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LMSC-HSV TR F320789-II 

Phase II 

Design Definition of the 


Laser Atmospheric Wind Sounder 

(LAWS) 

' Contract NAS8-37590 


DR-20 

Vol. II: FINAL REPORT 


November 1992 


Prepared for 

GEORGE C. MARSHALL SPACE FLIGHT CENTER 
MARSHALL SPACE FLIGHT CENTER, AL 35812 


'^Lockheed 

Missies & Space Company 


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4800 Bradford Blvd., Huntsville, AL 35807 


LMSC-HSV TR F320789-II 


DR-20 
PHASE II 

DESIGN DEFINITION OF THE 
LASER ATMOSPHERIC WIND SOUNDER (LAWS) 

VOL II: FINAL REPORT 

November 1992 

CONTRACT NAS8-37590 
Prepared for 

National Aeronautics and Space Administration 
George C. Marshall Space Flight Center (MSFC) 
Marshall Space Flight Center, Alabama 35812 


Prepared by ' • i 

D. J. WHson 

LAWS Deputy Program Manager 
Chief Engineer 


Date 



3 


Approved by 



W. E. Jones 

LAWS Program Manager 


Date /// z//7' 


Submitted by 


^Lockheed 

Missies & Space Company 


4800 Bradford Blvd., Huntsville, AL 35807 



LMSC-HSV TR F320789-II 


FOREWORD 


This document presents the final results of the 21-month Phase II Design Definition and 18- 
month laser breadboard efforts for the Laser Atmospheric Wind Sounder (LAWS). This work was 
performed for the Marshall Space Right Center (MSFC) by Lockheed Missiles & Space Company. 
Inc.. Huntsville, Alabama, under Contract NAS8-37590. The study was conducted under the 
direction of R.G. Beranek, NASA Program Manager, and R.M. Baggett, LAWS Instrument 
Project Office, JA92. The period of performance was 24 August 1990 to 30 June 199 . 
Subcontractors contributing to this effort are Textron Defense Systems - Everett, and Itek Optical 

Systems. 

The complete Phase H Final Report consists of the following three volumes: 

Volume I Executive Summary 

Volume II Final Report 

Volume HI Program Costs. 

Major contributions to this contract at Lockheed-Huntsville were made by T.K. Speer, G.R. 
Power Dr. S.C. Kurzius, Dr. W.W. Montgomery, Dr. W.R. Eberle, F.R. Davis, P.G. Porter, 
A J Condino, D.M. Tilley, R.E. Joyce, K.R. Shrider, W.M. Harrison, G.B. Washburn, B.J. 
Audeh, Dr. F. Wang, W. Dean, Z.S. Karu, J. Dyar, T.L. Sonnenberg, A.S. Stewart, J.T. 
Steigerwald, T.G. Larson, D.D. Coulter, J.C. Frost, R.G. Raney, and W.S. Johnson. 

At Textron Defense Systems-Everett, S. Ghoshroy, PM, was supported by Dr. H.P. Chou, 
F. Faria-e-Maia, I. Moran, H. Stowe, G. Crawford, M. Fava, M. Nguyen, and T. Christiano. 

Itek Optical Systems contributors were S.E. Kendrick, PM, C.M. Ullathome, and C. 
Robbert. 

Major contributions were also made by Dr. Carl Buczek, Laser Systems & Research Corp., 
and Dr. C. DiMarzio, Northeastern University. 


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LMSC-HSV TR F320789-II 


CONTENTS 

Section Pag e 

Foreword it 

Contents iii 

Illustrations v i 

Tables x 

Acronyms and Abbreviations xi 

1 INTRODUCTION AND SUMMARY 1-1 

2 SYSTEM ENGINEERING AND ANALYSIS 2-1 

2.1 Requirements 2-1 

2.2 Analysis and Trades (Phases I and II) 2-2 

2.3 Error Budget 2-3 

2.4 Risk Assessment 2-6 

2.5 Specification Requirements 2-12 

2.6 Interface Definition 2-14 

2.7 Reliability 2-14 

2.7.1 Parts Cost Consideration 2-14 

2.7.2 Manufacturing/Test Cost 2-14 

2.7.3 Summary 2-14 

3 PRELIMINARY DESIGN 3-1 

3 . 1 Overall Configuration and Accommodations 3-2 

3.1.1 Baseline LAWS 3-2 

3.1.2 Downsized LAWS 3-14 

3.2 Trades and Analyses 3-14 

3.3 Subsystem Designs 3-19 

3.3.1 Laser Subsystem 3-19 

3.3.2 Optical Subsystem 3-29 

3.3.3 Receiver/Processor Subsystem 3-40 

3.3.4 Structures and Mechanical Subsystem 3-50 

3.3.5 Attitude Determination, Scan Control, and 

Lag Angle Compensation 3-60 

3.3.6 Thermal Control Subsystem 3-72 

3.3.7 Electrical Power Subsystem 3-84 

3.3.8 Command and Data Management Subsystem 3-90 

3.4 Verification (Test and Evaluation) 3-97 

3.4.1 Development Test Plans 3-97 

3.4.2 Qualification Test Plans 3-97 

3.4.3 Acceptance Test Plans 3-100 


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CONTENTS 

(Continued) 

Section 

3.4.4 Prelaunch Validation Test Plans 

3.4.5 Documentation 

3.5 Operation Requirements and Scenarios 

3.6 Performance Analysis 

4 WORK BREAKDOWN STRUCTURE 

5 ENVIRONMENTAL ANALYSIS 

5 . 1 Actions and their Alternatives 

5 2 Environmental Impact ot the Actions and their Alternatives 

5.2.1 Prelaunch Phase 

5.2.2 Launch Phase 

5.2.3 On-Orbit Operations Phase 

5.2.4 De-Orbit Reentry 

5.3 Areas of Controversy 

5.4 Issues Remaining to be Resolved 

5.5 Conclusion 

6 LASER BREADBOARD 

6.1 Requirements 

6.2 Design 

6.3 Test Results 

6.3.1 Test Sequence 

6.3.2 Test Facility 

6.3.3 Results 

REFERENCES 

7 CONTAMINATION ANALYSIS ••••• 

7 . 1 Surface Contaminants Parameters 

7.2 Contamination Budgets 

7.3 Contamination Sources and Degradation Etfects 

7.3.1 Particulate Contaminants 

7.3.2 Molecular Contaminants 

7.3.3 Contamination Analysis 

7.4 Contamination Prevention and Containment Scheme.. 

7.4. 1 Design Considerations 

7.4.2 Personnel Training 

7.4.3 Operational Constraints/Guidelines 

7.4.4 Contingency Measures 


Page 

3-101 

3-102 

3-102 

3-102 

4- 1 

5- 1 
5-1 
5-1 
5-1 
5-1 
5-1 
5-2 
5-2 
5-2 

5- 2 

6 - 1 
6-1 
6-1 
6-8 
6-8 
6-8 
6-8 

6-23 

7-1 

7-1 

7-3 

7-6 

7-6 

7-6 

7-7 

7-7 

7-7 

7-8 

7-8 

7-8 


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LMSC-HSV TR F320789-II 


CONTENTS 

(Concluded) 

Section Page 

7.5 Contamination Analysis 7-8 

7.5.1 Launch Phase Contamination Concerns 7-9 

7.5.2 Orbital Operations Contamination Concerns 7-11 

APPENDIX A A-l 

APPENDIX B B-l 


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LMSC-HSV TR F320789-II 


ILLUSTRATIONS 


Section 

1 - 1 LAWS Science Requirements and Constraints 

1- 2 LAWS System Diagram 

2- 1 System Engineering Process 

2-2 LAWS System Functional Row Diagram 

2-3 LAWS Instrument Data Collected for Processing with the 

Science Team Algorithm 

2-4 Laser Frequency Variations Introduce LOS Wind Velocity Errors 
2-5 LOS Pointing Errors Introduce Errors into Wind Velocity 

Vector Measurements 

2-6 Signal-to-Noise Ratio Equation Used to Evaluate LAWS 

Instrument Performance 

2-7 Contribuung Factors for Maximized Signal-to-Noise Ratio 

2-8 Risk Assessment Process 

2- 9 ARTS Requirement Hierarchy 

3- 1 LAWS Subsystem Assemblies 

3-2 LAWS Baseline Design Flight Configuration 

3-3 Right Covers Removed 

3-4 LAWS Package on Bus Assembly 

3-5 LAWS Baseline Configuration 

3-6 LAWS Baseline Dimensions 

3-7 LAWS/POP in Atlas HAS Large Fairing 

3-8 Structure, LAWS Medium Base 

3-9 Opucal Bench Configurauon 

3-10 LAWS Optical Bench and Schematic 

3-11 LAWS Telescope Assembly 

3-12 LAWS Environmental Cover (Optical Bench) 

3-13 LAWS Signal Flow 

3-14 LAWS Baseline Current Mass Properties 

3- 1 5 LAWS in Delta Large Fairing 

3-16 LAWS Telescope with 0.75 m Diameter Mirror in Delta Fairing 

3-17 LAWS Instrument Fit-Check in Delta Fairing 

3-18 LAWS Downsized Mass Properties (6 April 1992) 

3-19 Selecuon of Pulse Repeution Frequency to Minimize Error in 

Wind Velocity Averaged over a Grid Square 


Page 

1-1 

1-2 

2-1 

2-2 

2-3 

2-4 

2-5 

2-6 

2- 7 
2-11 

2- 13 

3- 1 
3-3 
3-3 
3-4 
3-4 
3-5 
3-5 
3-7 
3-7 
3-8 
3-8 
3-9 

3- 10 
3-11 
3-13 
3-15 
3-15 
3-16 

3-17 


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LMSC-HSV TR F320789-II 


ILLUSTRATIONS 

(Continued) 


Section 

3-20 Effect of Pulse Repetition Frequency on Error in Averaged Wind 

Velocity with Variation in Wind Speed over Grid Square 

3-2 1 Trade Between Laser Pulse Energy and Telescope Diameter to 

Maximize SNR within Weight Constraints 

3-22 Laser Transmitter 

3-23 Laser Subsystem Block Diagram 

3-24 Resonator Optics Layout 

3-25 LAWS Discharge/Flow Loop, End View 

3-26 Two-Electrode Configuration, Side View 

3-27 Energy Discharge Processes 

3-28 Preliminary Layout of Pulsed Power Section 

3-29 Resonator Cavity Matching Control 

3-30 Auto-Alignment Functional Diagram 

3-31 Overview/Summary of the Laser Transmitter Subsystem 

3-32 Optical Subsystem Functional Flow Diagram 

3-33 Primary Mirror Design 

3-34 Reaction Structure Design 

3-35 Metering Structure Design 

3-36 Overview/Summary of the Optical Subsystem 

3-37 Alignment System Concept 

3-38 Two-Mirror Afocal Split Field Design 

3-39 LAWS Receiver/Processor Subsystem Block Diagram 

3-40 Receiver/Processor Layout 

3-4 1 Receiver/Processor Components - Side View 

3-42 Test Data 

3-43 Cryocooler Concept 

3-44 Vacuum Dewar with Cold Fingers, Detectors, and Pre-Amps 

3-45 Overview/Summary of the LAWS Receiver/Processor Subsystem... 

3-46 Overview/Summary of the Structures and Mechanical Subsystem 

3-47 Typical Mode Shapes 

3-48 Transient Response at Detector Due to Laser Firing Acoustic Shock. 

3-49 Transient Response at Detector Due to Laser Firing Acoustic Shock 

3-50 Transient Response at Telescope CG Due to Laser Firing 

Acoustic Shock 

3-5 1 Transient Response at Telescope CG Due to Laser Firing 

Acoustic Shock 


Page 


3-17 


3-18 

3-20 

3-20 

3-21 

3-23 

3-23 

3-24 

3-25 

3-26 

3-26 

3-28 

3-30 

3-31 

3-31 

3-32 

3-33 

3-34 

3-36 

3-42 

3-43 

3-43 

3-44 

3-46 

3-47 

3-49 

3-51 

3-55 

3-58 

3-58 

3-59 

3-59 


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LMSC-HSV TR F320789-II 


Section 

3-52 

3-53 


3-54 

3-55 

3-56 

3-57 

3-58 

3-59 

3-60 

3-61 

3-62 

3-63 

3-64 

3-65 

3-66 

3-67 

3-68 

3-69 

3-70 

3-71 

3-72 

3-73 

3-74 

3-75 

3-76 

3-77 

3-78 

3-79 

3-80 

3-81 

3-82 


ILLUSTRATIONS 

(Continued) 


LAWS Telescope Attitude/Balance Sensitivity 

Attitude Determination, Scan Control, and Lag Compensation 

Implementation 

Vlos Pointing Factor Errors 

Attitude Determination Functional Diagram 

Star Tracker View Traced Out Over an Orbital Period 

Attitude Determination Hardware Performance Tradeoft 

Preliminary Hardware Specifications tor Attitude Determination 

Summary of Components for Attitude Determination Preliminary Design.... 

Lag Compensation Functional Diagram 

Transmit-Receive Error Budget Tree 

Acceptable Boundary for Platform Attitude Jitter PSD 

Alignment Loop Representation for Stability Analysis 

Overview/Summary of the LAWS Thermal Control System 

LAWS Power and Thermal Load Schedule 

LAWS Active TCS Schematic 

LAWS Coolant Pump Package Schematic 

Coolant Line Layout 

Thermal Radiation Model Plot of LAWS Instrument Showing Passive 

TCS Surface Coatings 

Thermal Radiation Model Plot of LAWS Instrument Telescope 

Showing Passive TCS 

Overview/Summary of the LAWS Thermal Control System 

LAWS PDS 

Overview/Summary of the Electrical Power Subsystem 

LAWS Functional Hierarchy 

LAWS Flight Data Management Functional Hierarchy 

LAWS System Functional Flow Diagram 

LAWS Software Tree 

Overview/Summary of the Command and Data Management Subsystem — 

Vehicle Qualification Tests 

Component Qualification Tests 

Right Unit Acceptance Tests 

Signal-to-Noise Ratio Equation Used to Evaluate LAWS 

Instrument Performance 


Page 

3-60 

3-63 

3-63 

3-64 

3-65 

3-65 

3-66 

3-66 

3-68 

3-69 

3-69 

3-71 

3-73 

3-75 

3-77 

3-77 

3-79 

3-81 


3-81 

3-83 

3-85 

3-89 

3-91 

3-92 

3-93 

3-94 

3-98 

3-99 

3-100 

3-101 

3-105 


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LMSC-HSV TR F320789-II 


Section 

3-83 

3-84 


3-85 

3-86 


4-1 

6-1 

6-2 

6-3 

6-4 

6-5 

6-6 

6-7 

6-8 

6-9 

6-10 

6-11 

6-12 

6-13 

6-14 

6-15 

6-16 

6-17 

6-18 

6-19 

6-20 

6-21 

6-22 

6-23 

6- 24 

7- 1 
7-2 
7-3 
7-4 
B-l 


ILLUSTRATIONS 

(Continued) 

Contributing Factors for Maximized Signal-to-Noise Ratio 

Survey Mode Shot Pattern Showing Forward Looking and 
Aft Looking Shots 

Maximum Power Available for LAWS Experiment, Sun Synchronous 

Orbit, 0600 Descending Node Crossing 

Shot Schedule for Design Mode with Instrument Power 

Limited to 4800 W 

WBS 1.0 LAWS Instrument 

Breadboard Test Configuration 

Breadboard Test Configuration/Resonator Layout 

Pulse Power System Integration and Checkout 

Integration and Checkout of Flow Loop/Discharge/Pulse Power Units. 

Integration Plan for Resonator/Injection Assemblies 

Integrated LAWS Laser Breadboard 

System Ground Plane 

Test Plan Schedule 

Laser Breadboard System 

Breadboard Test Results Compared to TDS Code Predictions 

10 Hz Single Mode Operation, 50 Pulses Superimposed 

Chirp Measurement from Fast Fourier Transform, Measurement A 

Chirp Measurement from Fast Fourier Transform, Measurement B 

Beam Jitter from Pulse-to-Pulse: Much Less than Our 80 pm 

Measurement Resolution 

Single Pulse Energy Monitored by Scientech Joule Meter 

Typical Discharge Current and Voltage Waveforms 

Schematic Diagram of Experimental Apparatus 

Decay Rate of Small Signal Gain 

Deactivation Rate Constant on Temperature 

Temporal Variation of Gain: Comparison of Experimental Data to 
Code Prediction 

Energy Extraction Data for 12 C 18 02 Mixture 

Single Mode (L&T) Pulse at 9.11 pm 

Heterodyne Beat Signal at 19 kV 

Current Voltage Pulse from Pulse Forming Network 

Typical Particulate Contamination Budget Allocation 

Typical Molecular Contamination Budget Allocation 

Vibroacoustically Induced particle Redistribution 

LAWS Contamination Math Model 

10 J Output Demonstrated at 10 Percent Efficiency 


Page 

3-106 

3-108 

3-108 

3-109 

4-2 

6-2 

6-3 

6-4 

6-4 

6-5 

6-6 

6-7 

6-7 

6-9 

6-11 

6-11 

6-12 

6-12 

6-13 

6-14 

6- 15 

6- 17 
6-18 
6-18 

6- 19 
6-20 
6-21 
6-21 
6-22 

7- 4 

7- 5 

7- 11 
7-12 
B-3 


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ro to 


LMSC-HSV TR F320789-II 


TABLES 

Table Pa 2 e 

2- 1 Probability of Failure - Maturity 2-9 

-2 Probability of Failure - Complexity 2-9 

-3 Probability of Failure - Dependency on Other Factors 2-9 

2- 4 Consequences of Failure (Cf) 2-10 

3- 1 Potential Launch Vehicles 3-12 

3-2 Optical Design Characteristics 3-30 

3-3 Optical Component Data 3-32 

3-4 Receiver Channel Wavefront Error (WFE) Sensitivities (Rigid Body 

Alignment Errors) 3-37 

3-5 Transmitter Channel WFE Sensitivities (Rigid Body 

Alignment Errors) 3-37 

3-6 Receiver Channel LOS Error Sensitivities (Rigid Body 

Alignment Errors) 3-38 

3-7 Transmitter Channel LOS Error Sensitivities (Rigid Body 

Alignment Errors) 3-38 

3-8 Orbital Thermal Analysis Summary 3-40 

3-9 LAWS Natural Frequencies and Mode Shapes Telescope Motor 

Bearing Supported (Caged) 3-54 

3-10 Interface Reaction Loads 3-56 

3- 1 1 Static Deflections 3-56 

3- 1 2 Critical LAWS Attitude Pointing and Stabilization Requirements 3-61 

3- 1 3 Features of Attitude Control Preliminary Design 3-62 

3-14 Active vs. Passive Control of Platform and LAWS Jitter 3-68 

3-15 LAWS Electrical/Thermal Load Summary 3-74 

3-16 Results, LAWS Active TCS Coolant Temperatures 3-78 

3-17 PDS Commands 3-88 

3-18 Operating Modes, Mission Phases, and Support Requirements 3-103 

3- 19 LAWS Operational Characteristics Constrained by Available Power 3-110 

6-1 Laboratory Support Equipment 6-5 

6-2 Test Parameters 6-17 

6- 3 Summary of (001) Vibrational Relaxation Rate Constants 6-19 

7- 1 LAWS Optical Elements 2-2 

7-2 Total Transmission Efficiency for Several Loss Factors 7-3 

7-3 Tentative Particulate Contamination Budget 7-5 

7-4 Tentative Molecular Contamination Budget 7-6 

7-5 LAWS Contamination Evaluations 7-9 

7-6 Analytical Tools for Contamination Analysis 7-14 


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LMSC-HSV TR F320789-II 


ACRONYMS AND ABBREVIATIONS 


A/D 

ADP 

AGC 

ARTS 

ASE 

BDU 

BW 

CART 

CCAM 

CCSDS 

C&DH 

CEI 

CG 

CIL 

CPCI 

CPDP 

CTE 

CVCM 

DPA 

DVT 

EEE 

El 

EMC 

EO 

EOS 

EPS 

ESC/ESD 

EU 

FFT 

FMEA 

FOSR 

GFE 

GIDEP 

GIIS 

GSE 


analog-to-digital 

acceptance data package 

automated gain control 

automated requirements traceability system 

airborne support equipment 

bus data unit 

bandwidth 

condition of assembly at release and transfer 

collision/contamination avoidance maneuver 

Consultative Committee for Space Data Systems 

command and data handling 

contract end item 

center of gravity 

critical items list 

computer program configuration item 
computer program development plan 
coefficient of thermal expansion 
collected volatile condensable materials 
destructive physical analysis 
design verification test 
electrical, electronic, and electromagnetic 
equipment item 
electromagnetic compatibility 
electro-optical 
Earth Observation System 
electrical power subsystem 
electrostatic compatibility/electrostatic discharge 
engineering unit 
fast fourier transformer 
failure mode effects analysis 
flexible optical solar reflector 
Government furnished equipment 
Government-Industry Data Exchange Program 
General Instrument Interface Specification 
Ground Support Equipment 


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LMSC-HSV TR F320789-II 


ACRONYMS AND ABBREVIATIONS 

(Continued) 


HOSC 

Huntsville Operations Support Center 

H&S 

health and status 

HST 

Hubble Space Telescope 

IARM 

input annular reference mirror 

IAS 

integrated alignment sensor 

ICD 

Interface Control Document 

IMU 

inertial measurement unit 

LAEPL 

LAWS Approved EEE Parts List 

LAWS 

Laser Atmospheric Wind Sounder 

LC&DH 

LAWS Command and Data Handling 

LO 

local oscillator 

LOS 

line-of-sight 

MA 

multiple access 

MAPTIS 

Material Processing Information System 

MCS 

Manufacturing Control System 

MLI 

multi-layer insulation 

MUA 

Material Usage Agreement 

NSPAR 

Nonstandard Part Approval Request 

OARM 

output annular reference mirror 

OR 

obscuration ratio 

PA 

product assurance 

PCP 

platform command processor 

PDS 

power distribution system 

PDT 

product development team 

PFN 

pulse forming network 

PIND 

particle impact noise detection 

PLF 

payload fairing 

PMP 

program management plan 

POCC 

Payload Operations Control Center 

PRACA 

parts problem reporting and corrective action 

PRF 

pulse repitition frequency 

PRL 

program requirements list 

PSATS 

parallel spacecraft automated test system 

PZT 

pezio-electric transducer 


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LMSC-HSV TR F320789-II 


ACRONYMS AND ABBREVIATIONS 

(Concluded) 


RCS 

reaction control system 

rms 

root mean square 

R&RR 

range and range rate 

SA 

single access 

SBA 

scan bearing assembly 

SLM 

single longitudinal mode 

SMS 

structures and mechanical subsystem 

SN 

space network 

SNR 

signal-to-noise ratio 

SQU 

Structural Qualification Unit 

STDN/DSN 

Spaceflight Tracking and Data Network/Deep Space 
Network 

ST&LO 

system test and launch operations 

STV 

structural test vehicle 

TAP 

transportation adapter plate 

TCS 

Thermal Control System 

TDRSS 

Tracking and Data Relay Satellite System 

TML 

total mass loss 

TWG 

test working group 

ULE 

ultra-low expansion 

VCRM 

verification cross reference matrix 

WFE 

wavefront error 

WSMC 

Western Space and Missile Center 


xiii 

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LMSC-HSV TR F320789-II 


Section 1 

INTRODUCTION AND SUMMARY 

Lockheed personnel, along with team member subcontractors and consultants, have 
performed a preliminary design for the LAWS Instrument Breadboarding and testing of a LAWS 
class laser have also been performed. These efforts have demonstrated that LAWS is a feasible 
Instrument and can be developed with existing state-of-the-art technology. Only a commitment to 
fund the Instrument development and deployment is required to place LAWS in orbit and obtain the 
anticipated science and operational forecasting benefits. 

The LAWS Science Team was selected in 1988-89 as were the competing LAWS Phase I/II 
contractor teams. The LAWS Science Team developed requirements for the LAWS Instrument, 
and the NASA/LAWS project office defined launch vehicle and platform design constraints. From 
these requirements and constraints, several of which are listed in Figure 1-1 , the Lockheed team 
developed LAWS Instrument concepts and configurations. A system designed to meet these 
requirements and constraints is outlined in Figure 1-2. 


Constraints 


Platform 

• Power Resources 
Average 

Peak 

• Thermal Resources 

- Cooling & 

Exposure 

• Instrument Resources 
attitude & position 

• Orbit Parameters 
Altitude & Inclination 

• Structure 

- Envelope &Mass 
• Vibration Spectra 

- Deformation 

• Contamination 

F312SS0-OWB42 


Atmosphere 

• Aerosol seeding (lO '^m' 1 sr -1 ) 

• Attenuation 

« Turbulence effects 

- Coherence decorrelation 
time (1 , 2, - - 5 ps) 

- Velocity variability 
over grid 


i — 

Science Requirements 

• Tropospheric winds 

• > 6 pulses/horizontal 
100 x 100 km 

• < 1 km vertical res. 

• System error 
contribution limits <1 m/s 
line-of-sight (<5 m/s for 
low aerosols) 

• Global coverage 

• Eye safe 

• 5 yr. life 

• 10 9 Laser pulses 



Figure 1-1. LAWS Science Requirements and Constraints 


1-1 

LOCKHEED- HUNTSVILLE 





LMSC-HSV TR F320789-II 


F312599-TKS- 02 


Laser Subsystem 


Optical Subsystem 



Comm 


Thermal 

Cooler 




JComm 





Artrtude/Poertion 

Signal 

Prooeaaor 


Flight 

Computer 

— ► 

Determination 




Command 
4. Data 


I 


r— i 

1 



r 

„ — . — l L — — 

, — — — — 





Structures & Mechancat Subsystem 



Electrical Power Subsystem 

Ebdncai Power 
Distribution 
■ Cable Hameta 
• Junction Boxss 

i 

I Ptatlorm Power 

Low Level Reference Output 

High Powsr Laser OutpU 
Backacatter Return Signal 
Electnca^Electronic Signal 
Distant Star 

Electromagnetic Data Link 


Figure 1-2. LAWS System Diagram 

Figure 1-2 identifies the LAWS primary subsystems and interfaces - laser, optical and 
receiver/processor - required to assemble a lidar. The figure also identifies the support subsystems 
required for the lidar to function from space: structures and mechanical, thermal, electrical, and 
command and data management. The Lockheed team has developed a preliminary design of a 
LAWS Instrument system consisting of these subsystems and interfaces which will meet the 
requirements and objectives of the Science Team. 

This final report provides a summary of the systems engineering analyses and trades of the 
LAWS (Section 2). Summaries of the configuration, preliminary designs of the subsystems, 
testing recommendations, and performance analysis are presented in Section 3. Sections 4 and 5 
discuss environmental considerations associated with deployment of LAWS. Finally, the 
successful LAWS laser breadboard effort is discussed in Section 6 along with the requirements and 

test results. 

The Lockheed team baseline LAWS Instrument meets all Science Team requirements. The 
Instrument design is compatible with the Atlas HAS and, with minor modifications, the Delta 
launch vehicles. It is also compatible with the MSFC strawman orbital platform. 


1-2 

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LMSC-HSV TR F320789-II 


The team has also investigated the downsized LAWS Instrument, i.e., from a 15 J to 
5 J/pulse laser and from a 1.6 m to a 0.75 m aperture telescope. This Instrument can be 
developed and orbited at a somewhat reduced cost from the baseline LAWS. Our laser breadboard 
has already been operated at this reduced energy output, and wall plug efficiency, pulse frequency 
chirp, and performance have been demonstrated to meet these downsized Instrument requirements. 

The Lockheed team is ready to proceed with an aggressive program to orbit a LAWS 
Instrument in the near future. After performing these analyses, design studies, and laser 
breadboard development, we foresee no technical challenges to disrupt the early deployment of 
LAWS. We recommend an aggressive 18-month effort in testing the laser breadboards and 
optimizing detector performance, followed immediately with a Phase C/D program leading to an 
early year 2001 launch. 


1-3 


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Section 2 

SYSTEM ENGINEERING AND ANALYSIS 

The complexity and sophistication of the NASA LAWS Instrument, including its integration 
with the satellite platform bus, booster, and supporting ground and space systems, require a 
systematic application of a sound system engineering process. This process was applied by 
Lockheed to develop the LAWS Instrument configuration during Phases I and II as shown in 

Figure 2-1. 


ERROR BUDGET - DR-13 

CEI-OR-IO 

IRD - DR-9 




TRADES 


PREPARE 

SPECIFICATIONS 


INPUT 

REQUIREMENTS 




i FUNCTIONAUvi: 
ANALYSIS 


SYNTHESIS)-* OB 


EVALUATION i 
AND |-*<? 

DECISION ! 


J>1 


DESCRIPTION OF 
SYSTEM ELEMENTS 


PREPARE 

SUPPORTING 

PLANS 


0 


PRELIMINARY 

DESIGN 

DOCUMENT 

DR-8 


<7 


j. COST 
D ESTIMATE 
OR -6 


<7 


SE&I - DR-7 

WBS/DICTIONARY - DR-5 
PROJECT IMPLEMENTATION - DR-4 
SCHEDULE - DR-9 

ENVIRONMENTAL ANALYSIS - DR-17 
REQUIREMENTS/CONFIGURATION - DR-14 


Figure 2-1. System Engineering Process 

2.1 REQUIREMENTS 

Performance requirements, established by NASA and the Science Team, were analyzed by 
Lockheed and its subcontractors. These requirements were organized, flowed down, and allocated 
to different LAWS System functions as shown in Figure 2-2. As these requirements were 
accumulated, identified, and quantified, they were entered into a Lockheed developed computerized 
data base system known as the Automated Requirements Traceability System (ARTS). ARTS 
permits easy access for updating existing requirements and for adding new requirements as they 
are identified. Specification formats, compatible with the requirements of MM 8040. 12A, are 
included in this data base program; these formats can be selected as the requirements are printed as 
different types of specifications and interface control documents. Requirements collected by this 
process are listed in the Prime Equipment Detail Specification (DR- 10). 


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Platform 

Transceiver 


Attitude/Position 

Determination 


LEGEND 


Communicate Data 
and Commands 


Obtain Orbital 
Parameters, Time, 
and Attitude 


Flight 


Allocation 



Store 

Data 


— 


— 

L 


r 

Process Data and 
Commands 

1 

i 

Monitor Health, 
Safety, and Status 
Data 




— 





res 

Control Instrument 
Temperature 


EPS Distribution 


Laser 


To All 


Scan Drive 


Generate Laser 
Beam 


Control Scan Position 
and Alignment 


Telescope 

Expand, Direct, and 
Receive Beam 


Laser Output 
Beam _ 


Backs carter 
Return 


Detector 


Distribute 


u 

Detect and Process 

Electrical Power 


- 

Return Signals 


TDRSS 

SateHit&Ground 

Transceivers 


►To All 


POCC 


Transmit and 
Receive Data 


Decode and Monitor 
data quality 


Payload 

Developer 


Monitor 

Instrument 

Performance 


Science 
Team 

Evaluate Instrument 
Science Data 


Figure 2-2. LAWS System Functional Flow Diagram 

2.2 ANALYSIS AND TRADES (PHASES I AND II) 

The functions shown in Figure 2-2 were individually analyzed to identify each internal 
subfunction performed to achieve each assigned performance requirement. Interfaces with other 
functions were analyzed to determine how each function could best be accomplished. These 
analyses also allowed identification and evaluation of available approaches that could be 
synthesized by proven hardware and/or software techniques to implement the requirements. 

The results of these analyses were evaluated to determine performance compatibility and to 
establish requirement limits which were entered into the ARTS data base record. When multiple 
approaches were identified, trade studies were conducted to select the one best suited to perform 
the required function. 


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2.3 ERROR BUDGET 

The LAWS Instrument will collect large amounts of data from a satellite platform in a sun- 
synchronous orbit about the Earth. Additional atmospheric phenomena data will be collected from 
other sensors. These data will be processed with algorithms developed by the Science Team. 

Interaction between the collection and processing of these data to produce wind information 
is shown in Figure 2-3. Each block of the error tree is assigned an identification number. These 
identification numbers allow each of the parameter variation effects to be traced from the bottom of 
the error tree to the top, where the results of all effects are integrated. 

The LAWS Instrument errors are represented by laser frequency factors, pointing factors, 
and signal-to-noise factors as shown in Figure 2-3. Statistical data, produced by selected shot 
management modes of operation, are also recorded for input to the statistical sampling algorithm. 
Two types of data are supplied for input to the velocity algorithm. These data are related to 
pointing errors and to factors that affect the signal-to-noise ratio (SNR). 


1.0 



Attenuation 312500-29 

Clouds 

Turbulence 

Shear 

Refraction Effects 
Speed of Light 
Beam Bending 


Figure 2-3. LAWS Instrument Data Collected for Processing with the Science Team Algorithm 


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Errors introduced by undesired variations in the laser frequency are shown in Figure 2-4. 
These errors appear as incorrect shifts in the LOS doppler velocity measurement values. Pointing 
factor error sources are shown in Figure 2-5. The angular error values and the equivalent velocity 
errors are also shown in this figure. 

The ability to extract doppler shifted velocity information from low level signals that contain 
high levels of noise provides a useful measure of LAWS system performance. Because of the low 
signal levels expected to be received by the LAWS Instrument from suspended aerosols, design 
efforts are required to maximize the effective SNR. An SNR equation, recognized by NASA and 
members of the Science Team, is shown in Figure 2-6. This equation includes LAWS Instrument 
parameters which can be controlled by design to maximize the Instrument SNR. Factors which 
contribute to the maximization of the SNR are shown in Figure 2-7. The LAWS Instrument Error 
Budget Report was delivered to NASA as DR-13. Note that the SNR equation presented in Figure 
2-6 contains the pulse length (which is controlled by the contractor) and not the processing 
bandwidth (controlled by the Science Team). As such, this narrow band SNR is -14 dB greater 
than the wide band SNR. 



312500-30 


Figure 2-4. L as er Frequency Variations Introduce LOS Wind Velocity Errors 


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2-5 

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Figure 2-5. LOS Pointing Errors Introduce Errors into Wind Velocity Vector Measurements 


















LMSC-HSV TR F320789-II 


„ 1 »D 2 , CT . [Absorption Effects] |Efficiwlcjesl 

SNR = h7 7 r 2 J T" P [Turbulence Effects] 1 


Where: 


h v = Photon Energy = 2.18E - 20 J (for 9.11 urn) 

2 

rc D . = Aperture Area 

4 

j = Pulse Energy 

CT = Pulse Half Length (for Distributed Target) 

2 

R = Range to Target 
0 = Backscatter Coefficient (Given) 

Absorption Effects (Given) 

Turbulence Effects (Small Number at these Ranges) 
= Combined Efficiencies 

For LAWS 


t| = t| Transmit 
Optics 


. ti Receiver 
Optics 


. r\ Heterodyne 
Efficiency 


ti Effective 
Quantum 
Efficiency 


^ ^ , 312500-32 

Reference: EB23/W. Jones, November 1990, Modification for Turbulence to 
D. Emmitt's October 1990 memo. 

Figure 24. Signat-to-Noise Ratio Equation Used to Evaluate LAWS Instrument Performance 

2.4 BISK ASSESSMENT 

Lockheed is very sensitive to risk factors involved in the development, fabrication, testing, 
and extended, unattended operation of the LAWS Insmiment in space. Because of this 
Lockheed has selected a risk assessment technique that has proven to be effective on other 

successful Lockheed space programs. 

Three interrelated elements associated with program risks for the LAWS Program are 
technical performance, cost, and schedule. Recognition and identification of potermal program 
risks are the first steps required to circumvent or minimize problems that could seriously 
outcome of the program. This analysis begins with three steps: 

. Identification of potential hardware, software, and support system risk elements using a 
structured approach to ensure that all system areas have been considered 
. Quantitative assessment of the risk and ranking of items to determine those of most concern 
. Definition of alternate paths to minimize risk and establish criteria for .nutation or 
termination of these activities. 

The LAWS Instmmen. Work Breakdown Structure (WBS) is used for evaluation purposes to 
identify possible development risks for every element of the program down to 
(other than elements listed under Project Management). The risk assessment employed considers 
two factors: probability of failure (Pf) and consequence of failure (C F ). 

2-6 

LOCKHEED- HUNTSVILLE 



Maximized Signal to Noise 
Ratio Factors 



Figure 2-7. Contributing Factors for Maximized Signal-to-Noise Ratio 




























































LMSC-HSV TR F320789-II 


Pp considers the technical risks associated with a hardware or software item s potential 
failure to achieve technical performance specification requirements due to the item s state of 
maturity, degree of complexity, or dependency on interfacing items. Hardware and software 
designs are evaluated to determine whether potential technical problems exist, and the extent of 
these problems. Pf is obtained from the ratings given in Tables 2-1, 2-2, and 2 -3 for the five 
different problem categories as follows: 


P F = 


P «. + P M.. + P Q.+ P C^ P D 


where 

p M« 

Pmsw 

PCh 

p csw 

Pd 


= Probability of failure due to degree of maturity of hardware 
= Probability of failure due to degree of maturity of software 
= Probability of failure due to degree of complexity of hardware 
= Probability of failure due to degree of complexity of software 

= Probability of failure due to dependency on other items. 


