed design prescription Is

ltsted tn Table 3. Geometric raytraclng of this design both on axis and at

full field (Figure 4) confirms that thls design fulftlls the OPTRANSPAC image

quality requirements over the field of view. Diffraction analysis Indicates

this telescope's design wavefront quallty Is = _/30 RMS at a wavelength of

k = 532 nm.

3.0 OPTICAL SCHEMATIC

Following the definition of the optical system performance requirements and

selection of a telescope type, the optical paths need to be defined. The optlcal schematic (Figure 5) shows the general arrangement of the Cassegraln

telescope and the imaging optics that interfaces the telescope with the

system's transmitter laser source and various detectors. Table 4 lists the

components In the optlcal schematic.

Figure 6 breaks down the schematic Into the individual optlcal paths. The

telescope section serves as the common antenna for both transmit and receive

channels. The prime and redundant common optical path are also shared by the

transmit and recelve channels. These paths incorporate the fine beam steering

function. The chosen beam steering mechanlsm Is a torque motor beam steerer

(TMBS). Selectlon between the prime and redundant common path is accomplished

by means of rotary mirror mechanisms. The final component In the common optical path Is the transmlt/recelve beamsplltter. This Is a dlchrolc beamsplitter set to transmit 532 nm light while reflecting the I064 nm upllnk

llght and broadband visible light (earth tracking).

z_

The transmit path consists of Prlme and redundant lasers with a "pop mirror"

select mechanism, a focal beam expanding optics, and a commandable focus.

C-34
1

1

¢ I

The receive path is subdivided into a communication and a tracking section.

The communication section consists of prime and redundant avalanche photodiodes selected by the "pop mirror" mechanism, 25 A band pass filters centered on the 1064 nm upltnk for background rejection, and a dtchrotc beamsplitter to reflect 1064 nm light and pass the shorter wavelengths used in

tracking. The tracking section is somewhat more complex. The broadband solar

energy reflected off the Earth is used for tracking. Prime and redundant

charge transfer detectors (CTD) selected via "pop mirrors" are used to track

the Earth and hence the low Earth orbit terminal. The point ahead function

Is accomplished via offset tracking. A small amount of the transmit 532 nm

light is continuously leaked by the transmit/receive beamsplttter into the

receive/tracking path. A 532 nm bandpass filter reflects the broadband light

and admits the 532 nm light to a deviating prism which steers this transmit

light to an offset point on the CTD.

i

4.0 FIRST ORDER LAYOUT OF OPTICAL PATHS

In design of the telescope and imaging optics for a spaceborne system one must

keep the size and number of elements used to the minimum necessary to perform

all system functions within the required system performance levels. This maximizes system optical transmission while minimizing system size, weight, and

procurement cost. In diffraction llmlted optical systems such as this, one is

forced to multiple elements less components with moderately slow optical speed

to meet image quality requirements. Field of view and spectral range requirements further complicate the lens components. This exacts a penalty on optics

transmission, optical system envelope and optical system weight.

4.1 FIRST ORDER LAYOUT OF TELESCOPE - The elements and their first orders

properties for the telescope section are given in Table 5. A thin lens ray

trace through this section both on axis and at full field is shown in Figure

7. This section ls designed to be afocal. It maps the 0.5" (12.7 mm) diameter axial bundle of the imaging optics paths (i.e. common path) to the ll"

(279.4 mm) diameter of the primary mirror. The system aperture stop and

C-35
entrance pupil are at the primary mirror. This pupil is relayed by the telescope field lens to a convenient location within 30.5 mm of the first colllmator lens in the common path of the imaging optics.

The Cassegraln telescope optics are designed to produce an f/lO output cone.

This allows for a compact package whose optical speed is still slow enough to

ensure both good image quality and readily manufacturable components. The

f/I.75 concave hyperbolold primary and f/2.0 convex hyperbolold secondary

mirror are well within the state of the art for aspherlc fabrication.

4.2 FIRST ORDER LAYOUT OF COMMON PATH (PRIME/REDUNDANT) OPTIC% - In the

common path throughout the imaging optics It Is desirable to use a few basic

lens grouping in a repetitlve way to simplify the design and procurement of

the optics. Such repetitive lens groupings must relay the telescope pupil

through the imaging optics with minimal vlgnetting. This requires that the

sections of collimated space used for locating TMBS mirrors or beamsplltters

be rather short to minimize the need for large (> SO mm diameter) optics.

This is because the local field of vlew covered by the imaging optics components is 88 mrad (_ 2.5"), a factor of IB times the telescope field. A

typical grouping of collimator - field lens - collimator ("CFC Group") is

shown in Flgure 8. The two collimators symmetrically flanking the field lens

work at f/5. Thls f/# was selected as the best compromise between the two

extremes of fast f/#, compact, many element lens components and slow f/#,

space wasting, few element lens components.

The elements and first order properties for the prime and redundant common

path optics are given in Table 6. A thin lens ray trace on axis and at full

field through the common path are glven in Figure g.

4.3 FIRST ORDER LAYOUT OF TRANSMIT PATH - The optics in the transmit path are

considerably simplified by the fact that they are handling monochromatic (532

nm) light which is of small dlvergence and remains on axis (all steering of

the transmit beam is accomplished in the common optics path). The prime and

redundant lasers are selected by a "pop mirror" mechanism. The laser light

passes through an afocal relay. One of the components of this relay can be

shifted axially to introduce specific amounts of focus error to partially

C-36
compensate for wavefront errors introduced by component shifts. These component alignment shifts can arise from the forces during launch, from stress

relief of structural parts over time, or from long term temperature changes

during the course of the mission. In a diffraction limited system even such

minor variations can have a serious effect on performance. After leaving the

focus mechanism relay the beam enters an afocal expander to increase the beam

to 12.7 mm diameter for entry into the common optics path.

Table 7 lists the transmit path optics properties.

trace of the emerging transmit beam.