Where no software is involved, those two factors are omitted, and the denominator becomes 3. 

The Cf factor considers the impact on the LAWS Instrument system if an item fails to meet 
technical, cost, or schedule requirements. The Cf is determined by using values given in Table 2- 
4 for the three factors (technical, cost, and schedule) and calculating the average of these factors. 


Cp+Cp +Cp 
CF= S 4 — ^ 


where: 


Cfj =Consequences of failure due to technical factors 
Cp c = Consequences of failure due to changes in cost 

CFs = Consequences of failure due to changes in schedule. 

The Risk Factor (Rf) is calculated using the equation: 


r f = Pp + Cf - Pf x Cf- 

The risk evaluation process is shown in Figure 2-8. Risks are ranked from minimal to high 
according to established criteria, as in the following example: 

Rf < 0.3 risk is low 

R f > 0.3 < 0.7 risk is medium 

Rf > 0.7 risk is high. 


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Table 2-1 . Probability of Failure - Maturity 



Rating 

Hardware (Pm^) 

Software (PMg W ) 


0.1 (low) 

Off-the-shelf items; no new hardware 
required 

Existing, proven software or no new 
software required 

— 

0.3 

Minor redesign of proven hardware 

Some slight change in existing S/W; minor 
change in modules/lines of code 


0.5 

Technical feasibility established: change in 

Major change in existing S/W 

— 

design and performance requirements of 
existing hardware 

modules/lines of code 

— 

0.7 

Undergoing exploratory development; 
complex design and performance 
requirements; technology available 

New software; software similar to existing 
programs 


0.9 (high) 

Very limited experience; some research 
performed; significant change in state-of- 
the-art 

New software; programs pushing state-of- 
the-art 



Table 2-2. Probability of Failure - Complexity 


Rating 

Hardware (Pch) 

Software (Pcsw) 


0.1 (low) 

Simple design; no changes required or 
not applicable 

Simple design; no changes required, or 
not applicable 


0.3 

Minor increase in complexity or 
performance requirements 

Minor change in program complexity 


0.5 

Moderate increase in complexity or 
performance requirements 

Large increase in program complexity 

— 

0.7 

Significant increase in complexity 

Significant increase in program complexity; 
major increase in modules 

— 

0.9 (high) 

Extremely complex system 

Highly complex program; very large data 
bases and complex, rapidly operating 
executive programs 




Table 2-3. Probability of Failure - Dependency on Other Factors* 


Rating 

Description 

— 

0.1 (low) 

Independent of system/facility or associate contractor's performance or schedule efforts 


0.3 

Dependent upon the schedule for modification of existing system or facility to meet 
requirements 


0.5 

Dependent upon the performance, capacity or interface of system or facility to meet 
requirements 

— 

0.7 

Dependent upon the schedule for assembly and test of other items or the system to 
meet requirements 


0.9 (high) 

Dependent upon the performance of hardware/software or of interfaces of the system 
to meet requirements 


* Factors include other group hardware/software performance, interfaces, schedule, and availability. 


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Table 2-4. Consequences of Failure (Cf) 


Rating 

Technical (Cf t ) 

Costs (Cf c ) 

Schedules (Cf s ) 

0.1 (low) 

Minimal or no 

consequences; 

unimportant 

Budget estimates not 
exceeded; some transfer 
of monies 

Negligible impact on 
program; slight 
development schedule 
change compensated by 
available schedule slack 

0.3 

Some problems 
anticipated but easily 
corrected 

Cost estimates exceed 
budget by 1 to 5 percent 

Minor slip in schedule 
(less than one month); 
some adjustment in item 
milestones required 

0.5 

Some reduction in 
technical performance 

Cost estimates increased 
by 5 to 20 percent 

Moderate item 
development schedule 
slip (1 to 3 months); 
impact on item milestones 
with potential for impact on 
segment milestones 

0.7 

Significant degradation in 
technical performance 

Cost estimates increased 
by 20 to 50 percent 

Item development 
schedule slip in excess of 
3 months 

0.9 (high) 

Technical goals cannot be 
achieved 

Cost estimates increase in 
excess of 50 percent 

Large schedule slip that 
impacts segment 
milestones and/or has 
possible impact on system 
milestones 


Risk abatement activities for moderate and high risk items will then be established based on 
the above evaluation. These activities may include the following: 

• Initiation of parallel development activities 

• Initiation of extensive development testing 

• Development of simulations to develop performance predictions 

• Use of consultants and specialists to review design 

• Intensified management review of the development process. 

A risk management program will be developed which identifies risk abatement activities to be 
undertaken, balancing the risk level against the resulting cost and schedule impact on the program. 
A final review of the selected items and alternatives will be made against current state-of-the-art 
knowledge and recent experience on other programs to ensure that the development risk for any 
item has not been under-evaluated. 

Inherent in the monitoring and review process is the evaluation of predicted performance 
against specified requirements. Appropriate performance parameters for risk monitoring purposes 
are established at the top level, together with their contributors (or allocations) at the lower levels. 

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R I SK ANALY5 I S 

IDENTIFY POTENT I A l RISK ITEMSl 

i 

IDENTIFY R EQU I R EMENTS/ FACTORS | 
ASSOCIATED WITH RISK ITEMS 


i 


i 


DETERMINE POTENTIAL 
OF FA I LURE ( P F ) 






i 


i 


DETERMINE CONSEQUENCES 
OF FA I IURE IC F ) 


C f - 


WS 


T I 

COMPUTE RISK |R f ) 

tf = f f + C F ‘ V C F 



1 


1 ) RISK REPORT 
2} RISK ABATEMENT 
PLAN 

3 ) SPECIAL REVIEW 
TEAM 


MEDIUM RISK 

1) RISK REPORT 

2) RISK ABATEMENT 
PLAN 

3) FOLLOW AS ACTION 
ITEM 


LOW RISK 

1 ) REGULAR REVIEW TO 
ASSURE CONTINUED 
LOW STATUS 

2) MONITOR ACTIVITY 
IN PROGRAM STATUS 


Figure 2-8, Risk Assessment Process 


2-11 

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Risk items are continually monitored by Systems Engineering and actions recommended, 
such as initiation or termination of activities when the capability of an item (or its alternate) to meet 
performance requirement is established. 

A Risk Management Plan will be prepared for moderate and high risk items. The plan shall 
include the following as a minimum: 

• A statement of the risk 

• Assessment of the risk and assessment rationale 

• Consequence of failure 

• Alternatives considered and risk associated with each alternative 

• The recommended risk abatement actions 

• Implementation impact statement (cost, schedule, technical) 

• Implementation start date and key milestone schedule 

• Criteria for tracking and closure. 


2.5 SPECIFICATION REQUIREMENTS 

A Contract End Item Specification was prepared and delivered to NASA as a Contract Data 
Report (DR- 10). This specification was prepared in accordance with the requirements of 
MM 8040. 12A, Standard Contractor Configurations Management Requirements. 

To ensure compliance with higher level requirements and compatibility with LAWS 
interfacing requirements, this specification was prepared using the Lockheed developed Automated 
Requirements Traceability System (ARTS). ARTS creates a requirement hierarchy as shown in 
Figure 2-9. From the LAWS CEI level, requirements are allocated to lower level subsystems. The 
requirements matrix resulting from systems design requirements documents (SDRDs) ensures 
traceability and compliance through all program levels. ARTS is maintained by current data 
revision. 

The CEI specification and lower level SDRDs are maintained by configuration management 
(CM) and controlled by the LAWS configuration control board (CCB). This CEI specification and 
SDRDs are maintained by data revision to text in a CM data base and issued as either page revision 
or as a complete reissue, whichever is most cost effective, to reflect approved program changes. 
All changes to this specification are processed in accordance with the requirements of 
MM 8040. 12A. 


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LEVEL 1 

LEVEL 2 
LEVEL 3 


LEVEL 4 
LEVEL 5 

LEVEL 6 


ICDs - 
WBS - 


Statement of Work 
(SOW) 


LAWS 

CEI 

Specification 


Physical 
Reliability 
Maintainability 
Availability 
Safety 
Environment 
Transportability 
Storage 
Electrical 
Mechanical 
Materials 
Right Missions 


LAWS Project 
Elements 


Performance 

Operational Concepts 

Contamination Control 

Coordinate Systems 

Identification & Marking 

Facilities & Facility Equipment 

Human Performance / Human Engineering 

Workmanship 

Maintenance 

Supply 

Verification 

Quality Assurance 


Laser 

Subsystem 


Optical 

Subsystem 




Receiver/ 

Processor 

■ Transmitter 
- Master 
Oscillator 

• Isolator 

• Other 

| 

• Telescope 

• Beam Scanner 

Assembly 

• Interferometer 
- Lag Angie 

Compensation 

• Beam Isolation 

* Detector 

Assembly 
’ Cooler 

Assembly 

• Signal 

Processor 
■ Other 


L Other 


Command 
& Data 
Handling 


Electrical 

Power 

Distribution 


Mechanical 

Support 

Subsystem 

• Right 

• Electrical 

Instrument 

Computer 

Cable 

Platform 

> Right 

Harness 

Thermal 

Software 

* Power 

Control 

- Attitude 

Conditioning 

System 

Determination 

» Circuit 

Vehicle 

Assembly 

Protection 

Interface 

- Other 

• Other 

- Other 


• Product Specifications * Design Verification Requirements 


• Acceptance Test Requirements 


Figure 2-9. ARTS Requirement Hierarchy 


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2.6 INTERFACE DEFINITION 

Three types of LAWS Instrument interface control documents (ICDs) have been identified. 
These documents will be prepared in accordance with the requirements of MM 8040. 12A. Copies 
of each of these ICDs will be delivered to the MSFC-NASA Project Manager. 

One ICD is required to define and control the design of the interfaces between the complete 
LAWS Instrument assembly and the NASA supplied EDS Platform. This ICD will address all 
physical, functional, and procedural interfaces. All software, data, and commands will also be 
addressed to ensure compatible exchanges. This ICD will be mutually approved and controlled by 
the MSFC-NASA LAWS Project Manager and the Lockheed LAWS Program Manager. 

Two ICDs of a slightly different type are required to define and control the design of the 
interfaces between major subcontractor supplied components of the LAWS Instrument. One of 
these ICDs will address the physical, functional, procedural, and software interfaces between the 
LAWS Instrument and the LAWS laser subsystem. The second ICD will address the physical, 
functional, procedural, and software interfaces between the LAWS Instrument and the LAWS 
optical subsystem. 

Both of these ICDs will address the physical, functional, procedural, and software interfaces 
between the laser subsystem and the optical subsystem. The Lockheed Program Manager will 
resolve all design incompatibilities if any are found during the Instrument assembly, integration, 
and test operations. Both of these ICDs will be prepared and controlled by the Lockheed Program 

Manager and approved by each of the affected subcontract managers. 

The third type of ICD will address the LAWS Instrument Software, Data, and Command 
interfaces. All software interfaces, both internal and external, will be included in this ICD. The 
Lockheed Program Manager or his authorized representative will initiate, coordinate, and/or 
approve all changes to this ICD with the Lockheed subcontractors and with the MSFC-NASA 
LAWS Project Manager. 

2.7 RELIABILITY 

The LAWS Instrument has been given a Class B mission designation by the MSFC LAWS 
Program Office. This designation is based on a 5-year mission life and the fact that the payload 
will be installed on a free flyer spacecraft which will not be retrievable by use of the Space 
Transportation System. Lockheed has extensive experience with this type of payload and has 
determined that a combination of Class S and Class B parts may be acceptable depending on the 
assurance that system reliability goals are met. Significant cost and schedule savings may be 
achieved by using Class B parts. 


2-14 

LOCKHEED- HUNTSVILLE 


LMSC-HSV TR F320789-II 


2.7.1 Parts Cost Consideration 

Class S parts procurement costs are typically 2 to 8 times higher than Class B parts. 
Typically, Class B parts have a failure rate of 2 times that of Class S parts. Initial cost typically 
increases 1.5 times if Class S pans are used, and the reliability increases 1.25 times. e 
manufacturing cost includes materials procurement, fabrication, assembly, quality assurance, and 
test. Typically, only 15 percent of the manufacturing cost is for electronic/electncal parts for a high 
density electronic box. This explains the apparent discrepancy in the increased reliability of only 
25 percent if all parts used are Class S. 

2.7.2 Manufacturing/Test Cost 

Manufacturing costs would increase due to the higher number of failures of Class B parts. 
As stated above. Class B failure rates are approximately twice Class S rates. Therefore, early 
failures in manufacturing could be twice the Class S rates. Associated costs include addruonal 
failure analysis of failed parts, corrective action, rework, retest, and possible schedule slippage. 

2.7.3 Summary 

With the Class B mission designation, a mix of Class S and Class B parts will be used. 
Reliability analyses will be conducted to determine which components can use lower grade parts 
and still meet LAWS program reliability goals. 


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Section 3 

PRELIMINARY DESIGN 

The LAWS Instrument preliminary design has been subdivided into the following six 
primary subsystems: optical, laser, receiver/processor, command and data management, structures 
and mechanical (including the thermal control system), and electrical power. Figure 3-1 identifies 
the element of these subsystems, which are described in the following pages. 

Section 3. 1 provides information on the overall system configuration and accommodations, 
including the overall layouts, envelope drawings, mating with the bus and earner vehicle, and 
structural design. Section 3.2 reviews the trades and analyses which were used in defining the 
system concept/configurations. Section 3.3 presents the preliminary design of the six primary 
subsystems, as well as of the thermal control system and the attitude determination system. 
Section 3.4 describes our test and evaluation plan, and Section 3.5 defines LAWS operation 
requirements and scenarios. 


OPTICAL SUBSYSTEM 

LASER SUBSYSTEM 

Telescope Assembly 
Momentum Compensator 
Azimuth Scanning System 
Interferometer Assembly 
Lag Angle Compensator 

Transmitter Laser 
Local Oscillator 
Seed Laser 

Laser Subsystem Interface 

RECEIVER-PROCESSOR SUBSYSTEM 

COMMAND & DATA MANAGEMENT 
SUBSYSTEM 

Photo Detector Array 
Active Cooling Assembly 
Analog-Digital Converter 
Down Converter 
Preamplifier/Bias Electronics 
Interfaces 

Right Computer 
Software Module 

Attitude and Position Determination 
Transceiver Interface Modules 
Subsystem Interfaces 

STRUCTURE & MECHANICAL SUBSYSTEM 

ELECTRICAL POWER 
SUBSYSTEM 

Base Structure 
Attach Mechanisms 
Satellite Bus Accommodations 
Component Support Structures 
Thermal Control System 

• Active 

• Passive 

Power Distribution Unit 
Platform Electrical Power Interface 
LAWS Electrical Power Interfaces 
EMI Control 

F31 2599-DWb-06 


Figure 3-1. LAWS Subsystem Assemblies 


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3.1 OVERALL CONFIGURATION AND ACCOMMODATIONS 

3.1.1 Baseline LAWS 

The LAWS Instrument baseline design is illustrated in Figure 3-2. The velocity vector is 
depicted with the telescope bearing on the leading side of the Instrument platform, the laser on the 
trailing side, and the telescope rotating about nadir. Dual Star Trackers are shown on the cold side 
of the Instrument in close proximity to the inertial measurement unit (IMU). This configuration 
meets all packaging requirements for the Atlas IIAS launch vehicle and can be accommodated by 
the Titan vehicle and, with minor changes, the Delta vehicle. It is designed with clear access for 
assembly, installation, checkout, and removal of all components. Components are located either 
around the perimeter of the Instrument base or on the optical platform. The laser tank and 
telescope bearing are mounted to the Instrument base with critical optical components mounted to 
the optics bench, which can be isolated from the base. The base is, in turn, kinematically mounted 
to the spacecraft. 

Figure 3-3 depicts the Instrument with the environmental covers removed. Smaller optical 
elements, including the redundant local oscillator lasers and the redundant receiver coolers, are 
shown in the figure. The optical bench provides a thermally and structurally stable platform for 
mounting and alignment of critical optical elements. The telescope motor-bearing assembly and 
laser pressure vessel are mounted directly to the base structure through cut-outs in the optical 
bench. 

LAWS, in an orbiting configuration, is illustrated in Figure 3-4. The solar panels are 
deployed in the orbital plane. The radiators are deployed facing deep space. The spacecraft closely 
resembles the generic LAWS spacecraft designed by MSFC personnel. 

Figure 3-5 depicts three views of LAWS. Components are located for optimal passive 
thermal control. Two of the views show the 1.67 m aperture telescope. The telescope secondary 
mirror is tripod-mounted with spacing for the f: 1.5 primary mirror. 

The dimensions of the LAWS Instrument are shown with three views in Figure 3-6. 
Instrument volume is optimized with a 2.5 x 2.9 x 3.6 m package size. The 1.67 m aperture 
telescope provides approximately 1 dB additional SNR over a 1.5 m aperture version. LAWS is 
packaged as a single integrated Instrument and can be assembled and checked out either with or 
without the spacecraft. 

LAWS is shown within an Atlas IIAS fainng in Figure 3-7. A 0.16 m clearance is provided 
between the telescope spider and the fairing for clearance during launch shock and vibration. A 
2.3 m available (longitudinal) space is allowed for the spacecraft envelope. The IIAS payload 
adapter interface is also shown. 


3-2 

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X 

IVCIOCITTI 


Figure 3-2. LAWS Baseline Design Flight Configuration 



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LMSC-HSV TR F320789-II 


The Instrument base structure shown in Figure 3-8 is constructed of graphite epoxy with 
metallic fittings where necessary. Several base thickness values were analyzed and modeled, with 
the 0.35 m thick base selected as optimum from weight and stiffness standpoints. Kinematic 
mounts connect the base to the spacecraft on one side and support the optical bench on the opposite 
side. 

The optical bench is outlined in Figure 3-9. The bench is notched for location of the laser 
pressure vessel and telescope scan bearing. Bench thickness is 0.2 m. 

Figures 3-10 and 3-11 depict the optical path layout including redundant local oscillators, 
seed lasers, and detectors. The transmitter laser pressure vessel is mounted to the base structure 
and isolated from the optical bench. The telescope bearing assembly is also mounted to the base 
rather than the optical bench. Only low mass components are mounted to the optical bench. In 
Figure 3-10 , the local oscillator and seed laser outputs are mixed and the seed laser is controlled 
with a specified off set. The seed laser is injected into the transmitter laser and used to control the 
cavity length prior to transmitter oscillator firing. Output of the transmitter is directed across the 
optical bench toward the telescope bearing. The 4 cm beam is directed along the scan bearing axis 
{Figure 3-11 ) and deflected by a pair of mirrors to enter the telescope at an off-set. Prior to 
entering the rear of the primary, the beam traverses a field corrector lens assembly. The 
transmitted beam travels to the secondary, fills the primary, and is directed toward Earth. 

The returned beam is collected by the telescope approximately 5 ms after transmission. By 
this time, the telescope has traveled ~ 0.2 deg and the beam is received near on-axis, dependent 
upon orbit altitude (a variable) and scan rate. The primary condenses the beam onto the secondary, 
which in turn directs the beam axially through the primary toward a pair of mirrors; these mirrors 
direct it down the scan bearing, this time parallel to the bearing axis and off-axis. The periscope 
follows at the lower end of the scan bearing, is driven by an encoder/phase lock-loop, and brings 
the beam back on-axis where it is directed onto the optical bench again via fixed mirror {Figure 
3-10). A three element (refracture) pupil relay is inserted in the receive optical assembly as Eli, 
EI2, and EI3, with the pupil coincident with the dynamic lag-angle compensator tip-tilt mirror. The 
receive beam is directed off a beamsplitter toward the detectors. The local oscillator beam is also 
fed through the beamsplitter to combine with the received radiation at the detectors. Cryocoolers 
driven by redundant compressors chill the detectors to the 80 K operating temperature. 

Figure 3-12 depicts the environmental cover which assists in the control of the optical bench 
environment. With partitions and vents, this cover helps to stabilize component temperatures and 
protects from contamination. 


3-6 

LOCKHEED-HUNTSVILLE 


LMSC-HSV TR F320789-II 




DIMENSIONS IN METERS 
(INCHES) 



Figure 3-8. Structure, LAWS Medium Base 



Figure 3-9. Optical Bench Configuration 


3-7 

LOCKHEED-HUNTSVILLE 


0 1 0 :0t0 




LMSC-HSV TR F320789-II 



Figure 3-12. LAWS Environmental Cover (Optical Bench) 

LAWS signal flow through the laser, optics, and receiver/processor subsystems is shown in 
Figure 3-13. Tip-tilt mirrors are depicted for low bandwidth adjustment of the local oscillator 
beam- higher bandwidth adjustment is required for the dynamic lag angle compensanon. Telescope 
internal alignment is maintained by an out-of-band alignment assembly. Focus/de-focus capability 
at the receiver provides increased field-of-view for initial acquisition. Optical paths are dashed, 
while electrical paths are shown as solid lines. The components shown with a "2" have been 
tentatively selected for redundancy. 

A condensed baseline mass properties table is depicted in Figure 3-14. The weight values are 
based on design analyses or vendor data for selected hardware elements. The weight budget of 
800 kg is met, but little contingency is presently available. A major emphasis will be placed on 
weight reduction in the following months. The CG is located close to the longitudinal (X) 
centerline. The telescope rotating mass has been minimized to 161.5 kg. The telescope mass CG 
is located on the axis of rotation for minimum inertia effects. The momentum compensator is 
included to compensate for telescope rotational momentum. 


3-9 

LOCKHEED-HUNTSVILLE 


LMSC-HSV TR F320789-II 



3-10 

LOCKHEED-HUNTSVILLE 


Figure 3-13. LAWS Signal Flow 




















System CG Location (m) 

System Contents Weight (kg) X Y Z 

Structure Base, Bench, Environmental Cover, Mounts 138.0 0.03 0.57 -0.34 


LMSC-HSV TR F320789-II 


CM 05 

co co 
d o 

i i 


CM 

o> 


CM 

O 


00 

CM 


CM 


CM 

CO 


O 

to 


CO 

o 


oo 

00 

o 


CM 

GO 


o 

o 


o 

CM 



o 

o 





o 

o 


to 

o 


05 

CM 


CM 

CO 


O 

d 


o 


s 

o 


to 

05 

CM 

05 


05 

CM 

CO 

O 

O 

05 

o 


o 

CM 

CM 

CM 


CM 

CM 

CM 

to 




o 

05 


© 

-Q 

CO 


o 


1 


TZ 




5 


© 




© 

© 


i 

H 


© 

55 

z> 

2 


© 



© 

© 


3 



E 

8 

E 

3 

C 

© 

E 

o 


o» 

c 


o 

a. 



3-11 

LOCKHEED-HUNTSVILLE 


Figure 3-14. LAWS Baseline Current Mass Properties 


LMSC-HSV TR F320789-I1 


Table 3-1 shows our LAWS baseline configuration can be accommodated by Atlas HAS, 
Delta, and Titan vehicles with minor changes. Titan load factors were used during preliminary 

analyses for conservatism. 

The LAWS Instrument with telescope can be fitted into a Delta (large) fairing (shown in 
Figure 3-15) by reducing the telescope aperture from 1.67 m to 1.60 m diameter. This size 
reduction results in a signal-to-noise loss of approximately 0.5 dB. 

Conclusions for the LAWS configuration are listed below: 

. LAWS configuration fits in the Atlas HAS payload fairing with adequate room for 
spacecraft accommodation 

• Configuration is easily adaptable to Delta or Titan vehicles 

• LAWS is within weight and volume allocations 

. LAWS configuration interfaces with preliminary MSFC Orbiting Platform design and other 
similar spacecraft configurations 

• LAWS configuration provides a one piece integrated unit for instrument 
validation/calibration 

. LAWS packaging provides easy access to all components for maintenance and calibration 
after platform/launch-vehicle integration 

• All GnS interface requirements are met 

• Weight reductions are possible with dedicated LAWS spacecraft. 

Table 3-1 . Potential Launch Vehicles 


LAUNCH 

VEHICLE 

FAIRING : ' 
DIAMETER 

DESIGN 

LOAD 

FACTORS | 

<9) 

LAWS 

CONFIGURATION 

Atlas HAS 

4.19 large 

6.0 axial 

2.0 lateral 

Baseline 

Delta 

3.0 large 

6.3 axial 
3.0 lateral 

Reduces telescope 
diameter & base 
mount height 

Titan 

5.08 

6.5 axial 

3.5 lateral 

Baseline 

F312594-49 


3-12 

LOCKHEED-HUNTSVILLE 



SOLAR ARRAY 


LMSC-HSV TR F320789-II 




a 


3-13 

lockheed-huntsville 


Figure 3-15. LAWS in Delta Large Fairing 


LMSC-HSV TR F320789-II 


3.1.2 Downsized LAWS 

NASA Program personnel have indicated that with the overall Earth Observation System 
budget reductions, a downsized LAWS may be more appropriate for the initial LAWS system 
rather than the more optimized baseline LAWS. The downsizing presented to the contractors by 
NASA has been from a 20 J/pulse laser to a 5 J/pulse laser and from a 1.67 m aperture telescope 
to 0.75 m aperture. These reductions degrade SNR by approximately 13 dB. 

Figures 3-16 and 3-17 depict the downsized LAWS Instrument. In developing the 
downsized configuration, Lockheed has left much of the baseline configuration intact and reduced 
dimensions and weights of the transmitter laser and telescope. The thermal control system weight 
along with instrument power requirements have also been reduced accordingly, since with less 
energy per pulse and similar pulse repetition rates, energy consumption and dissipation rates are 
reduced. Figure 3-18 shows the mass budget of the downsized LAWS. 

A cross section of the reduced size LAWS Instrument is shown in Figure 3-16 within the 
Delta fairing. This configuration allows 3.2 m for the bus (platform) compared with 2.3 m in the 

Atlas/baseline configuration of Figure 3-7 and 1.7 m in the Delta/near baseline configuration of 
Figure 3-15. 

For the downsized laser shown in Figure 3-17, we have reduced the tank dimensions from 
Figures 3-6 and 3-10, but left the resonator intact along with seed laser and local oscillator. 

3.2 TRADES AND ANALYSES 

The most fundamental system level trades are the selection of laser pulse energy and the 
selection of telescope diameter. Selection of laser pulse energy is a trade between many pulses of 
low energy and few pulses of high energy, within constraints of laser weight and maximum pulse 
energy which can be developed with reasonable technical risk. Selection of telescope diameter is a 
trade of allocation of available mass into the laser or the telescope within the physical constraints of 
the launch system and the maximum diameter which can be manufactured. 

Initially, in the program, a trade to determine optimal pulse repetition frequency (PRF) was 
conducted. The objective is the minimization of 

CO 2 = (<*v 2 + ar^/N 

where 

co = characteristic velocity in a 100 km by 100 km grid square 
Cy = standard deviation of measurement error for a single shot 
c r = standard deviation of wind velocity 
N = number of shots in a 100 km by 100 km grid square. 


3-14 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 


- LAWS / OP 

interface 


% 

FAIRING 



DIMENSIONS IN METERS 


Figure 3-16. LAWS T elescope with 075 m Diameter Mirror in Delta Fairing 


ENVIRONMENTAL 



Figure 3-1 7. LAWS Instrument Fit-Check in Delta Fairing 


3-15 

LOCKHEED-HUNTSVILLE 




LMSC-HSV TR F320789-II 


System 

Contents 

System 
Weight (kg) 

X 

C. G. Location (m) 

1 z 

Structure 

Base, Bench, Environmental Cover, Mounts 

127.5 

•0.57 

0.03 

•0.34 

Power 

Distributor, Cable 

13.6 

•0.88 

0.77 

-0.32 

Thermal 

Pump Package 15.5 Heaters, Cable, 
EOS Cold Plates 12. Lines, Misc. 

71.35 

•0.82 

-0.2. 

-0.38 

Telescope 

Mirrors, Reaction & Metering Structures, 
TCS, Motor/Beanng, Misc. 

125.9 

0.0 

0.0 

0.59 

Laser 

Laser & Power Supp., Oscillators & Power 
S„pp., Seed Lasers & Power Supp., Misc. 

134.1 

-1.20 

-0.05 

0.21 

Data 

Computer, Cables 

20.4 

-0.79 

0.29 

•0.28 

Receiver / Detec. 

Electronics, Cryo Cooler, Controller, 
Compressors, Displacers, Bias, Preamp, Misc. 

52.0 

-0.26 

0.62 

-0.24 

Momentum Comp. 

Momentum Compensator, Heat Exchanger 

12.9 

0.0 

0.0 

•0.62 

Pointing 

IMU, Star Trackers 

41.0 

0.17 

0.97 

-0.50 

Total 


598.7*kg 

-0.54 m 

0.12 m 

-0.03 m 


* Could be replaced by platform pump if LAWS goes on dedicated platform. 

"Could be replaced with 5 kg heat exchanger if LAWS goes on dedicated platform. 
A Telescope downsize saves 178.75 kg without telescope contamination cover. 


Figure 3-18. LAWS Downsized Mass Properties (6 April 1992) 

In a power- limited system, both ov and N are functions of the laser pulse energy. A low 
PRF gives relatively good velocity measurement for each pulse, but does not allow averaging over 
a large number of pulses. Conversely, a high PRF gives relatively poor velocity measurement for 
each pulse, but allows more averaging over a large number of pulses. The results of this trade are 
shown in Figure 3-19. The abscissa shows the pulse repetition rate. The ordinate shows the 
statistical expectation of standard deviation of velocity measurement (using the Cramer-Rao 
velocity estimator) for n pulses in a 100 km by 100 km grid square. The left side of the figure is 
limited by laser pulse energy (with laser power less than the maximum available), and the right side 
of the figure is limited by power available to the laser (with pulse energy less than the maximum 
acceptable). The figure shows that velocity measurement error is minimized when both maximum 
laser pulse energy and maximum laser power are used. In Figure 3-19 , there is no variance in 
wind velocity. The analysis was extended to the situation in which there is natural variance of 
wind velocity in the grid square and it is desired to determine a single value of velocity which is 
representative of the wind velocity within the grid square. Figure 3-20 shows these results. The 
figure shows that for good backscatter (low altitude), overall velocity error is decreased by 
increasing PRF, allowing more averaging of the natural atmospheric variance. For poor 
backscatter (high altitude), a lower PRF is preferable. As compared with a high PRF, the 

3-16 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 


improvement in measurement accuracy for each pulse more than offsets the advantage of averaging 
over many pulses for natural atmospheric variance. Therefore, the optimal PRF is approximately 
5 Hz. During the study, this trade led to the selection of 3 pulse pairs per grid as the appropriate 
shot density for the survey mode. 


WAVELENGTH - 9. 1 1 MICRON OPTICS Dl A - 1 .67 METER 
PULSE LENGTH >32 MICR05EC NADIR 3 45 DEG 



Figure 3-19. Selection of Pulse Repetition Frequency to Minimize Error in Wind Velocity 
Averaged Over a Grid Square 


WAVELENGTH - 9,1 I MICRON OPTICS DIA 3 1 67 METER 



Figure 3-20. Effect of Pulse Repetition Frequency on Error in Averaged Wind Velocity with 
Variation in Wind Speed Over a Grid Square 


3-17 

LOCKHEED-HUNTSVILLE 


LMSC-HSV TR F320789-II 


The conclusion that a high pulse energy, low PRF system is preferable to a low pulse 
energy, high PRF system results from the fact that much of LAWS operation is in marginal 
backscatter conditions. If backscatter were significantly larger, a low pulse energy, high PRF 
system would be preferable. 

Figure 3-21 shows the trade between laser pulse energy and telescope diameter. The 2. 1 m 
limit in telescope diameter is that which can be manufactured with available facilities. The 20 J 
limit in laser pulse energy is a judgment of the maximum which can be developed with acceptable 
technical risk. Given the requirement of 3 shot pairs per 100 km by 100 km grid square for the 
survey mode, the laser pulse energy is also limited by the 2200 W average power for the survey 
mode. However, this limit is less constraining than is the 20 J maximum pulse energy. Lines of 
constant instrument mass and lines of constant narrow band SNR are shown. The lines of 
constant mass indicate that instrument mass is a function of both pulse energy and telescope 
diameter, and these two parameters must be traded to achieve constant mass. 


PARAMETERS AFFECTING BOUNDARIES PARAMETERS AFFECTING PERFORMANCE 

Laser Pulse Eff. = 6% Wavelength = 9.11 x 10" 6 m 

3 Shot Pairs per 100 km Grid Pulse Length = 3.2 x 10-« s 

Power (Excl. Laser Pulse = 609 Watt) Backscatter = 1 xlO-^nri 

Weight (Excl. Laser & Telescope = 373 kg) Design Point S/N a 0.56 dB (N Band) 



Telescope Diamter, (m) 


F 320 707 -01 
F3207M 


Figure 3-21. Trade Between Laser Pulse Energy and Telescope Diameter to M aximize SNR 
within Weight Constraints 


3-18 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 


The figure shows that for an instrument mass of 800 kg, SNR is maximized by using a 20 J 
laser and maximizing the telescope diameter within the 800 kg mass limit. Therefore, the 20 J 
laser and 1.67 m diameter have been selected as the baseline design point. The chart permits 
evaluation of sensitivity of instrument mass and SNR to other candidate design points. 

3.3 SUBSYSTEM DESIGNS 

3.3.1 Laser Subsystem 

This section addresses the design of the laser subsystem, which consists of the following 
assemblies: 

• Optical resonator 

• Electrical discharge 

• Pulse power supply 

• Pressure vessel structure 

• Gas flow loop 

• Controls and instrumentation 

• Injection laser 

• Local oscillator. 

The physical layout of the transmitter laser subsystem is shown in Figure 3-22. Its general 
configuration is fundamentally that proposed in Phase I. Modifications of note are removal of the 
resonator optics from the pressure vessel, the addition of a contraction to the flow loop, and 
relocation of the catalyst beds upstream of the heat exchangers. The functional interactions 
between the transmitter laser assemblies are outlined in Figure 3-23 and discussed in the following 
paragraphs. 

3. 3. 1.1 Optical Resonator 

The resonator configuration, shown in Figure 3-24, closely resembles that of the breadboard 
design. Although some design parameters were modified to accommodate the interface of the 
transmitter with LAWS optical bench, care was taken to ensure that performance parameters such 
as mode discrimination and sensitivity to misalignment were not adversely affected. The key 
resonator parameters are listed below: 


Type 

unstable 

Equiv. ffesnel no. 

1.56 

Magnification 

2.25 

Cavity length 

3.0 m 

Gain length 

1.5 m 

Beam size 

4x4 cm. 


The resonator is of the unstable type with a conventional concave primary mirror and a 
lens/grating combination acting as the feedback mirror. A folded cavity configuration was chosen 
for compactness, with both folding mirrors partially reflecting. 


3-19 

LOCKHEED-HUNTSVILLE 


LMSC-HSV TR F320789-II 




Figure 3-22. Laser Transmitter 


PLATFORM 
POWER 
120 VOC 


PLATFORM 
POWER 
l 20 VOC 


COOLING 

INTERFACE 


GAS 

HANOLING 


INJECTION 
LAGER 
P S 


injection 

LASER 
PS * 2 


H V POWER 
SUPPLY 


INJECTION 
LASER M 


PULSE POWER 


PfN 


THYR. 


GAS 

RESERVOIR 


LASER HEAOf 


INJECTION 
LASER "7 


RESONATOR 


DISCHARGE 


FLOW LOOP 


CATALYST 


LOCAL 

OSCILLATOR 


LOCAL 

OSCILLATOR 


T 


COOLING 

SYSTEM 


OUTPUT 

BEAM 


BEAM 

OETECTOft 


TO 

TELESCOPE 



1 



COMPUTER 

INTERFACE 

CAVITY MATCHING 


AL IGNMENT 


ELECTRONICS 


LASER 


UNIT 


FLIGHT PROCESSOR 


♦ 


D*f*ns« Systems 


autoalignment 

SOFTWARE 

MODULE 


FAILURE 
HANOLING 
SOFTWARE MOOULE 


DATA HANDLING 
MODULES 


SIGNAL ANO 

COftlANO 

MOOULES 


Figure 3-23. Laser Subsystem Block Diagram 


3-20 

LOCKHEED-HUNTSVILLE 











LMSC-HSV TR F320789-II 



The seed laser light is injected through one of the folds with the single longitudinal mode 
(SLM) detector monitoring the light transmitted through the same fold. The intensity of light 
transmitted through the opposite fold is measured by a cavity matching detector. Information from 
the "finesse" curve thus obtained is used by the cavity matching electronics to adjust the piezo- 
electric transducer (PZT) drive on which the feedback assembly is mounted. Use of a dithering 
system instead of the ramp function used in the resonator design verification test (DVT) and also in 
the breadboard will be considered. 

Laser output energy is extracted by a scraper mirror located near the feedback assembly and 
measured by a pyrodetector located between the scraper and the telescope. 