Figure 10 shows a ray

4.4 FIRST ORDER LAYOUT OF RECEIVE COMMUNICATION PATH - The 1064 nm upllnk

from the terminal In low Earth orbit is collected by the telescope, and relayed through the common path optics, reflected by the dichrolc transmlt/recelve beam splitter, relayed by a CFC lens group, reflected by the dichroic

communlcatlon/tracklng beamsplitter, and f11tered by a 25 A bandpass

filter. The prime or redundant detector are selected by a "pop mirror"

mechanism. The elements and first order properties for the receive path from

the transmlt/receive beamsplltter through the communication detectors are

given In Table 8. On and off axls ray traces are shown in Figure If. Note

that the avalanche photodlodes are operated In a slightly defocussed position

to keep the spot slze large enough that the variation in detector responsivlty

over its surface has a negligible impact on communication llnk margln.

4.5 FIRST ORDER LAYOUT OF,RECEIVE TRACKIN _ PATH - To accommodate the tracking

and point ahead functions on the same detector, an offset Is introduced

between the broadband reflected sunlight from earth and the 532 nm transmit

spot retroreflected from the transmit beam. A dlchrolc beamsplltter separates

the broadband and 532 nm light into two separate channels. A wedge prism deviates the 532 nm light prior to entering the tracking detector lens. The

broadband light enters the tracking detector lens undevlated yielding the

angular offset required. Note that the tracking detector lens axls is decentered wlth respect to the input receiver axls to accommodate the two spatially

separated Earth track and polnt-ahead offset track channels. The elements and

first order properties of the receive path from the communlcatlon/tracklng

C-37
beamsplltter through the tracking detector are shown in Table 9. On and off

axis ray traces for the broadband visible light are shown In Figure 12. On

and off axis ray traces for the 532 nm light are shown In Figure 13.

S.O M_CHANICAL LAYOUT,.PACKAG!NG, AND ENVELOPE

Using the component distances and sizes given In Secton 4.0 the optical paths

can be layed out to determine the envelope of the resulting package. Figure

14 Illustrates an Isometric view of this layout. Figure 15 shows the same

layout In orthographic projection. The optlcs are divided into two planar

sections above and below a mounting baseplate. To t_ts samebaseplate the 11"

telescope is attached in a cantilevered configuration. When baseplate, optics

covers and electronics boxes are lncluded a more complete picture of the

system envelope emerges (Figure 16). The entire system fits within a 2'x2'x4'

envelope yet offers uncrowded, ready access to the optlcal paths for alignment, adjustment and performance characterization.

6.0 TRANSMISSION .BUDGETING

A summary of overall path transmlsslon is given in Table I0. A detailed

breakdown on the contributers to transmission In each path is included in thls

Section. These numbers are for beginning of llfe. The Indlvldual component

performance estimates contained hereln are readily manufacturable specification values and are based on the measured performance of optical components

built and used in Lasercom systems over the past eight years. The component

substrates and coating materials are restricted to materials wlth demonstrated

radiation resistance. Radiation resistance test Indlcate element transmittance degradation Is typically small ( = 0.2% per element at = lO5

Rad (Sl)). Equivalent end of llfe transmissions based on this are also shown

In Table 10.

l

6.1 TELESCOPE PATH

6.1.1 TRANSMIT (532 nm) USAGE

C-38
8

Telescope Collimator (4 Element) - - T1 = (0.998)

Field Lens/Refractive Corrector Lens Set (3 Element) - - T2 . (0.998)

2

Secondary Mirror/Primary - - T 3 = (0.98)

Telescope Transmission At $32 nm= T1T 2 T3 = 0.93

6.1.2 RECEIVE (1064 nm) USAGE

Telescope Transmission at 1064 nm= Telescope Transmission at 532 nm

6.1.3

= 0.93

RECEIVE (BROADBAND 400 nm- 100 nm) USAGE

p

e. f

Unlike the monochromatic performance numbers in Section 6.1.1 and 6.1.2, the

anttreflectlon coating performance w111 be substantially poorer over the

broadband visible and near infrared band of reflected sunlight. The same will

be true over the blue-green end of the spectrum for the high reflectance

coatings on telescope primary and secondary mirrors. This Is due to the inherent llmttattons In multllayer interference coatings whlch can only function

at their maximum performance over spectral bandwidths much less than the required 700 nm.

Telescope Collimator (4 Element) - - T1 = (0.99) fl

Field Lens/Refractive Corrector Lens Set (3 Element) - - T 2 = (0.99) 6

Secondary Nlrror/Prlmary Mirror - - T3 = (0.97) 2

Telescope Transmission (400-1100 nm)= TIT2T 3 = 0.82

C-39
6.2 COMMON (PRIME/REDUNDANT) PATH

6.2.1 TRANSMIT (532 rim) USAGE

Transmission of Basic Collimator Field Lens - Collimator Group on "CFC" Group

(Consisting of 9 Elements) = (0.998) 18 = 0.965 = Tcfc

Reflectance of Fold Mirror Or TMBS Mirror = (0.98) = Rm

Transmittance of Dichrolc At 532 nm= (0.98) = Td

The Common Path (Prime or Redundant) consists of three CFC groups, two

prime/redundant select mechanism mirrors, two TMBS mirrors, and two passive

fold mirrors. The resulting transmission is given by

t

I

T of Common Path at 532 nm= (Tcfc)3 (Rm) 4 Td

= 0.81

6.22 RECEIVE (1064 nm) USAGE

Common Path performance at I064 nm is similar to that at 532 nm except for the

dlchrolc transmit/recelve performance.

So, we have

RD = Dlchrolc Reflectance at 1064 nm= 0.95

T of Common Path at I064 nm= (Tcfc)3 (RM) 4 RD

= (0.965) 3 (0.98) 4 (O.9S)

l,

= 0.79

C-40
6.2.3 RECEIVE (BROADBAND VISIBLE) USAGE

Broadband Transmission of Baslc CFC Group (9 Elements) = (0.99) IB

.0.83. TCF c

Broadband Reflectance of Fold Mirror or TMBS - (0.97) = Rm

Broadband Reflectance of Olchrolc . (o.go) = RD

The resulting Common Path broadband transmlsslon Is given by

T Common Path Broadband = (0.83) 3 (0.97)

= 0.46

4

(o.9o)