The primary and scraper mirrors will be made either of copper or dielectrically coated silicon 
substrates, while the folding mirrors and pressure vessel windows will be made of ZnSe to allow 
alignment in the visible regime. 


3-21 

LOCKHEED-HUNTSVILLE 





LMSC-HSV TR F320789-II 


3. 3. 1.2 Flow/Discharge Subsystem 

These assemblies have been defined as one subsystem because of their high level of 
mechanical integration and functional interdependency. As Figures 3-25 and 3-26 indicate, the 
layout closely matches that of the breadboard. Differences arise primarily in the choice of materials 
and addition of redundant components wherever failure mode analysis and breadboard lifetime 
tests indicate a need. 


A UV preionized self-sustained discharge scheme was chosen with four electrode pair 
modules (two per side) providing redundancy and eliminating alignment and current distribution 
problems associated with long electrodes. A modified Ernst profile was chosen for the cathode 
based on extensive electrostatics code calculations substantiated by the DVT results. A flat anode 
profile was chosen for flow compatibility and compactness. Preionization is achieved through 
holes in the anode utilizing a dielectric/corona bar assembly. The dielectric material chosen for the 
preionizer housing can be machined and is impermeable. The relevant operating parameters of the 
discharge are listed below: 


• Gas mixture 

• Gas pressure 

• Discharge dimensions 

• Pulse length 

• Specific energy loading 

• Discharge voltage 


3:1:1 He:C0 2 :N 2 
0.625 atm 

4.2 x 4 x 150 cm 

3.2 - 4.0 \ls 
86J/L 
21-23 kV. 


The flow loop is designed to accommodate the discharge assembly described in the previous 
section. It provides fresh gas to the discharge and moves the used hot gas at the appropriate speed 
to prevent arcing. This gas is subsequently reconditioned by the catalyst bed, where recombination 
of CO and O into C0 2 dissociated during the discharge occurs. Subsequently, the thermal energy 
resulting from the inefficiencies inherent in the laser kinetics processes is removed by a fan and 
tube heat exchanger. The sidewall mufflers, located in both sides of the cathode, attenuate the 
acoustic waves generated by the discharge in order to maintain the homogeneity of the lasing 
medium in the cavity below the levels dictated by beam quality and cavity matching requirements. 


3-22 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 



Figure 3 - 25 . LAWS DischargelFlow Loop, End View 


END FAN SUPPORT 


CENTER COUPUNG & FAN SUPPORT 

IMPELLER OUTER SHROUD 

CROSS FLOW BLOWER 


PREIONIZER 


WINDOW 

ASSEMBLY 


CATHODE 


PERFORATED PLATE 

TUBE/FM HEAT EXCHANGER 


' (b) Side View PRESSURE VESSEL 

24.12 OUTSIDE DtA 


VACUUM CONNECTION 

Figure 3-26. Two-Electrode Configuration, Side View 


MOTOR A 

MAGNETIC 

COUPLING 



H.V. FEEDTHROUGH 


FLOW LOOP 
SUPPORTS 


3-23 

LOCKHEED-HUNTSVILLE 





LMSC-HSV TR F320789-II 


Key features of the flow loop design are the dual tangential fans chosen for both compactness 
and modularity, the contractions upstream of each discharge which assist in restoring flow 
uniformity, and the skewed positioning of the catalyst bed and heat exchanger which provides 
compactness and causes a gradual equilibration and cooling of the hot gas. This last feature 
minimizes density perturbations to the laser medium which could otherwise affect the medium 
homogeneity in the cavity. The relevant flow parameters are listed below. 


• Mass flow rate 

• Flow velocity in cavity 

• Fan speed 

• Cavity flush factor 

• Available catalyst volume 

• Porosity of muffler wall 


23 g/s 
1.26 m/s 
1700 rpm 
3.0 at 10 Hz 
12.6 L 

3 percent - no packing resistance 
Backup Design: 30 percent - 1 cgs rays/cm. 


3. 3. 1.3 Pulse Power 

The pulse power system in a discharge pumped CO2 laser is formed by three primary 
components: a high voltage dc-dc converter, a pulse forming network (PFN), and a thyratron. 
The function of the high voltage power supply is to step up the 120 Vdc prime power mput to the 
40 kV charge voltage required by the PFN. The PFN in turn is charged by this power supply and 
upon switching by the thyratron, generates a pulse with the desired length as well as voltage and 
current characteristics. The pulse energy is subsequently discharged into the gas by the discharge 
assembly described in the previous section. A functional diagram of these processes is shown in 

Figure 3-27. 



120 V* 

PlAironi 

PC**R 


Figure 3-27. Energy Discharge Processes 


3-24 

LOCKHEED-HUNTSVILLE 












LMSC-HSV TR F320789-II 


The configuration of the PFN, shown in Figure 3-28, will be an E-type, thyratron switched 
scheme similar to that utilized in the breadboard. Primary differences arise in the choice of lighter 
weight, space qualified components, particularly capacitors, and the use of redundant critical 
components such as the thyratron, capacitors, and diodes. Also, because operating the PFN in a 
pressurized environment would result in a considerable weight penalty, vacuum operation is 
anticipated. This requires mounting components on a coldplate and active cooling of the thyratron. 
The operating parameters of the pulse subsystem are listed below: 


Total energy stored in PFN 

264 J 

PFN charge voltage 

40 kV max 

PFN current 

<2.5 kA 

Total capacitance 

400 nF 

Pulse length 

4.5 |is max 

PRF 

10-15 Hz. 




Figure 3-28. Preliminary Layout of Pulsed Power Section 


3-25 

LOCKHEED-HUNTSVILLE 












LMSC-HSV TR F320789-II 


3. 3. 1.4 Controls and Instrumentation 

The controls and instrumentation units needed for the space device have been identified, as 
have their functions. A computer interface unit will be provided to handle signal and command 
flow between the transmitter components and LMSC's flight computer. In addition, electronics 
units for the auto-alignment feedback loop, cavity matching loop, and fan drive will be 
implemented. Functional diagrams for these loops are shown in Figures 3-29 and 3-30, 

respectively. 



Figure 3-29. Resonator Cavity Matching Control 



Figure 3-30. Auto-Alignment Functional Diagram 


3-26 

LOCKHEED-HUNTSVILLE 













LMSC-HSV TR F320789-II 


3- 3. 1.5 Parts Requiring Complex Manufacturing Techniques 

Some of the complex components of the space laser device are expected to be similar, if not 
identical, to those of the Phase II risk reduction laser breadboard. The following breadboard 
drawings may be used for preliminary specifications: 

• Contour machined components 

- Muffler Assembly L232 1 9 

- Cathode Housing LA W232 15, LA W23285 

- Preionizer Assembly LAW23250, LAW23236 

• Cathode Bar LA W232 14 

• Shroud Brackets LAW2328 1 

• Contraction Duct LAW2323 1 

• Corona Bar LAW23229 

• Support Plates LAW23234. 

3. 3. 1.6 Software Systems 

Software required to operate and monitor the transmitter consists of modules designed to 
handle I/O of signals and commands via the computer interface unit. These modules are 
implemented for the operating system and in the programming language specified by NASA for the 
flight computer. 

These modules monitor control input bits from on/off sensors, control on/off - open/close 
devices, monitor analog signals sensors, and control analog actuators and voltage controlled 
functions. One module provides the logic needed to undertake steps outlined in the failure mode 
analysis for those failure modes which cannot be automatically handled by mechanical or electrical 
switching. Another module implements the algorithm for autoalignment of the transmitter optics. 
Finally, additional modules perform such functions as data partitioning, communications, and 
compression as well as flagging and linking of routines. 

3. 3. 1.7 Laser Subsystem Summary 

Figure 3-31 summarizes development efforts for the laser subsystem. This summary 
includes the overall development schedule (A), the required subsystem equipment and 
implementation/verification plan (B&C), trade studies to establish the baseline (D), and a summary 
of subsystems risk (E). 


3-27 

LOCKHEED-HUNTSVILLE 



LASER TRANSMITTER PLAN OVERVIEW 


TASKS 


MAJOR MILESTONES ZLa ^ 

ATP PRR PDR 


Engineering Unit 

Design & Spec Prep ^ 

Procurement 

Long Lead Times 
Assembly/Construction & 
Check Out of Components 
Functional Tests of Sub- | 
Assembly Comp Chk. Out 
Shock/Vibration Tests 
Subsystem Integration 
Laser Tests @ TDS 
Eng Unit Laser Tests 
© LMSC 


Qualification unit 

Design & Spec Prep 
Procurement 
Construction 
Functional Testing of 
Laser subassemblies 
Subsystem integration 
Laser Qualification Tests 
@ TDS 

Qual. Unit Laser Tests 
O LMSC 


Right Unit 

Procurement Long Lead 
Items 

Fabricate, Assemble, Wire 
Funct Testing 
Subassemblies 
Subsystem Integration & 
Laser Testing @ LMSC 



[b] required subsystem equipment 

COMPONENT* 

SOURCE 

QUANTITY/UNIT 

ENG. UNIT 

QUAL. UNIT 

FLIGHT UNIT** 

Pulsed Power Laser 

TDS 

1 

1 

1 

1 

Discharge Cavities 

TDS 

2 

2 

2 

2 

Flow Loop/Fans/Catalyst 

TDS/VOP 

1/2/2 

1/2/2 

1/2/2 

1/2/2 

Pressure Vessel 

TDS 

1 

1 

1 

1 

Pulse Forming Network 

TDS 

1 

1 

1 

1 

Thyratrons 

TDS 

2 

2 

2 

2 

Pulsed Power Supply 

ALE 

1 

1 

1 

1 

Optical Resonator/Bench 

TDS/LMSC 

1/1 

1/1 

1/1 

1/1 

CW Injection Laser 

MPB 

2 

2 

2 

2 

Single Mode PZT Controller 

BURLEIGH 

1 

1 

1 

1 

CW Local Oscillator Laser 

MPB 

2 

2 

2 

2 

Controls and Instrumentation 

TDS 

1 

1 

1 

1 

Alignment Laser and Mechanism 

ITEK 

1 

1 

1 

1 

Laser Thermal Control System 

TDS 

1 

1 

1 

1 


Fan Failui 


Individual 


Catalyst C 



Thyratron 


PFN Cap, 


Feedback 


Mirror Da 


Window [ 


FOLDOUT FRAME 





























LMSC-HSV TR F320789-II 




2000 

| 2001 


^ Launch 


LAWS/Bus 


C REQUIREMENT IMPLEMENTATION/VERIFICATION j 

KEY REQUIREMENT 

IMPLEMENTATION 

VERIFICATION 

Operational Life and Reliability 

• 5 yr on orbit 

• 10 9 Shots 

Extended Life Tests 

• Components to > 10 9 

• System to > 3 x 10 8 

Design for; 
Robustness and 
Key Component 
Redundancy 

Performance 

• 9.11 nm (C’9 02) 

• 20 J/Pulse 

• Single mode pulses 

• 3 pp FWHM pulse length 

• <200 kHz CHIRP 

• 4.67 Hz scan mode 1 xjc/2 

• 1 0 Hz design mode J (max PRF) 

Performance Validation Test 

• Breadboard 

• Eng Unit 

• Qual Unit 

• Flight Unit 

Interfaces and Software 
Functions 

• Flight Processor 

- Auto alignment SW 

- Failure handling SW 

- Data handling SW 

- Signal & Command SW 

• Telescope control system 

• Platform Power Control System 

• Platform Thermal Control System 

• Beam Detector System 

• Gas Handling System 

Simulation and lest 


E 


PLANNED TRADE STUDIES 


TRADE ITEM 


Discharge Parameters 

• Gas mixture composition 

• Gas pressure 

• Electrodes/Preionizers materials 

• Cavity dimensions/Gain length 

• Voltage/Energy Loading 

• Flush factor 
Resonator Parameters 

• Magnification 

• Scraper geometry 

• Cavity reflectivity 
Flow Loop Parameters 

• Catalyst configuration 


BASELINE DESIGN 


• He : C ia 0 2 : N 2 = 3:1:1 

• 0.625 atm 

• Proprietary 

• 4.2x4 cm/1 50 cm 

• 35 kV/80 J/L 

• 3.0 

• 2.25 

• Square (square vs. circular) 

• Uniform (uniform vs. graded) 

• Dual in-line Beds, 400 cells/in 2 


RISK SUMMARY 


RISK ITEM 

RISK LEVEL 

RISK REDUCTION APPROACH 


Moderate 

1 . Dual fans provide redundancy; laser can operate on one fan. 

discharge Arc or Preionizer Failure 

Moderate 

1 . Redundant preionizer and associated discharge modules. 

2. Operation at lower discharge voltages to reduce probability of arcing and failure 

ntamination 

Probably Low (not yet 
established) 

1 . Catalyst reactivation heaters, flushing of pressure vessel laser gas refill, 
plus pre-launch clean room and bake out procedures. 

ailure 

Moderate 

1. Backup provides redundancy. 

;itor Short Circuit 

Moderate 

1 . Isolate faulty capacitor, switch in backup spare. 

4irror PZT Drive Failure 

Low 

1. Robust design is essential. 

age or Contamination 

Risk not yet established 

1 . Robust design and cleanroom and bake out procedures to minimize effects. 

mage or Contamination 

Risk not yet established 

1. As above; also addition of lasing mixture, cope with a small crack. 


Figure 3-31. Overview/ Summary of the 

Laser Transmitter Subsystem 

3-28 FOLDOU 

LOCKHEED-HUNTSVILLE 


)' FRAivi^ 



LMSC-HSV TR F320789-II 


3.3.2 Optical Subsystem 

3 3 2 1 Optical Subsystem Baseline Design 

' ’ The LAWS optical subsystem has two major functions. First, it acts as a transmitter in the 

lhe LAW a op / nf the 9 11 micron laser and forming a 1.67 m 

role of a beam expander, taking e c .. oss ^ Earth's atmosphere. Second, it 

diameter beam which is scanned via a bearing asse y scattered energy from the 

performs the function of a receiver, acquiring * » w S 1lTv“the transmitter relay 

Performs dynamic lag ang.e 

compMMtiort^ ^ functioi^Jlo^tt^m configwration^^ratii^ 

baseline design for the te . es ~ p * * flts with in the current packaging envelope. The 

With a F/1.5 primary objec, space in older to remove the course lag 

transmit optic axis is onented o y ,th and the round trip time for each transmitted 

ang,e which is due to the telescope scanning : ™tha^d round mp urn from 

for second order dynamic lag angle compensation. 

The periscope follower is a ~ror " 
telescope in order to fold the receive ra anon a ^ location and orientation of the 

additional active sensors and maintaine y actua o preliminary baseline design 

m^o^e^Clm c2 ” due to a low sensitivity design, thealignmen, 

”mce sysiem, and the use of ULE will, its virtually zero CTE and vanauon of CIE. 

in Figures 3-33, 3-34, 3-Ji, respecu cy & lhe only n0 nplanar 

^"meTs' » ^“include die three receive channel relay elements and 

the two transmit channel compensator elements. 

lead time items are the ULb mamcs ior u.c ^ 

approximately 9 months to fabncate. . 

Subassemblies for the various units are listed in ^ 3-S6 <B>. Key re q uiremems, 
implementations, and verification approaches are shown in Figure 3-36 (C). 


3-29 

lockheed-huntsville 



LMSC-HSV TR F320789-II 


FROM 

TRANSMITTER 


Transmitter 
Compensator 
Optics 


Transmit/ 

Receive 

Telescope 


Tip/Tilt 
Lag Angle 
Compensator 


Periscope 

Follower 


Receiver 
Pupil Relay 
Optics 


Alignment 

Sensor 


RECEIVER 


Figure 3-32 . Optical Subsystem Functional Flow Diagram 


Table 3-2. Optical Design Characteristic 



Parameter 


Aperture (cm) 

Magnification 
Primary F-N umber 
FOV (circular) 

Pri-Sec Spacing (cm) 
Obscuration (area) 
Wavelength (pm) 

Optical Quality (RMS WFE) 


Value 


Tran.: 13 prad 
Rec.: 0.6 mrad 


Trans.: 0.018 1 X 
Rec.: 0.003 1 X 


3-30 

LOCKHEED-HUNTSVILLE 





LMSC-HSV TR F320789-II 



F312539-««k-12 


Superb CTE variation minimizes 
wavefront errors from temperature 
soaks 

— 53 °C soak change yields 0.007 
rms 

Lightweight mirror design meets 
weight requirements while 
maintaining its durability 

— 1.5 m diameter LPMA withstands 
18 g's 

Detailed weight estimate (kg) 


Facesheet 

37 

Closure wall 

7 

Ribs 

27 

Fillets & 

4 

Parasitics 


Total 

75 


Figure 3-33. Primary Mirror Design 


2-54mm / 0.1 in 




Mass Properties (kg) 
Outer closure ring 4 

Inner closure ring l 

1.9 m Facesheets (front and back) 13 
Main ribs 1 1 

Minor ribs 3 

Hardware, mounts & clips 3_ 


Total 


35 


1.25mm/ .0« io^l^_ L25mm/ M in 


• All graphite epoxy structure 

Egg crate construction yields high stiffness with a low density 

G/E coefficient of thermal expansion well matched to that of ULE 

. Primary mirror kinematically mounted to structure through three bipods 

F312539-lt«k-13 


Figure 3-34. Reaction Structure Design 


3-31 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 






All graphite epoxy structures with invar hardware inserts 
Weight summary (kg) 


Metering structure 
Reenforced tube 22.0 

Hardware 3 

Metering legs 

Legs 0.5 

Hardware 1.5 

Secondary mirror 
assembly 

Mirror 0.5 

Bezel & hardware 1.5 

TOTAL 29.0 


F312539-»«k-14 


Figure 3-35. Metering Structure Design 


Table 3-3. Optical Component Data 


Component 

Radius 

Conic Constant 

F-Number 

Diameter 

Primary 

502.920 CC 

-1.00000 

1.50 

167.6 

Secondary 

12.040 cx 

-1.00000 

1.50 

4.0 

Compen 1 
(Trans.) 

4.382 CC 
59.172 CC 

t=0 .38 1 

— 

3.6 

Compen 2 
(Trans.) 

12.656 CC 
4.969 cx 

t=0.635 

— 

5.1 

Relay 1 
(Rec.) 

130.176 cx 

oo 

t=1 .270 


10.9 

Relay 2 
(Rec.) 

6.975 cx 
5.630 CC 

t=1 .270 


3.8 

Relay 3 
(Rec.) 

53.273 CC 
34.724 cx 

t=1 .270 

— 

5.1 


* Linear units are centimeters 


3-32 

LOCKHEED-HUNTSVILLE 













OPTICAL SUBSYSTEM OVERVIEW 


TASKS IS 

MAJOR MILESTONES K A 

ATP PRR 


Design & 

Development 

Preliminary design Z 
Final design 

Fabrication 

Engineering Unit 
Qualification Unit 
Flight Unit 

Integration 

Engineering Unit 
Qualification Unit 
Flight Unit 

Test Support 

Engineering Unit 
Qualification Unit 
Flight Unit 

Engineering Support 

Sustaining Engineering 
Bus Integr. Support 
Launch Support 
On-Orbit Calibration 
& Align Support 


4 /97 6/p 7 


REQUIRED SUBSYSTEM EQUIPMENT 


COMPONENT* 

Primary Mirror Assembly 
Secondary Mirror Assembly 
Metering Structure 
Reaction Structure 
T ransmit Relay Optics Set 
Receive Relay Optics Set 
Fold Optics Set 
Thermal Control System 
Azimuth Scanning System 
Tip/Tilt Mirror 

Telescope Alignment System 
Mechanical, Thermal, Electrical, 
and Optical Interfaces 


SOURCE 

Litton-ltek Optical Systems 
Itek 
Itek 
Itek 
Itek 
Itek 
Itek 
Itek 
Itek 
Itek 
Itek 
LMSC 


QUANTITY/UNIT ENGINEERING. UNIT 


* “S” Parts 

"Engineering unit components used for spares 


FOLDOUT FRAME 


LMSC-HSV TR F320789-II 


2000 




2001 


LAWS/Bus A ALaunch 



fc] REQUIREMENT IMPLEMENTATION/VERIFICATION | 

KEY REQUIREMENT 

IMPLEMENTATION 

VERIFICATION 

Operational Life 

Maximize Heterodyne 
Efficiency 

5 years on orbit 

• Wavefront error < 0.07 
waves RMS 

• Rat field over receive FOV 

• Obscuration < 3% 

• Round trip pointing stability 
< 1.5 farad 

• Magnification: 42X 

Comparison and test 

Analysis, simulation, and 
test 

Lag Angle Compensation 

• Format: 2 points separated 
in field by 0.185° 

• Dynamic tip/tilt mirror 

Analysis and simulation 

... L 


1 

JAL. UNIT 

FLIGHT UNIT** 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

1 

5 

5 


0 


PLANNED TRADE STUDIES 


Baseline Design 


Alternate Design 

Fully Passive 


Partial Heater 1 

Control 


Control I 


Tube/ 

Tripod 


Athermalized 

Truss 


Metering Rods/ 
Metering Tube 


ULE 


SiC 


Fused 

Silica 


Graphite 

Epoxy 


ULE 

(Rods) 


1 _ sic I 

(Rods) | 


Metal 

Matrix 


Passive 

Radiator 


Fully Passive 
No Radiator 


Active Diode 
Heat Pipe 



RISK SUMMARY ~~ ”1 

RISK ITEM 

RISK LEVEL 

RISK REDUCTION APPROACH 

Motor/Bearing/Encoder 

Low 

Space qualified and demonstrated unit 

Optical Coating Fatigue 

Low 

Risk reduction testing with LAWS laser 
breadboard 

Telescope Alignment System 

Low 

Risk reduction testing with alignment 
system breadboard 


F320789-07 


Optical Subsystem 


3-33 

LOCKHEED-HUNTSVILLE 


FOLDOUT FRAME 























LMSC-HSV TR F320789-II 


3 . 3 . 2. 2 Optical Subsystem Alignment System 

An active alignment system is required to correct for positional changes due to gravity release 
when the unit is first placed on orbit and for changes which result from thermally induced 
contractions and expansions. The latter are relatively slow changes, so high bandwtdth responses 
will not be required. The alignment system provides an error signal to control the secondary 

mirror position in tilt and despace. 


The basic alignment approach is shown in Figure 3-37. The concept consists of three major 
optical paths. The first path is a sample of the transmit beam direction represented by a coaligned 
laser beam centered in the main transmit beam. This is folded into the integrated alignment sensor 
(IAS) via a penta fold mirror. The sample from the transmit beam is offset in angle from the 
receiver axis. This process is used to separate this return from the receiver beams The transnut 
beam surrogate is spatially separated and focused onto a CCD detector. This provtdes the angular 
coordinates of the beam in primary mirror coordinate space. 



Itek Optical Systems 


Figure 3-37. Alignment System Concept 


3-34 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 


The return beam from the input annular reference mirror (IARM) is reflected from a beam 
divider to provide angular coordinates of the IARM in IAS space. This allows the IAS to move 
since the measurements are made relative to this measurement. This division of the main beam is 
made possible by locating the divider at an image of the IARM -- a pupil. The beam which passes 
through the IARM image provides a measurement of the tilt and defocus of the telescope. 

The second path originates from the IAS. The annular collimated beam is folded onto the 
receiver axis by a combining mirror - a perforated flat mirror. It then autocollimates off an IARM 
which is kinematically mounted to the primary mirror assembly. This returns a beam representing 
the receiver axis coordinate system. 

The third beam is formed by allowing a portion of the IAS beam to pass through the IARM. 

It traverses the optics of the telescope. The beam from the primary (output space) then 
autocollimates from the output annular reference mirror (OARM). This sample provides both tilt 
and defocus information of the beam expander. 

The alignment is carried out by means of surrogate beams which are co-aligned with the 
LAWS optic axes. The alignment sequence is as follows: 

( 1 ) The transmit beam and transmit reference beam are coaligned 

(2) Transmit reference is folded into IAS via a penta fold 

(3) The IAS reference is inserted onto the receiver optical axis via the combining mirror 

(4) The IAS reference beam is divided at the IARM 

(5) Part of the beam is autocollimated off the IARM - this represents the coordinate system 
of the primary mirror 

(6) The transmitted portion of the beam traverses the telescope to the OARM 

(7) The three returns (transmit beam reference, IAS reference, and telescope reference) enter 
the LAS. 

The IAS transmits the alignment data to the flight computer, which then provides signals to 
the actuators that control the orientation of the secondary mirror. 

3 . 3 . 2. 3 Design Trades and Sensitivity Analyses 

Various configurations for the telescope have been evaluated against the requirements. These 
requirements changed during the course of the system development when the nadir angle 
specification was changed from variable to fixed. With a fixed nadir angle, the dominant 
contribution to the lag angle is also fixed. This eliminated the need for some of the flexibility 

initially considered. 

The preliminary trades reduced viable telescope configurations to two candidate options: a 
two-mirror afocal and a three-mirror afocal system. With the adoption of a fixed nadir angle, a 
split field telescope was considered the most applicable approach and the inherent design flexibility 
afforded by the three-mirror was no longer demanded. Wavefront error sensitivities for tilts and 
displacements were calculated and determined to be comparable. The strongest advantage of the 
three-mirror system is the existence of a real pupil. With a real pupil, second order lag angle and 

3-35 

LOCKHEED-HUNTSVILLE 



IMSC-HSV TR F320789-II 


other dynamic corrections can be accommodated with a single tip/tilt mirror. However, with 
additional optics, a real pupil can be created in a two-mirror telescope. This approach allows a 
reasonable compromise of simplicity while still allowing the very small secondary mirror of the 
two-mirror design. 

The two-mirror afocal design is shown schematically in Figure 3-38. The large primary 
operates at F/1.5, and both mirrors are parabolas. The split field allows the static lag angle due to 
telescope rotation during the round trip time of the laser pulse to be accommodated. Sensitivities to 
tilts and displacements have been analyzed for this as well as the three-mirror design. The results 
of these calculations for the two-mirror system are shown in Tables 3-4 through 3-7. 


167.6 cm 


L 


T 


Secondary 


8-34° 

T = 


Transmit 

channel 


tVL 


Compensator 

Receive 

channel 

Pupil relay 
element 1 


f 


245.5 cm 


F/1.5 Primary: Mag = 41.8 Parabola/Parabola 


Figure 3-38. Two-Mirror Afocal Split Field Design 


3-36 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 



Table 3-4. 


Element 


Primary 

(W/Refocus) 


Secondary 

(W/Refocus) 


Relay 1 
(W/Refocus) 


Relay 2 
(W/Refocus) 


Relay 3 
(W/Refocus) 


Receiver Channel Wavefront Error (WFE) Sensitivities (Rigid Body 
Alignment Errors) 


Displacement 


Alpha 


0.030 















Gamma 


0.000 


.000 


.000 


0.000 


.000 


0.000 


0.000 


0.000 


0.000 


0.000 



AH entries are rmp WFE for wavelength = 9.11 nm 
A X,Y,Z = 0.001 inches (Primary and Secondary) 

A X,Y = 0.005 inches (Relay 1,2,3) 

A Z s 0.002 inches (Relay 1,2,3) 

A Tilt = 0.0001 radians (Primary and Secondary) 

- 0.0010 radians (Relay 1,2,3) 


Primary Tilt is most sensitive error 


TIR (inches) 
PRI 
SEC 
REL 1 
REL 2 
REL 3 


0.0066 

0.0002 

0.0045 

0.0015 

0.0020 


Table 3-5. Transmitter Channel WFE Sensitivities (Rigid Body Alignment Errors) 


Element 


Primary 

(W/Refocus) 


Secondary 

(W/Refocus) 


Compen 1 
(W/Refocus) 


Compen 2 
(W/Refocus) 
Comp 1 & 2 
(W/Refocus) 



.013 


.011 


.002 


.002 


.009 


.009 


.004 


0.004 

0.002 

0.002 



.018 


.007 


.003 


.003 


.009 


.009 


.004 


.004 


0.002 

0.002 


Gamma 


0.000 


0.000 


0.000 


.000 


0.000 


0.000 


.000 


0.000 


0.000 

0.000 



0.005 


0.004 


0.004 


.004 


.020 


.020 


0.020 


.020 


0.000 

0.000 


Displacement 

Y 

0.006 


.005 


.005 


.004 


.022 


.020 


.021 


.020 


0.000 

0.000 



All entries are rms WFE for wavelength = 9.11 |im 

A X.Y.Z = 0.001 inches (Primary and Secondary) 

A X,Y = 0.005 inches (Compensator 1 & 2) 

AZ = 0.002 inches (Compensator 1 & 2) 

A Till = 0.00001 radians (Primary) =* 

= 0.0001 radians (Secondary) 

: 0.0010 radians (Compensator 1 & 2) 

Primary Tilt is most sensitive error 


TIR (inches) 

PRI C 

SEC C 

REL 1 ( 

REL 2 « 

REL 3 ( 


0.0066 

0.0002 

0.0045 

0.0015 

0.0020 


F312511-*^22 



.118 


.004 


.121 


.004 


.000 
0.000 


3-37 

LOCKHEED-HUNTSVILLE 
















































































































LMSC-HSV TR F320789-II 


Table 3-6. Receiver Channel LOS Error Sensitivities (Rigid Body Alignment Errors) 


Element 

X-Tilt 

Y-Tilt 

X-Dec 

Y-Dec 

Z-Dec 

Primary 

o.m 

0.199 

-0.010 

-0.010 

0.00000 

Secondary 

-0.005 

• 0.005 

0.010 

0.010 

0.00000 

Relay 1 

-0.0002 

-0.0002 

-0.0033 

-0.0033 

0.00000 

Relay 2 

-0.0006 

•0.0006 

-0.0012 

-0.0012 

0.00000 

Relay 3 

0.0005 

0.0005 

0.0045 

0.0045 

0.0000 


A X.Y.Z = 0.001 inches (Primary and Secondary) 

A X,Y = 0.005 inches (Relay 1,2,3) F3i25tv»ek-23 

= 0.002 inches (Relay 1,2,3) 


A Till = 0.0001 radians (Primary and Secondary) 

3 0.0010 radians (Relay 1,2,3) 

. Primary lilt is most sensitive error 

Note: LOS displacement in object space (mrad) 


Table 3-7. Transmitter Channel LOS Error Sensitivities (Rigid Body Alignment Errors) 


Element 

X-Tilt 

Y-Tilt 

X-Dec 

Y-Dec 

Z-Dec 

Primary 

-0.020 

-0.020 

0.010 


0.00003 

Secondary 

0.005 

-0.005 

-0.010 

-0.010 

-0.00080 

Relay 1 

-0.0010 

•0.0010 

-0.0555 

•0.0555 

-0.00036 

Relav 2 

0.0018 

0.0018 

0.0550 

0.0550 

0.00023 

lESnZHH 

0.0113 

0.0113 

• 0.0005 

-0.0005 

0.00000 


F31251 l-Hsk-24 

A X,Y,Z = 0.001 inches (Primary and Secondary) 

A X,Y = 0.005 inches (Compensator 1 & 2) 

A 2 - 0.002 inches (Compensator 1 & 2) 

» 

A Tilt = 0.00001 radians (Primary) 

= 0.0001 radians (Secondary) 
a 0.0010 radians (Compensator 1 & 2) 

. Primary tilt is most sensitive error 

Note: LOS displacement is object space (mrad) 


3-38 

LOCKHEED-HUNTSVILLE 
























LMSC-HSV TR F320789-II 


Table 3-4 lists receiver channel wavefront errors induced by tilts and displacements by the 
individual optical elements. The relay elements refer to the relay optics which transfer the received 
beam through the scan bearing and form the real pupil, which is used for second order lag angle 
correction. Refocus refers to the telescope's ability to adjust secondary-primary mirror spacing for 
best focus. The optic axis for the telescope is labeled the Z-axis. 

Table 3-5 is a tabulation of similar data for the transmitter channel. The two compensator 
lenses refer to the device which adjusts the curvature of the transmitter beam before it impinges on 
the secondary mirror. In this table, as before, wavefront error is listed in fractions of the operating 
wavelength. 

Tables 3-6 and 3-7 are tabulations of the effects of tilts and displacements on the lines-of- 
sight of receiver and transmitter channels, respectively. In Table 3-7, relay 1, relay 2, and relay 3 
refer to transmitter compensator lens 1 , compensator lens 2, and compensator lenses 1 and 2 acting 
as a unit, respectively. 

Other trades were performed to ensure that the optical subsystem meets its requirements and 
still remains within weight, volume, and power restrictions. A list of the selected baseline 
approaches and alternates is shown in Figure 3-36 (D). In each case, we selected the baseline 
design adequate to meet requirements for the lowest weight or power. Ultra low expansion (ULE) 
glass has been chosen for the material for the primary mirror because it exhibits virtually zero 
coefficient of thermal expansion (CTE) and a very low variation of the coefficient of thermal 
expansion within the mirror blank. The low values of CTE and the low variation of CTE help 
minimize the sensitivity of the optical subsystem to changes in thermal soaks and gradients. Errors 
related to the non-zero CTE typically result in focus type errors which can be compensated by 
adjusting the primary to secondary mirror spacing. On the other hand, the variation of CTE within 
the mirror results in random wavefront errors which cannot be corrected. The results of an 
analysis of worst case orbital thermal effects on the optical subsystem are tabulated in Table 3-8. 

An assessment of risk areas is summarized in Figure 3-36 (E) along with approaches for risk 
reduction. 


3-39 

LOCKHEED-HUNTSVILLE . 



LMSC-HSV TR F320789-II 


Table 3-8. Orbital Thermal Analysis Summary 



AT °C 

RMS WFE 
@\=9.11 pm 

(WAVES) 
with refocus 

LOS Pointing 
Error (prad) 

Primary mirror 





Gradient 

2.5 

0.080 

0.008 


Soak (ACTE) 

53.0 

0.007 

0.007 


Soak (radius of curvature) 

53.0 

0.080 

0.008 


Metering Structure 





Despace 

53.0 

0.051 

0.001 

0.006 

Decenter (grad) 

8.0 

0.0005 

0.0003 

0.9 

Decenter (ACTE) 

53.0 

0.001 

0.0006 

2.0 

Tilt (Grad) 

8.0 

0.003 

0.0007 

4.9 

Tilt (ACTE) 

53.0 

0.006 

0.002 

10.8 

Notes: 





• Worse case orbital thermal analysis provides thermal soaks and gradients for each of the 

major components 

• Thermal perturbations used in conjunction with the optical alignment sensitivities, 
produced from the lens design, to yield the corresponding LOS and wavefront errors. 


3.3.3 Receiver/Processor Subsystem 

The receiver/processor subsystem baseline is summarized as follows: 

• Redundant HgCdTe photovoltaic detector arrays with 52 percent effective quantum 
efficiency at 100 MHz and 43 percent at 1300 MHz (47.5 percent average) 

• Mixing efficiency of 0.33 for uniformly illuminated annular aperture with ratio of inner to 
outer diameter of 0.44 

• Signal aligned on central element of array with exterior elements for alignment monitoring 

• Local oscillator beam tailored for central (signal) element for shot noise limited operation 
with phase front matched to signal beam; spill over to alignment elements 

• Redundant Split Stirling Cycle cryogenic coolers to optimize detector operating temperature 

• Redundant Split Stirling Cycle cryogenic coolers to optimize preamp operating temperature 

• Bias supply and preamplifiers space-qualified versions of standard units 

• Automatic gain control for wide dynamic range between aerosol and ground returns 

• 10 bit 75 million samples per second analog-to-digital (A/D) converter for adequate wind 
signal frequency response and dynamic range. 


3-40 

LOCKHEED-HUNTSVILLE 








LMSC-HSV TR F320789-II 


The LAWS receiver/processor subsystem consists of a wide bandwidth photo detector array, 
active cooling for the photo detector, bias circuitry, preamplifiers, and on-board signal processing 
electronics. For each of these components, several options were considered. These options will 
be outlined below, along with the logic for selection of the baseline receiver/processor subsystem 
components. 

Figure 3-39 is the receiver/processor subsystem block diagram, and Figures 3-40 and 3-41 
are views of the physical arrangement. The local oscillator optical source (upper left hand comer 
of Figure 3-39 ) from the master oscillator is expanded to match the 4 cm diameter of the beam 
received from the telescope before being focused on the photo detector. The Doppler signal is 
received from the telescope and optical train, superimposed on the local oscillator, and directed 
toward and focused on the photo detector array. Cooling is provided for the detectors. Outputs 
from the detectors are amplified and frequency shifted to the frequency/amplitude range of the A/D 
converter. The "zero" Doppler (relative to the ground) is set for the center of the 0 to 30 MHz 
baseband to minimize A/D frequency span requirements. The levels of each channel from the 
detector array are measured to monitor the received optical signal spot location upon the detector 
array for optimal alignment. The output of the A/D is buffered and telemetered to the platform data 
interface. 

3.3.3. 1 Photo Detector 

The LAWS photo detector is a critical element of the overall system. The detector detects the 
returned signal (Doppler shifted radiation) which is mixed with the local oscillator (LO) radiation at 
a controlled frequency to produce the Doppler shifted beat signal. 

The line-of-sight Doppler signal of the tropospheric winds as measured from the orbiting 
satellite will vary from +(2 A)(V S ±1 V w )Sina to -(2/X)(V s ±1 V w ) Sina. As the LAWS telescope 
traverses the conical scan, the satellite velocity either adds to or subtracts from the wind velocity 
component. For a cone half angle (oc) of 45°, a laser wavelength (X) of 9. 11 x 10 ^ m, and a 
satellite velocity (V S ) of 7.5 km/s, this satellite velocity bias varies from approximately 5.3 km/s to 
-5.3 km/s or ±1.16 x 10 9 Hz. (The wind velocity adds only ±15 MHz to this number for 150 kn 
winds.) Thus, if the detector sees a purely homodyned signal with no LO offset, it must be 
capable of efficiently detecting signals with a bandwidth of approximately ±1.2 GHz. 

Single element detectors have been built and tested with 70 to 80 percent effective quantum 
efficiency for bandwidths of less than 0.3 GHz, 35 to 45 percent for bandwidths up to 1 GHz, 
and to 35 percent for bandwidths up to 2 GHz. Figure 3-42 presents test data. Optical 
preamplifiers can lead to increasing these efficiencies, as has been demonstrated with low pressure, 
low bandwidth, optical preamplifiers for low bandwidth requirements. However, for the above 
GHz bandwidths, the optical preamplifier requires a high pressure, low electrical efficiency design, 
and is thus not included in this baseline. 


3-41 

LOCKHEED-HUNTSVILLE 



F312S99-DW-02 


LMSC-HSV TR F320789-II 



Figure 3-39. LAWS Receiver! Processor Subsystem Block Diagram 





LMSC-HSV TR F320789-II 



Figure 3-40. ReceiverlProcessor Layout 



Figure 341. ReceiverlProcessor Components - Side View 


3-43 

LOCKHEED-HUNTSVILLE 








LMSC-HSV TR F320789-II 


Quantum efficiency is stressed here because a 1 dB improvement in receiver efficiency is 
equivalent to a 26 percent increase in laser energy or telescope aperture area. Potentially the 
highest quantum efficiency could be achieved via heterodyning ™th a controllable local oscillator 
signal i e an LO which can be programmed to provide a known frequency output as a function 
of conical scanner position to compensate for the gross Doppler shift due to the satellite velocity. 
The following two methods have been discussed to offset the local oscillator frequency. 

• Shift the frequency of the LO laser with cavity length tuning 

• Externally modulate the frequency with either an acousto-optical or electro-optical (EO) 
modulator. 

The desired frequency shift of the LO is a controlled +0.9 GHz to -0.9 GHz for the 45 deg 
cone half angle. The resulting beat signal of the optical signal on the detector would be below 
+0 3 GHz. This bandwidth reduction would allow us to maximize detector performance and 
receiver efficiency. However, it has not been demonstrated as a compact, space qualifiable device, 
and is thus eliminated from our baseline design. 


F312W0W-05 


c 

o 


© 

Q- 

O 


Oi 

s 0.54 

5 


© 

> 

c 

o 

O 

I 

til 

P 


0.44 

0.38 

0.33 







' 













































■ r 









_ 

- 








* 

. 






















L 





^-Max L 

>WS Irec 
1 

yuency 

J 



Frequency (GHz) 

Note: NEP/B of 1.88 x lO^w/Hz equivalent to ideal effective heterodyne quantum efficiency 
(includes idea pre-amplifier noise figure) at 10.6 urn 


o 

X 


5 

CL 

IU 

z 


Figure 3-42. Test Data 


3-44 

LOCKHEED-HUNTSVILLE 




LMSC-HSV TR F320789-II 


Thus the LAWS detector baseline configuration requires a high bandwidth detector with a 
non-shifted LO. 

A two dimensional detector array of elements is selected over a single element to simplify 
system alignment. Matched optics are used to optimize LO distribution upon the detector elements. 
Typical detector arrays have some losses due to physical (line width) separation between e 
elements; optimal performance is achieved when the signal is directed to the single sign e ement. 
The elements will be physically arranged to allow optical alignment of all received signals upon the 
central element. Ground returns will be used to aid in this alignment process. Defocusing o t e 
receiver will allow acquisition of the ground returns from non-optimally aligned optics. 

3. 3. 3. 2 Detector Cooling 

Photo detectors operating in the 9 to 12 pm range have optimum performance when cooled 
to approximately 77 K. For long-term satellite operation, two types of cooling are potential y 
available to achieve operation at these temperatures: passive or active. 

Passive cooling is practical on satellites for low energy heat loads where free-space look 
angles are available to the detector cold finger. The cold finger must be kept short in length to 
minimize heat leaks into the detector which would raise the detector s temperature. e passi 
cooler is ruled out for LAWS baseline because of the geometries involved, the low polar or it, 
the overall cooling requirements. 

The active theimal cooler proposed for many of the other EOS Facility payloads is adequate 
and is selected for the LAWS baseline. Lifetime of the cooler is a consideration and is being 
tested/enhanced for these other programs. Vibration is a consideration which is important with the 
LAWS Instrument. Lockheed/Lucas are developing a very low vibration cryocooler assem y. 
Care must be taken in designing the mechanical fixtures and providing vibration isolation where 
required. Views of the cooler arrangement are shown m Figures and ^ 
schematic is shown in LAWS DR-8, Preliminary Design Document (LMSC-HSV TR rjlaby 

3. 3. 3. 3 Bias and Preamplifiers 

Bias and preamplifiers for the LAWS receiver are very similar to those used for conventional 
coherent lidar systems, but the LAWS device must be space-qualified and operates over a very 
widt dynamic rLge, with f> varying from 10r» to !0« nr* sr* (plus speckle) and ground returns 

varying up to 10*^ (plus speckle). 


3-45 

lockheed-huntsville 



IMSC-HSV TR F320789-II 






LMSC-HSV TR F320789-II 



3-47 

LOCKHEED-HUNTSVILLE 


Figure 344. Vacuum Dewar with Cold Fingers, Detectors, and Pre-Amps 









LMSC-HSV TR F320789-II 


To provide this wide dynamic range and incorporate very low noise preamplifiers, an 
electronic switch (cooled with a 0.1 dB noise factor) is used in the signal channel to (1) switch 
between preamplifier frequency ranges and (2) switch in a shunt when the preamplifiers become 
saturated. The preamplifiers - cooled only where required - are switched between frequency 
spans as a function of scanner azimuth angle. If saturation occurs over 50 ns, the shunt is 
switched in (gallium arsenide preamplifiers recover in this time) and Earth returns are measured 
with unsaturated amplifiers. Shorter saturations due to speckle do not activate the switch. Actions 
of the switch are monitored and entered into the data stream for subsequent amplitude data 

processing. 

Less concern about preamplifier noise is applied to the outlying alignment detectors. Wide 
dynamic range is also a requirement. Thus the signal is split prior to the first preamplifier, with the 
low level signals receiving 30 dB more gain than the higher level signals. A less than 3 dB loss is 
incurred in this split. A single preamplifier is used to span the entire frequency range, with less 
stringent control of preamplifier noise than for the signal channel. Knowledge of scanner azimuth 
angle and satellite velocity are again used to reduce the A/D conversion frequency requirement to a 
modest 3 MHz bandwidth. A/D output is fed to the computer, where sum and difference alignment 

computations are made. 


3. 3. 3. 4 Signal Processor 

The signal processor receives the preamplified signal from the preamplifiers, provides gain to 
the signal appropriately for input into the A/D converter, and performs any required addition! I on- 
board signal processing. A signal amplitude detector (i.e.. a track and hold and narrow band A/D) 
is required for each detector element for alignment purposes under conditions of strong returns. 
For baseline configuration, a frequency synthesizer is used to convert the 0 to 1.2 GHz signal into 
a 0 to 30 MHz signal analog bandwidth. The 0 to 30 MHz allows measurement of line-of-stght 
wind velocities from -150 to +150 kn or over any selected 300 kn span (e.g., from -50 to +25 


kn). 

Discussions by the Science Team have revealed a potential requirement for real time wind 
velocity (frequency spectra) data to be downlinked directly from the LAWS Platform. To meet this 
requirement, an optional on-board FFT processor is offered. To provide ±100 kn winds with 
1 m/s resolution (0.2 MHz), a 512 point FFT processor is selected for 256 point frequency 
resolution. This will be a miniaturized version of the unit we have operating in the laboratory 

today. 


3. 3.3. 5 Summary 

Figure 3-45 provides an overview of the receiver/processor subsystem development, 
including schedule, component quantities, requirement, implementation and verification, planned 
trade studies, and risks. 


3-48 

LOCKHEED-HUNTSVILLE 



0 


RECEIVER / PROCESSOR PLAN OVERVIEW 

i 


TASKS 




1994 


MAJOR MILESTONES Jfc A A 
ATP PRR PDR 

Design & Devel. 

Detector Arrays 
Electronics 
Bias, Amps, SW 
A/D, Controls 
Optics 

Cryo Coolers 
Interlaces 
Software 


Fabrication 

Components 

Interfaces 

Integration 

Eng. Unit 
Qua). Unit 
Flight Unit 

Test Support 

Eng. Unit 
Qual. Unit 
Flight Unit 

Engr. Support 

Bus. Integration 
LV Integration 
Launch Support 
Orb. Verification 
Att. Deter. Simulation 


1995 


A 

CDR 







9/94 

r 

4/95 


7/95 


1996 


9/95 1 1/95 


1997 


1/97 6/97 


1998 


1999 


LAWS Ship 


5/98 12/98 


W 2/9 

r 


E 


REQUIRED SUBSYSTEM EQUIPMENT 

COMPONENT* 

SOURCE 

QUANTITY/UNIT 

ENG. UNIT 

QUAL. UN 

Detector Array 

RP 1 

2 

2 

2 

Support Optics 

RP 2 

1 set 

1 set 

1 set 

Support Electronics 
Bias ckt, Amps 

RP3 

1 set 

1 set 

1 set 

A/D Conv., Controls 

RP4 

2 sets 

2 sets 

2 sets 

Cryo Cooler Assembly 

LMSC 

4 

4 

4 

Cables 

LMSC 

2 sets 

2 

2 


* "S" parts 

" Engineering unit components used for spares 



FOLDOUT FRAME 

















LMSC-HSV TR F320789-II 


2001 


A Launch 
LAWS/Bus 


1/01 2/01 
1 2/01 


fcl REQUIREMENT IMPLEMENTATION/VERIFICATION 

KEY REQUIREMENT 

IMPLEMENTATION 

. 

VERIFICATION 

Operational Life 

• 5 yr on orbit 

• No single point fail 

Comparison and test 
Analysis and test 

Performance 

• A/C quantum effect 

• Closed loop tracking 

• Acceptable aging 

• Temperature control 

• Data handling/control 

Measurement 

Analysis, measurement 
and simulation 

Measurement, analysis, 
comparison 

Measurement and analysis 
Simulation 

Interfaces and Software 
Functions 

• Ground return alignment 

• Automated gain control 

• Data digitization & storage 

• System performance 
monitor 

Simulation and test 


PLANNED TRADE STUDIES 



TRADE ITEM 


Cooled vs. uncooled Amps 
Number of pre-amps for signal detector 
Redundant vs. nonredundant 
Adjustable focus vs. fixed miniscus lens 
Dual tip-tilt vs. single for L.O. adjustment 
Number of array elements 


BASELINE DESIGN 


Cooled where noise figure is improved 
Baseline is four switched pre-amps 
Redundant detectors and coolers 
Adjustable focus 
Dual 

Four alignment plus central 


E 

RISK SUMMARY 

RISK ITEM 

RISK LEVEL 

RISK REDUCTION APPROACH 

Detector Failure 

Moderate 

1 . Produce several batches of detectors and perform 
accelerated aging tests. 

2. Design with redundant detectors. 

Loss of S/N from 
misalignment 

Moderate/Low 

1 . Design for graceful S/N loss from misalignment. 

2. Design for low BW on orbit alignment correction. 

Cooler failure 

Low 

1 . Lockheed developing/qualifying under EOS-A 
contracts. 


[FJ PERFORMANCE ENHANCEMENT TOOLS 


POTENTIAL ENHANCEMENT 


Detector A/C Perform 18 to 30 month development/test effort; anticipate 30 to 60 % 

quantum efficiency performance improvement. 


































LMSC-HSV TR F320789-II 


3.3.4 Structures and Mechanical Subsystem 

This section describes key analyses, trades, and verification plans for the LAWS structures 
and mechanical subsystem (SMS). (The thermal control system, which is pan of the structures 
and mechanical subsystem, is discussed in paragraph 3.3.6.) The major structural elements of the 
SMS are the base platform, the telescope mounting pedestal, the optical bench, and the telescope 
structure. The SMS mechanism is the telescope motor/bearings with V-band caging device for off- 
loading the bearings during ascent 

The base structure design is structural edge beams with internal cross beams covered by top 
and bottom face sheets. All components are constructed from graphite epoxy for light weight, high 
strength, and low thermal coefficient of expansion. Three kinematic mounts provide the structural 
interface between the LAWS Instrument and the spacecraft. All components are sized for the 
launch loads with the prescribed safety factors. 

The base structure is the mounting platform for the laser, telescope, and majority of other 
subsystem components. The subsystem components are mounted around the perimeter on the 
edge beams. The location is based on thermal requirements to take maximum advantage of passive 
heating or cooling. 

The optical bench is attached to the base structure by three kinematic mounts. The optical 
bench is a honeycomb structure with face sheets, and is made of graphite epoxy material for 
minimum distortions and light weight. The seed laser, local oscillator, detectors, and all relay 
optical system elements are mounted on the bench. 

3.3.4. 1 Requirements and Design Margins 

The SMS baseline design was developed by combining directly specified requirements and 
derived requirements from the spacecraft, telescope, and optics system levels to their respective 
structure and mechanical subsystems. A summary of key SMS requirements and respective 
verification methods is given in Figure 3-46 (B). 

3. 3. 4. 2 Analyses and Trade Studies 

NASTRAN structural math models were developed to perform deformation, stress, and 
modal analyses. Preliminary sizing of the structural components was based on these analyses for 
stiffness and launch loads. Modes, frequencies, and structural response to the launch loads and to 
the laser pulses were determined and the structure sized for these load environments. 


3-50 

LOCKHEED-HUNTSVILLE 



J3 

TASKS 

MAJOR MILESTONES 
Qualification Unit 
Base-Design, Fab, StrTest 
Bench - Design, Fab, Str Test 
Mounts - Design Fab 
Base Assembly 
T eiescope 

Telescope Motor Bearing 
Telescope Assembly 
Mass Simulators 
SMS Assembly 
SMS Align, Balance, Test 


STRUCTURES & MECHANICAL DDT&E PLAN OVERVIEV 


1994 


1995 


~ ks z r 

ATP PRR PDR 


“ 2 T~ 

CDR 




Tfc 


“W 


1996 


1997 


1998 


1999 


— 25 “ 

LAWS Ship 


Flight Unit 

Base - Fab. Struct T est 
Bench - Fab. Struct Test 
Mounts - Fab 
Base Assembly 
Base Instruments 
Base Subsystem Assembly 
Laser Subsystem 
Bakeout & Assembly 



LAWS Assembly 
Alignment Tests, Bal, Wt 
Telescope 

Telescope Motor Bearing 
Telescope Assembly & Bakeout 
Mirror 

Telescope Subsystem Assembly 
Alignment Check 


3/96 

c 


B 

SMS REQUIREMENTS/VERIFICATION 

KEY REQUIREMENT 

IMPLEMENTATION 

Launch Vehicle Interface 

• Shroud Envelope 

* Interlace Loads 

Designed to meet envelope for max ascent loads 

Contamination 

Contamination shield material selection 

Strength/Dynamic Characteristics 

Designed for positive margins with adequate factors of safety & inert structural 
frequency/stiffness requirement 

Operational Life 

Motor bearing design 

Redundancy Management 
Allignment/Stability 

Redundant motor bearings 

* Telescope Rotation 

Telescope dynamically balanced 

* Laser Pulse 

Structure stiffness/shock mounts 

* Thermal Deflections 

Thermal covers/control 

• On Orbit Dynamics 

Structural stiffness design 

• IG/OG Distortion 

Structural stiffness design 


ATTITUDE D 


/ 

frame 


foldout 


LMSC-HSVTR F320789-II 


2000 




LAWS Bus 


2001 
Launch 


C DESIGN ANALYSES & TRADE STUDIES | 

ITEM 

ANALYSES 

Optical Bench' 

To determine weight/stiffness/strength optimum for Honeycomb or 
multiple truss core 

Base Structure' 

To determine weight/stiffness/strength optimum for GE member 
size and layup 

Telescope Pedestal 

To determineweight/stiffness/strength optimum for material 
trade and design 

Laser Mounts* 

To determine laser pulse effects on telescope pointing 

SMS* 

To determine sensitivity of telescope imbalance on telescope . 
attitude & optics alignment 

SMS 

To determine the effect of gravitational field alignment at on orbit 
conditions 

SMS 

To determine changing structural design effects on dynamic 
modes & natural frequencies (thereby, attitude control) 

SMS 

To determine space platform effects on attitude control 

SMS 

To determine thermal distortion effects on attitude control and 
optics alignment 


•Ongoing analyses begun in Phase B. 


[d] PLANNED TRADE STUDIES 1 

TRADE ITEM 

BASELINE DESIGN 

Telescope Support Pedestal 
Optical Bench Core 
Base Thickness 

Titanium vs. Graphite Epoxy 
Honeycomb vs. Multiple Truss 
Thick, Thin, Medium (completed) 


VERIFICATION 

Test & Analysis 

Test & Analysis 
Test & Analysis 

Test & Analysis 
Test 

Test 

Test & Analysis 
Test & Analysis 
Test & Analysis 
Test & Analysis 

TERMINATION SUBSYSTEM 


E 

SMS VERIFICATION SUMMARY 








£ 

> 

£ 


CO 

o 

3 

o 

35 




i 

o 



• 

> 

« 

JC 

1 



c 

3 

£ 

s 

! 

i 

« 

i 

i? 

(0 

3 

8 

(0 

S 

>* 



LL. 

<0 

* 

ft 

H 

< 

CL 

SMA Qualification Structure 

X 

X 

Q 

Q 

Q 

Q 

Q 

w/Mechanism 









SMS Flight Structure 


X 

X 

A 

A 

A 

A 

A 


X 

_ 

Same Levels Qual/Flight 





Q 

= 

Qualification Test Levels* 





A 

- 

Acceptance Test Levels* 





* 

= 

Levels Per Mil-Std-1540B 





312594-MT-FO 

Figure 3-46. OverviewlSummary of the Structures 
and Mechanical Subsystem (1 of 2) 


3-51 

LOCKHEED-HUNTSVILLE 


pm nm fT 


FPfiMF 



[F] dynamic test plan/features 

• Free-Free Modal Test 

• Measure dynamic stiffness of spacecraft interface via impedance test 

• Combine results of these two tests to produce fixed base mode shapes and natural frequenr 

- Test article suspended by air bearings 

- All suspension system modes below 2 Hz 

- Pure random excitation 

- -50 + acceleration measurements 

- Modal curve fitting techniques extract mode shapes, natural frequencies, and modal 



G 


RISK SUMMARY 

RISK ITEM 

RISK LEVEL 

RISK REDUCTION APPROACH 

Structural Assembly Failures 

Low 

- Large strength margins 

- Early identification and control of fracture critical items 

Motor/Bearing Failure 

Low 

- Redundant motor wiring 

- Similarity with other flight proven units 

SMS Attitude Control Failure 

Low/Med 

- Dynamic analyses with respect to space platform pertur 

- High rev dynamic balance of telescope 

- Deflection analyses supported by tests 

Optics Alignment Failure 

Low/Med 

- Thermal deflection analyses and testing 

- Dynamic analyses with respect to space platform pertur 


— 

E 

REQUIRED SUBSYSTEM EQUIPMENT 


COMPONENT 

SOURCE 

HERITAGE 

FLT 

QUAL 

MO 

— - 

Base 

LMSC 

New 

1 

1 



Bench 

LMSC 

New 

1 

1 


— 

Telescope Mount 

Vendor 

New 

1 

1 



Motor Bearing 

Vendor 

Modified Flight Proven 

1 

1 



Telescope 

Vendor 

New 

1 

1 



Mirror 

Vendor 

New 

1 

0 



Test Hardware: 







Mass Simulators 

LMSC 

New 


1 ea 



Test Fixture 

L * 

LMSC 

New 


1 



/ 

FOLDO’JT FRAME 


F 


LMSC-HSV TR F320789-I! 


ies 


lamping I 


ances 


ances 



CD 

PLANNED SMS ANALYSES 



ANALYSIS TYPE 

ALL SMS 
EQUIPMENT 

ALL SMS 
STRUCTURES 

SMS 

MECHANISMS 

Strength 

X 



Dynamics 


X 


Thermal 

X 



Mass Properties 

X 



Producibility 

X 



l ife Cycle Cost 

X 



FMEA 

X 



Reliability 

X 



Venting 

X 



Stress Controls 



X 

Performance 



X 

Math Model Verification 


X 

nr-n < ■ it r- rv _ l n 


312594-MT-FO-2 of 2 


m PHASE B 
^ STRESS/DYNAMICS 


Equipment packages are 
reproduced as point masses 
(not plotted). 

454 Grids 
1034 Elements 



Figure 3-46. Overview! Summary of the Structures 
and Mechanical Subsystem (2 of 2) 


Fou 


Si 

ldgut frame 


3-52 

LOCKHEED-HUNTSVILLE 





LMSC-HSV TR F320789-II 


The thick/thin base trade study is complete and results included in the 0.34 m thick baseline 
design. This study is presented in detail in Lockheed Engineering Memorandum LMSC-HSV EM 
F3 12479, 22 April 91. 

Key planned design analyses and trade studies for Phase C/D are given in Figure 3-46 (C). 
These studies will complete the studies initiated in Phase II for SMS optimization. The analyses in 
Phase H are discussed in paragraph 3.3.4.8. A trade study summary is given in Figure 346 (D). 

3. 3. 4. 3 Flow Network/Schedule 

The schedule for major SMS DDT&E activities, shown in Figure 346 (A), includes efforts 
directed at risk reduction and a natural flow and integration of hardware consistent with specified 
program milestones. 

3. 3. 4. 4 Verification Process 

A summary of the verification approach for major SMS components is given in Figure 3-46 
(E). Qualification and acceptance test levels are in accordance with MIL-STD-1540B. A 
pyroshock test is required for the telescope motor bearing decaging device. A functional test is 
required for all subsystems after each environmental test for both qualification and flight units. 
Static tests on base and bench are performed to verify calculated structural stiffness (the math 
models) and to verify manufacturing/design integrity. Corrections required due to these tests 
(performed soon after fabrication) will have minimal schedule impact and ensure compliance with 

critical performance requirements. 

The dynamic test plan and features are given in Figure 3-46 (F). These tests will further 
verify the math model, provide a basis for the coupled load analysis, and qualify/accept the 
respective units per MIL-STD-1540B. 

3. 3. 4. 5 Risk Reduction 

All elements of the SMS will be analyzed for high risk identification. Each major assembly 
of the SMS has appropriate risk reduction actions defined. Figure 346 ( G ) summarizes SMS risk 
reduction. The motor bearing mechanism considered will be similar to a flight proven model. The 
base and bench structures will be both statically and dynamically tested and modeled. 

3. 3. 4. 6 Equipment Summary 

SMS hardware to be produced during Phase C/D is categorized in Figure 346 (H). 

3. 3.4.7 Planned SMS Analyses 

Key planned SMS analyses are summarized in Figure 346 (I). Strength, dynamics, thermal 
deflection, and stability analyses will be made with a continuously updated math model. This 
model will incorporate design changes and will be verified by qualification static, dynamic, and 
weight measurements. The Phase II math model is shown in Figure 346 (J). 


3-53 

LOCKH EED-HU NTSVILLE 



LMSC-HSV TR F320789-II 


3.3.4. 8 Phase II SMS Analysis Summary 

Five SMS analyses were initiated in Phase II and will be ongoing as the LAWS design 
matures and more accurate data are incorporated. These analyses will be discussed in the 
following paragraphs. 

The LAWS SMS platform thickness trade study has been completed and is documented in 
LMSC-HSV EM F3 12479. 

The current modes and frequencies summary for free and constrained conditions is presented 
in Table 3-9. Typical mode shapes are given in Figure 3-47. These data are a result of the latest 
mass and motor bearing stiffness data. Current caged and uncaged effective stiffness are almost 
equal, which means that constrained on orbit and constrained lift-off modes are very similar. 

Interface reaction loads and key deflections under launch and staging conditions are given in 
Tables 3-10 and 3-11. These data result from worst case static plus dynamic design load factors 
obtained from the Titan IV User's Handbook. Maximum lateral loading occurs at liftoff, while 
maximum axial loading occurs at stage 1 burnout. The telescope deflections reflect the latest 

motorbearing caged stiffness. 

Table 3-9. LAWS Natural Frequencies and Mode Shapes Telescope Motor Bearing 
Supported (Caged) 


MODE DESCRIPTION 

FREE-FREE 

(PRIMARY MOTION) 

MODE 

FREQUENCY (Hz) 

Rigid body modes 

1-6 

Approximately 0.0 

Telescope pitch (X rotation) 

8 

19.17 

Telescope yaw (Z rotation) 

7 

14.97 

Telescope roll (Y rotation) 

- 

— 

Laser Y translation 

- 

— 

Laser roll (Y rotation) ! 

9 

24.62 

Telescope sun shield breathing 

10 

30.2 

Optical bench warp 

11 

37.31 

Base/bench warp 

12 

39.27 

Telescope sun shield ringing 




CONSTRAINED 


MODE 


NA 

1 

2 

3 

4 

5 

6 

7 

8 
9 


FREQUENCY (Hz) 


NA 

13.0 

13.8 

14.6 

23.3 

24.4 

34.6 

37.2 

40.2 

42.5 


3-54 

LOCKHEED-HUNTSVILLE 






LMSC-HSV TR F320789-II 



FREE -FREE 
FREQUENCY = 14.97 Hz 



CONSTRAINED 
FREQUENCY = 13.8 Hz 



FREE - FREE 
FREQUENCY = 19.2 Hz 



FREQUENCY = 13.0 Hz 


Figure 3-47. Typical Mode Shapes 


3-55 

LOCKHEED-HUNTSVILLE 


LMSC-HSV TR F320789-II 


Table 3-10. Interface Reaction Loads 


REACTIONS (Norton*) 


-1 57E+4 1 12E+4 1 57E+4 -1 12E+4 

•1 53E+4 *€ 62E+3 -1306+4 -6 56E+3 

4 2E+1 -1 95E+4 1 60E+2 -i 52E+4 


1 oads m shown an ’wont cam’ (static ♦ 
dynamic} daatgn load factors obtainad from 
tha Titan /V Uaars Handbook. Maximum 
iatarai loading occurs at lift-off whda max 
axial loading occurs at stags 1 burnout. 



Rz 


-2 81E+4 -9 65E-1 

-5 49E-5 1 32E+4 

9 206-5 -179E+4 



Table 3-1 1 . Static Deflections 


LOAD 

LOCATION 

DISPLACEMENT (M) 

3.5 Gx 

Secondary Mirror 

6.45E-3 (X) 

3.5 Gx 

Spin Bearing CG 

1.33E-3 (X) 

3.5 Gx 

Laser Power Supply 

3.06E-3 (X) 

3.5 Gx 

Optical Bench (Detector) 

1.17E-3 (X) 

3.5 Gy 

Secondary Mirror 

9.74E-3 (Y) 

3.5 Gy 

Spin Bearing CG 

1.01 E-3 (Y) 

3.5 Gy 

Laser Power Supply 

5.25E-3 (Y) 

3.5 Gy 

Optical Bench (Detector) 

2.99E-4 (Y) 

6.5 Gz 

Secondary Mirror 

1.03E-3 (Z) 

6.5 Gz 

Spin Bearing CG 

8.48E-4 (Z) 

6.5 Gz 

Laser Power Supply 

1.69E-3 (Z) 

6.5 Gz 

Optical Bench (Detector) 

9.06E-4 (Z) 

" F312fl®MIT-04 


3-56 

LOCKHEED-HUNTSVILLE 










LMSC-HSV TR F320789-II 


The laser pulse analysis was performed to ensure that the acoustic shock of laser firing does 
not propagate through the structure and disturb the operation or performance of the LAWS 
Instrument. The analysis also ensures that the firing frequency of the laser does not couple with a 
natural frequency of the structure to produce instability. The following assumptions are made: 

• One percent of discharge (2 J) goes to acoustic impulse 

• Load profile is sinusoidal over 0. 1 ms 

• Laser fires every 62.5 ms (16 Hz) 

• Transient model was run for 2 s (32 laser firings). 

The following conclusions were reached: 

• Steady state response is reached within 1 s 

• Maximum deflection at detector = ± 0.12 pm 

at telescope = ± 0.08 pm 

• Maximum rotation at detector = ± 1 prad 

at telescope CG = ±0.15 prad 

• No coupling of lower modes with laser firing frequency. 

Transient response plots are given in Figures 3-48 through 3-51 . This analysis will be 
continued with continued laser mounting definition and enclosure design. 

A study was made to determine the sensitivity of static and dynamic telescope balance to its 
attitude stability. The results are as follows (assuming 188 kg mass rotating at 8.3 rpm as shown 
in Figure 3-52): 

• For static imbalance only: 0.1 16 prad/m (or 0.1 16 prad telescope deflection for one mm 
CG off axis of rotation); the variation from x direction to y direction is 34 percent 

• For dynamic imbalance only: 1.5 prad/N»m with x/y direction variation of 2 percent. 

Note: As modeled, with no provisions for dynamic balancing, the telescope exerted 5.26 N*m 
moment at 8.3 rpm. 

We concluded that, since telescope attitude budget is 50 prad per resolution, telescope 
balance is not as critical as anticipated. However, design and testing to minimize imbalance loads 
will be undertaken. 


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LMSC-HSV TR F320789-II 


m 



TMC (SECONOS) 


Figure 3-48. Transient Response at Detector Due to Laser Firing Acoustic Shock 



Figure 3-49. Transient Response at Detector Due to Laser Firing Acoustic Shock 


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LOCKHEED-HUNTSVILLE 







LMSC-HSV TR F320789-II 


M 

2 


CL 

<A 

O 


Q 

< 

GC 


O 

GC 



Steady State Deflection 
Approximately ±0.06 jam 





rtilKOM ca 


Steady State Rotation 
Approximately ±0.12 jarad 


Figure 3-50. Transient Response at Telescope CG Due to Laser Firing Acoustic Shock 



0.6 j 


0.6 



b 

0.4 

X 


2 

0.2 

GL 

tn 

Q 

0.0 


-0.2 


-0.4 


-0.6 


n 












i 




- — 






— : 





/ 

\ 



1 — 


t| — - 


\ 


■ 


\ 

\ 



1 


> ! 


\ 





v_ 

Li 



2 



\ 

-t — 

/ 



\ 

f 

1 



4 

6 





1 c 


is 

ISA 

1.1 

KO 


TIME (SECONOS) 


M 

o 

< 

s 


o 

GC 


1.5 

1.0 

0.5 
0 O 
-0.5 


- 1.0 















f 





i 

_i 

4 



/ 

/ 

'\J 

i 

i 

\ 

/ \ 


i\ 

/ 




T 

“V 



/ ^ 

\ 

r 

4 

J 


L 











\j 




TUIKOM ca 


1.910 


1.920 


TIME (SECONOS) 


Figure 3-51 . Transient Response at Telescope CG Due to Laser Firing Acoustic Shock 


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. Due to telescope/base deflection with respect to space platform 
. For 188 kg mass rotating at 8.3 rpm 


STATIC IMBALANCE 
SENSITIVITY: 

95.98 n rad/m off axis of rotation 
(or .096 urad/mm CG off axis) 
Variation X to Y direction = 34% 


DYNAMIC IMBALANCE 
SENSITIVITY: 

1.5 prad/N«m of dynamic imbalance 
(as modeled telescope has 5.26 N*m 
dynamic imbalance) 

Variation X to Y direction = 2% 


Figure 3-52. LAWS Telescope Attitude! Balance Sensitivity 

3.3.5 Attitude Determination, Scan Control, and Lag Angle Compensation 
3.3.5. 1 Introduction 

The successful operation of the LAWS Instrument dictates that attitude knowledge, attitude 
accuracy, and transmit-teceive alignment be controUed/maintained within acceptable limits. 

Attitude knowledge is maintained by the onboard attitude determination subsystem, which 
determines the LOS of each outgoing laser pulse. An accurate accounting of this parameter^ 
required in order to permit the resolution of the orbital and Hard, rotanon velocty component 
along the LOS of the laser pulse. These velocity components must then be removed from th 
measured LOS velocity in order to determine the true wind velocity. 

The attitude accuracy involves the control of space platform attitude and scanner position 
such that the desired shot placement results. 

The transtnit-receive alignment involves control of the receiver LOS such that it is properly 
oriented in space and with respect to time when the returned energy arrives at the LAWS 
Instrument This control requires compensation for scanner modon, space platform mooon, and 


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misalignment and jitter of the Instrument structure due to disturbance forces and thermal 
influences. The requirements for each of these parameters are presented in Table 3-12. 

Table 3-12. Critical LAWS Attitude Pointing and Stabilization Requirements 


CATEGORY 

DEFINITION 

PARAMETER 

AFFECTED 

MAJOR ERROR 
CONTRIBUTORS 


REQ’MT 1 

Transmit- 

receive 

angular 

misalignment 

(Jitter) 

Misalignment between 
transmit LOS when laser 
is fired and receiver 
LOS when backscattered 
energy returns 5 ms later 

S/N 

• Platform jitter 

• Telescope C.G. offset 
& bearing wobble 

• Lag compensation errors 

• Laser disturbance 

• Static misalignment 

Vv-B 

1 

Attitude 

knowledge 

error 

The error in determining 
the LOS of the outgoing 
laser 
energy 

LOS velocity 

measurement 

error 

• Attitude reference (IRU) 
unit errors 

• Structural flexibility 

• Scanner bearing runout 

• Static misalignment 

100 jirad 
per axle 
one 
elgma 

Attitude 

control 

The difference In the 
actual and desired attitude 
of the LAWS Instrument 

Shot 

placement 

• Platform attitude error 
« LAWS attitude knowledge 
error 

8 mrad 
per axis 
one 
eigma 


F312S84-RJ-10 


3. 3. 5. 2 Overview of Design and Design Drivers 

In order to meet the requirements for attitude/pointing described in Table 3-12, provision is 
made for control of five elements: 

• Attitude control accuracy 

• Attitude determination 

• Lag compensation 

• Platform jitter compensation 

• Transmit-receive alignment. 

Attitude control accuracy is determined by space platform attitude accuracy and LAWS 
Instrument attitude knowledge. The requirement is 8 mrad per axis. Since the LAWS Instrument 
attitude knowledge accuracy is 100 prad per axis and the quoted EOS-B accuracy is 50 arc-s 
(-250 prad) per axis, this requirement is met. 

The implementation of the four remaining attitude control elements and the primary design 
drivers is summarized in Table 3-13. A schematic representation of the implementation illustrating 
the primary hardware and computational interfaces is shown in Figure 3-53. In this figure, the 
software element for the transmit-receive alignment is combined with the lag angle compensation 
element since both functions are mechanized by action of the tilt-tip mirror. Scan control consists 
of determining a scan rate which will satisfy a combination of shot placement and static lag 
compensation requirements. The scanner drive control will then maintain the constant scan speed 


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“ — t0le ™ Ce ' ^ b0d5, - fiXed S “ Trackm MU « mounted on and par, of 
the LAWS Instrument and form the heart of the attitude determination system. 

Table 3 ' 13 - Features of Attitude Control Preliminary Design 

f ' I 


ATTITUDE CONTROL 
CATEGORY 

Attitude Determination 


Lag Compensation 


Platform Jitter 
Compensation 


Transmit-Receive 

Alignment 


3 0 _ 0 

.3.5.3 Attitude Determination 

The implementation of the attitude determination function is driven directly by the allowable 
enms m die measured LOS velocity due to etTors in the attitude knowledge. The budget for the 
LOS velocity error due to uncertainty in the LOS of the output pulse is presented in Figure 3-54 
and is approximately 100 (trad per axis, as previously stated. Of this total, 63 7 trrad 
(corresponding to an LOS velocity measurement error of 0.42 m/s) is budgeted for the Instrument 

Insmmenf" 6 " 05 ’ ^ rCqUirem ' nt is met by locatin S a ” attitude reference at the LAWS 

Figure 3-55 shows a functional diagram of the prelimintuy design for the LAWS attitude 
determination. Included are an IMU and two Star Tracker units. Software is provided to 
implement a strapdown attitude reference with the gyro readings and to provide compensation for 
attitude reference drift. The scanner encoder output is then utilized to determine the estimated 
utpu pu se in inertial space. In addition, the position in orbit resolves the LOS in Earth- 
fixed space permitting the determination in applicable coordinates such as latitude and longitude of 
tile illuminated area. This information is time tagged to conespond to each User pulse such that the 
location of the returns may also be catalogued, along with the LOS data. 