6.3 TRANSMIT PATH

6.3.1 TRANSMIT (532 nm ONLY) USAGE

8

Transmission of Focus Compensator Group (4 Elements) = (0.998) = 0.984

= TFC

Transmlsslon of Beam Expander (4 Elements) = (0.998) 8 . 0.984 = TBE

Transm_sslon of Transmit Path at 532 nm = TFC TBE = 0.97

6.4 RECEIVE COMMUNICATION PATH

6.4.1 RECEIVE COMMUNICATION (I064 nm ONLY) USAGE

Transmlsslon of CFC Group at I064 nm = (0.965) = TCF c

Reflectance of Passive Fold Mirror = (0.98) = RM

Reflectance of Communlcation/Track Olchrolc = (0.95) = RD

Bandpass Filter Tranmsmlttance . (0.70) = TBp F

C-41
6

Detector Lens Transmittance (3 Element) = (0.998) = (0.988) = TDL

Transmission of Receive Communtcaton Path at 1064 nm

= TCFC RH RD TBpF TDL = (0.62)

6.5 RECEIVE TRACKING PATH

6.5.1 BROADBAND VISIBLE USAGE

CFC Group Transmlsslon • (0.99) 18 = 0.83 = TCFc

i

Passive Fold Mirror Reflectance = (0.97) = RM

Dlchrolc Transmittance . (0.90) = TD

Variable Attenuator Base Transmittance . (0.99)

2

• (0.98) = TVA

532 nm Bandpass Filter Reflectance • (0.90) = RBp F

Tracking Lens Transmittance (3 Element) = (0.99) 6 = (0.94) - TTL

Total Transmlsslon In Broadband Visible Light For Tracking

= (TcF C (RM)2 (TD) (TvA) (RBp F) (TTL)

= 0.59

6.5.2. TRANSMIT (532 nm) USAGE FOR OFFSET TRACKING

Reference Transmit Path Trasmittance At 532 nm

. (0.97) = TTp

i

•Transmittance" of Xmlt/Recelve Dichrolc (Leakage to Corner Cube and

Retransmlsslon into Tracking Path) = (0.005) (0.98) - 0.00490 = TDp A

|

=

Transmittance of CFC Group at 532 nm= (0.965) = TCFC

Reflectance of Fold Mirror - (0.98) = RM

C-42
Transmittance of Tracklng/Comm Olchrotc Beamsplltter -- 0.98 = T D

2

Variable Attenuator Transmittance = (0.998) = TVA

Band Filter Transmlttace = (0.70) = TBpF

Wedge Transmittance = (0.998) 2 = (0.996) . Tg

Tracking Lens Transmittance (3 Element) = (0.998)

Total Transmission For 532 nm Offset Tracking

= (0.988) = TTL

- (TTp) (Top A) TCFC) (RH)( TO) (TvA) (TBp F) (T N) (TTL)

= 0.0030

7.0 IMAGE QUALITY AND WAVEFRONT BUDGETING

The most critical requirement on image quality for the OPTRTANSPAC system Is

that for the transmit path. In order to meet llnk margin requlrements at the

10 AU maximum range a near diffraction limited wavefront needs to be radiated

from the transmitter. Both telescope and imaging optics are each Indlvldually

required to maintain the transmlter wavefront to _ x/20 RMS (k = 532 rim).

The total system performance resulting from combined telescope and imaging

optics wlll then be =_/14.1 RMS.

For the telescope to maintain ),/20 RMS what Is required of the individual

optical surfaces within the telescope? Recall that the total RMS wavefront

error of a system Is given by the root sum square of the individual contributors:

w T =

i 1

For each individual refracting surface the RMS wavefront error contributor is

given by:

t

Wi, Refracting = (N2 - Nl) ERM S

C-43

i
where N2 - N1 . index difference across refracting surface

= 0.5 for "alr R spaced elements

and ERMS - RMS surface error specified for surface (In waves)

For each lndlvldual reflecting surface the RMS wavefront error contributor Is

given by:

2 Erm s

W i, Reflecting -

COS O

where ERMS , RMS surface error specified for surface (In waves)

e . incidence angle on mirror

It can be seen from these two expressions that the contributions to wavefront

error from mlrrors are far larger than for lens elements specified to the same

amount of figure error. Strain free mounting of the optics Is an important

requirement on the mechanical packaging of the system, particularly with regard to mirrors.

In the telescope part of the system there are two mirrors (primary and secondary) at near normal incidence, six refractive elements located away from the

telescope focal plane (the 4 element telescope collimator and the two element

refractlve corrector) and one lens In or very near the focal plane (telescope

field lens). The Impact of the field lens on transmit wavefront error can be

ignored due to the very small section of the clear aperture actually used.

Thls section has figure errors far smaller than the value specified over the

entire ciear aperture. In the root sum square process producing the total RMS

wavefront error of the system thls very small contributor Is swamped by the

C--44
much larger contributors from the collimator lenses and mlrrors. In the

lmaglng optlcs paths that carry the transmit beam to the telescrope there are

four mirrors at 45 ° tncldance, and 33 lens elements (not including 3 single

element field lenses).

If we designate E = a11oted RMS surface error then we have

Wrms,Tota I = E V Mo ° 2 + M45,2Vr'2 -+ L

where Mo = # of mirrors at near normal incidence

- 2 for telescope and Imaging optics

together

M45 = # of mirrors at 45 ° incidence

= 4 for telescope and imaging optics together

L = # of lens elements

= 39 for telescope and imaging optics together

Yielding WRMs, total = (7.37) E.

To meet a requirement of WRMs, total = X/14.1 suggests that surfaces

would need to be specified to above x/lO0 RMS (1.e. =x/20 Peak-toValley). This requirement Is within the state of the art for today's optical

fabrication technology. Performance like this has been successfully demonstrated on the Laser Crossllnk program. The fabrication and testing of surfaces

to this quality level Is considerably more labor intensive (and therefore

costly) than for the more commonly encountered k/4 or k/8 peak-to-valley

surfaces.

The above analysis does not address the other potential sources of wavefront

error. These other sources include glass tnhomogenetty, thermal gradients

over optical components, mounttng stresses, mtsallgnments and residual uncorrected destgn aberrations.

C-45
61ass lnhomogenelty effects are controlled by using only optical glasses

available In the better homogeneity grades. Fused slltca and the radiation

resistant (1.e, cerium doped glasses) are available In these better homogenelty grades for the small (_ 60 mm) lenses used In the OPTRANSPAC optlcal

system. In fact most of the lenses are < 25 mm In diameter. Homogeneity for

the large telescope mlrrors Is Inconsequential since the llght never transmits

through them. This Is one benefit of the reflective telescope over a refractive design.