PRELIMINARY OESIGN 
METHODOLOGY/FEATURES 


Dedicated inertial reference unit and star 
trackers. 


• Static - Optics has fixed angular 
offset between transmit & receive. 

• Dynamic - Programmed motion of 
tilt-tip mirror. 


Passive techniques are used. Included 
are mechanical design including use of 
isolators and dampers. 


Low bandwidth active control using the 
multi-element detector as sensor and 
tilt-tip mirror as corrector. 


PRIMARY DESIGN DRIVER 

! OR JUSTIFICATION 




Space platform attitude reference 
output not accurate enough to 
satisfy velocity accuracy 
requirement. 


Scan and orbital motions can be 
predicted with sufficient accuracy 
for this implementation. 


Active systems require sensing and 
actuating bandwidths that extend 
the state of the art. 


A closed loop system is necessary 
to maintain the S/N at maximum. 


F312SM-nj-01 


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STAR 

STAR ANGLE A 
MAGNTTUOE ^ 

I HAOfxtno 
(2) 



DELTA ANGLES, 
PITCH, YAW, A ROLL 


SCAN 

SCANNER AZIMUTH 
READOUT 

ANOLt 

ENCOOER 



PULSE FIRING 
INDICATOR 


ONBOARD 

TIME 

REFERENCE 

CLOCK 



i-bQOlT 

GROUND 

PARAMETERS 

DATA LINK 



ATTITUDE 

DETERMINATION 

SOFTWARE 


LAG ANGLE 
COMPENSATION 
SOFTWARE 


SCAN CONTROL 
SOFTWARE 


LASER LOS VECTOR, 
LONGITUDE, LATITUDE. 
A ALTTTUOEOf 
MEASUREMENT 


MIRROR 

COMMANDS 


OOMMANOEO 
SCAN RATE 


LAWS INSTRUMENT 

COMPUTER uwwwr 


TO SCIENCE 
’ DATA ARCHIVAL 


TO TILT-TIP 
MIRROR 


, TO SCANNER 
DRIVE CONTROL 


Figure 3-53, Attitude Determination , Scan Control , and Lag Compensation Implementation 

4 To 1.1 


V L0S Errors 
Pointing Factors 

98 4 tirad / 0.64 m/s 


i.i, 2.2.1 r — 

Instrument Attitude Knowledge 
, Errors w x 


63.7 (irad 
0.42 m/s 


Attitude 
63.7 urad. 


Velocity 
0.42 m/s 


1.1. 2.2.2 


0.6 / .004 m/s 


1.1 .2.2.3 1 

Telescope Alignment 
Errors 

Attitude Velocity 

743 iirad / .48 m/s 


1.1 .2.2.1. 1 I 


Inst rum. Attitude 
Reference Unit (ARU) 




1.1. 2.2.1. 1 I 


ARU Alignment 
WRT Optical Bench 



* Pointing errors are per axis values, 1o unless otherwise 
specified. Velocity errors are total LOS velocity errors, la 

Figure 3-54 . Pointing Factor Errors 


F312511-RJ-06 


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Figure 3-55. Attitude Determination Functional Diagram 

A schematic of the IMU and Star Tracker mounting with respect to the Instrument system is 
shown in Figure 3-56. As shown, the Star Trackers view the celestial sphere in a plane normal to 
nadir and toward the cold side of the sun-synchronous orbit. The area that the Star Trackers will 
view (for an 8 deg field of view) during an orbit is also shown in Figure 3-56. 

The event shown is for a typical day approximately three months after launch. Due to orbital 
precession, the right ascension will cycle through 360 deg in one year. The bounds of the 
declination angles that the Star Trackers will view remain essentially constant throughout the yearly 
cycle. An analysis of the star field appearing within the star tracker field of view indicates that at 
least ten Star Tracker updates per orbit are practical. The proposed Star Tracker units can acquire 
stars as dim as +6 magnitude. 

A trade of permissible IMU drift rate uncertainty vs. scale factor error for 5 and 10 Star 
Tracker updates per orbit are shown in Figure 3-57. For the case of 10 updates orbit and Star 
Tracker error of 5 arc-s, an IMU with drift rate uncertainty less than 0.01 deg/hr and scale factor 
error less than 75 PPM is satisfactory. The IMU and Star Tracker design specifications are 

summarized in Figure 3-58. 

The specifications for the attitude reference are indicated in Figure 3-59. The IMU and Star 
Tracker will be procured, and interfaces (brackets, cables, etc.) will be built or procured by 

Lockheed. 


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LMSC-HSV TR F320789-II 



Right Ascension (deg) 

Figure 3-56. Star Tracker View Traced Out Over an Orbital Period 



IMU Drift Rate Uncertainty, (deg/h) 

(per axis, one sigma) 

Figure 3-57. Attitude Determination Hardware Performance Tradeoff 


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LMSC-HSV TR F320789-II 


IMU 

• Drift Rate Uncertainty: 

< 0.01 deg/b 

• Scale Factor Error: 

< 75 PPM 

Star Tracker 

• FOV > 8° x 8° 

Error < 5 arc-s 
Sensitivity * +6 magnitude 

F3125SH-RJ-06 


Figure 3-58. Preliminary Hardware Specifications for Attitude Determination 


COMPONENT 

QUANTITY 

WEIGHT 

VOLUME 

POWER 

IMU 

1 

17 kg 
37.5 lb 

33 cm x 30 cm x 28 cm 
(13"x12"x 11") 

22.5 W 

Star 

Tracker 

2 

8.2 kg ea 
18 lb ea 

18 cm dia x 30 cm long 
(7" dia x 1 2" long) 

12 W ea 



* “S" Parts 

“Engineering Unit Components Used for Spares 


Figure 3-59. Summary of Components for Attitude Determination Preliminary Design 


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3.3.5. 4 Lag Angle Compensation and Transmit-Receive Alignment 

Lag compensation is performed to account for motions of the LAWS Instrument during the 
interval between transmit and receive (approximately 5 ms for a 525 km orbit). These motions 
arise from scanner motion (approximately 6 rpm) and orbital motion (nadir tracking). The required 
compensation angle consists of a static or bias component and a much smaller dynamic component. 
Included are also components along and normal to the direction of scan. Typical static and 
dynamic lag angle components for a 525 km orbit and 6 rpm scan are approximately 2300 jirad and 
90 (irad, respectively. 

The static lag compensation is implemented by a fixed angular offset between the transmit 
and receive optics. The implementation of the dynamic lag compensation is shown in Figure 3-60. 
The dynamic lag compensation is accomplished by slewing of the tilt-tip mirror located on the 
optical bench. The lag compensation commands due to scanner motion and the compensation for 
orbital motion are combined vectorially to form the final slewing commands as shown. 
Adjustments in the static lag angle are necessary to correct for orbital altitude variations resulting 
from orbit decay and reboost. This adjustment is accommodated through onboard 
hardware/software by periodic resetting of the tilt-tip mirror "zero" position. 

The lag compensation is open loop in that it accounts for transmit-receive LOS differences 
due to known motions only, e.g., scan and orbital motions. Two other sources of transmit-receive 
LOS error are also of concern. The first of these sources is platform jitter. The two options 
considered as solutions were active and passive compensation. Active control consists of sensing 
jitter motion and compensating with a high bandwidth gimballed mirror in the receive optical path. 
Passive control consists of utilizing isolating mounts and damping where applicable to attenuate 
space platform jitter disturbances. The passive technique was selected for the preliminary design. 
The trades illustrated in Table 3-14 were the basis of this selection. The primary disadvantages of 
the active approach are the large bandwidths anticipated for sensors and actuators (estimated to be 
on the order of 1 kHz). 

The locations of isolators, if required, are anticipated between the space platform nnd LAWS 
base to attenuate the platform jitter. The characteristics of the isolators are dependent on the power 
spectral density of the platform jitter. The error due to platform jitter is budgeted at 0.7 (irad (see 
Figure 3-61). Acceptable power spectral density boundaries for residual platform jitter at the 
LAWS Instrument have been determined. These boundaries are shown in Figure 3-62. Evaluation 
of space platform jitter characteristics (when available) will permit the evaluation of the need for 
isolators and the required isolator characteristics. 

A second source of transmit receive alignment error is the misalignment due to zero-g and 
thermal cycling. These misalignments will be monitored during orbital flight and corrected using 
the multi-element detector and the tilt-tip mirror capability. 


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Figure 3-60. Lag Compensation Functional Diagram 
Table 3-14. Active vs. Passive Control of Platform and LAWS Jitter 


CATEGORY 

ACTIVE* 

PASSIVE** 

Within State-of-the-Art? 

No; high bandwidth sensors and 
actuators 

Yes 

Weight 

Lightest 

Heavier 

Risk 

Higher 

Lower 

Reliability 

Lower 

Higher 

Cost 

Higher 

Lower 

Selection 


✓ 


* Jitter measured by sensors and corrected by gimballed mirror 
"Structural design, dampers, and isolators used 


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* Total solid angular error, lo 0.95 <w 


Figure 3-61 . Transmit-Receive Error Budget Tree 


F31251 1RJ 03 



Figure 3-62. Acceptable Boundary for Platform Attitude Jitter PSD 


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3. 3. 5. 5 Scan Control 

The scan control is described above. The implementation consists of selecting a commanded 
scan rate which is optimized for shot placement. The scanner is commanded to rotate at a constant 
rate for a given orbit. The regulation of scan rate is critical to the proper operation of lag 
compensation. Based on the allocation of 2.0 urad for lag compensation error (see Figure 3-61) 

the speed should be regulated to within 0.0055 rpm. This represents a speed regulation of 0.09 
percent for a 6 rpm scan rate. 

3. 3. 5. 6 Software 

The software modules required for the attitude determination, lag compensation, and scan 
control functions consist of that software required to interface with the various hardware units 
including commands and readouts where applicable. Other functions include algorithms required 
for attitude reference propagation, coordinate transformations, and compensation for the various 
lag angle phenomena. 

3. 3. 5. 7 Structural Dynamics and Component Math Models 

General 

Math models are maintained for all major components associated with attitude pointing, 
attitude determination, and attitude stabilization of the LAWS Instrument. The significant 
parameters include dynamic characteristics, frequency response, performance, error models, etc. 

Structural/Dynamic Models 

The structural design refinement process of Phase C/D will require periodic dynamic analyses 
using an updated model. Modes and minimum natural frequencies will be considered with respect 
to attitude control error budget for structures and mechanisms. 

A telescope dynamic balance study will determine the sensitivity of both static and dynamic 
imbalance on telescope attitude and optics alignment. This study will use the model to determine 
the degree of accuracy required in balancing the telescope about its axis of rotation. The 

structural/dynamics model also will be used to determine the effect of space platform perturbances 
on attitude control. 

The accuracy of the mathematical model used in these analyses will be verified by the static 
and modal structural tests. 

3. 3. 5. 8 Simulations 

A computer simulation will be used for partial verification of the attitude pointing, attitude 
determination, and attitude stabilization concept, performance, hardware component parameter, and 
software algorithms. 


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The simulation will be based on existing 6-DOF models where possible and will incorporate 
the rigid and flexible body dynamics of the LAWS Instrument It will also account for disturbance 
forces and orbital effects and will be based on the component math models discussed above. 

3. 3. 5. 9 Stability Analysis 

In Phase C/D, models of all closed loop systems will be maintained and periodically updated. 
Stability analyses will be performed to verify that response characteristics are adequate and that 
stability margins are within accepted limits for orbital space pointing/stabilization systems. 

Typical stability analyses to be performed are illustrated by the baseline preliminary design 
for the transmit-receive alignment loop shown in Figure 3-63. Analyses will be utilized to verify 
that gain margins of 6 to 10 dB and phase margins on the order of 25 to 30 deg are achieved. 



Figure 3-63. Alignment Loop Representation for Stability Analysis 


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3 3 5 10 Mode Definition and Description 

Various modus will be provided for power up/down, checkout, initialization, and nominal 
on-orbit operation of the attitude determination, lag compensation, scanner, an a tgnmen 
operations These modes will be under the control of the LAWS Instrument computer w„h 
appropriate ground inhibit/override and status monitoring as per TBD requirement. 

3.3.5.11 Summary 

Figure 3-64 outlines the program plan, equipment list, verification approach, planned trade 
studies, and risk reduction plan for the attitude determination subsystem. 

3.3.6 Thermal Control Subsystem 

' This section discusses the LAWS thermal control subsystem (TCS). design 

reouirements in the form of thermal loads and timelines are presented and explained. Next the 
3" active (fluid loop) subsystem and the passive (radiation cooled) subsystem are 
discussed. These subsystems are shown to meet all of the stated requtrements. Fmally an 
overview chart is shown for the combined TCS. 

3.3.6. 1 Design Loads and Conditions 

" Table 3-15 summarizes the LAWS thermal loads to be dissipated by the TCS. Electrical 
powerhs also shown. The baste requirement is that the LAWS Instrument wtU not exceed an 

orbital average power of 2200 W. 

Each component which uses power and generates heat is shown. These components are 
further broken down into variable power thermal load and constant power thermal load groups. 
Thesetoads are also broken down into survey mode, 4.61 Hz average PRF, and destgn mode, 
10 0 Hz average PRF. Maximum allowable temperature for each component is also given. 

As seen in Table 3-15, the total variable thermal loads are 1444 W and 3133 W for the 
survey and design modes, respectively. The total constant load is 569 for both operating modes 
The total variable plus constant loads are 2013 W and 3702 W for the survey and design modes 
respectively The thermal load is less than the electrical load because some of the energy is 
dissipated by other means, for example that which goes out in the laser beam. 
mental load s (i.e„ solar UV, albedo, and Earth IR) are not included in the i values in Ta * e i _ 
because most of these components are either under the thermal cover or on the cold side of LAWS^ 
S components are assumed to be mounted on thermal isolators to prevent heat transfer to die 

base. 


3-72 

lockheed-huntsville 


ATTITUDE DETERMINATION PLAN OVERVIEW 


TASKS 


MAJOR MILESTONES 


Design & Devel. 

Inertial Reference Unit 
Star Tracker 
Interface Design 
Software Req'ts 
Software Development 


Fabrication 
Mech. I/F 
Elec. I/F 

Thermal Protection 
Alignment IF 

Integration 
Eng. Unit 
Qual. Unit 
Flight Unit 


Test support 
Eng. Unit 
Qual. Unit 
Flight Unit 


Engr. Support 
Bus. Integration 
LV Integration 
Launch Support 
Orb. Verification 
Att. Deter. Simulation 



COMPONENT* 


Inertial Reference Unit 
Star Tracker 
Mechanical Interface 
Cables 


REQUIRED SUBSYSTEM EQUIPMENT 


SOURCE 



quantity/unit 

ENG. UNIT 

1 

1 

2 

2 

3 

3 

6 

6 


QUAL. U 


1 



* “S” Parts 

** Engineering Unit Components Used for Spares 


foldout frame 













LMSC-HSV TR F320789-II 


2000 


20 01 _ 
2?T1\ Launch 


LAW Sr! 


Bus 


1/01 2/01 

3 

-, 2/01 


2/M 


C REQUIREMENT IMPLEMENTATION/VERIFICATION | 

KEY REQUIREMENT 

IMPLEMENTATION 

VERIFICATION 

Operational Life 

5 yr on Orbit 

Companson and Test 

Performance 

• Attitude Knowledge: 

100 prad/Axis, One Sigma 

• Receive Transmit Align: 

3 prad/Axis, One Sigma 

• Pointing Accuracy: 

8 m rad/ Axis, One Sigma 

Analysis and 
Simulation 

Interfaces and Software 
Functions 

• IRU Attitude Update 

• Star Tracker Update 

• Lag Compensation 

• Receive-Transmit Alignment 
Loop 

Simulation and Test 


| [d] PLANNED TRADE STUDIES ~~ j 

TRADE ITEM 

BASELINE DESIGN 

No. of Star Updates Per Orbit vs. IRU 
Performance 

On Orbit Recalibration Procedures for 
Attitude Determination 

Methodology of Compensating for Space 
Platform Jitter; Active vs. Passive 

10 Updates/Orbit, Scale Factor Error < 75 PPM, 
Gyro Drift Rate Uncertainty < 0.01 deg/hr 

Use Hard Target Return to Recalibrate LOS ol 
Outgoing Laser Beam 

Passive; Use Isolators Between Base Assembly 
and Optical Bench as Required (Need Goddard 
to Supply Jitter PSD of Space Platform) 


E 

RISK SUMMARY | 

RISK ITEM 

RISK LEVEL 

RISK REDUCTION APPROACH 

IRU Failure 

Low 

Space Qualified and Demonstrated Unit 
with Built-in Double Redundancy for 
Each Attitude Axis 

Star Tracker Failure 

Low 

Space Qualified and Demonstrated Unit 

Misalignment Due to Zero g 
and Launch 

Moderate 

Develop Methodology to Recalibrate 
Using Hard Target Returns 


Figure 3-64. Overview! Summary of the Attitude 
Determination Subsystem 



FOLDOUT FRAME 

3-73 - 

LOCKHEED-HUNTSVILLE 












LMSC-HSV TR F320789-II 




Electrical 

„ Power. W 

Maximum 

Operation 

Temp°C 461 Wx 


Variable Power/Thermal Load 
Laser 

Power Supply 
Laser Energy Loading 

Total Variable Load 

Constant Power/Thermal Load 
Laser 

Laser Fans 
Thyratron Filament/ 
Reservoir 
Local Oscillator 
Power Supply 
Energy Loading 
Seed Laser 
Power Supply 
Energy Loading 
Receiver 

Det Bias/Preamp 

Cryocoolers 

Electronics 

Optics 

Azimuth Drive 
Moment Compensator 
Telescope Thermal Control 

Electrical 

Power Distribution 
Thermal 

Thermal Control I 

Comm, and Data Handling 
Flight Computer 

Attitude and Position Ref 
IMU 

Star Tracker 

Total Constant Load 

Total: Variable + 

Constant Load 


40 66 


50 23 

50 25 


Thermal 
Load, W 

4.61 Hz 10 Hz 
RefPRF AvgPRF 
(Survey {Oeeign 
Mode) Mode) 

Central System 
Heat Rejection. W 

4.61Hz 10H2 

RetPRF AvgPRF 
{Survey (Oe$igrt 
Mode) Mode) 

384 

833 

384 

833 

1060 

2300 

1060 

2300 

1444 

3133 

1444 

3133 

40 

40 

40 

40 

135 

135 

135 

135 

8 

8 

8 

8 

22 

22 

22 

22 

27 

27 

27 

27 

73 

73 

73 

73 

10 

10 

_ 

_ 

50 

50 

50 

50 

20 

20 

— 

— 

30 

30 

30 

30 

15 

15 

15 

15 

10 

10 

- 

- 

66 

66 

66 

66 

15 

15 

- 

- 

23 

23 

. 



25 

25 

- 

- 

569 

569 

466 

466 

2013 

3702 

1910 

3599 



103 

F31 2594-33 


* Required for heaters on mirrors and telescope, therefore not included in thermal load 


3-74 

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LMSC-HSV TR F320789-II 


Figure 3-65 shows a typical power and thermal load timeline or schedule for one orbital 
period. The shot frequency is managed to yield an orbital average power consumption of 2200 W. 
The shots are scheduled to prevent overlap along the ground track as the orbit approaches and 
passes over the poles. This reduced frequency conserves energy which can then be used m 
operating at the 10 Hz design mode for some time without exceeding the 2200 W orbital average 
limit In this case, 879 seconds of design mode operation were obtained. The orbital conditions 
used for this case are also shown in Figure 3-65. Again, note the difference in the electrical and 

thermal loads. 

The life expectancy requirement for the LAWS TCS is from 5 to 7 years operation in orbit 
without maintenance. 

The following orbital parameters are planned for LAWS: 

• Sun synchronous orbit 

• Orbital altitude = 525 km 

• Orbit inclination = 97.497 deg 

. 6:00 a.m. (14 hr GMT) launch due south from Vandenberg AFB. 



Figure 3-65. LAWS Power and Thermal Load Schedule 


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LMSC-HSV TR F320789-II 


These conditions yield an annual beta angle range of approximately 59 to 90 deg. These two 
beta angles were used as upper and lower limits in the passive system design. This orbit results in 
an occultation period (i.e., not full sunlight for the entire orbit), which gives the percent time in the 
sun for each day of the year. This percentage ranges from 77 to 100 percent. Both the power 
system and the TCS were designed to accommodate these conditions. 

3. 3. 6. 2 Design Approach 

The design approach used for the LAWS TCS is to dissipate as much waste heat passively 
(i.e., by radiation to space) as possible. However, many of the larger loads, particularly the laser 
gas cooling, have to be taken out by using a convective heat exchanger. This then dictates the use 
of an active pumped coolant loop system. Table 3-15 shows which components are cooled 
actively and which are cooled passively. 

3. 3. 6. 2.1 Active Thermal Control System Description 

Figure 3-66 is a schematic of the LAWS active TCS. A two-stage centrifugal pump is used 
to flow the coolant through all components to be cooled and then through two 1,000 W coldplates 
which are mounted back-to-back with the spacecraft central system coldplates. The coolant being 
used is a 30/70 ethylene-glycol/water mixture with a freezing point of approximately -18 °C (0 °F). 

Following the flow path of Figure 3-66, after the coolant leaves the coldplates it begins its 
circuit to pick up heat, going first through the components to be maintained at the lower allowable 
temperature and then proceeding to the higher allowable temperature components. The cryocooler 
compressor and expander thermal/mechanical mounting flanges are cooled first, then the seed 
lasers and local oscillators. A flow divider device then splits the flow into two equal parts which 
flow in parallel through the laser gas convective heat exchangers. The flow then combines into a 
single line and cools the laser power supply and laser fan motors. Vibration isolation loops are 
provided between the optical bench and laser components to reduce vibration transmission from the 
laser pulses. 

The coolant then flows through the seed laser and local oscillator power supplies, the azimuth 
drive motor, the momentum compensator motor, and back to the pump. Filters are provided 
before the flow enters the pumps, the pump check valves, the diverter valves, the pump bypass 
valve, and the flow dividers to prevent contamination from interfering with their operation. 

Figure 3-67 is a schematic of the LAWS coolant pump package. This package contains two 
redundant pumps. Only one pump runs at a time. Check valves prevent backflow through the 
non-operating pump. Each of the pumps is capable of meeting the stated life expectancy 
requirement. This provides a factor of 2 margin on pump life. 

The pumps chosen are existing space qualified pumps which have been used on the Space 
Shuttle Orbiter TCS for a number of years. 


3-76 

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LMSC-HSV TR F320789-II 



ComoonMW Motor 


Azimufri Drtva 
Motor 


Local Oaattator 
Powar Supply 2 


Local Oaciftafor 
Powar Supply 1 


Laaar Powar 

Supply 


Laaar Powar StppV 
(Othar Componama) 


Laaar Fan Motor 1 


| Laaar Fan Motor 2 | 

f 

H 

Vferafion 

taolabon 

1 nrv^ 

Saad Laaar 

Powar Supply 1 



Saad Laaar 

Powar Supply 2 


Compraaaor Flanga i 
| £x P* nd>r F 1 « f V» 1 
Compraaaor Flanga 2 
Expandar Flanga 2 
Saad Laaar 1 
f Saad Laaar 2 
) Local OaoHator 1 

P== 


Figure 3-66. LAWS Active TCS Schematic 


Temperature 

Sensor Delta P Backup 

Filter \ Switch Pump/Motor 


Pressure 


Quantity 

Sensor 


Flowmeter 



Accumulator 


Check 

Valves 


Primary 


Pump/ 

\ \ Motor 

FHter Temperature 
Sensor 


Pump 
Inlet 

Pressure 

Sensor 


Pump 


Temperature 

Sensor 

F3i2see-ao-o« 


Figure 3-67. LAWS Coolant Pump Package Schematic 


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LMSC-HSV TR F320789-II 


i awc^ ^ 16 Sh ° WS the cooIant temperatures which result as the active TCS cools each of the 
UWS components. Inlet and oude, ten«s ate shown a*l compared ,o allow" e al u fo 

8 : ^ ° Perati0 "' fl ° W i* 23* leX' 0 GPM 

, 6S TT toOUg dl componen,s “ "te order shown on the layout of Figure 

3-68 and then dumps the heat to the platform coldpla.es. It then returns to the original 5 °r 

emperature. Companson of these temperatures shows they are within the allowable ltoits. 


Table 3-16. Results, LAWS Active TCS Coolant Temperatures 


COMPONENT 

CRYOCOOLERS 

#1 COMP FIANCE 
#1 EXPANOER FLANGE 
n COMP FLANGE 
#2 EXPANOER FLANGE 
SEED LASER #1 
SEED LASER M2 
LOCAL OSCILLATOR #1 
LOCAL OSCILLATOR M2 
FILTER NO 1 
FLOW DIVIDER M a \ 

LASER GAS HT EX 1,2 
LASER POWER SUPPLY 
LASER FAN MOTOR #1 
LASER FAN MOTOR M2 
SEED LASER PR SUPP M 1 
SEED LASER PR SUPP #2 
LOCAL OSCIL POW S #1 
LOCAL OSCIL POW S M2 
A2MUTH DRIVE MOTOR 
MOM COMPENSATOR MOTOR 
FILTER NO 2 
PUMP 

PUMP MOTOR 
FILTER NO 3 
DIVERTER VALVE 
FILTER NO 3 
FLOW DIVIDER M2 


LAWS/EOS C PLATES 1,2 


SURVEY MOOE 


0 OOT 

T IN 

WATTS 

0EG C 

22 

15.00 

3 

15.10 

22 

15.12 

3 

15.22 

36.5 

15.23 

36.5 

15.40 

11 

15.57 

11 

15.62 

0 

15.67 

0 

15.67 

1060 

15.67 

519 

20.61 

20 

2 3.02 

20 

23.12 

13.5 

23.21 

13.5 

23.27 

4 

23.34 

4 

23.35 

30 

23.37 

15 

23.51 

0 

23.58 

0 

23.58 

66 

23.58 

0 

23.89 

0 

23.89 

0 

23.89 

0 

1910 

23.89 

-1910 

23.89 


DT 
DEG C 


T OUT T ALLOWABLE 
D EG C DEG C 


0. 10 

15.10 

0.01 

15.12 

0.10 

15.22 

0.01 

15.23 

0.17 

15.40 

0.17 

15.57 

0.05 

15.62 

0.05 

15.67 

0.00 

15.67 

0.00 

15.67 

4.93 

20.61 

2.42 

23.02 

0.09 

23.12 

0.09 

23.21 

0.06 

23.27 

0.06 

23.34 

0.02 

23.35 

0.02 

23.37 

0.14 

23.51 

0.07 

23.58 

0.00 

23.58 

0.00 

23.58 

0.31 

23.89 

0.00 

23.89 

0.00 

23.89 

0.00 

23.89 

0.00 

23.89 


20 

20 

20 

20 

20 

20 

20 

20 

40 

40 

27 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 


*8.89 15.00 


DESIGN MOOE 


Q OOT 
WATTS 

22 

3 

22 

3 

36.5 

36.5 
11 
11 

0 

0 

2300 

968 

20 

20 

13.5 
13.5 

4 

4 

30 

15 

0 

0 

66 

0 

0 

0 

0 

3599 


T IN 
DEG C 

15.00 

15.10 

15,12 

15.22 

15.23 
15.40 
15.57 
15.62 
15.67 
15.67 
15.67 
26.38 
30.88 
30.98 
31.07 
31.13 
31.20 

31.22 

31.23 
31.37 
31.44 
31.44 
31.44 
31.75 
31.75 
31.75 
31.75 


DT 
DEG C 

0.10 

0.01 

0.10 

0.01 

0.17 

0.17 

0.05 

0.05 

0.00 

0.00 

10.70 

4.51 

0.09 

0.09 

0.06 

0.06 

0.02 

0.02 

0.14 

0.07 

0.00 

0.00 

0.31 

0.00 

0.00 

0.00 

0.00 


T OUT T ALLOWABLE 
DEG C DEG C 


*3599 31.75 -16.75 


15.10 

15.12 

15.22 

15.23 
15.40 
15.57 
15.62 
15.67 
15.67 
15.67 
26.38 
30.88 
30.98 
31.07 
31.13 
31.20 

31.22 

31.23 
31.37 
31.44 
31.44 
31.44 
31.75 
31.75 
31.75 

31.75 

31.75 


15.00 


20 

20 

20 

20 

20 

20 

20 

20 

40 

40 

27 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 

40 


NOTES: 

1 Ethylene glycol/water (30/70) 

2. Mass flow rate, 238 kg/hr (534 ttvhr) - 1 gp m 

3. CP - 0.778 Cal/g-K, (0.778 Bfu/lbm ®R) 

4. Density - 1 .04 g/cm3 (65.4 tyft 3 ) 

5. All temperatures within allowable limits 


3-78 

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LMSC-HSV TR F320789-II 


& 


0 

© 

© 

© 

® 

® 

© 

© 


Oh SEQUENCE: 

::olers 

seed lasers 
:sc:..a'3P5 

L 3SER MEA t EX . 14 2 

L 

- l PSER PQWER supplies 

LASER RAN MOTORS 

SEE3 LASER POWER 5UPP . 

lOCAL OSCILLATOR POWER 5UPP • 

q7M[ T H DRIVE MOTOR 

MOMENTUM COMPENSATOR MOTOR 

p ijMP PACKAGE 

lAWS/EOS BACK TO BACK 
COLD PLATES 1 4 2 


Figure 3-68. Coolant Line Layout 



3 . 3 . 6 . 2. 2 Passive Thermal Control System Description 

On-orbit thermal control for the LAWS Instrument is achieved by a hybrid form of thermal 
control system. An active fluid loop as described above is used to transport the heat from high 
powered components such as the main laser, oscillator, seed laser, and azimuth drive. The heat is 
transferred through interfacing coldplates to be rejected to space via EOS central thermal bus 
radiators. Heat is also rejected passively by radiation from external surfaces of all components 
with an adequate field-of-view to space. Components are placed on the LAWS Platform such that, 
in combination with conventional passive thermal techniques augmented with electrical heaters, 
they are controlled effectively to within their allowable temperature limits during operational and 
non-operational (survival) modes. Passive thermal control is achieved by use of multilayer 
insulation (MLI), thermal coatings and tapes, thermal covers, and thermal isolation materials. The 
passive TCS is based on HST TCS design with a wide application of low a/e atomic oxygen 
resistant Ag FOSR (Flexible Optical Solar Reflector: Teflon with vapor deposited silver) designed 
fora 15 yr lifetime. 


3-79 

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LMSC-HSV TR F320789-II 


The planned baselined orbit for LAWS is a sun-synchronous orbit, with a 6:00 a.m. launch 
due south from Vandenberg AFB, attaining an attitude of 525 km with an orbit inclination of 
97.497 deg. This results in the orbit beta angle varying between 59 and 90 deg and an occultation 
period for some 100 days of the yearly cycle that begins when the beta angle becomes less than 68 
deg. 

The planned attitude for the LAWS Instrument is with its X-axis in the velocity vector, i.e., 
with the telescope leading. The combination of attitude and sun-synchronous orbit results in one 
side (-Y side) always toward the sun. Therefore, this configuration is used to position the receiver 
electronics, flight computer, power distribution unit, cryocooler controller, and Star Trackers on 
the cold side, facing deep space, since these components generate a significant amount of heat (see 
Table 3-15). These components, with the exception of the Star Trackers, are effectively controlled 
with a lightweight thermal cover with a combination of thermal coatings inside and outside. The 
Star Trackers require a thermal coating of SiOx on vapor deposited aluminized Kapton taped on 
them for maintaining design allowable temperatures. The IMU and detector bias and preamps are 
positioned on the +X side of the Platform. The pump package, which is inherently self-cooling, is 
positioned on the hot (-Y) side since it is part of the active fluid loop central bus heat rejection 
system. The pump package is protected from freezing in case of active system shutdown by being 
placed on the hot side of the Platform. The passively controlled components discussed are shown 
in Figure 3-69 with their designed thermal coatings/covers. 

The telescope is also passively controlled using A1 teflon tape. Surfaces along the optical 
path are painted black. Varying total absorbed orbital fluxes as the telescope resolves were 
considered in the TCS design and evaluation of required thermal coatings. This is depicted in 
Figure 3-70. A similar telescope assembly was also evaluated for the downsized 5 J laser. The 
primary mirror diameter is approximately half that for the 20 J laser. The mirror, made of Coming 
ULE material, was analyzed to predict temperature gradients along the surface. This was 
necessary for thermal stress and deformation evaluation and to study the effect on optical 
performance. 

The graphite epoxy base structure on which the laser is mounted, the graphite epoxy 
honeycomb structure for the optical bench, and the telescope mount are anticipated to be covered 
with MLI to reduce temperature gradient and structural distortion within these structures. A1 FOSR 
is applied over the environmental/thermal cover for the optical bench. Although a temperature 
gradient can be expected on the cover, it is a nonstructural part and is maintained cold with the low 
a/e coating so it can be used as a contamination collector for the optics mounted on the optical 
bench and under this cover. 


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LMSC-HSV TR F320789-II 


The laser power supply is mounted on the laser tank, which in turn is mounted on the base 
structure. Acoustical/electrical induced vibrations are isolated from the optical bench, which is 
mounted on kinematic mounts attached to the base. Active cooling is required for the high internal 
heating components such as the thyratron and dc-dc converter of the power supply. To reduce the 
active cooling load, these components are placed on a coldplate with their surfaces exposed to 
space. The relatively low internal heating PFN capacitors are controlled by using a small cover 
coated with aluminized teflon (see Figure 3-69). 

The passive TCS design philosophy has been to perform analysis for a hot design case using 
+3 sigma orbital environment, end-of-life optical properties, minimum altitude, minimum MLI 
effectiveness, and maximum component duty cycles for the range of (5 angles expected. Maximum 
component temperatures are maintained well below their upper operational allowable temperatures. 
Then the cold design case with -3 sigma fluxes, maximum altitude, beginning-of-life optical 
properties, maximum MLI effectiveness, and minimum component duty cycles is performed for 
the range of p angles to ensure that component responses are above their lower operational limit. 

By using a worst case combination of fluxes, optical properties, MLI effectiveness, duty 
cycles, and P angles, a good TCS design margin is provided. Heaters, when designed for the 
telescope, are sized with a 50 percent margin at minimum bus voltage to ensure adequate 
capability. MLI design will incorporate net spacers to obtain better performance. The number of 
layers and net spacers will be based on thermal-vat (TV) tests, and the blankets will be baked out 
separately or after installation to minimize contamination. TCS requirements are verified by 
analysis, component level TV tests, and LAWS systems TV tests. 

3. 3. 6. 3 Laws Thermal Control Subsystem Overview 

Figure 3-71 shows an overview of the LAWS TCS. Part A shows the schedule from 
January 1994 through launch in 2001. The first nine quarters are used for preliminary and detailed 
design after one quarter of finalizing requirements. We propose to start pump life testing at the 
very beginning of this schedule and continue throughout the design and qualification period. Two 
pumps are expected to be sufficient to meet the 5 to 7 year LAWS life requirement. However, 
design "scars" will be provided so that additional pumps can be added to the pump package if 
required as a result of these life tests. Up to four pumps can be easily used in the existing design 
package. 

Functional and development tests are planned at both the component and integrated levels. 
These will provide inputs directly to the design. Passive and active testing will be conducted 
separately at first, and then these systems will be combined for continued integrated testing. Both 
component level and integrated testing will be conducted on both the qualification and flight 
hardware. Thermal support will be provided for LAWS-to-bus integration and checkout and for 
bus-to-vehicle integration. 