Assessment of the effects of thermal gradients on the optics is not possible

wlthout a representative thermal model of the system. Typically, lmaglng optics subsystems can be kept at a fairly constant temperature, this is accomplished passively by the thermally conductive (typlcally aluminum) box enclosure. Thls tends to evenly distribute the heat throughout the sub-system.

If necessary one can also emplace an active heater and Insulation over the

imaging optics baseplate to ald the process (at the cost of the power needed

to run and control the heater). Imaging optics wavefront performance can be

degraded by the heat generated Internally by mechanisms or electronics boxes.

Thls can be kept under control by ensuring an adequate heat flow out of these

source locations coupled with Insulation of critical components from the heat

loading. The OPTRANSPAC lmaglng optics are Inherently Insensitive to

thermally or structurally induced mlsallgnments of the transmit and receive

paths due to the basic optical design which allows continuous polntlng readJustment based on offset tracking information.

Typically, telescope optics are more often affected by thermal gradients than

are imaging optics. The large aperture optics In telescopes can have far

larger temperature differences over their clear apertures than one would find

over a typical 25 mm diameter lmaglng optics lens cell. This can perturb the

figure of the mirror surface and degrade the transmitted wavefront. Passive

thermal control ls not always effectlve since the telescope faces directly out

lnto space (there must be a hole In any thermal shrouding to allow the transmitted and received beams to pass). The telescope prlmary to secondary mirror

alignment Is generally critical to the maintenance of a good transmit wavefront. Only slight (say 3 wm) motions can severely degrade the image

quality of a near diffraction limited system. Thls can happen from very small

-i

C-46
thermal gradients even when the telescope structure Is made from low expansion

materials such as ]nvar. If passive thermal control means are inadequate to

maintain the required image quallty active heating may be required. The dectslon _¢an only be made on the basis of a thermal model whose results are fed

back tnto the optics model to determine opttcal performance In the baseline

environment and configuration.

Mounting stresses can be mlmlmlzed by good mechanical design practices.

Stresses coupled lnto the opttcal component wlll distort surface quallty and

induce tndex changes In the glass. Both effects degrade the system image

qualtty. Stress free mountlng techniques lnclude component retention wtth

elastomertc agents, maintenance of sufficient clearance between cells and elements over thermal excursions, seating of components on either 3 point or

lapped multtpotnt pads, and the application of compressive rather than tensile

loads when loading of glass element Is required.

g

¢

Component mlsallgnment effects can only be calculated when the structural and

thermal modeltng can provide reasonable estimates of component motions over

environment. These motions can be inserted back tnto the optics model for an

assessment of the impact on transmit wavefront quality. Component mtsaltgnments can arise from differential expansion with temperature excursions or

from thermal gradients across the optical system structure. Mlsaltgnments can

also result from component shifts as a result of launch loads, spacecraft Internal vibrations, or mechanism vibrations coupled tnto the optical system

structure. Some mtsaltgnments can be butlt into a system whtle In the alignment process due to a lack of sufficient adjustment range or resolution of

motion on the optics mounts. The optlcs mount designs are based on a tradeoff

of the confllcttng requirements of adJustablllty and stability under launch

loading. During the alignment process the test equipment used to monitor

alignment must have sufficient resolution to measure in or out of spec alignment conditions. For instance the OPTRANSPAC optics that carries the near

diffraction ltmlted transmit beam will have to be aligned using an tnterferometer.

While exact budgeting of the wavefront error associated with mtsaltgnment must

await a better definition of the mechanical structure and thermal envlronment

C-47
one can st111 make some conclusions based on alignment sensltlvltles. It ts

anticipated that OPTRANSPAC, like most diffraction limited optical commun_cat_on systems, w_11 be sensitive to m_sal_gnments _n the telescope. The telescope mlsallgnment sensitivities are shown In Table 11. It can be seen that

this telescope Is fairly sensitive to despaclng of the prlmary and secondary

mlrrors. To maintain the specified wavefront quality the despace should remain well below _3_m. Should the despace exceed this value the active

focus wtll need to be uttlized to compensate for this added focus error. The

other m_sallgnments cause smaller effects and the solution Is the same.

Residual uncorrected wavefront errors can be left In the optlcal system by the

design process ltself. This wt11 leave a residual wavefront error In the optical system regardless of the fabrication process errors or lack thereof.

The way to control this ls to continue to Increase the complexity of the optlcal system and keep opttmlzlng unttl either all wavefront requirements are

withtn spec, the number of elements start to Increase weight and decrease

transmission to unacceptable levels, or the time or money spent exceed the

levels the contract w_11 suppport. The amount of residual design aberrations

and thelr Impact on wavefront error w111 have to awalt the performance of a

thick lens design and optimization of the OPTRANSPAC optical system.

The Image quality requlrements In the OPTRANSPAC receive paths are far less

stringent than those for the transmit path. Receive _mage quality ls drlven

by the tracklng requirements. Figure 17 Illustrates the variation _n the size

of the Earth Image on the tracklng detector as a function of range during the

m_sslon. Thls lmage slze vs range curve Is plotted parametrically as a function of receive path Image quality. Diffraction 11mlted performance Is not

necessary to meet the required tracklng uncertalnlty. An optics resolution In

the 15 - 25 urad range Is acceptable. This allows optlcs In the recelve

only portion of the lmaglng optics to have much less stringent surface quallty

requirements, say x/8 or _/10 peak-valley flgure error (wtth _ , 500 or

600 nm, I.e. st_11 measured In the visible).

C-48
8.0 FUTURE WORK

This design note represents the results of work completed under the current

contract toward design of the OPTRANSPAC optical system. To carry on the

development of this system under future contracts the following tasks will

need to be accomplished next:

(a) Thick Lens Oestgn and Optimization - The current thin lens design defined

herein w111 need to be converted to a thick lens prescription using the

optimization features of ACCOS V, Code V or an equivalent optical design

code. Thls is necessary to determine the number of optical elements

needed, determine nomlnal opttcal performance, determine required manufacturlng tolerances, determine alignment requirements, and define the

system In sufficient detatl for detailed mechanical design to proceed.

(b) Thermal-Structure-Optical Analysis - Thermal and structural analysis of

the optical and mechanical design will need to be conducted to determine

the expected range of component motions and anticipated thermal

gradients. These results ulll be folded back into the optical model to

calculate optical performance of the perturbed optical system. The wavefront budgets and alignment budgets can then be fully computed.