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THERMAL CONTROL SUBSYSTEM DEVELOPMENT, TEST AND EVALUATION PLAN OVERVIEW 


1994 
/\PRR APDR 


1995 


1996 


Acdr 


Reqmts Def. & Allocation 

I 


Laser Gas Flowloop Analysis & Des ign 


Coolant Loop Analysis & Design 


Telescope & Mount Analysis & Design 

i ~~i 


Optical Bench Thermal/Alignm ent Analysis 


Telescope Mirror Analysis 

i i 


Passive Controlled Compon ents Analysis 

Pump Life Test (Qual) 


T 


Pump Lite Test (Development) 


1997 


Pump Functional Test 


C apacitor/Heat Exc hanger Test 


Inte grated R uid Loop Test 


Passive TV Test 


3 


Integrated Actrve/Passive 

c 


Qual Test Fab 


Component Level Qual Test 


TV TCS Complete System Test[ 


1998 


1999 




SS 


Ship 


Qual Test of all Systems 

I 


2001 

— 

Launch 


Right Unit to Bus Integration & Checkout 


Flight Unit Fab 


Fliqht Unit TV TCS Test 

i c 


Flight Component Test 


D 



Bus to Veh Integratic 



REQUIRED THERMAL CONTROL SUBSYSTEM EQUIPMENT 


COMPONENT 

SOURCE 

MATURITV7HERITAGE 

LIFE 

TESTING 

ENGINEERING 

MODEL 

QUAL 

UNIT 

FLIGHT 

UNIT 

SPARES 

Pump Package 

Supplier 1 

Space Qual/Space Lab 

1 

1 

1 

1 

1 

Pumps 

Supplier 1 

Space QuaUSpace Lab 

3 




1 

Cold Plates 

TBD 

Same as EOS 


2 

2 

2 


Diverter Valves/Controllers 

Supplier 1 

Space Qual/Shuttle 

1 

1 

1 

1 

1 

Heat Exchangers 

TBD 

Modified/Breadboard 


2 

2 

2 

2 

Ag FOSR 

Supplier 2 

Off the Shelf/HST 


TBD 

TBD 

TBD 


MLI 

Supplier 2 

Off the Shelf/HST 


TBD 

TBD 

TBD 


Heaters, Kapton 

Supplier 3 

Off the Shelf/HST 


TBD 

TBD 

TBD 






fOl.DOUT 





LMSC-HSV TR F320789-II 


|~C~| TCS REQUIREMENT IMPLEMENTATION/ VERIRCATIOn| 

KEY 

REQUIREMENTS 

IMPLEMENTATION 

VERIFICATION 

1 . Maintain laser gas temp 
at all PRF 

Convective heat exchanger 
within active cooling loop 

Analysis, test 

2. Five year life on active 
cooling system 

Redundant pumps 

Life test 

3. Maintain temp limits of 
components for all 
mission phases 

Active cooling and passive 
Aq FOSR outer surfaces, 
MLI 

Analysis, TVT 

4 Control of thermally 
sensitive optical bench, 
telescope & mirror/ 
supports 

Controlled by FOSR, heaters, 
ULE optics, gold coatings 

Analysis, TVT 

5. Minimize contamination 
of optics 

Optical bench thermal cover as 
contamination collector and 
spatial separation of fluid lines 
from optics 

Analysis, TVT 

6. Design for 5 year atomic 
oxygen environment 

Teflon Ag FOSR, Act = 0.012 
per year based on flight data 

Analysis, 
LDEF data 

7. Decouple optics from 
orbit environment 

Thermal covers, Ag FOSR outer 
surfaces, low a/e external, 
thermal isolators 

Analysis, TVT 

8. No single point failure 

Redundant pumps, valves 
heater systems 

Analysis 

9. Maintain hardware and 
components above 
survival temperatures 

Safe mode developed with 
heaters/thermostats to maintain 
component above lower survival 
limits 

Analysis 

10. Maintain struct temp 
gradients and changes to 
meet pointing require m"ts 

Thermal cover, Ag FOSR and 
heater system 

Analysis, TVT 




TCS TRADE STUDIES AND ANALYSIS 


1 . Position of passively controlled avionics components on the base. 

2. Compact convective heat exchanger versus back-to-back cold plates for 
EOS/LAWS active thermal control interface. 

3. Pumped loop versus heat pipe active TCS. 

4. Redundant loops versus single loop with redundant pumps for active TCS. 

5. Existing space qualified pump packages versus new development 
long life pumps. 

6. Passive versus active cooling of main laser power supply. 


m 


TCS RISK REDUCTION SUMMARY 


RISK 

1 . Five year pump life 


LEVEL RISK REDUCTION APPROACH 

High Life testing & redundant pumps 



Test Verified TCS 


G 


DESIGN MARGINS AND GROWTH 


1 . Using hot and cold design cases with 3o fluxes 


2. Five year valve life 

3. Five year life of other TCS compon. 

4. Contamination due to outgassing 

5. Contamination due to coolant leaks 

6. Contamination due to 
biological growth in coolant 


Med 

Low 

Low 

Low 

Low 


Cyclic testing 
Stable materials 

Material selection, bakeout & design 

Use of brazed joints and leak 
containment devices 
Sterilization system 


2. Range of equipment duty cycles and MLI/thermal coating performance 

3. Heaters sized 1 .5X required at minimum bus voltage 

4. Controlling to levels well within requirements 

5. 2.0 x pump life 

6. Redundant pump controller and power circuits 

7. Redundant heaters & heater thermostats 312SM . 11 


Figure 3-71 . Overview! Summary of the LAWS 


Thermal Control System 



3-83 f0LDGur FRAME 
LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 


Figure 3-71 (B) shows the major TCS components, the intended source, the 
maturity/heritage and quantities of each component required for life testing, engineering units, 
qualification units, flight units, and spares. The suppliers selected are all well qualified with a 
wealth of experience in their particular areas. The maturity/heritage for these components shows 
either already qualified or off-the-shelf availability. Heritage derives from the Shuttle, HST, and 
Space Lab or EOS. 

Figure 3-71 (C) identifies the steps or design features incorporated to implement the key 
requirements planned for verification of each requirement. Part D summarizes some of the trades 
and analyses either completed, planned, or on-going. 

Figure 3-71 (E) shows the TCS risk reduction summary. Pump-life risk is minimized by 
actual life testing in our labs to verify the pump performance. Design scars will be left in order to 
incorporate the number of pumps needed to meet the 5 to 7 year life required with a factor of 2 
margin. At present, it is felt that two redundant pumps meet this goal. If not, additional pumps 
will be added as required. 

Figure 3-71 (F) shows the logic to be used during the process of the LAWS program to 
produce a verified design/subsystem. Part G shows the design margin, again showing the planned 
factor of 2 on pump life. 

3.3.7 Electrical Power Subsystem 

3. 3. 7.1 Overview 

The block diagram of the LAWS power distribution system (PDS) is shown in Figure 3-72. 
The spacecraft's two 120 Vdc (GIIS- specified) power buses are labeled Platform +120 Vdc bus 1 
and Platform +120 Vdc bus 2 in Figure 3-72. The PDS derives two redundant 28 Vdc power 
buses from the spacecraft's two 120 Vdc power buses. Each of the two buses are capable of 
supplying all power required by the LAWS Instrument. Since both buses are active 
simultaneously, each bus supplies half of the LAWS power load. For clarity, the redundancy of 
individual components in the PDS is not shown. The PDS supplies 120 Vdc to the transmit laser 
and 28 Vdc to the other LAWS subsystems. Only power distribution to the transmit laser, 
computer, and receiver is shown. Power distribution to other LAWS subsystems is similar. 

In Figure 3-72, circuit breaker 1 and circuit breaker 2 protect the spacecraft 120 Vdc power 
bus from faults in the LAWS system. Circuit breakers 3 and 4 protect the PDS dc/dc converters 
from faults occurring in the individual LAWS subsystems. These circuit breakers are remotely 
resettable. If a circuit breaker trips, it can be reclosed by commands from the flight computer or 
spacecraft Commands and health and status monitors pertaining to the PDS are discussed in later 

sections. 


3-84 

LOCKHEED-HUNTSVILLE 



Command Command Command 


LMSC-HSV TR F320789-II 



3-85 

LOCKH EED-HU NTSVILLE 


Figure 3-72. LAWS PDS (Commands Indicated) 




LAWS 
+ 120V 


IMSC-HSV TR F320789-II 



Figure 3-72. LAWS PDS (Monitors Indicated) 




LMSC-HSV TR F320789-II 


Relay 1 and relay 2 disconnect the LAWS Instrument from the spacecraft s 120 Vdc bus. If 
the spacecraft 120 Vdc bus 1 is utilized, relay 1 is closed. If the spacecraft’s 120 Vdc bus 2 is 
used, relay 2 is closed. Relays 3 and 4 disconnect the converters from the +120 Vdc power buses. 
The dc/dc converters convert the 120 Vdc buses to two redundant 28 Vdc power buses. Relays 5 
and 6 disconnect the 28 Vdc power buses from the LAWS subsystems. All relays in the figure are 

the latching type. 

The spacecraft's 120 Vdc power buses are filtered by filters 1 and 2. In addition, the 28 Vdc 
power to the individual subsystems is filtered at the PDS output connectors. 

3.3.7. 2 Commands & Monitors 

Location of the commands is shown in Figure 3-72 (1 of 2). The command labeled 1 closes 
circuit breaker 1 if the breaker trips. Commands 3 and 4 open relay 1 and close relay 1, 
respectively. The commands are summarized in Table 3-17. 

Figure 3-72 (2 of 2) shows the monitors in the PDS. The monitors in the PDS are used to 
monitor the PDS status and isolate PDS faults. Monitors 2, 4, 14, 16, 20, and 22 are current 
monitors. All others monitors are voltage monitors. Monitors 1 and 5 are used to monitor the 
voltages at the points where they are located and to determine the status (open or closed) of circuit 

breaker 1. 

3. 3. 7. 3 Redundancy 

The individual circuit breakers and relays shown in Figure 3-72 represent four circuit 
breakers and four relays. Placing two relays in series protects against a short circuit failure. 
Placing two relays in parallel protects against an open circuit failure. For example if relay A does 
not close, the path can be closed by closing relays C and D. Thus, the configuration functions if 
any one relay becomes stuck closed or open. 

3. 3. 7. 4 Cabling 

Cabling for the LAWS system is shown in Figure 4-73 of DR-8. The cables are summarized 
in Table 4-19 of DR-8. Power cables from the PDS to the individual LAWS subsystems and data 
cables from the computer to the LAWS subsystems are included. 

3.3.7. 5 EMI/EMC 

The generation and control of electromagnetic radiation have been considered in the overall 
design of the LAWS system and in the design, fabrication, and testing of the laser breadboard. 
The generation of short (i.e., a few |is) energy pulses used to power the laser, if not properly 
isolated and shielded, provides a relatively high energy source of EMI. In designing the laser, we 
considered the control of this specific radiation source. Control of EMI is required to prevent 
contamination of laser cavity match electronic controls as well as overall LAWS Instrument and 

platform operations. 


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Table 3-17. PDS Commands 


Command 

Description 

1 

Recloses circuit breaker 1 

2 

Recloses circuit breaker 2 

3 

Opens relay 1 

4 

Closes relay 1 

5 

Opens relay 2 

6 

Closes relay 2 

7 

Opens relay 3 

8 

Closes relay 3 

9 

Opens relay 4 

10 

Closes relay 4 

11 

Recloses circuit breaker 3 

12 

Recloses circuit breaker 4 

13 

Opens relay 5 

14 

Closes relay 5 

15 

Opens relay 6 

16 

Closes relay 6 


3.3.7.6 Subsystem Summary 

Figure 3-73 provides a summary of the electrical power subsystem development. 


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MAJOR MILESTONES 

Design PDU 

PDU Procurement 

Fabricate PDU 
Engineering Unit 

Initial PDU Test 

Assemble LAWS 
Engineering Unit 

Test LAWS 
Engineering Unit 

LAWS Engineering Unit 
Support Operations 

Fabncate PDU 
Qualification Unit 

PDU Qualification Test 

Fabricate PDU 
Flight Unit 

Assemble LAWS 
Flight Unit 

LAWS Flight Unit Tests 

LAWS/Spacecraft 
Integration and Tests 

LAWS/Spacecraft/Vehicle 
Integration and Tests 

Launch 

Orbital Verification 

Design Software 

Implement and Test 
Software 

Maintain Software 

Write PDU test plans 
and procedures 

Design PDU Special 
Test Equipment (STE) 

Fabricate PDU STE 
Initial STE Test 



















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IRVIEW 1 


2000 

2001 



1A52 


1/01 



VO 

1 3/01 

ZZ3 

3/01 1 

0 
3/01 

6/01 





i 

NIT 

QUAL. UNIT 

FLIGHT UNIT 


1 

1 


28 

28 


F320789-02 


fcl REQUIREMENT IMPLEMENTATION/VERIFICATION I 

KEY REQUIREMENT 

IMPLEMENTATION 

VERIFICATION 

28 V ± TBD Vdc 

dc/dc Converter output 
voltage = TBD Vdc 

ATT 

TBD W of power 

dc/dc Converter output 
voltage = TBD W 

A/T 

Energy storage 

Batteries 

A/T 

Circuit protection 

Remotely resettable circuit 
breakers 

A/T 

Redundancy 

! 

Multiple parallel components in 
power path; two redundant 
isolated power busses 

A/T 



PLANNED TRADE STUDIES 


Distributed vs. centralized power distribution units 


E RISK SUMMARY | 

RISK ITEM 

RISK LEVEL 

RISK REDUCTION APPROACH 

PDU Failure 

Low 

Space qualified parts, redundancy, system testing 


|Tj VERIFICATION SUMMARY 

Acceptance, development, and verification testing per MIL-STD-1540 



SI ACCOMMODATION 


Standard power control and distribution interface 


[~H~] DESIGN MARGIN AND GROWTH 

Multiple power buses rated for 20% growth in loads 


Figure 3-73. Overview! Summary of the 
Electrical Power Subsystem 

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3.3.8 Command and Data Management Subsystem 

The C&DM subsystem baseline design is summarized as follows: 

* Hardware implementation 

- Flight computer 

- Communication links 

- Star Trackers 

- Inertial reference unit 

• Software modules 

- System management 

- Shot management 

- Communication management 

The flight computer, applying associated software, provides autonomous direction to the 
LAWS Instrument, controlling when the laser is to be fired to achieve measurements for selected 
wind components. The flight computer also receives and executes commands from the spacecraft 
via the BDU and exercises stored math models to compute the time associated with the telescope 
pointing angles for the laser pulses. Star Trackers (2) are located on the LAWS Instrument 
baseplate. Outputs from these Star Trackers to the LAWS Instrument are managed by the attitude 
and position determination elements of this subsystem. The command and data transceiver 
assembles and transfers data from the LAWS Instrument to the spacecraft for transmission via data 
relay satellites as depicted in Figure 3-74. 

All communications with the LAWS Instrument, to and from the spacecraft, and with the 
NASA control centers are directed through the LAWS C&DM subsystem via the BDU. The few 
interfaces not controlled by this subsystem are related to the LAWS spacecraft electrical, thermal, 
and mechanical interfaces. These interfaces, however, are monitored and reported by the health 
and status instrumentation sensors. 

The flight computer controls laser shot management firing commands, computes orbital 
Platform position location, collects telescope line-of-sight azimuth angle values for each laser shot, 
provides short time storage of wind data for transmission to the spacecraft data management 
system and formatting of data into CCSDS format, and performs other command and data 
management functions. 

Decisions for flight hardware and software (command, communication, and control of the 
system) based on requirements analysis and definition of the associated functions to be 
implemented and their interrelationships have been completed. The C&DM subsystem 
encompasses all functions associated with system control, data processing, and communication 
control. The system operation concept described above shows how this subsystem provides the 
control and communication management. This subsystem controls system operation and 
communicates data and commands (see Figure 3-74 for the location of these functions in the 
system functional hierarchy). 


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Figure 3-74. LAWS Functional Hierarchy 


3.3.8. 1 Requirements Analysis 

The following system requirements govern this subsystem design: 

1 . Provide continuous on-board operation 

2. Provide a control system 

3 . Employ shot management to conserve laser life and obtain optimal measurements of the 
wind vector components 

4. Monitor and report Instrument health and status 

5 . Report measured wind data in Level 0 format 

6. Append Platform ephemeris data, ground calibration data, and time to level 0 to create 
Level 1 A data 

7 . Perform calibration and alignment checks 

8 . Accept commands from BDU 

9. Provide safing control. 

Requirement 1 dictates that the LAWS operation be in real-time. Requirement 2 is an all 
encompassing requirement that says a separate and distinct control must be provided. Requirement 
3 is based on analysis conducted in Phase 1. Requirements 4 through 8 are derived from analysis 
of the "LAWS Data System Preliminary Requirements Review," dated 6 December 1989. The 
creation of level 1 A data is included as an option. Requirement 9 is applicable for all EOS Platform 
operations. 


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3. 3. 8. 2 Flight Software Definition 

Figure 3-75 identifies the functions to satisfy the operations of the LAWS system and meet 
the system requirements as identified. These functions have been classified as related to system 
management, shot management, and communication management. All system management 
functions are associated with control and implementation of the system operations. Shot 
management controls the laser pulse operation. Attitude/position determination is a function that 
supports shot management. It provides Instrument attitude and position data required to correctly 
fire the laser for a given beam location during a telescope scan. The timing of each laser pulse is 
derived from logic based determination of attitude, time position in space, and position in the scan. 
Communication management is concerned with communication between the LAWS Instrument and 
its host Platform and between the LAWS hardware components. All communications (i.e., 
commands received from or data transmitted to the ground) to and from the ground station are 
assumed to be handled by the host Platform. Therefore, the LAWS design communication 
interface between the Instrument and host Platform is through the BDU. 



• DETERMINE HEALTH 
AND STATUS 

• STORE DATA 


DETERMINE 

REFERENCE 

ATTITUDE 


• PROVIDE 

• PERFORM DATA PROCESSING PLATFORM 

EPHEMERIS 

• PERFORM POWER-UP 
SEQUENCE 


• COOE/TRANSMIT 
PROCESSED OATA 

• PERFORM SUBSYSTEM 
COMMUNICATION 
MANAGEMENT 


• PERFORM POWER-DOWN 
SEQUENCE 

• DETERMINE DATA QUALITY 


CONTROL CALIBRATION 
AND ALIGNMENTS 


• PERFORM SAFING OPERATION 


Figure 3-75. LAWS Flight Data Management Functional Hierarchy 


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3. 3. 8. 2.1 Control of/and Data Flow from Subsystems 

Both hardware and software are required to implement the functions identified in Figure 3- 
75. Figures 3-74 and 3-76 present the LAWS Instrument from a top level systems viewpoint and 
show the first level of allocations to the hardware components. Figure 3-76 also indicates the 
overall flow of signals through the Instrument. 

3.3. 8. 2. 2 Flight Computer Functions 

The flight computer implements all functions associated with system management, shot 
management, and communication management. The actual functional implementation is via the 
flight software identified in Figure 3-77. It is assumed the flight software will be a single 
configuration end item. As shown in Figure 3-77 , the flight software configuration end item 
consists of three subelements: the system management module, shot management module, and 
communication management module. Brief descriptions of these major modules and their 
submodules are given below. 



Figure 3-76. LAWS System Functional Flow Diagram 


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Figure 3-77. LAWS Software Tree 

System Management Module. The system management module provides the overall 
control for operation of the LAWS Instrument. This module is activated at system start-up and 
operates continuously until the Instrument is powered down. The clock provides the system time. 
Provisions are included to update the time from either the host Platform or the ground. The time 
accuracy is currently TBD. Data storage is provided to store system control parameters and 
Platform ephemeris, and to temporarily store ancillary data and processed data. 

System Executive. This module is the system real-time monitor and schedules the 
activation of other modules to execute the appropriate function. The system executive module 
accepts ground commands for Instrument status determination. A status message is generated for 
transmission to the ground receiving station. 

Power Management. This module has two functions: (1) initiate and manage the 
Instrument power-up sequence, and (2) initiate and manage the Instrument power-down sequence. 

The modules execute via a preprogrammed sequence for each mode (i.e., power up or power 
down). When power up is complete, a ready status flag is generated to indicate that the Instrument 


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is ready for operation. During the Instrument deployment operation, this module manages all 
operations required to deploy the LAWS Instrument (i.e., the telescope). It also manages locking 
the telescope in position for reboost. 

Sating Management. This module initiates and controls operations required to bring the 
LAWS to a condition compatible with the Platform requirements. 

Health and Status Management. This module maintains the current health and status of 
the LAWS Instrument. It executes in a background mode on a predefined schedule and polls 
hardware component status sensors to determine the operational status of each component. 

Data Formatting Management. This function generates two data strings: Level 0 data 
and Level 1A data. All data strings are encoded with the proper "hand shaking" for transmission. 
Level 0 data includes all Instrument data, which are the digitized data stream. Instrument 
performance data, and status information. The status information to the Level 0 data is a status 
indicator. The status indicator denotes routinely and upon command. 

Calibration and Alignment Management. This module initiates and controls 
calibration and alignment checks performed by various hardware elements. Lag angle 
compensation (tip-tilt) is performed under this function. 

Attitude/Position Determination. This module provides the current attitude and 
position. The reference attitude is obtained from the attitude and position determination system. 
The Platform ephemeris is obtained from the host Platform and stored for use. The telescope 
azimuth angle is obtained from the beam scanner assembly. 

Laser Pulse Manager. This module contains the logic to compute the timing sequence 
necessary to correctly generate a laser pulse at the appropriate times. 

3.3.8. 3 Other LAWS Software 

Figure 3-77 identifies three categories of software required for the LAWS Instrument: 
system support software, flight software, and support software. Flight software is discussed 
above. System support software includes GSE software. GSE software is any software that will 
be developed for the GSE. Support software includes any software required to support 
development of the flight, GSE, or support mission operations." Development support software is 
primarily the set of case tools used in design of the flight software. Operations support are data 
bases and software used in Instrument performance evaluation. System simulation is any software 
used in simulating the Instrument operations. Mission support software is any software developed 
by the prime contractor to support mission operations. Test support software includes all software 
used to checkout and verify the flight and GSE software. 


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3. 3. 8. 4 LAWS Computer Hardware 

The computer subsystem will be sized from a detailed analysis of the required computational, 
interface, and storage functions. Interface functions are delineated above. The computational 
functional requirement is sensitive to shot management and on board alignment functions. The 
computer memory requirement is dependent upon the above stated requirements to acquire and 
store data from such sources as the ephemeris and to reformat from Level 0 to 1A. (This function 
could be performed on the ground.) If the option selected by the LAWS team is to broadcast 
frequency spectra data direct from the Platform, the storage requirement could increase 
significantly. If the LAWS Instrument instead of the Platform is required to provide data storage 
for down link to EOS facilities, the data storage requirement increases from fractions of a second to 
several minutes. 

The selection of a candidate flight processor was driven not only by computational criteria but 
also by environmental data. The GIIS Section 11.2, "Right Environments," was used to provide 
baseline orbital environments. Previous programs have indicated that the requirements for 
radiation hardening, single event upset, and single event latchup can be the more important drivers 
in selecting a flight computer. For this reason, it was desirable to find a previously flown or soon 
to be flown computer. The unit understudy will fly on an MIT experiment before LAWS. A 
different generation was flown by LMSC on a Shuttle experiment. It is modular in form and can 
be configured to meet the LAWS requirements. It has the following features: 

• Modular based microcomputer 

• Incorporation of fault tolerant and fault recovery circuits 

• Radiation harness 

- Total dose 

> 10 6 rads SI 

> 10 14 neutrons/cm 2 

- Transient 

10 9 rads/s functional 

10 12 rads/s survival 

- SEU 

< 10* 10 errors/bit/day 

- Latchup immune 

• High reliability with "S" level parts. 

The operating temperature is -55 to 70 °C or - 175 to 70 °C with the use of thermofoil electric 
heaters. The design accommodates a vibration environment of 30 g’s rms in a vacuum of 
IE-8 torr. 

Special test equipment will be purchased from the manufacturer to support integration and 
test tasks. 


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3. 3. 8. 5 Command and Data Management Subsystem Summary 

Figure 3-78 provides a summary/overview of the C&DM subsystem development, including 
a top-level schedule, study and verification plans, risk identification, and related planning 
information. 

3.4 VERIFICATION (TEST & EVALUATION) 

3.4.1 Development Test Plans 

A complete test program for the NASA LAWS Instrument hardware and software includes 
plans for development, qualification, acceptance, and prelaunch testing. NASA scientists, assisted 
by members of the Science Team and Lockheed operations engineers, will also develop plans for 
tests to be conducted after the LAWS satellite is launched and operating in space. 

Development tests have been conducted during the breadboard laser development phase to aid 
in the selection of suitable materials, components, and assemblies for use in building the operating 
laser. Additional tests will be performed to validate the use of other materials, hardware 
components, and assemblies as new tests are performed during the Phase II Extension period. 

Development test plans will also be prepared to validate the design of components, 
assemblies, and software modules developed during the CD Phase. The purpose of these tests is 
to ensure that the hardware components and software modules produced by these designs meet the 
qualification test limits imposed by MIL-STD-1540B and perform the measurement functions 
required by the LAWS Instrument CEI Specification. 

3.4.2 Qualification Test Plans 

Test plans will be prepared and conducted in Lockheed owned and operated test facilities to 
demonstrate that the LAWS Instrument hardware and software fully meet the qualification test 
margins imposed by the requirements of the NASA approved LAWS Instrument CEI Specification 
and MIL-STD-1540B. The sequence of tests to be conducted is shown in Figure 3-79. 
Component qualification tests will be conducted to verify that components and assemblies, built in 
accordance with the approved LAWS Instrument design, can withstand the rigors of qualification 
tests as individual elements. 


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C&DM REQUIREMENTS/VERIFICATION 


MAJOR MILESTONES 


Design and Development 
Eng Support 


Software 

SW Reqmt Specification 
SW Architecture 
SW Development 
SW Test & Implementation 
SW Maintenance 


Engineering Unit 
Fab & Assembly 
Sub Sys I & T 
Sys I & T 
OPS Support 


Quantity Unit 
Fab & Assembly 
Sub Sys I & T 
Sys I & T 
Qual Test 


Flight Unit 
Fab Assembly 
Sub Sys I & T 
Sys I & T 
Bus Int 
Bus/ Veh Int 


Suport Launch/Orbit Fit 
Eng Operations 



REQUIRED COMMAND AND DATA MANAGEMENT SYSTEM EQUIPMEI 


Component 


Source 


Maturity/heritage 


Bread- 

boards* 


Development 

Units 


Flight processor 
Observatory bus interface unit 
Oscillator 

So Atlantic anomaly detector 
•Number of cards to be bread boarded 


Modified NASA/ESs 


Lockheed 


Modified HSI 
Modified/H EAO-2 












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REQUIREMENTS 
C IMPLEMENTATION/ 
VERIFICATION 

Requirement Implementation Verification 

Merge ENG and SCI data 

FP, BDU 

T, S 

Selectable fixed and 
programmable telemetry 
formats 

FP 

T 

Command decoding 
with error detection 

FP 

T 

Digital processing with 
100% margin 

FP 

A, 1 

Timing accurate to 10 0 
in 24 hr, time coding 
to within 100 nsec 
of UTC 

FP. BDU 
Oscillator 

A, S 

High energy protect 

MCU, OBS 
BDU, SAAD 

T, A 

* A = Analysis/simulation, 1 
S = Similarity, T = test 

= inspection, 




|~d 1 planned trade studies 


|~F ~1 VERIFICATION SUMMARY 


Development teste 

FP development test 

Purpose: establish functional FP operation 
Equip required: development unit, MCU 
development caids, test equipment 

Integrated avionics test 

Purpose: establish functional CDMS 
operation of the MCU with BIUs via the serial 
bus 

Equip required: tested MCU dev unit, a 
tested OBS BIU development unit, a tested 
SI BIU development unit, a non-flight-item 
oscillator, a vehicle systems simulator, and 
the MCU test equipment 
Environment: ambient 
SAAD development test 

Purpose: establish functional SAAD 
operation 

Equip required: SAAD dev unit and standard 
digital test equipment 
Environment: ambient 
Qualification/acceptance teats 
On units shown above 

Purpose: individual equipment qualification 
Equipment required: per unit as shown 
above 

Environment: ambient, thermal vacuum, 
thermal cycle, vibration, and EMI 


Structured vs. object oriented tech. 
ADA vs C language 


SCIENCE INSTRUMENT 
ACCOMMODATION PLAN 


light 

inlta 

Spares 

1 

1 

1 


2 

1 

2 0 


m 

RISK SUMMARY 

Risk Item 

Risk 

Level 

Risk Reduction Approach 

Command 

processing 

Low 

Utilize existing designs as 
applicable. Engineering 
Specialist (ES) to monitor 
process flow. 

TLM format 
and rates 

Low 

Utilize existing formats as 
available, provide hardwired 
contingency format. ES to 
monitor process flow. 

Computer 

processes 

Medium 

New S/W design - ES to 
evaluate HW/SW design. 

Subsystem 

integration 

Low 

Identified hardware/ software 
test facility. Critical path 
monitored by ES. Assure QA 
surveillance of parts used. 

Safe mode 
control 

Medium 

Minor modification to existing 
design. ES to monitor 
standard process flow. 


South Atlantic anomaly detector provides warning to 
instruments based on software selectable thresholds. 

Safe mode power control commands backup primary 
science instrument power switching system. 


Figure 3-78. 


1-77-1 DESIGN MARGINS 

L£U AND GROWTH 

• : i - ; 


Processor sized to ensure 100% margin in worst 
case: average processor margin is 240% 

Memory sized to ensure 100% margin In worst case: 
average margin provided is 260% 

Serial bus provide 320% margin at 1 MHz 
Bus design allows for additional BIUs 

BIU design allows command and telemetry to t>e 
added in discrete increments by adding appropriate 
cards 

Modular design allows the incorporation of new 
technologies 


Overview/Summary of the Command and Data 
Management Subsystem 

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Figure 3-79. Vehicle Qualification Tests 

After all of the components have been tested as shown in Figure 3-80, and have satisfactorily 
met the component qualification level test criteria, the components and assemblies will be 
assembled into a complete LAWS Instrument qualification test assembly. This assembly will be 
mechanically and optically aligned and functionally tested before formal qualification tests proceed. 
These functional tests will be repeated after scheduled test sequences are completed to verify that 
the test results show no degraded performance characteristics due to stresses imposed by the 
qualification tests. 


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ENCODER 

SIMULATOR 



THERMAL-VAC 
SHOCK & VIB 
EMI 

PERFORMANCE 


F320761-GP-02 


Figure 3-80. Component Qualification Tests 


343 Acceptance Test Plans 

Acceptance tests will be conducted on the LAWS Instrument flight hardware as shown tn 
Figure demonstrate dte flight-worthiness of dte Insmtmen, hardware and softwam. 

These tests will rigorously exercise all night software controlled «V“-" ces 

computatiOTts! as well as die electrical ^ ^^^^^ate^all^sn^Mnt 

space by the Instrument or the space platform wtH ^uPh ^ to s^u ^ 

Acceptance Tests are conduct. Records of Utese tests will be collected for companson wtdt da. 
collected when the Instrument is integrated with the Space Platform. 


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Figure 3-81 . Flight Unit Acceptance Tests 

3.4.4 Prelaunch Validation Test Plans 

Test data, resulting from tests conducted on both the Space Platform and the LAWS 
Instrument, will be closely examined for interconnection compatibility before the two are 
physically and electrically interconnected. The planned test sequence will progressively verify all 
physical clearances for the operational modes before powered drive sequences are initiated. 

Every operational command sequence will be exercised and data transfer links will be 
operated as they will be operated in space. Remotely commanded optical and mechanical alignment 
of the LAWS Instrument will be tested and calibrated. All health and status sensors and transducer 
circuits will be checked for validity and calibration. Software self-test sequences will be tested and 

verified. 

Launch stowage conditions will be checked, and the programmed sequence required to begin 
the operation of the Instrument in space, after the satellite orbit has been established, will be 
verified. The Instrument stowage and recovery operation required for the reboost operation will 
also be verified. 


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3 4.5 Documentation 

’ Test procedures will be prepared for all tests to be conducted to prepare the NASA LAWS 
Instrument for launch and operation in space. These tests will document component development, 
qualification, acceptance, prelaunch validation, and software tests. 

A Contaminadon Control Plan will be prepared to control the accumulation of paiticulant and 
non-volatile residue contamination on Instrument flight critical surfaces during fabrication 
assembly, test, integration, launch, and operation in space. A Safety Plan will be prepared 
assure the safe operation of the Instrument during all test operations. 

3 5 OPERATION REQUIREMENTS AND SCENARIOS 

The EOS Oils defines the mission phases and the EOS Platform modes of operation. The 
services provided to the experiments during each mode of operation are discussed in the EOS 
GIIS The primary operating constraints of the LAWS Instrument are driven by the services 
provided by the EOS Platform in each operating mode. Table 3-18 shows each mission phase, 
Le, platform supplied support (LAWS constraints), LAWS mode, LAWS activate , 
and LAWS support requirements. Two platform modes of operation are not included, boost- 
orbit adjust mode. I. is assumed .ha, during these phases the LAWS Instrument 

will enter into the survival mode. 

3 6 PERFORMANCE ANALYSIS 

There are two aspects of LAWS performance: signal-to-noise ratio (SNR) performance and 
scanning performance. 

SNR Performance. Figure 3-82 shows the SNR equation used to evaluate LAWS 
Instrument performance. The equation is narrowband SNR. The rationale for selection of 
telescope diLeter and laser pulse energy is presented in Section 3.2. Given these selections, die 
primary effect of system design on SNR performance is contained m the optical effi “ ency ' 

3-S3 shows how die elements which contribute to optic efficiency are built up into tire overeU optic 

efficiency. The contributors to transmit and receive optics efficiences are Cleary • 

derivation of the receiver-related elements (i.e., mixing efficiency and effective quantum efficiency) 
are discussed in Section 3.3.3. 

Scanning Performance. Scanning performance is controlled by the scanning 
requirements and by the limitations on power to the Instrumenti In the following 
fc, discussion is related to scanning performance during die portion of die year when die orbit 

not occultated. 


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Table 3-18. Operating Modes, Mission Phases, and Support Requirements (1 of 2) 


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<N 



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SNR » — L. n D \ j -2_L |j ^ b ^° rpti ° n - — [Efficiencies] 

° h v 4R 2 2 p [Turbulence Effects] 


Where: 


h u « Photon Energy* 2.18E - 20 J (for 9.11 jam) 
2 

- 2 * Aperture Area 

4 

J * Pulse Energy 

* Pulse Half Length (for Distributed Target) 

R * Range to Target 
p * Backscatter Coefficient (Given) 

Absorption Effects (Given) 

Turbulence Effects (Small Number at these Ranges) 


r\ * Combined Efficiencies 


For LAWS 


n = x] Transmit . ri Receiver . r\ Heterodyne 
Optics Optics Efficiency 


ti Effective 
Quantum 
Efficiency 


1 - 312500-32 

Reference: EB23/W. Jones, November 1990, Modification for Turbulence to 
D. Emmitt's October 1990 memo. 


Figure 3-82. Signal-to-Noise Ratio Equation Used to Evaluate LAWS Instrument Performance 


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Figure 3-83. Contributing Factors for Maximized Signal -to -Noise Ratio 






















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Survey Mode. The basic scanning mode is the survey mode, which requires three shot 
pairs per 100 km by 100 km grid square, with the grid aligned along the satellite ground track. In 
order to give a good distribution of shots throughout each grid square, the scan rate has been 
selected to give two scans per 100 km of ground track at the satellite altitude of 525 km. This 
results in a scan rate of 8.43 rpm or 7.12 s per scan. The shot schedule was then selected to give a 
uniform distribution of shots in the lateral direction. Figure 3-84 shows the shot pattern for the 
survey mode. The first scan for each grid square makes shots at 50, 117, 183, 250, 317, 383, 
450 and 517 km from the satellite ground track. The second scan for each grid square makes 
shots at lateral distances of 17, 83, 150, 217, 283, 350, 417, 483, and 55 km from the satellite 
ground track. The figure shows that each shot has a ground track of approximately 24 km for 
measurements from 20 km altitude to the earth’s surface. For each 100 km of ground track for the 
survey mode, the instrument takes 66 shots in 14.24 s, resulting in an average pulse repetition 
frequency (PRF) of 44.63 Hz. The maximum PRF is 7.27 Hz for forward and aft shots near the 

satellite ground track. 

Design Mode. The design mode requires a scan-average PRF of 10 Hz. For a uniform 
distribution of shots across the swath, a maximum PRF of 15.7 Hz would be required, requiring a 
maximum power of 5844 W. Figure 3-85 shows a limitation of 4800 W during the non- 
occultation period. Therefore, the power limitation of 4800 W imposes a non-uniform distribution 
of shots across the swath. The design mode has been selected to operate at a PRF of 12.57 Hz 
(maximum achievable with 4800 W of power) in the center portion of the scan and to provide 
uniform distribution of shots in the outer portion of the scan so that the overall average PRF is 10 
Hz This produces a shot density of 10.4 shots per grid at the center of the swath and a shot 
density of 13.1 shots per grid at the edges of the swath. Figure 3-86 shows the times and lateral 

distances at which shots are made for a half scan. 

Operation During Occultation Portion of the Year. Table 3-19 summarizes LAWS 
scanning performance constrained by available power. As discussed above, scanning performance 
during the days when occulation does not occur are shown. For the days in which occulation 
occurs, the survey mode is not contrained during darkness, and some design mode operation is 
possible. Figure 3-85 shows the maximum available power during daylight during days in which 
occultation occurs. During operation in light, the survey mode is constrained as shown in the 
table, and design mode operation is not possible. 