(c)

(d)

Breadboard Offset Tracking Configuration - To reduce risk and gain useful

engineering data it would be worthwhile to procure optics and CTO camera

and breadboard the offset tracking conflguraton. Thts wtll valldate performance of the optics and provide a useful test bed for evaluation of

electronics, computer Interfaces, and tracking software.

Baffle Design and Stray Light Analysts - The Cassegratn telescope baffle

system Is not yet designed. It needs to be designed and its performance

estimated using a stray light code such as APART or 6UERAP (or, at least

MIni-APART) to ensure tn specification solar rejection.

c-4g
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C-50
TELESCOPE CONFIGURATIONS

=,,.APERTURE STOP

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SPHERICAL

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/

(f) MAKSOUTOFF

4_F FOCAL

POINT

C-51

FIGURE
FIGURE 3

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C-70
TABLE 3 TELESCOPE PRESCRIPTION

EXAMPLE RITCHEY CHRETIEN WITH CORRECTOR LENSES

BASIC LENS DATA

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(

SURF RD

0 0.000000

1 0,000000

2 -32.5919597

3 -6.785295

4 -1,309862

5 -91.157784

6 -.619316

7 -,808265

8 2.395433

9 -4,736221

lO 0.000000

11 0.000000

REFRACTIVE INOICES

SURF NI

4 1.452590"

6 i.452590

8 1.452590

TH

• 755700£+10

13. 593)96

-13.593396

13 049691

271074

259634

543737

O. 000000

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3. 520722

-4. 885582

O. 000000

MEDIUM RN

AIR

AIR

REFL

REFL

SILICA 1.452S90

AZR 1.4B2590

SILICA 1.452590

AIR

SILICA 1.452590

AIR

AIR

_R

CC AND ASPHERIC DATA

SURF CC

2 -.10214E+01

3 -.25397E+01

DF

-52.364

-52.364

-52.36_

-52.364

N2 N3 N4 N5

I. 000000 1.0_}000 1. 000000 =1. 000000

1 . 000000 1. 000000 1 . 000000 i ,000000

1. 000000 1. 000000 1 •000000 I. 000000

AD AE AF AG

CLEAR APERTURES AND OBSTRUCTION

SURF TYPE

1 CIRCLE

1 (08) CIRCLE

2 CIRCLE

2 (08) CIRCLE

3 CIRCLE

6 CIRCLE.

7 CIRCLE

8 CIRCLE

9 CIRCLE

REF OB3 HT REF SURF

I

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.2920

4

0.00000

1.0000

IMG SURF

11

5

•00000

1. 0000

-.157700£+08 ( .12 OG)

EFL BF

145.4814 -4.8856

WAVL NBR 1

WAVELENGTH .53200

SPECTRAL N'r 1.0000

CAY CAX

s,4999

.9868

5.4999

.9428

.9568

.4321

.6168

.6796

.6824

REF AP HT 089 SURF

5,49990 0

F/NBR LENGTH

10.39 18.0652

2 3

1.06400 .e3500

1.0000 1.0000

NO APERTURE STOP

LENS UNITS ARE INCHES

EVALUATION MOOE IS FOCAL

C-71
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APPENDIX D

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INTERFACEDOCUMENT

z

D-]
OPTRANSPAC INTERFACE REQUIREMENTS

OPTRANSPAC STRUCTURE

• Defined in Figures 1 through 10

LOOK DIRECTION AND FIELD-OF-VIEW

• System pointed nominally towards earth +2 mrad

• Minimum unobstructed FOV: +2 mrad

MASS PROPERTIES

• Total weight less than 115 Ibs. (see Figure 11)

• Center of Gravity defined in Figure i

• Moving Parts

- Torque Motor Beam Steerers (4)

- Moment of Inertia: <5 x i0 -_ oz-in-se_ 2

- Maximum Angular Ac_leration: <2 x i0 D rad/sec 2

- Control Loop Bandwidth: 20 HF

INSTALLATION

• Mount: 3-point

° Mounting Surface Finish:

• Mounting Surface Flatness:

see Figure 1

see Figure 1

I

ALIGNMENT

• Installation Alignment: < ÷ 0.17 mrad

• Alignment Stability: < +--._0 _rad, variations assumed slow. Use of auto

coITimators to indlcat_d'_ift errors. Biases removed by control system.

ELECTRICAL POWER INTERFACES

• For details not given here, see OPTRANSPAC Design Note, "OPTRANSPAC

Electronics", by H. J. Mingo

° Bus Power: 28+6 volts direct current

• Total Power Requirement: 57 watts (see Figure ii)

• Telemetry - Analog, Serial, and Bi-level

• Command Control Data - Discrete and Serial

• Communication Data - <i00 Kbps Pulse Position Modulation (PPM)

{)-2
l

i

_--i_ _ ̧

THERMAL CONTROL

• S/C Interface - Adiabatic

• Operating Temperatures

Electronics

Electro-optics

Median

Operating

Temperature

30°C

20°C

• Survival Temperatures (Non-operating)

Electronics

Electro-optics

• Heat Rejection

Subassembly

Electro-optics

Detectors

Electronics

Variation

From

Median

+40°C

I

+3°C

_ Survival

Temperatures

35 o + 90°C

0° + 35°C

Power Dissipated (Watts)

6.0

5.0

46.0

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PATENTCONTRACT

EIC[e'T_'_ S _E _O

OF POOR QUALITY APPENDIX E

/

MCDONNELL DOUGL_

4

CORIIOI_ATION

EMPLOYE DISCLOSURE RECORD

TYPEWRITER

i A CONTINUOUSLASERCOMOPTICAL ALIGNMENTDEVICE

L. Casey Rt. I,

-O'Ji[ CLOCK 0_ =-O_,g

_DE_T E_ NO

"°301A/231 Lf3_-

Box 124,Marthasville,M0 63357 E.413 064552 8445

,2,

3

William

i

0% _V_AT DATE DI_ YO_ FIRST T'_INK OF T_'tlS I_VE%TION ON ITEM _ iWP_AT _ECOROS SHOW THIS _

INFORMATION 08 Feb 1985 Monthly Status Report MDAC-STL to CITIJPL 0PTRANSPAC

AND DATES Study - Attachment (I)