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LMSC-HSV TR F320789-II 



Figure 3-84. Survey Mode Shot Pattern Showing Forward Looking and Aft Looking Shots 


No stated limit on orbit average power 



Figure 3-85. Maximum Power Available for LAWS Experiment, Sun Synchronous Orbit, 0600 
Descending Node Crossing 


3-108 

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LMSC-HSV TR F320789-II 


600 
500 
400 

Lateral 

distance (km) 300 

200 
100 
0 

Time from forward look (sec) 

Figure 3-86. Shot Schedule for Design Mode with Instrument Power Limited to 4800 W 



3-109 

LOCKHEED-HUNTSVILLE 




LMSC-HSV TR F320789-II 


Table 3-19. LAWS Operational Characteristics Constrained by Available Power 


Non-occultation 


Survey 


Design 


Occultation (=100 days) 


Dark 


Light 


Survey 


Design 


Survey 


Design 


Maximum Instantaneous Power 
Available (watts) noted) 


4800 


4800 


6497 


6497 


2200 


Scan Average Power (watts) 


2146 


3924 


2146 


3924 


1920 


Maximum Instantaneous Power 
(watts) 

Heat Rejection Rate (watts) 

Orbit Average Heat Rejection 
with No Design Mode (watts) 


3022 


4800 
note (2) 


3022 


2013 

1722 


3702 


2013 

1722 


5844 

3702 


2200 
note (35 

1995 
I 703 


a 

m 

o 

a. 


Orbit Average Heat Rejection with 
1615 sec Design Mode (watts) 


2200 


2200 


Orbit Energy Balance 
T ime per orbit (nrs) 

Avail at LAWS (watt-nrs) 
(assumes 4800 watts in sun) 
Used for laws (watt-nrs) 
Avail for Storage (watt-hrs) 

Used from storage (watt-hrs) 
(amp hrs @ 26 8 volts) 
Battery storage cap (amp hrs) 
Charge Current (amps) 


I 59 


CD 
CD 
>* C 
-Q ro 


TD 

CD 

C 


<TJ 

_Q 

CD 

L_ 

CD 

c 

CD 


0 36 
0 

773 

0 

773 

28.8 

1 30 
-80 


<D 

4 — > 

o 

c 


2185 


1 23 
5904 

2362 
2504 
note (5) 


23 4 


Notes 

( 1 ) Maximum instantaneous power defined by 'bow r chart 

(2) For uniform distribution of shot across swath, max. power is 5844 watts. Max. available power 

imposes non-uniform distribution of snots across swath Average PRF of 10 Hz required for 
design mode is met 

(3) For uniform distribution of shots across swath, max. power is 3022 watts. Max available power 

imposes non-uniform distribution of shots across swath. Shot density is 3.9 shots/grid at center 
of swath and 6 snots/ grid at edges of swath 

(4) Orbit energy balance is defined by survey mode during occultation Some design mode operation is possible 

(5) Effic’ency of solar to battery to load process is 0 707 that of direct solar to load process 


3-110 

LOCKHEED-HUNTSVILLE . 



LMSC-HSV TR F320789-II 


Section 4 

WORK BREAKDOWN STRUCTURE 

Figure 4-1 provides the top-level work breakdown structure for the LAWS Instrument. A 
complete WBS depicting hardware and software to be developed and produced, services to be 
performed, and data to be submitted during the Phase C/D contract is provided in Volume III of 
this final report and separately as DR-5. These documents provide a WBS tree to the required 
levels, a WBS index, and a WBS dictionary. 


4-1 

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LMSC-HSV TR F320789-II 



4-2 

LOCKHEED- HUNTS VILLE 


Figure 4-1. WBS 1.0 LAWS Instrument 












Section 5 

ENVIRONMENTAL ANALYSIS 


LMSC-HSV TR F320789-II 


This section briefly identifies LAWS program suggested actions and alternatives, and their 
environmental effects. 

5.1 ACTIONS AND THEIR ALTERNATIVES 

The LAWS Instrument will be transported into low Earth orbit via an Atlas HAS. This LAWS 
Instrument will not be returned to Earth for any reason other than bum-up during its de-orbit to 

Earth. 

During the orbital mission, a small quantity of benign gases will occasionally be vented to 
exoatmosphere. These gases include helium, nitrogen, and carbon dioxide. No other viable 
alternatives to this program have been identified at this time. 

5.2 ENVIRONMENTAL IMPACT OF THE ACTIONS AND THEIR 
ALTERNATIVES 

For the issue of this document, possible areas of concern will be identified, and initial analysis 
performed. 

5.2.1 Prelaunch Phase 

During the manufacture, assembly, verification, transportation, and launch integration of the 
LAWS Instrument, care will be taken that no environmentally harmful substance is used or 
generated by LMSC or its subcontractors. The ony currently identified environmental concern is 
the possible health hazard related to the ground testing of the laser subsystem. This issue is 
resolved by proper protection and procedures. 

5.2.2 Launch Phase 

No environmental effects due to the LAWS mission payload, carried into orbit by an Atlas 
HAS launch vehicle, have been identified. A possible concern is the possibility of a crash or 
accident during launch through the atmosphere. It is possible that a small amount of benign gas 
material (CO 2 , helium, and nitrogen) contained in the laser subsystem is released. We have 
concluded that this is not an environmental issue. 

5.2.3 On-Orbit Operations Phase 

Other than (1) normal outgassing of the LAWS Instrument components in the low Earth orbit 
environment, and (2) occasional release of the benign gas material (CO 2 , helium, and nitrogen) 
used in the laser subsystem, the LAWS payload will appear to the environment as a passive, non- 
interacting object. 


5-1 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 


The use of a laser system raises the concern of eye and skin safety. The effects of lasers on 
human eyes and skin have been investigated extensively, and industrial standards for the safe use 
of lasers have been established. Maximum permissible energy (MPE) loading due to the 
operations of different types of lasers are documented in ANSIZ136.1-1986. 

An analysis of the space operation of the proposed LAWS laser indicates that there is no eye 
or skin safety concern. 

5.2.4 De-Orbit Reentry 

The LAWS Instrument is designed to stay on orbit for a five-year mission life. There is no 
plan to return the LAWS Instrument to Earth, so there would be no environmental impact for its 
Earth return. At the end of the mission, the LAWS Instrument together with the space platform 
will be de-orbited and reenter the Earth’s atmosphere. The aerodynamic heating of the reentry will 
break up the LAWS Instrument and bum the majority of its components. The only hardware that 
could pose a danger as reentry debris is the 1.67 m diameter primary mirror of the telescope. 
Considerations in designing and material selection will enhance the break-up and bum-up of the 
LAWS Instrument. Controlled reentry maneuvering will restrict the dispersion of reentry debris to 
an unpopulated region of the Earth. Analysis will be performed to investigate the reentry integrity 
of the LAWS Instrument. Mission analysis will also be performed to predict the scattering 
footprint of the reentry debris. Results of these analyses will be reported in an update to this 

document. 

The only other concern during reentry is the possible release into the atmosphere of a small 
column of benign gas material (C0 2 , helium, and nitrogen) contained in the laser subsystem. This 
release is not an environmental issue. 

5.3 AREAS OF CONTROVERSY 

At this time no areas of controversy have been identified. 

5.4 ISSUES REMAINING TO BE RESOLVED 

Two issues remain to be resolved: 

• Dispersion of reentry debris during the end-of-life de-orbit reentry of LAWS Instrument in 
the atmosphere 

• Eye and skin safety during ground testing of LAWS Instrument 

5.5 CONCLUSION 

At this time, LAWS is viewed as an environmentally passive object in low Earth orbit. As 
such, no major environmental concerns have been identified. During the design phase of the 
program, this issue will continue to be analyzed and this report updated for further reviews. 

No requirement for an environmental impact statement has been found at this time. 


5-2 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 


Section 6 

LASER BREADBOARD 


6.1 REQUIREMENTS 

A specification for the laser breadboard was developed jointly by LMSC and TDS using 
MSFC laser requirements integrated with LAWS system requirements as a baseline. This 
functional specification, included as an appendix to this report, includes such parameters as 
repetition rate, output energy, maximum energy in gain switch spike and tail, maximum frequency 
chirp over pulse length, minimal extraction efficiency, mode purity, beam quality, physical 
envelope, and other performance indicators. 

6.2 DESIGN 

The laser breadboard design was developed to duplicate the requirements of the LAWS laser 
flight unit with respect to the resonator, flow control, catalytic operation, cavity/pressure vessel 
size and output parameters: energy - 15 to 20 J/pulse, repetition rate - 20 Hz, lifetime, and pulse 
fidelity parameters. Since the pulse forming network (PFN) and energy storage devices offered 
less of a challenge to the state-of-the-art, commercial/laboratory grade devices were used to 
perform these functions with plans to upgrade these assemblies with space traceable, reduced 
volume hardware downstream in the program. 

Figures 6-1 and 6-2 depict schematics of the laser breadboard layout including control and 
diagnostic instrumentation. Instrumentation is depicted for monitoring alignment, frequency 
variation as a function of time, pulse power out as a function of time (pulse energy), mode purity, 
output pulse spatial profile, laser line, and other pertinent laser performance characteristics. 

Figure 6-3 presents the flow for checkout and integration of the pulse power subsystem from 
component test through the subsystem testing. Likewise, Figure 6-4 depicts the integration and 
testing of the flow loop/discharge/pulse power units into an assembly. In Figure 6-5, the 
components are first tested for the resonator and injection assemblies; next they are assembled and 
tested as subassemblies; finally they are brought together and tested as an integrated assembly. 
Table 6-1 lists components of the laboratory support equipment used to support the LAWS laser 
breadboard tests. 

Figure 6-6 presents end and side views of the LAWS laser breadboard. The power supply 
PFN is located beneath the optical bench which supports the resonator optics and test 
instrumentation. The power supply/PFN can be rolled under the bench or removed for 
diagnostics. Care was taken in designing the breadboard to control ground loops and associated 
electromagnetic interference. Figure 6-7 depicts the ground plane arrangement used for the 
breadboard. 

Figure 6-8 provides an overview of the test schedule for the LAWS laser breadboard. 
Fifteen months were expended from the go-ahead to TDS to develop the laser breadboard until first 
light was extracted from the laser. 


6-1 

LOCKHEED-HUNTSVILLE 



HeNe ALIGNMENT 
V LASER 


LMSC-HSV TR F320789-II 



6-2 

LOCKHEED-HUNTSVILLE 


Figure 6-1. Breadboard Test Configuration 
















LMSC-HSV TR F320789-II 



Figure 6-2. Breadboard Test Configuration/Resonator Layout 












LMSC-HSV TR F320789-II 


COMPONENT COMPONENT TEST 


SUBSYSTEM 

ASSEMBLY 


SUBSYSTEM TEST 



TO SUBSYSTFM TEST 
Willi TLOW LOOP 
UISCMATTGE 


Figure 6-3. Pulse Power System Integration and Checkout 


COMPONENT COMPONENT 
TEST 


SUBSYSTEM 

ASSEMBLY 


SUBSYSTEM TEST 



Figure 6-4. Integration and Checkout of Flow Loop/Discharge/Pulse Power Units 



6-4 

LOCKHEED-HUNTSVILLE 








LMSC-HSV TR F320789-II 


COMPONENT TEST 


SUBASSEMBLY TEST 




SUBSYSTEM TEST 
ASSEMBLY 



Figure 6-5. Integration Plan for Resonator/Injection Assemblies 
Table 6-1. Laboratory Support Equipment 


Laser Flow & Gas Control 


Output Diagnostics Alignment 


Gas Chromatography 
Mass Spectrometer 
Spectrum Analyzer 
Cooler 16 kW, E.l. # 39343 
Vacuum Pump & Valves 
Gas Bottles, Gages 
Vac/Press Pump System 

Data Acquisition 

Oscilloscope, Lecroy #9400 (2) 
Oscilloscope, Textronix #251051 
Oscilloscope, Textronix #2430 
Visicorder # 1858 
Visicorder # 1508B 
Pulse Generator, Datapulse 
Rack 19" x 22" x 5' 


Detectors 

Lens 

Beam Splitter (3) 
Fold Mirrors (3) 
Attenuators (2) 
Calorimeter 
Spatial Filter 
Spectrometer 
Detector, Waveform 
Beam Dump 
Kinematic Bases 
Optical Mounts 


HeNe Laser 
Alignment Telescope 
Dichroic 
Laser Mount 
Fold Mirror 
Optical Mounts 


Power Supp ly 

HVPS Ale# 1 53L (2) 
Current Transformer, 
Pearson #310&301(2) 
Current Transformer, 
Pearson #410 
H.V. Probe Tex #6015 
Cap Divider, Pearson 
#V305A 

Rack 19" x 22" x 5' 


6-5 

LOCKHEED-HUNTSVILLE 








LMSC-HSV TR F320789-II 


nil SF PlIVFR 


SUPPUR T framf:/ 
F.HI FNCI.MSIJRF 


Figure 6 - 6 . Integrated LAWS Laser Breadboard ( 1 of 2) 




0 
L ■ 


5 * 

SCALE 


m* 

i J 



Figure 6-6. Integrated LAWS Laser Breadboard (2 of 2) 


6-6 

LOCKHEED-HUNTSVILLE 





LMSC-HSV TR F320789-II 



DISCI IARGE 
CANISTFR 


RESONATOR 

TUNING 

ELECTRONICS 

BOLTED DOWN BARS FOR 
MESH TIE DOWN 

OPT ICAl HR ICO I 




I'l N GRI II INI' ( yj |vr N 
n A,( WIRE ( 

Ml SI I 


IUJS I IIP DA I A ACUII1SI I HIM AND 
CUM I Rill RACK GROUNDS, II 1 1 HR 
EQUIPMENT I GROUNDS, PRIMARY 
POWER GROUND, AND TARO I 
GROUND 


Figure 6-7. System Ground Plane 


CY 1991 

CY 1992 


IDDDDDDDDBODDDDDDDDDDBI 


BREADBOARD ACCEPTANCE 
TEST PUN (SDRL Q-1) 

BREADBOARD ACCEPTANCE 
TEST DOCUMENT REPORT 
(SDRL Q— 2) 

PREIONIZER LIFE TEST 

BREADBOARD CHECKOUT TESTS 

BREADBOARD 10.59 mM 
PERFORMANCE TESTS 

TRANSMITTER MODIFICATION 

BREADBOARD 9.11 M 
PERFORMANCE AND LIFE 
TESTS 

FINAL REPORT 


"I I I I I I I I I I 
V S? X7 

REVIEW REVIEW APPROVAL 


I I I I I 


REVIEW 


V V 


Figure 6-8. Test Plan Schedule 


6-7 

LOCKHEED-HUNTSVILLE 





LMSC-HSV TR F320789-II 


6.3 TEST RESULTS 

The LAWS breadboard laser was developed and tesled for LMSC. The primary 0 

the breadboard tests was to demonstrate acceptable performance parameters for the laser m "° 
"p ” CO, gas mixtures. The other objective was to conduct life tests to determtne 

component and system reliability. 

Tests were carried out at the Textron Defense Systems facility in Everett, Massachusetts, in 
accordance with the LAWS Laser Breadboard (1) Statement of Work, (2) Funcuonal Spectfica , 
and (3) Acceptance Test Plan. 

6.3.1 Test Sequence 

Acceptance tests were carried out in die sequence described in the Breadboard Acceptance 
Test Plan referenced above. In general, performance measurements were conducted first tut no 
“tores at the specif, ed 10 Hr prf and 10.6 pm wavelength. Thts was followed by Ufe test 

in the same at an accelerated prf of 20 Hz. 

The breadboard system was then modified for operation at 9.1 1 pm wavelength in isotopic 
CO, Zgen m mixture These tests were limited because of lack of availability of suffteten. 
isotopic gas and because the catalyst could not be labeled with oxygen- 18 pnor to the tests because 
of very long lead times for acquisition of the labeling gas. 

0 3 2 Test Facility 

The laser breadboard was assembled and tested in a designated area at the TDS-Everett 
faci.it? Spedaitre was taken to separate the assembly and checkout of the resonator and the 
to^lMp subsystems to avoid possible contamination of the optica, components. F, g ure 6-9 
shows a photograph of the system during the acceptance tests. 

6 ' 3 ?h" describes the results from dre acceptance res. ~ 23^rU 1^ 

and 2 July 1992. The procedures for conducting these tests were de 
Acceptance Test Plan referred to earlier and will not be repeated here. 

6. 3. 3.1 Performance Tests at 10.6 pm 

_r j in in 1 2c 1^0? mixtures to evaluate the laser 
Acceptance tests were performed at 10 Hz in u 2 mixiu 

transmitter. The test parameters measured are summarized below. 


0-8 

lockheed-huntsville 



LMSC-HSV TR F320789-II 


OKIG'NAL i-A'-if. 

BLACK A U'u WriiTE PHOTOGRAPH 



Figure 6-9. Loser Breodboord System 


6-9 

LOCKHEED-HUNTSVILLE 


LMSC-HSV TR F320789-II 


Average Pulse Energy 

Pulse energy was measured at the output of the laser with a Scientech 360401 laser power 
meter under two pressure conditions and several energy loadings as listed in Figure 6-10. The 
tests were conducted in the specified 3:1:1 (He:C02:N2) gas mixture. Also plotted in the figure are 
results of the TDS performance prediction code. Two different pump pulse times are depicted: 4.5 
(is for the lower energy levels, and 3.75 (is at the higher levels. Although energy loadings have 
not yet been extended to the design point, over 19 J/pulse are projected at the design point. 

Pulse Shape 

A photon-drag detector was used to obtain output pulse temporal characteristics as shown in 
Figure 6-11. The pulse width (FWHM) is 3.6 (is and the output decays to 10 percent of peak 
within 2 pulse widths, and to very nearly 0 in 3 pulse widths. (Note turn on at +1/5 major 
division.) This figure depicts several seconds of data at 10 Hz prf, indicating highly stable pulse- 
to-pulse operation. 

Intrapulse Chirp 

Intrapulse chirp was monitored by recording the heterodyne beat signal against an offset local 
oscillator and performing a Fast Fourier Transform (FFT) of the recorded data. Figures 6-12 and 
6-13 are typical examples of these measurements. These figures are a composite of three traces: 
(1) the lower trace is the beat signal in the time domain at 1 |is/horizontal division; (2) the fine 
grain central trace is the FFT at 5 MHz/horizontal division and 10 dB/vertical division; and (3) the 
more coarse central trace is the expanded FFT at 200 kHz/horizontal division and 5 dB/vertical 
division. 

In the example case in Figure 6-12, half (3 dB) of the pulse energy spectrum lies within 
±55 kHz. In the example case in Figure 6-13, half (3 dB) of the pulse energy spectrum lies within 
±82 kHz. Likewise, for each of the example cases, three-quarters of the pulse energy spectrum 
lies within ±120 kHz and ±127 kHz, respectively. These measurements have been made without 
an attempt to fully optimize the laser pulse forming network (PFN) impedance match. As the 
energy of the laser is increased toward 20 J/pulse, spectral frequency spread within the pulse is 
expected to increase. However, with adjustment of the PFN, the chirp at 15 to 20 J/pulse is 
expected to remain within specification. 

Energy Output Repeatability 

Repeatability of the output energy was measured under repped mode at 10 Hz. Under 
normal operating conditions, the laser was activated every morning. From a cold start, the energy 
meter immediately displays 7 J/pulse. After a 30 minute warm up period, the energy meter levels 
off at 6.6 J/pulse and remains at that level throughout the test period. During testing over several 
days with a single gas fill (500,000 pulses), no energy degradation was observed. 


6-10 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-1I 


20 


16 


a) 

Z3 

CL 


12 


>N 

O) 

a) 

c 

LU 

3 

Q. 

3 

O 


1 1 1 1 

Comparison of Breadboard Test Results 

and TDS Code Prediction 

i i i i 



^ 



f 

0 

at 

r 

utput ene 
different 

r 

r gy at 10.6 

specific er 

r 

pm 

lergy load 

ngs 



4 










1 

1 

1 



4. 

PL 

j j 

£ 

5 (is 

imp pulse 






1 

1 

1 

U 




rV 

3.75 ps 

— ni imn ni 

jlse 

_X Test 
O Code 

Data 

» Predictio 

a r\ 4 

1 

1 

1 B 

n 1 d 

1 

V 

i on 1 

ireadbos 
esign pc 

]/ 

QA 1 A 

ird 

>int 

lC\ 

X 



put 1 Ip pi 


Specific bnergy Loaaing in u/L-atm — 

Figure 6-10. Breadboard Test Results Compared to TDS Code Predictions 



6-11 

LOCKHEED-HUNTSVILLE 





T/dlv 1 jje 

Figure 6-13. Chirp Measurement from Fast-Fourier Transform, Example Measurement B 


6-12 

LOCKHEED-HUNTSVILLE 






LMSC-HSV TR F320789-II 


Beam Jitter 

Beam jitter was measured by relaying the far-field spot onto a strip of bum paper attached to 
the rim of a rotating cylinder with a fixed 0.5 mm wire cross-hair in front of the cylinder, as shown 
in Figure 6-14. The beam jitter angle was below our 80 jirad measurement resolution. 

Laser Efficiency Measurement 

The laser efficiency, defined as the ratio of near-field laser energy to the electrical energy 
stored in the PFN, was calculated at three different operating points using the laser output energy 
and PFN charging voltage measurements: 

• 5.8 percent @ 54 J/L-Atm 

• 6.4 percent @ 73 J/L-Atm 

• 7.4 percent @88 J/L-Atm*. 



Successive burn patterns on moving, thermally sensitive paper. 

F32078d-ll-06 


Figure 6-14. Beam Jitter from Pulse-to-Pulse: Much Less Than Our 80 jlm 
Measurement Resolution 


This 7.4 percent efficiency was increased to 10 percent on an ID program subsequent to the initial 
measurements above using the same hardware (see Appendix B). 


6-13 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 


The extraction efficiency, which is more strictly defined as the ratio of energy output to the 
energy deposited in the discharge, is higher than the displayed ratio. Further, it is obvious from 
the values above that the efficiency continues to increase as the energy loading approaches the 
design point (~ 120 J/l-atm). 

6. 3. 3. 2 Life Tests 

Life testing was carried out in oxygen- 16 based mixture to one million pulses. The 
parameters measured are summarized below. There were no major component failures. Most of 
the shots were accumulated at 10 Hz. 6,000 shots were accumulated at 20 Hz to demonstrate 
capability of the breadboard to operate at the accelerated prf. 

Average Power 

The following average power readings were obtained at several different prfs: 

• 52 W @ 10 Hz 

• 63 W @ 12.5 Hz 

• 73 W @ 15 Hz 

• 100 W @ 20 Hz. 

Figure 6-15 depicts pulse energy for a single pulse as monitored by the Scientech joule 
meter. 

Discharge Voltage and Current 

The voltage and current waveforms are shown in Figure 6-16, which is a representative shot 
taken at a PFN charge voltage of 22 kV at 320 Torr gas pressure. As can be seen from the voltage 
(upper trace), there is a mismatch between the PFN impedance and the discharge impedance 
resulting in post-pulse reflections. The match improves as the design point is approached. 
Improvement of this match will improve laser efficiency as well. 


5/11/92 



Figure 6-15. Single Pulse Energy Monitored by Scientech Joule Meter 


6-14 

LOCKH E ED-HU NTSVILLE 





LMSC-HSV TR F320789-II 


Figure 6-16. Typical Discharge Current and Voltage Waveforms 



6. 3. 3. 3 Performance at 9.11 urn 

The selection of wavelength for LAWS involved consideration of both atmospheric 
transmission and back scattering properties of aerosols in air. According to the LAWS Science 
Team, the recommended wavelength is 9.1 1 pm. Since the R-branch of the 00°l-02 0 mode of 
12 q 1802 has a relatively strong transition at a wave number of 1097.15 cm 1 (9.1145 pm), the 
ideal gas mixture for LAWS transmitter laser will contain isotopic 12 C 18 02 as the active molecule. 

Kinetic information for radiative transitions of the 00°1-(10°0, 02°0) CUD bands of isotopic 
species of * 2 C 18 C>2 was obtained and analyzed by different authors (references 1 through 3). 
However, previous gain and extraction measurements were mostly made under low-pressure 
continuous wave (cw) pumping conditions. Since wind sensing Doppler Lidar requires a pulsed 
coherent laser output, additional data regarding the collisional deactivation rate of upper laser level 
as well as temperature dependence of the rate constant were needed to construct and validate a 
reliable model to predict and optimize the performance of a candidate laser. 


Design/Validation Experiment 

A single-pulse, closed volume, UV-preionized, self-sustained discharge in the isotopic laser 
mixture was used to determine the kinetic characteristics of the gain media. The discharge test 
section had a 1.22 x 4.2 x 20 cm 3 volume and a gain length of 3 x 20 cm since three passes of the 
probe beam were made. A simple two-mirror stable resonator was built to study the laser energy 
extraction. A concave copper mirror with radius curvature of 16.8 m and a flat output coupler 
were used to construct the resonator, which produces a 1.2 x 1.2 cm 2 multimode square output 
beam. The following gases were used: He (Liquid Carbonics, 99.99 percent), N2 (Liqui 


6-15 

LOCKHEED-HUNTSVILLE 



LMSC-HSV TR F320789-II 


Carbonics, 99.998 percent), and 12 C 18 0 2 , the 18 0 2 isotopic content of which was better than 95 
percent (Isotec Inc.). Figure 6-1 7 shows the layout of the experimental setup. All test parameters 
used to determine the kinetic data are listed in Table 6-2. The measured decay rates of gain under 
various gas mixtures and pressures are shown in Figure 6-18. The unamplified probe signal Io 
was determined by chopping a grating tunable CW laser beam (MPB C0 2 Laser Model GN -802- 

GGS). The laser output power (TEMoo mode pattern) at 9.11 pm of the 2 C 0 2 line was 9 

The beam was 7.1 mm in diameter. The amplified probe signal I was detected by using a Cd-Hg- 
Te detector and recorded on a Tektronix 7104 oscilloscope. The small signal gain coefficient g 0 
was calculated by using the expression exp (g 0 L) = I/Io, where L is the effective length o e 
discharge region. The measured gain was found to vary within 10 percent from shot to shot. One 
gas fill lasted approximately 40 to 50 shots without significant change of output. The energy 
extraction was measured by a power meter (Gentec Joule meter) with two different output couplers 
(12 percent and 40 percent). The mixture composition and the pumping energy, which in turn 
determine the temperature of laser gas, affect the overall decay rate of gain. 

Applying multiple regression analysis (reference 4) to the measured decay rates, a set of de- 
activation rate constants is determined, given in Figure 6-19. 

A summary of (001) vibration relaxation rate constants is given in Table 6-3. Since no gas 
temperature information was reported in reference 3, a direct comparison with our measurements 
was made at room temperature (300 K). Good agreement was obtained for the rate constant of 
Kco2-C02, but significant discrepancy occurred at Kne-cm The known 0 2 relaxation rate 
constants for 12 c16o 2 gas mixture (reference 5) are also listed in Table 6-3 Both K N 2-C02 and 
Kro2-C02 for 12 C 18 0 2 have much higher deactivation rates than 12 C 16 0 2 . The new rate 
constants were subsequently incorporated into the kinetic code to enable prediction of performance 
of an isotopic 12C 18 0 2 laser. Figure 6-20 shows a comparison of code predictions with the 
experimental data for the performance of an isotopic 12 C 18 C>2 laser. More specifically. Figure 6- 
20 shows a comparison of code predictions with the experimental data for temporal variation oi 
gain. Comparison of energy extraction measurements with code predictions is shown in Figure 6- 
21. Good agreement is evident in both plots. 

Performance Tests of LAWS Breadboard 

The resonator was modified for 9.11 pm operation through insertion of a grating in the cavity 
and subsequent tuning for the 9.11 wavelength. The laser head was filled with a C0 2 ;N 2 :He 
mixture, with the C02 being the rare 12 C 18 0 2 . No preconditioning was performed with 0 2 as 
would be required for long term, full performance operation, because of the unavailability of 0 2 
due to the long lead time for the gas. (The isotopic preconditioning gas is scheduled for delivery 
over the next six months at 50 L per month.) More than 8 x I03 discharge shots were recorded 
with the mixture. The initial few shots were monitored with a spectrum analyzer at 9.21 pm, 
however, after grating adjustment, the 9.1 1 pm wavelength was achieved on the third shot and 
maintained through the tests. As these tests were limited in nature, additional testing is desirable to 
fully characterize the laser performance at 9. 1 1 pm when a full supply of both the C 0 2 and 
1*02 become available. Model results, validated experimentally as discussed earlier in this section, 
verify that 14.6 J at 9.11 pm are achievable with the current breadboard design with a 1:1:3 
mixture and 0.625 atmospheric pressure. 


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u.v. Hco wao 
CO 2 OISWAMC 



Figure 6-1 7. Schematic Diagram of Experimental Apparatus 


Table 6-2. Test Parameters 


• Small-signal gain and energy extraction measurements 


- Energy Loading 

- Gas Pressure 

- Gas Temperature 

- Gas Mixture 

He N 2 C0 2 

0 11 
0 2 1 

1 1 1 

2 1 1 

- Output Coupler 

- Gain Length 


40 J/L - 180 J/L 
150 TORR - 600 TORR 
300°K - 440°K 

He N 2 C0 2 

3 2 1 

3 1 1 

2 3 1 

3 3 1 

12%, 40% 

60 cm (Double) 


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Normalized Small Signal Gain (g 0 (+)/g 0 ) 


LMSC-HSV TR F320789-II 



Figure 6-18. Decay Rate of Small Signal Gain 


T(K«) 



0.12 0.13 0.14 0.15 

j -1/3 


Figure 6-19. Deactivation Rate Constant on Temperature 


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Table 6-3. Summary of (001 ) Vibrational Relaxation Rate Constants 


QAS 

K He * C0 2 

K N2 -C0 2 

Kc 02 ’ co 2 

TEMPEHATUHE 

(TORR' 1 S' 1 ) 

(TORR’ 1 S* 1 ) 

(TORR’ 1 S’ 1 ) 

340°K 

106 ±10 

384 ± 20 

1192 ±101 (MEASURED) 

390°K 

128 ±12 

649 ±50 

1052 ±120 (MEASURED) 

440°K 

170 ±15 

1147 ±100 

2380 ± 200 (MEASURED) 

3Q0°K 

80.5 

212.1 

815.6 (Interpolated) 

300°K 

54.8 

354.6 

773.7 (ST1) 

300°K 

85 

106 

350 ( 12 C 16 0 2 ) 


C02:N2:He 1:1:1 — 225 torr - 153 J/L - Atm 



0 2 4 6 8 10 12 14 16 18 20 


Time (mlcrosec) 

Figure 6-20. Temporal Variation of Gain: Comparison of Experimental Data to Code Prediction 

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Figure 6-21. Energy Extraction Data for I2 C I8 02 Mixture 


Figure 6-22 depicts a single mode (transverse and longitudinal) pulse at 9.1 1 pm. Injection 

seeding with the 9. 1 1 pm seed laser was used to maintain single mode operation for the 9. 1 1 pm 
tests. 

Figure 6-23 depicts the current pulse from the PFN and the heterodyne beat signal for these 
9.11 pm tests as the laser output is beat against the local oscillator. The 2.2 ps delay between 
initiation of the current pulse and the laser output is apparent in the figure. The same low chirp 
performance of the laser operating at 9. 1 1 is expected as was measured at 10.6 pm ( Figure 6-12). 
In additional tests the detector output must be digitized and analyzed (as depicted in Figure 6-13) to 
further validate the chirp characteristics in extended testing. 

Figure 6-24 depicts the current/voltage (I/V) pulse out of the PFN (into the laser). The 
ringing displayed in the figure is again indicative of a non-ideal impedance match between the laser 
and the PFN. Laser efficiency improvement is achievable with a better impedance match. 


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Figure 6-22. Single Mode (L&T) Pulse at 9. 11 [lm 



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7/1/91 



Figure 24. Current Voltage Pulse from Pulse Forming Network 


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References 

1 . Freed, L. E., Freed, C„ and O'Donnell, R. G., IEEG Journal of Quantum Electronics, QE- 
18, 1229 (1982). 

2. Starovoitov, V. S., Et. Al„ Journal of Quantum Spec. Radiat., Transfer 41, No. 2, 153 
(1989). 

3. Fisher, C. H„ Et. Al„ Final Report GL-TR-89-0292, Geo. Lab., Hanscom AFB, 
Massachusetts. 

4. Jeong, K. M„ Et. Al., Journal of Physical Chemistry, Vol. 93, No. 3, 1 145 (1989). 

5. Witteman, W. J., "The CO 2 Laser," Springer-Verlag (1986). 


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Section 7 

CONTAMINATION ANALYSIS 

7.1 SURFACE CONTAMINANTS PARAMETERS 

To analyze the effects of surface contaminants on the transmission of laser intensity through 
the optical elements, it is assumed that a loss factor due to surface contaminants, ai, can be 
specified for each optical element. Depending on the number of surfaces, n[, of the optical 
element, the efficiency of transmission of each element can be defined as 

T], = (1 - aft (7-1) 

The total efficiency of the optical train due to surface contaminants can thus be obtained as 

n= fi rii= ri o-ao^ 

i = i i=i . 

Here N equals the total number of optical elements. 

The LAWS Instrument has the following optical train. 

The transmitting optics has 7 elements including 6 mirrors and one doublet, giving a total of 8 
surfaces. The receiving optics has 15 elements, including 11 mirrors, three lenses, and one 
window, giving a total of 18 surfaces. A list of all optical elements and the approximate angles of 
incidence is given in Table 7-1 . 

By assuming a constant loss factor for all the optical elements, total transmission efficiency 
can be obtained. Table 7-2 gives the results for several assumed values of ai. It can be seen that 
in order to keep total efficiency above 90 percent, average loss factors cannot exceed 0.3 percent. 

Surface contaminants can be divided into particulates and molecular, and their effects on 
optical system performance can be treated separately. 

Under ideal situations, molecular deposition of surfaces can be assumed to be uniform. The 
effects resulting from this molecular deposition are changes in total transmissivity and reflectivity. 
Loss of reflectivity due to deposition of common spacecraft outgas sources has been measured by 
Woods, et.al. (AEDC-TR-87-8), and results are given in terms of complex index of reflections. 
The optical system performance can thus be calculated knowing the thickness of the molecular 
deposition. In real situations, however, the molecular deposition could be quite nonuniform. In 
this case, measurements are needed to obtain actual degradation in optical properties. We use 
either the transmissivity or the reflectivity degradation at the required wavelength as a measure of 
the contamination effects from molecular species. 


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Table 7-1. LAWS Optical Elements 


Element 

Type 

Surface 

Angle (deg) 

Transmitting 

Optics 



1 

Primary Mirror 

1 

90 

2 

Secondary Mirror 

1 

90 

3 

Doublet 

2 

90 

4 

Mirror 

1 

67.5 

5 

Mirror 

1 

45 

6 

Fixed Mirror 

1 

45 

7 

Mirror 

1 

45 

Receiver Optics 

1 

Primary Mirror 

1 

90 

2 

Secondary Mirror 

1 

90 

3 

Mirror 

1 

67.5 

4 

Mirror 

1 

45 

5 

ELI 

2 

90 

6 

Driven Mirror 

1 

45 

7 

Driven Mirror 

1 

45 

8 

Fixed Mirror 

1 

45 

9 

EL2 

2 

90 

10 

EL3 

2 

90 

11 

Mirror 

1 

l 

CD 

O 

12 

Mirror 

1 

~60 

13 

Mirror 

1 

~20 

14 

Mirror 

1 

~45 

15 

Window 

1 

90 


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Table 7-2. Total Transmission Efficiency for Several Loss Factors 


Loss Factor aj(%) 

Individual Efficiency 

Total Transmission 
Efficiency 

0.1 

0.999 

0.9743 

0.2 

0.998 

0.9493 

0.5 

0.995 

0.8778 

1.0 

0.99 

0.7700 

2.0 

0.98 

0.5914 

5.0 

0.95 

0.2635 

10.0 

0.90 

0.0646 


The effect due to particulate contaminants is expressed in terms of obscuration ratio (O.R.). 
This parameter defines the percent of actual area of the optical surface blocked by the particulates, 
and can be measured directly. The preferred method of measurements is the imaging method. 
Other methods which can be used include solvent wash and particle counting, tape lifting from 
fallout witness samples. 

The relationship between the O.R. and optical system performance degradation has been the 
subject of investigation. Dependence of transmissivity loss on the wavelength and particle size 
distribution needs to be established. Scattering effects may also be of importance. At present, we 
are only concerned with the loss of signal. 

7.2 CONTAMINATION BUDGETS 

To ensure the performance of the LAWS subsystems from excessive degradation due to 
contamination, contamination budgets will be used to guide the establishment of contamination 
control requirements. The contamination parameters identified in the previous section will be used, 
and each will be given a total not-to-exceed limit. An analysis of the flow of hardware from 
cleaning/assembly through integration/launch to the end of mission will be performed. By 
analyzing the activities of all mission phases, a contamination budget can be established. Using 
this budget as a guideline, contamination control requirements for the different phases of the 
program can be defined. With proper planning and control, the state of cleanliness of the system 
can thus be maintained. 

Experience from previous space flight indicates that the largest particulate contamination 
accumulation comes from acoustic testing and during launch phase of the mission. The largest 
contribution of molecular contaminants comes from thermal vacuum testing and during launch and 


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early phase of orbital operations. Hardware design and operations control will be used to reduce 
contamination buildup during those critical periods. 