CONCERNING _,_E _ATE _ A_= _DE_-_,_v EA__,i[s'r S_TC_ O'_ o,_A,_,_,_

THIS ITEM Illustrated in above Monthly Status Report

WHE'_ W_ERPE A%_ TO /_HOM _ YO_ MAKE T_E _'I¢_ST OISC_.OS_E TO OTHERS 0_" _E _TEM

NK(_rD IN 1"HE [VI'N 'r Or 't' CON- El TH ER O_ A _L y _F:_ IN _VRI TIN G._

TEST OF" ¢_NlOff_?'# _I" INVleN'TION

.... , _._. o,,,_, o,,,_=. ,_ In above Monthly Status Report

N[CONDS CIT[_ S_OULO I£ DAY£D

AN n SrGN[D IV _SPONSIIL[

[_PL0_I[fS). _ND 0AT[D _ND

51GN[O Iv _wO INDEm£ND£_T

WI_NESSIES W_'IO I._AV£ IIEAD _ND

SALE OR

PUBLICATION

_BU,C _SE _ll _.t $1_CE _10

S:JC- DA_£

POTENTIAL MARKET

INFORMATION

,_£{__,[: r:m _OS_II_[ ,v_,l. II,,irT.

l;.t"_°°"£....... :£'"" "°

[mS

CONTRACT

INFORMATION

} (_E} HEREBY ASSIGN MY ;O_,J

_ESC=_ _IE CETA, LSO_" ANy _/O1_. O¢_ TESTS OO'_E TO _O_JCE OR OI=>ERATE T_E ITEM

C,_vE DATES A_,O _,_T'_ESSES '_SE O_'_EFI I=A.GES IF NECESSARY)

Filter technolomv/alimnment of laser to recei.ve path _nd.use of ,

qalvanome_er drf_en mfrrors _or aJl(_nment and po]nt-aneao _o DaCK many

OESCRIRE ANI_ _ VE :A_'ES (_1" ANY OTHI[R SKETCHI[S. ORAW{NGS OR RF.I_ORTS _E_'_INIF.NT ,TO _{S _TI[M.

years in Lasercom.

0nly sketches in above Monthly Status Report

IF" A _Ev_CE _AS BEE_ SOLO O_ _SEO _'OR PROF'" ' _EN A_O TO WHOM OI[LIVERE_ O¢1 _HE_ A%E_

MOW _SEE)'

Not yet sold or delivered

_A5 A OR_NTE_ OESCq_=T_ON OF" T_IS I*'I[M BEEN _AC_F AVAILABLE TO ¢I[RSONS OU_'SIP_E THE COMPANY'

me SO _O_ A_ ?'_E_ A_O _AS USE REST_ICTED_

Yes - To CIT/JPL in above referenced Monthly Status Report

, _o_E._,T: In Lasercom systems for space/air

:O_ME_:,AL In Lasercom systems for space/air

C _EN_IF'Y A_Y KNOW_ _I_MS O_ VENdOrS W_O _AY _E I_TE_ESTE_ IN TME ITEM

.... ,,, ,

F" THIS ITEM _AS F_ST CONCI[IVI[D OR CONST_CTI[O IN CO_NECTION _ITP_;

A. A GOVERNME%T CONTRACT GIvE GOVI[RNMENT CONTRACT NLJMBI[R AND SHOiI= OROI[R

B AN IRAD 0_ CO_APAN v D¢_OJEP_T GIVE PROJECT IDENTIF'ICATION N:JMIIER

_E,,,E,, _o.. E_L,,,, CITIJPL Contract No. 957601 for Design and

An 1 si S y. of _ac r f _.tic 1 TransceiverDackaqe.

O TATED CONTRACT OR SHOP_OROER?

Yes

I ENTI_E INTEREST IN Tt-_E ABOCE ENTI'fLI:'E: ' ITEM oR INVENTION, AS SHOWN ANO 0ESCRtBEC ON T_E

ACCOMI_ANYI"*IG SHEETIS) TO MCOONNELL OOL.IG_.AS CORI_OI_ATION.

SiGNAT_.JRI[!SI OF" EMI_LOYE{S: INVENTOR[S}: IuSE BLACK INK)

(I) _ 12)

=,,E Ig July 1985 oATE

(3)

SOC,ALSECURI'rYNO lqP-PCJ-gF,;)] ..,OCIA_.sEcu.,_ NO

NOTE: ANY OMIT_EO INF'ORMATION BI[COMING AVAILABLE AT A LATER TIM I II,.,IOUI,.E) TMIrN _i[ _l_OVlOi[{:)

TO THI[ I_ATENT 0EPARTMI[NT

E-1

M_C 13S-1 101 JUL sg)

OAT[

.SOCIAL SECIJIilTY NO.

mAGI[ I OF"
/

MCDONNELL DOUGLA 

PRINT OR USE

TYPEWRITER

CORPORATION

EMPLOYE DISCLOSURE RECORD

(DRAWING AND DESCRIPTION SHEETI

PROVII_E THE FOLL.OWING INI=OIItMATION C:ONCI[RNING THE I_tSCLOSE_ ITEM ANt) iN THE INI_ICATE_ SEQUENCE:

I, _lSCUSS THE PROBLEMS WHICH THE ITEM IS DESIGNED TO SO(.V_, REFERRING TO ANY PRIOR DEVICES OF A SIMILAR _JAT_.R_" &IT_ _*_l_.H

YOU MAY BE FAMILIAR.

2. STATE THE: AI_VANTA_ES OF THE ITEM OVER PRESENTly KNOWN I:_EVICES, SYSTEMS OR PROCESSES.

3. SPECIF'ICALLY _ESCRIBE THE ITEM AN_ ITS OPERATION. YOU MAY USE ANO ATTACH COPIES OF SKETCHES, PRINTS, PHOTOGRAPHS,

AN[_ ILLUSTRATIONS. WHICH SHO(_LO mE SIGNE_), WITNESSED AND [_ATEO. USE NUMBERS AND _ESCRIPTIVE _AME5 IN [_ESCRIPTION5

AND DRAWINGS,

4. LIST ALL KNOWN AND OTHER POSSIBLE USES F'OR THE ITEM,

S. LiST THE FF. AT'_IRES 0 ¢ Till[ iTF.M THAT AR£ BELIEVE_ TO BE NOVEL.

USE AS MANY OF THESE SHEETS AS NECESSARY AND ATTACH TO COMPLETEO FORM MOC I3S-I I01 JUL 69)

i. In Lasercom systems, the narrowness of the transmit laser beam requires that the laser

needs to continuously be coaligned to the track detector in the presence of thermal/

structural bending. If continuous realignment is not performed, the transmit laser beam

divergence must be increased to accommodate the maximum misalignment between the track

detector boresight and the laser. If this is not done, the laser beam may be misdirected

and the peak laser power will not impinge on the lasercom terminal to which the laser

beam is directed. Conventional quadrant detectors are capable of only one track point

so two beams cannot be tracked simultaneously. Option D of Attachment (1) illustrates

the operation of conventional realignment techniques as follows; when not tracking another

terminal, the shutter in front of the alignment cube is pulled out of the way and laser

energy is then reflected from the beamsplitter to the cube and retroreflected back through

the beamsplitter toward the tracker. The tracker, a quadrant detector receives the laser

signal and supplies error signals to the align/point-ahead beam steering mirrors to null

the track point. When it is necessary to provide a point-ahead function, the point-ahead

beam steerers (TMBSs) are driven to the calculated position which offsets the laser beam

to compensate for the velocity aberration of light. When the realignment is complete, the

shutter is closed and the terminal may function as a transceiver. Another realignment cannot

be performed until tracking is discontinued.

2m The proposed device offers the following advantages: ........

a. Continuous alignment is possible even when the receive wavelength is very Close

(20 angstroms) to the transmit laser wavelength. If polarization isolation is used,

continuous alignment should still be possible since polarization separation can be

performed at the device instead of spectral separation.

SlGNATUIqEIS} OF [MPLOY[{S}'INVENTORIS): [:)ATE

PAGE _ O F"
PRINT OR USE

, TYPEWRITER

ORIGI_._AL FI_E [3

OF POOR QUALITY MCDONNELL DOUGL

CORPd) RATION

EMPLOYE DISCLOSURE RECORD

rDRAWlNGAND DESCRIPTIONSHEET_

PROVIDE THE FOLLOWING INFORMATION CONCERNING THE DISCLOSED ITEM AND IN THE INDICATED SEQUENCE:

I. DISCUSS THE PROBLEMS WHICH THE ITEM IS DESIGNED TO SOLVE, REFERRING 1"O ANY PRI0¢_ _EVICES OF A S[MFLAR NATJ_IE 4'_ _:C_

YOU MAY lie FAMILIAR.

2. STATE THE ADVANTAGES OF THE ITEM OVER PRESENTLY KNOWN DEVICES. SYSTEMS OR PROCESSES,

3. SPECIFICALLY DESCRIBE THE ITEM AND ITS OPERATION. YOU MAY tJSE AN[_ ATTAC_ CO¢_IES OF SKETCHES. PRINTS. PHOTOGRAPHS.

AND ILLUSTRATIONS. WHICH SHOULD BE SIGNED. WITNESSED AND DATED USE NUMBERS AN_ :ESCRIPTIvE NAMES IN DESCRIPTIONS

AND DRAWINGS.

4. LIST ALL KNOWN AND OTHER POSSIBLE USES FOR TIME ITEM.

S. LIST THE FEATURES O F THE ITEM THAT ARE _ELIEVED TO BE NOVEL.

USE AS MANY OI ¢ THESE SHEETS AS NECESSARY A_ AT_AC)-_ TO COMPLETED FORM MDC 13(5-1 {01 JUL 6g)

b. An alignment shutter/drive is not required since continuous alignment

and the alignment cube may be continuously utilized.

is performed

c. Weight, size, power and complexity should be reduced.

r

d. Continuous alignment is possible when wide band sources are tracked since the

laser wavelength is "sliced out I'of the wide band with only minor transmission loss

e. Diffraction limited optical quality beams can be accommodated with only minor..

degradation.

i

i 3. The device works as follows, referring to Figure 1: Collimated energy of wavelengths

),_ and >,2 enter the device at point (A), pass through the block and reach the filter (B)

where the two wavelengths are spectral ly separated. Lambdal is transmitted through the

filter (B) and >,_ is reflected from the filter (B) toward the mirror (C). Mirror (C)

reflects the X2 energy at an angle that is not parallel to the ),_ optical axis. This

angular offset permits X_ and _,2 to be tracked at two separate points on the CCD

detector (F). The amount of angular separation determines the physical track point

offset x and is a function of system requirements. Lambda two energy, following

reflection from mirror (C)exits the block at point (D). Both X_ and >,2 pass through

the lens (E) and are imaged onto the CCD track detector (F). With a polarization

separation of orthogonal linear polarizations a polarization beam splitter cube or

equivalent polarization element replaces the filter but the device functions the same

as for spectral separation. The CCD tracker track points for the two signals can be

extracted for alignment and/or point-ahead functions.

SIGNATURE(S) OF EM_LOYE(SI/tNVENTOR(Sh

i

(2]

(M

(EACH SHEET MUST |[ SIGNED, DATED AND WITNESSED. USE BLACK INK.)

HOe ,.-2 (0' JU,. --} E-3

DATE: WITNES|ED. READ AND UN OERSTOOD J ...... |V: DATE:

PAGE OF
,

MCDONNELL DOUGL_z

CORl_)nAT#Oltl

EMPLOYE DISCLOSURERECORD

I DRAWiNG AND DESCRIPTION SHEETI

PROVIC)E THE FOLLOWING INFORMATION CONCERNING THE E)ISCLOSED ITEM AND IN THE INDICATED SEQUENCE:

I.

PRINT OR USE

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_ATEN _ _E=T _OCKET %0

DISCUSS THE PROBLEMS "WHICH THE ITEM IS {SE51GNED TO SOLVE, REFERRING TO ANY =RIOR DEVICES OF A StMILa, R NAT_;E _rT_q _ C_+

YOU MAY mE FAMILIAR.

2, STATE THE ADVAN""AGES OF THE ITEM OVER PRESENTLY KNOWN DEVICES, SYSTEMS OR PROCESSES.

S. SPECIFICALLY DEItERIImE Tile ITEM A_ ITS OPERATION, VOtj MAY LISE AN: AT"AEH COPIES OF SKETCHES, PRqNTS, ;=ilOTOGRA_HS

AND ILLUSTRATIONS, WHICil SHOLJL_ BE SIGNED. WITNESSEE_ AND E_ATEE) _JSE NUMBERS AN: _.'ESCRI_mTIVE NAMES IN DESCRIPTIONS'

AND {_R AWIN GS.

4. LIST ALL KNOWN AND OTHER POSSIBLE USES FOR THE ITEM.

S, LIST T_E FEATURES O L- THE ITEM THAT ARE '::IELIEVEE) TO _E NOVEL.

USE AS MANY OF THESE SHEETS AS NECESSARY AND ATTACH TO COMPLETED FORM MOC 136-I (01 JUL iSg}

4. Known or possible uses for the device include:

a. Continuous alignment of two narrow wavelength sources in a lasercom system.

b. Continuous alignment of a wideband source with a narrow wavelength source

embedded in it in a lasercom system.

c. Continuous separation of more than two sources by cascading two or more devices

and imaging onto a single CCD for a lasercom system.

d. Continuous alignment for laser systems other than lasercom,

e. Continuous alignment of polarization isolation lasercom systems by use of

polarization separation in lieu of spectral separation.

b. Separation/alignment continuously in a lasercom system or elsewhere when two

sources are as close as 20 angstroms spectrally.

. Features believed

a.

b.

C,

d.

e,

f.

to be novel:

Continuous alignment of a lasercom system.

Close spectral separation (20 angstroms) alignment of a lasercom system.

Continuous point-ahead function with closed loop correction. Implementation

normally open loop based on calculated values.

Continuous alignment of a lasercom system using polarization isolation.

Continuous alignment of a lasercom system where the transmitter wavelength

is emSedded in a broadband track signal such as an earth-tracker.

Any of the above not in a lasercom system.

is

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CORPOnATIORI

EMPLOYE DISCLOSURE RECORD

,DRAWING AND DESCRIPTION SHEET,

pFIOVIOE THE FOLLOWING INFORMATION CONCERNING THE DISCLOSED ITEM AND IN THE INDICATED SEQUENCE:

1, DISCUSS THE PROBLEMS WHICH THE ITEM IS DESIGNIEO TO SOLVE, REFERRING TO ANIY PRIOR DEVICES OF A S!M_LAR NA_/=E h!-P( h_ C_

YOU MAY DE IrAMILIAR,

2. STATE THE ADVANTAGES OF THE iTEM OVER PFIESENITLY KNOWN DEVICES. SYSTEMS OR PRCCESSES,

S. SPECIFICALLY DESCRIBE THE ITEM AND ITS OPERATION. YOU MAY USE AN_'_ AT_ACI-_ COI) IES OF SKETCHES, PRINTS. P_'_OTOGRAI=_)S

AND ILLUSTRATIONS. WI'_ICH 5HOIJLO _E SIGNED. WITNESSED AND DATE_, vSE NUMBERS A_D _ESCRIRTIVE NAMES IN DESCRIPTIONS

AND DRAWINGS,

4. LIST ALL KNOWN AND OTHER I_OSSIBLE USES FOR TIlE ITEM

5, LIST THE f"EATUI_ES O r THE ITEM THAT ARE 9ELIEVEO TEl _E NEIVEL,

USE AS MANY ElF THESE SHEETS AS NECESSARY AND ATTACH TO COMPLETED FORM MDC 13(S*1 _..',1 J_L (5_l)

PRINT OR USE

TYPEWRITER

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L
APPENDIX F

ACRONYM LIST

I

T

b _'

Ik

Al

APD

AU

BER

CCD

CID

cm

dB

EORS

FWHM

Hz

IOA

JPL

KBPS

Kg

Km

KPPS

KRADS

LSI

MeV

MHz

nwn

msec

wsec

mW

NASA

NASA-JSC

Nd:YAG

Aluminum

Avalanche Photodlode

Astronomical Unlt

Bit Error Rate

Charge Coupled Device

Charge Injection Device

centimeter

decibel

Earth Orbiting Relay Station

Full Width Half Maximum

Hertz

Imaging Optics Assembly

3et Propulslon Laboratory

Kiloblts per second

Kilograms

K11ometers

Kilopulses per second

Kllorads

Large Scale Integrated

Mega Electron Volt

Megahertz

millimeters

millisecond

microsecond

m1111watts

National Aeronautics and Space

Admlnlstratlon

NASA Johnson Space Center

Neodymium Yttrlum Aluminum Garnet

F-I
APPENDIX F

ACRONYM LIST

nm

NRZ

OPTRANSPAC

PCU

PMT

PPM

PPS

RMS

RS

RSS

sec

Si

SNR

TMBS

VLSI

nanometer

Non-Return to Zero

Optical Transceiver Package

Power Conditioning Unit

Photomultlpller Tube

Pulse Position Modulation

Pulse Per Second

Root Mean Squared

Reed-Solomon

Root Sum Squared

second

Silicon

Signal To Noise Ratio

Torque Motor Beam Steerer

Very Large Scale Integrated

F-2
REFERENCES

l . 'Orbital Measurements Of The Earth's Radiation Budget During The First

Decade Of The Space Program", William R. Bandeen, Earth Radiation

Science Seminars, NASA Conference Publication 2239, June 1980 to October

IgBl.

i )

!7

, "Lasercom Background Noise Sources", Informal Memo, S. D. Simmons,

McDonnell Douglas Astronautics Company - St. Louis Division, 3 Nov IgB2.

. "PPM Demodulation For Reed-Solomon Decoding For The Optical Space

Channel", D. Divsalar, R. M. Gagllardl, And J. H. Yuen, TDA Progress

Report 42-70, May - June IgB2.

. 'Spacecraft Attitude and Articulation Control Systems for Future

Planetary Missions," L. F. McGllnChey (JPL) and R. E. Rose (TRW), AIAA,

79-1717.

#

I

D •

!

!

!

I

I

!

| \

.

.

"Space Optical Communication With The Nd:YAG Laser", M. Ross, etal,

Proceedings of the IEEE, Vol. 66, No. 3 March 1978.

"An Advanced Tracker Design For Pointing And Control Of Space Vehicles

Using The Charge Injection Device", C. Jones, J. Kollodge, Annual Rocky

Mountain Guidance and Control Conference, January 30 - February 3, IgB2.

FI-1
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