Since it is unrealistic to try to maintain all optical elements at the same level of cleanliness, it is 
our intent to keep internal optics at a higher cleanliness level than the exposed optics. Thus we 
plan to address the contamination requirements for the primary and secondary mirrors differently 
than those for internal optics. Measurements of cleanliness of the exposed elements will be used as 
a verification of contamination control. 

Typical contamination budgets for molecular and particulate contaminants for the Hubble 
Space Telescope (HST) are given in Figure s 7-1 and 7-2. 

A plan to conduct measurements of surface contamination accumulation at different phases of 
the mission will be established. This plan shall include the method of measurement, the data type, 
the frequency of measurement, the analysis to be performed, and the pass-fail criteria. Direct 
measurement of the critical surface is the preferred method, supplemental with indirect 
measurement data from environmental monitoring. Contingency measures will be used if the 
measured contaminant level exceeds allocated budget. 

Tentative contamination budgets for particulate and molecular contaminants for LAWS 
primary and secondary mirrors are given in Tables 7-3 and 7-4. These budget allocations will be 
updated as more data from measurements and/or analysis become available. 


5 uj A 

ss 4 

Jo 3 
<n (j 

ST ID 

2 

£ < 

< * 

2 ° 

- QC 4 

CC IE 1 
^2 


PERIOD 1 EVEHT 

INCREMENT / CUMULATIVE^ 
WITH PM CLEANING 

PRIMARY MIRROR BEFORE CLEANING 

2.4/ 2.4 

SECONDARY MIRROR, AS CLEANED 

0.1 / 0.1 

PRIMARY MIRROR AS CLEANED 

0.7/ 0.8 

FALLOUT DURING OPERATIONS 

0.1 / 0.9 

TRANSPORT TO SUNNYVALE 

0.1 / 1.0 

FALLOUT h CHIMNEY EFFECT 


BEFORE ACOUSTIC TEST 

1.3/ 2.3 

ACOUSTIC TEST 

1.3/ 3.6 

FALLOUT & CHIMNEY EFFECT 

0.5 /4.1 

REWORK & STORAGE 

0.2/ 4.3 

TRANSPORT & PRELAUNCH OPERATIONS 

0.1 / 4.4 

^LAUNCH 

0.6/ 5.0 j 


LEGEND 

— - BUDGET ALLOCATION 
# ACTUAL MEASUREMENT 
*-• PREDICTED 

5.0 




»| 1 HX rH — 34 ' 4 



3.6 < 

w^FALLOUTA 

CHIMNEY EFFECT 

! 

PICTfUrl 

STORAC 

;e prelau 

NCH 

Jr_ 

* ^ 

> 

^ ACOUSTIC TE 

rr 




/ \ ,A 3 CLEANED 

1 A 

m 



1.5 

e 

— 

A ■ L HST ASSEMBLY 

(snip j 

'-OTA ASSEMBLY j 






1984 


1985 


1986 


1987 


1988 


1989 


1990 


Figure 7-1. Typical Particulate Contamination Budget Allocation 


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10 


8 


>£ 

x co 

< CO 

§2 
O m 
o> 6 

uj rz 

Q UJ 
2 -i 

< u- 

5S 4 

< £ 

IS 

£ cc . 

*2 2 


r PERIOD /EVENT 

INCREMENT / CUMULATIVE 

GROUND OPERATIONS. DANBURY 

TRANSPORT TO SUNNYVALE 

GROUND OPERATIONS. SUNNYVALE 

HST TV TEST 

REWORK & STORAGE 

TRANSPORT 4 PRELAUNCH OPERATIONS 

STS LAUNCH 

IN-ORBIT 4 MAINTENANCE 

0.1 / 0.1 
0.1 (02 
0.1 / 0.3 
4.3/ 4.6 
0.2/4 8 
0.1 / 4.9 
3.6/ 8.5 

1.5/10.0 j 


LEGEND 

— BUDGET ALLOCATION 

® ACTUAL MEASUREMENT 
BELOW DETECTION LIMIT 

PREDICTED 


4.6 


TV TEST 


^TRANSPORT 
~A PRELAUNCH 


-REWORK & STORAGE^ 


DETECTION LIMIT 



8 5 

0 


LAUNCH 


r GROUND OPS OANBURY r-GROUNOOPS 
\ ^TAA«PORT ^SUNNYVALE 


thaws pc Frr 




I 


1964 


lies" 


i 


O 0-3 


JSL 


JSL 


J*L. 


J&SL. 


1986 


1987 


1988 


1989 


1990 


Figure 7-2. Typical Molecular Contamination Budget Allocation 
Table 7-3. Tentative Particulate Contamination Budget 


Operation Phase 

% Obscuration 
increment/cum. 

Primary Mirror Cleaning 

0. 1/0.1 

Secondary Mirror Cleaning 

0.1/0. 2 

Fallout during Operations at Itek 

0. 1/0.3 

Transport to Huntsville 

0. 1/0.4 

Fallout Prior to Acoustic Test 

1.0/1. 4 

Acoustic Test 

1. 0/2.4 

Fallout Prior to Shipping 

0.2/2. 6 

Transport to Launch Site 

0.05/2.65 

Prelaunch Operations 

0.05/2.7 

Launch/Deployment 

0.7/3. 4 

Orbital Operations 

0. 1/3.5 

Total Particulate Budget 

3.5% 


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Table 7-4. Tentative Molecular Contamination Budget 


Operation Phase 

% Reflectivity Loss 
increment/cum. 

Ground Operations, Itek 

0. 1/0.1 

Transport to Huntsville 

0. 1/0.2 

Ground Operations Huntsville 

0.1/0. 3 

LAWS TV Test 

2. 0/2. 3 

Storage 

0. 1/2.4 

Transportation to Launch Site 

0.05/2.45 

Prelaunch Operations 

0.05/2.5 

Launch/Deployment 

1. 5/4.0 

On-orbit Operations 

1. 0/5.0 

Total Molecular Budget 

5% 


7.3 CONTAMINATION SOURCES AND DEGRADATION EFFECTS 


7.3.1 Particulate Contaminants 

Sensor performance degradation can be caused by limiting optical throughput, scattering of 
off-axis radiation due to particle clouds, and enhancement of mirror scattering reflectance (i.e., the 
bi-directional reflectance distribution function measurements) due to surface particulate 
contaminants. Major sources of particulate contamination are 

• Airborne particulates settling on hardware surfaces during manufacturing, assembly, and 
test operations 

• Paint overspray, insulation shreds, clothing fibers, and other human induced substances 

• Particles generated from launch vehicle and payload enclosure material and redistributed 
during ascent 

• Particles dispersed by opening of payload enclosure and deployment of appendages (solar 
arrays, radiators, antennas, etc.) 

• Redistribution of particles trapped in internal surfaces and in crevices of the instrument 

• Materials released on-orbit by space vehicle, including products from upper stage, reaction 
control system (RCS), attitude control system, and orbit transfer rocket firing. 

7.3.2 Molecular Contaminants 

Deposition of outgassed products on LAWS optical mirrors, optical sensors, and critical 
optical surfaces may cause performance degradation (e.g., reflectance change). The contaminant 
sources are 


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• Lubricants, leaks, and exposed organics from which volatile condensables may emanate 
and be transferred to critical surfaces 

• Volatile condensable materials in the environment to which contamination sensitive critical 
surfaces may be exposed 

• Orbital molecular cloud generated from space vehicle and payload operations 

• Molecular flux returning to critical surfaces due to collision with ambient species or among 
outgas molecules 

• Interaction of spacecraft material with space environment, such as atomic oxygen, UV, 
high energy particles, and space debris. 

7.3.3 Contamination Analysis 

Contamination studies are needed in support of the LAWS contamination control effort. 
These analyses identify effects due to various contamination sources which contribute to the 
contamination budget during various phases of the LAWS mission. These analyses shall include 
but not be limited to the following: 

• Studies to predict outgassing effects from LAWS materials on critical optics 

• Studies concerning the redistribution of particulates and their effect on primary mirror 
obscuration. 

The basis of the LAWS contamination control plan shall be derived from the LAWS contamination 
analyses and shall indirectly be responsible for the LAWS contamination control requirements. 

7.4 CONTAMINATION PREVENTION AND CONTAINMENT SCHEME 

To achieve the contamination control requirement and to ensure that the contamination budget 
allocation will not be exceeded, a series of activities will be initiated. A contamination control plan 
will be developed which identifies the necessary steps to follow during the various phases of the 
program. It shall include the requirements for manufacturing, cleaning, verification, monitoring, 
personnel training, and material selection. 

7.4.1 Design Considerations 

To maintain cleanliness of the optical elements after the initial cleaning and assembly, a 
contamination enclosure is used to protect the optical train from external environments. It is 
designed so that contamination accumulations on the optical trains are minimized during testing, 
launch, and on-orbit operations. With this reduced degradation of the majority of internal optical 
elements, it is possible to allocate higher contamination budgets for the external optics, mainly the 
telescope primary and the secondary mirrors, which are exposed to the elements. 

Partitions will be used to isolate internal optical elements from potential contamination 
sources, thus reducing direct depositions during on-orbit activities. Venting paths are designed to 
avoid transport of contaminants toward critical optical elements. Materials selection guidelines will 


7-7 

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be established and followed, and the optical bench and the environmental cover will be thermal 
vacuum baked out to reduce the amount of material outgas on orbit. 

Other design features may be implemented as needed as a result of tradeoff studies and 
sensitivity analysis which identifies major contributors to the contamination budget. The use of 
purge gas and the use of an aperture opening cover are among the designs to be studied. 

7.4.2 Personnel Training 

Since the main source of contamination during ground operations is through human beings, it 
is critical to reduce the generation and transfer of ground contaminants during manufacturing, 
testing, and integration. A program will be initiated to train personnel working on the LAWS 
program on the contamination control requirements. 

7.4.3 Operational Constraints/Guidelines 

As a result of contamination analysis, constraints shall be established for orbital operations to 
reduce the possibility of contaminating the critical LAWS external surfaces. As an example, the 
analysis of contaminant transport during reboost phases will be used to establish constraints and 
procedures during such operations. 

7.4.4 Contingency Measures 

Contamination levels for the critical surfaces will be monitored at scheduled intervals during 
ground operations. The monitored level of contaminants will be compared with the contamination 
budget. If the measurements indicate the possibility of exceeding budget, contingency measures 
will be initiated. Such corrective measures shall include the identification of contamination 
sources, the effect due to the contaminations, suggested corrective actions, and verification of the 
success of the corrective actions. A revised contamination budget shall be established taking into 
consideration the results of all these actions. 

7.5 CONTAMINATION ANALYSIS 

Once the LAWS Instrument is integrated with the spacecraft, installed in the launch vehicle 
and ready for launch, the chance for further contamination monitoring and cleaning diminishes. 
However, activities that follow will add contaminants to the ones already accumulated on the 
critical surfaces. Analyses are used to establish the estimated contamination budget during the 
launch/deployment, orbital verification, and on-orbit operations, including reboost phases of the 
mission. Table 7-5 lists the critical surfaces, their contamination concerns, and the transport 
mechanism involved. Corresponding analysis will be needed to obtain level of contamination 
accumulated on the critical surfaces. Some of the omission phase contamination concerns will be 
discussed in the following sections. 


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Table 7-5. LAWS Contamination Evaluations 


Critical Surface 

Contamination Concern 

Transport Mechanism 

Primary mirror 

Exposed to ambient environment 
Direct view of telescope interior 
Ground and launch phase particulates 
Stray light 

Molecular deposition 
Return flux 
Redistribution 
Particle cloud 

Secondary mirror 

Exposed to ambient environment 
Direct view of telescope interior 
and spacecraft 
High laser energy flux 

Direct deposition 
Impingement 

Star Trackers 

Exposed to ambient environment 
Susceptible to spacecraft contaminants 
Susceptible to re- boost contaminants 
Stray light 

Return flux 
Plume backflow 
Particulate deposition 
Particle cloud 

Laser windows and 
transmitting optics 

High energy flux 

Laser internal contaminants 

LAWS internal contaminants 

Diffusion transport 
Molecular deposition 
Particle redistribution 

Detectors and 
receiving optics 

Low signal level 

LAWS internal contaminants 

Diffusion transport 
Molecular deposition 
Particle redistribution 

Thermal control 
surfaces 

LAWS external sources 
Space environmental effects 

Molecular deposition 
Return flux 

Cryogenic Surface 

LAWS internal sources 
Cold surface 

Molecular deposition 
Diffusion transport 


312599- FW-01 


7.5.1 Launch Phase Contamination Concerns 

The most noticeable flight-phase contamination events during launch operations that need to be 
carefully reviewed/addressed are identified below. 

• Pre-Launch Standby 

Inclement weather during the pre-launch standby period can induce contaminant ingestion 
into the payload fairing (PLF) interior through the peripheral vents. The ingestion rate and 
quantity will depend upon the balance between the external wind environment (gust speed 
and direction) and the PLF internal purge or air-conditioning flow rate. The resulting 
LAWS subsystem degradation will be affected by the external air quality, i.e., the 
contaminant contents, as well as the contaminant distribution on spacecraft surfaces. The 
wind-ingestion analysis will aid in the establishment of additional contamination protection 
requirements during the pre-launch standby phase. 

• Launch/Ascent 


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The vibroacoustic level during lift-off can cause particle suspension from the PLF interior 
surfaces and subsequent redeposition on various sensitive thermal control and optical 
surfaces. Since the LAWS telescope will be installed on top of the spacecraft with aperture 
opening pointing up in a launch/ascent configuration ( Figure 7-3), that vibroacoustically 
induced particulate contamination will have to be investigated. A parametric analysis, 
correlating surface particle deposition and surface area obscuration increase with initial PLF 
cleanliness level, is needed to establish PLF cleanliness and contamination control 
requirements. 

• Booster Motor Staging 

Of primary concern during booster separation is the upper staging motor plume which may 
recirculate over the core vehicle and enter into the PLF interior through various vent ports. 
Contaminant distribution on critical surfaces will occur due to internal flow 
diffusion/convection. However, for an Atlas IIAS launch vehicle, this shall not be a 
problem, since the Castor IVA booster motors are located far below the vent ports, and no 
rocket motors are used for separation. 

• Stage Separations 

Depending on the launch vehicle used, the stage separations may contain possible 
contamination events. The first is the retro-rocket firing, which could cause plume 
impingement, especially if particulate products are involved. This plume impingement 
phenomenon is affected by the separation trajectory (tipoff rate, misalignment effect, 
misfiring occurrence) and the firing duration. Secondly, the separation charge operation 
during stage separation will generate a particulate debris cloud. Inter-particle collisions and 
the aerodynamic drag of the debris particles could cause some debris particles to reach the 
spacecraft surface. 

Inasmuch as the present contamination analysis encompasses all events from PLF installation 
through the collision/contamination avoidance maneuver (CCAM), the following upper stage 
spacecraft integration sources, independent of launch vehicle operations, need to be addressed. 

• Propellant venting constraints have been imposed on post upper stage spacecraft separation 
maneuver operations (one of them being a vent inhibition distance of 500 ft) so that the 
spacecraft will not be subject to impingement by vented propellant gases. From the 
spacecraft molecular contamination view point, the main engine propellant vent problem 
may seem trivial, depending on whether the propellant gas is condensable on any 
noncryogenic spacecraft surfaces. On the other hand, venting of the hydrazine 
monopropellant (most likely in liquid form) for any upper stage RCS could cause 
condensation because of trace contaminants in the propellant. 

• Aside from the propellant venting concern voiced in the preceding paragraph, upper stage 
RCS firing and the attendant plume impingement or backflux during CCAM could cause 
spacecraft contamination, because trace contaminants in the propellant and from the catalyst 
bed could survive the chamber combustion environment and be present in the exhaust 
plume flowfield. Experience with the CCAM problem for the Shuttle launch systems may 
be used for assessment 


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LMSC-HSVTR F320789-1I 


Approach: 

• Estimate PLF interior GRMS based on 
available vibroacoustic analysis results. 

• Determine particle resuspension quantities 
from data. 

• Predict particle fallout on payload surfaces. 


31 2599 -FW -04 



Figure 7-3. Vibroacoustically Induced Particle Redistribution 

Other flight-phase contamination sources/events that warrant evaluation include possible 
contaminant ingestion due to reverse venting during terminal shock traversal, nonmetallic material 
outgassing during ascent flight, and debris dispersion during upper stage spacecraft separation. 
Although all the major launch phase contamination issues have been identified, pertinent data on 
source characteristics and certain flight operational details have not been completely acquired. 
Therefore, the scope of the study is not fully comprehensive. Future analysis updates shall be 
performed when up-to-date contamination source data become available. 

7.5.2 Orbital Operations Contamination Concerns 

The contamination concerns for this phase of the LAWS program include material outgassing 
during the early phase of the mission; particulate generation during the appendages deployment and 
spacecraft checkout phase; plume backflow from the orbital reboosts engine firings; and various 
on-orbit contamination sources due to operational maneuvering of the space vehicle. 

The contamination control approach for this orbital operations phase is to use preventive 
measures and constraints. Design features of the LAWS Instrument include the use of an 
environmental cover to protect the internal optics; the use of compartmentalizauon to isolate optics 


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from potential contamination source from ancillary equipments; and the use of venting path design 
to reduce the possibility of contamination deposition during pressure transients. Location of 
external critical surfaces will be chosen to minimize the impact from the external contamination 
sources. 

Contamination analysis will be performed to study the impact of various design options. 
Operational constraints will be established to reduce contamination impact during periods of high 
rate of contamination generation. Such measures as pointing the telescope away from 
contamination sources, or turning on the purge gas system, will be used to ensure that the end-of- 
life contamination budget will not be exceeded. 

A mathematical model for external contamination analysis during orbital operations has been 
established. Figure 7-4 depicts the LAWS Instrument external surfaces to be used in on orbit 
contamination transport analysis. This model will be updated when the details of the spacecraft 
configuration are made available. For overall system contamination control, ground operational 
events, such as particle fallout at various facilities and air conditioning (or purge air) flow 
recirculation (if the air cleanliness is substandard), should also be considered in contaminant 
buildup estimates and contamination control procedures development. 



Figure 7-4. LAWS Contamination Math Model 


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LMSC-HSV TR F320789-II 


The key analytical tools (computer codes) for performing the LAWS launch-phase and orbital 
phase contamination analysis are listed in Table 7-6. 

Similar to the situation for the launch phase analysis, pertinent data for the orbital operation 
phase analysis have not been completely established. Therefore, only preliminary study can be 
performed at this time. Future analysis update shall be performed when design features and source 
characterization data are made available. 


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Table 7-6. Analytical Tools for Contamination Analysis 


LMSC-HSV TR F320789-II 



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and redistribute by vibroacoustic excitation 



LMSC-HSV TR F320789-II 


Appendix A 

Functional Specification 
LAWS Laser Breadboard 


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LMSC-HSV TR F320789-II 


LMSC-HSV SPEC F3 12362 
Rev. B 
1 April 1991 


FUNCTIONAL SPECIFICATION 

LASER ATMOSPHERIC WIND 
SOUNDER (LAWS) BREADBOARD 



4800 Bradford Blvd, HunlaviM, AL 35807 


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LMSC-HSV TR F320789-II 


LMSC-HSV SPEC F312362 
Rev. B 

1 April 1991 


LAWS LASER BREADBOARD 
functional specification 


; / l 


1 




S.C. Kj*s«as y -J 

Laser 's8AW orisibl * e 
Eouipment Engineer 


APPROVED: 




D. JwWilson 

Deputy Program Manager/ 
Chief Engineer 


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LMSC-HSV TR F320789-II 


LMSC-HSV SPEC F312362 
Rev. B 
1 April 1991 


CONTENTS 


Section Page 

1 SCOPE 1 


2 APPLICABLE DOCUMENTS 

2.1 Government Documents 

2.2 LMSC Documents 

2.3 Subcontractor Documents 


1 

1 

1 

1 


3 

3.1 

3.1.1 

3.1.2 

3.2 

3.2.1 

3.2.2 

3.2.3 

3.2.4 

3.2.5 
3.3 

3.3.1 

3.3.2 

3.3.3 

3.3.4 

3.4 

3.5 

3.5.1 

3.5.2 


REQUIREMENTS 

Laser Breadboard Definition 

Breadboard Diagram 

Interface Definition 

Characteristics 

Performance 

Physical 

Maintainability 

Environmental Conditions 

Transportability 

Design and Construction 

Materials, Processes, and Parts 

Electromagnetic Radiation 

Nameplates 

Safety 

Documentation 

Furnished Component Characteristics 

LMSC Furnished Injection Laser 

Government Furnished Catalyst Characteristics 


1 

3 

3 

3 

6 

7 

7 
3 

8 
8 
8 
8 
3 
9 
9 
9 

10 


APPENDIX 

10 Phase I Selected Design Specifications 11 

Figures 

1 Laser Breadboard Block Diagram 2 

2 Laser Resonator Configuration 12 


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LMSC-HSV SPEC F312362 
Rev. B 
1 April 1991 


1. SCOPE 

This Laser breadboard functional specification establishes the performance, 
design, development, and verification requirements for the LAWS laser 
breadboard that will be used to test parts of the LAWS transmitter. 

2. APPLICABLE DOCUMENTS 

The following documents form a part of this specification to the extent 
specified herein. 

2 . 1 Government Documents 
Phase I Final Report. 

2 . 2 LMSC Documents 

LMSC/HSV SOW F312354 - LAWS Laser Breadboard SOW, January 1991, and the 
Rev. A applicable LMSC documents cited therein (SOW Para. 2.2) 

2 . 3 Subcontractor Documents 


None . 


3. REQUIREMENTS 

3.1 Laser Breadboard Definition . The LAWS laser breadboard is a frequency 
stable pulsed C0 2 laser system that will be used to demonstrate critical 
parts of the LAWS transmitter. 

3 . 1.1 Breadboard Diagram * The LAWS laser breadboard shall consist of the 
following systems, identified in Figure 1. 


1 

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LMSC-HSV TR F320789-II 


LMSC-HSV SPEC F312362 
Rev. B 
1 April 1991 



Figure 1. Laser Breadboard Block Diagram 


2.1.1. 1 Active Optical Control System . The optical coherence system shall be 
comprised of the following breadboard subsystems: 

a. Resonator 

b. Injection Laser 

c. Cavity Matching Electronics 

d. Beam Diagnostics 

e. Local Oscillator 

3. 1.1. 2 Pulsed Discharge and Gain Control System . The pulsed discharge and 
gain control system shall be comprised of the following breadboard subsystems: 

a. Discharge 

b. Pulse Power and Pulse Forming Network (PFN) 

c. Power Supply 

d. Instrumentation and Laser Controls. 


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LMSC-HSVTR F320789-II 


LMSC-HSV SPEC F312362 
Rev. B 
1 April 1991 

nl! r ~^.. Control System . The U»r Uo» apd gas control 

s ,, ste „ shall b. comprised of the follows breadboard subsystems: 

a. Flow Loop and Gas Supply 

b. Catalyst 

Ca Gas Contamination Monitor 
d. Cooling System. 


3.1.2 Interfac e Definition 


3. 1.2.1 Power Source Interfa ce, 
power sources of 220 ±10 


The LAWS laser breadboard shall operate with 


3. 1.2. 2 coolant In terface. T3D 
3 1.2.3 T.idar Interface., TBD 


3.2 characteristics. • This parasraph specifies 
User breadboard. Specifications of the Phase 
are appended in Section 10 for reference. 


the characteristics of the 
I selected design configuration 


3.2.1 Pprf ormance 

, 2 i.l asrffla,. the UWS U..r br.adboard shall op.rat. as sp.clflad b.reid 
with an overall sy.t.m warm-up time rot to exceed IS min. 

3 2 12 a m la muim . uus us " 6 " a<lh ° ard shiU ‘ 

discharge afUeiane, ceneietent -1th a LAWS system laser final d.stgb -a 
plug efficiency of not less than 5 percent. 

_ , rh- LAUS laser breadboard shall have a laser beam 

3 . 2. 1.3 gp-r^v oar Pulse. The LAWS laser 

output energy per pulse of not less than 15 J (goal: 20 J>. 


3 

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LMSC-HSV SPEC F312362 
Rev. B 
1 April 1991 

3. 2. 1.4 Puisewidth . The LAWS Laser breadboard Laser beam output puisewidth 
CFVHM) shall be within the range of 2 to 3 ys. 

3. 2. 1.5 PuLse Shape . The LAWS Laser breadboard Laser beam output pulse shape 

shall have Less than 10 percent of the energy in the gain switched spike and 

less than 20 percent of the energy in the tail. The tail shall be down not 

less than 20 dB from the main pulse intensity after two pulse widths. 

3. 2. 1.6 Laser Beam Mode . The LAWS laser breadboard laser shall have not Less 
than 95 percent of the output beam energy in a single longitudinal and single 
transverse mode. 

3.2.1. 7 Wave Length . With a puLse Laser working gas mixture containing 

I2 C 16 0.. the LAWS laser breadboard laser output beam shall have a wave- 

length of 10.59 ±0.01 urn. 

3. 2. 1.8 Wavelength with Isotone . With a pulse laser working gas mixture 

12 13 

containing the isotope 0 0^, the LAWS laser breadboard laser output 

beam shall have a waveLength of 9.11 +0.01 um. 

3. 2. 1.9 Chir? . The LAWS laser breadboard Laser output beam shall have Less 
than 200 kHz chirp. 

3.2.1.10 Beam Quality . The LAWS laser breadboard laser output beam quality 
ratio shall be less than 1.2 (goal: 1.1) relative to a plane wave of the same 
dimensions. Beam quality is defined as 


BQ = exp 



4 


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LMSC-HSV SPEC F312362 
Rev. B 
1 April 1991 


where 


OPD * rms optical path difference across the laser beam 
\ = wavelength. 

3.2.1.11 Beam Dimensions . The Laser breadboard output beam dimensions shall 
be square. 

3.2.1.12 Beam Polarization . The laser breadboard laser output beam 
polarization shall be not less than 95 percent linear. 

3.2.1.13 Beam Jitter . The Laser breadboard laser output beam jitter shall be 
less than 100 yrad (goal: 25 yrad) . 

3.2.1.14 Beam Energy Stability . The laser breadboard pulse-to-pulse laser 
output beam energy shall not fluctuate more than 10 percent. 

3.2.1.15 Divergence . The laser breadboard laser output beam divergence shall 
be less than 1.2 times the diffraction limit. 

3.2.1.16 Pulse Rate Frequency . In performance test mode, the j.aser 
breadboard laser beam output pulse rate frequency (PRF) shall be variable from 
1 Hz to not less than 10 Hz. 

3.2.1.17 Pulse on rnmman d Delay. Reserved. 

3.2.1.18 Tntracavitv Beam Mode: The laser breadboard resonator intracavity 

beam shall have not less than 90 percent of its energy in the lowest order 
cavity mode. 

3.2.1.19 Life Test Pulse Rate Frequency: The nominal laser breadboard laser 

beam PRF shall be designed with a goal of 20 Hz at an energy per pulse of 20 J 


5 


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LMSC-HSV SPEC F312362 
Rev. B 
1 April 1991 

while operating in a life test mode. In the life test mode, the LAWS laser 
breadboard need not meet performance specifications of paras. 3. 2. 1.1 through 
3.2.1.18 herein. 

3.2.1.20 Life . The laser breadboard shall have as a goal an operational life 
with maintenance of not less than 3 x 10 3 shots while maintained under the 
ground based operating environment specified in para. 3.2.4 herein. 


3.2.2 Physical 

3. 2. 2.1 Weight . Reserved. 

3 .2.2. 2 Dimensions . The laser breadboard head dimensions shall not be 
greater than 1.1 m x 2.2 m x 1.1 m. 

3. 2. 2. 3 Breadboard Dimensions . The laser breadboard volume shall not be 
greater than 10 m 3 , consistent with commercial transport requirements. 

3. 2. 2. 4 structural Characteristics . The laser breadboard shall have 
structural characteristics enabling it to meet operating and non operating 
environment requirements specified in para. 3.2.4 herein. 

3 . 2 . 2. 5 Material Compatibility . The laser breadboard shall contain only 
materials which are compatible with each other and with the environments 
specified in para. 3.2.4 herein. Specifically, the laser breadboard design 

shall minimize potential oxygen isotope contamination of working gas mixtures 

12 18 

containing the isotope C 

3. 2. 2. 6 Leakage . The laser breadboard flow loop leakage rate shall be less 
than 1 x 10 -2 torr per hour for a period of at least 30 days. 

3.2.2. 7 Connectors ■ Connectors shall preclude incorrect installation or 
application. When appiicable, connectors shall contain physical alignment 
guides . 


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LMSC-HSV SPEC F312362 
Rev. B 

1 April 1991 

3 .2.2.8 Guards . Critical and vulnerable items on the laser breadboard shall 
be located or shielded in accordance with standard laboratory practices. 

3.2.3 Maintainability 

3. 2. 3.1 Design Requirements 

3. 2. 3. 1.1 Corrective Maintenance . The laser breadboard design shall allow 
for easy access and corrective maintenance. 

3. 3. 3. 1.2 Protective Features . The laser breadboard design shall include 
protective features necessary to prevent a safety hazard for maintenance 
actions , 

3. 2. 3. 1.3 Verification . The laser breadboard design shall provide a 
capability for functional verification. 

3. 2. 3. 1.4 Maintenance Points . The laser breadboard design shall include 
maintenance points for the laser breadboard gas system, including those for 
filling or purging, in accessible locations. 

3. 2. 3. 2 Support Equipment 

3. 2. 3. 2.1 Safety . The use of support equipment shall not introduce a safety 
hazard . 

3. 2. 3. 2. 2 Verification of Status. The operational status of all support 
equipment shall be verifiable. 

3.2.4 Environmental Conditions . The laser breadboard storage and operational 
environments are those found in ground based offices and laboratories. 


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Rev. B 
1 April 1991 

3.2.5 Transportability . Shipping containers, packaging, and other safeguards 
shall protect the laser breadboard from normal risks incident to 
transportation, storage, and handling of scientific hardware. 

3 . 3 Design and Construction 

3.3.1 Materials, and Parts . Reserved. 

3.3.2 Electromagnetic Radiation . Reserved. 

3.3.3 Nameplates . Nameplates or product markings shall identify the laser 
breadboard and each of its major components. Identification shall include 
product name and fixed asset owner. 

3.3.4 Safety . The design of the laser breadboard shall address safe 
operational conditions such that failures which may occur will not cause major 
damage to interfacing equipment. 

3.:. 4.1 Hazardous Material . Materials which present toxic hazards to 
personnel shall be avoided in the design of the laser breadboard, Where use 
of toxic materials cannot be avoided, manufacturing and processing controls 
shall be implemented such that environmental limits specified by the 
Occupational Safety and Health Act (OSHA) shall not be violated. Identified 
carcinogenic materials shall not be used in any phase of development. 

Suspected carcinogenic material(s) shall be identified and require LMSC 
approval prior to use in any phase of development. 

3. 3. 4. 2 Dangerous Components 

3. 3. 4. 2.1 Covers . The laser breadboard shall protect personnel from 
accidental contact with potentially dangerous parts such as high voltage 
components . 


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LMSC-HSV SPEC F312362 
Rev. B 

1 April 1991 

3. 3. 4. 2. 2 Identification of Dangerous Components . The laser breadboard shall 
label dangerous components sufficiently to reasonably ensure safety from 
accidental contact. 

3.3.4. 2. 3 Safety Interlocks . Parts of the Laser breadboard which present a 
danger of electrocution shall have interlocks to prevent access when the part 
is energized. 

3 . 3 . 4. 3 Failure Criteria . The laser breadboard shall be designed such that 
no single failure or combination of two failures result In a catastrophic 
event capable of causing injury or loss of Life to personnel. The laser 
breadboard shall be designed such that no single failure results in a critical 
event capable of major damage to facilities or other breadboard components. 

3 . 4 Documentation 


Reserved. 

3 .5 Furnished Component Characteristics 

3 . 5.1 LMSC Furnished Injection Laser . The laser breadboard injection laser 
will be a continuous wave (cw) CO^ laser. 

3 . 5 .1.1 Injection Laser Power . The laser breadboard injection laser power 
output will be not less than 10 W. 

3. 5. 1.2 Injection Laser Beam Diameter . The laser breadboard injection laser 
output beam diameter will be 2.5 ±0.5 mm. 

3 . 5 . 1.3 Injection Laser Valves . The laser breadboard injection laser will be 
sealed . 


3. 5. 1.4 Injection Laaer Beam Mode . The laser breadboard injection Laser 
output beam will have 98 percent of its energy in a TEM^ 


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LMSC-HSV SPEC F312362 

Rev. B 

1 April 1991 

3 . 5 . 1.5 Injection Laser Line Selection . The laser breadboard injec- 
tion Laser will be able to select the 10.59 um line when utilizing a tube 
filled with a 12 C 16 0 2 mixture and the 9.11 um line when using a 

12 C 18 0 mixture. Two tubes shall be provided, one to operate at 10.59 
um and 2 one to operate at 9.11 um. A grating will be incorporated for line 

selection. 

3 . 5 .1.6 Injection Frequ ency Stability. Reserved. 

3 5.1.7 Tniection Laser Beam Po larization. The laser breadboard injection 
laser output beam will have a linear polarization of not less than 95 percent. 

3.5.2 ftnvemment Furnished Ca talyst Characteristics. Reserved. However, the 
breadboard flow loop design is to be based on a GFE catalyst impregnated on a 
monolith support structure in the main flow loop. i.e.. a design without a 
bypass flow loop. 


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LMSC-HSV SPEC F312362 
Rev. B 
1 April 1991 


APPENDIX 

10. PHASE I SELECTED DESIGN SPECIFICATIONS 

10.1 Scope . Specifications based on the Phase I selected design 
configuration are as suntnarized below. It is recognized however that elements 
of this Phase I design are subject to revision in Phase II engineering trade 
studies. Therefore, the specifications below are of use primarily as initial 

design points . 

10 . 2 Power Resonator Pe rformance 

10.2.1 P7.T Control . The laser breadboard resonator PZT cavity length tuning 
device will have a preprogrammed mirror acceleration/deceleration that 
minimizes feedback mirror relocation and have a maximum settling time of 5 ms. 

10.2.2 PZT Tuning Range . The User breadboard resonator PZT cavity length 
tuning range will be not less than 25 um. 

10 .3 Flow Performance . The laser breadboard flow loop will provide 

homogeneous gas flow within the discharge cavity. 

10.3.1 r.a, Temperature . The laser breadboard laser gas temperature will be 
293 +20 K prior to discharge. 

10.3.2 Gas Homogeneity . The relative density variation is not to exceed 1 x 
10 

10.4 Pulse Power . The laser breadboard pulse power unit will consist of a 
full voltage pulse forming network (PFN) and a thyratron discharge switch. 

The laser discharge voltage will not be greater than 40 kV. 


11 

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Rev. B 
1 April 1991 

10.5 R?son?tpr - The laser breadboard resonator will be a confocal unstable 
resonator with square mirrors configured as shown in Figure 2. 

Primary Mirror Ml 



10.5.1 Cavity Magnification. The laser breadboard resonator cavity 
magnification will be 2.25 ±0.25. 

1C. 5. 2 Equivalent Fresnel flumber . The laser breadboard resonator equiv- 
alent Fresnel number will be 2.4 +0.1. 

1C *5. 3 Cavity Length. The laser breadboard resonator cavity length will be 

2.2 ±.02 m . 

1C*5.4 Cavity Size. The laser breadboard resonator cavity beam size 

will be 4 ± . 1 cm x 4 +. 1 cm. 


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LMSC-HSV SPEC F312362 
Rev. B 
1 April 1991 

^•5.5 Primary Mirror Curvature . The laser breadboard resonator effective 
primary mirror curvature (grating + lens) will be 17.5 +.1 m. 

l0 - 5 ' 6 F-gedback Mirror Curvature. The laser breadboard resonator feedback 
mirror curvature will be -7.7 +.1 m. 

IO ' 5 ' 7 — ^ ror Construction . The laser breadboard resonator feedback and two 

turning mirrors will be copper plated and liquid cooled. 

Piezo Translation . The laser breadboard resonator cavity length will 

be tunable by a piezo-electric translation (P2T) device mounted on the 
feedback mirror. 

10.5.9 Grating . The laser breadboard resonator will have a blazed grating 
for beam wavelength selection. 

10.5.10 Windows. The laser breadboard windows will be anti-reflection coated. 

10 • 6 Discharge Cavity 

f 

Laser Excitation . The Laser breadboard power laser excitation will be 
via a surface corona ultraviolet CUV) pre-ionized glow discharge. ! 

* 0,6,2 Specific Loading. The laser breadboard specific loading will be 100 
to 175 Joule per liter atmosphere ( J/ liter-atm) . 

x0,6,3 Gflin Length « The laser breadboard resonator gain length will be 1.50 
±0.1 tn. 

10 . 7 Flow Loop 

Gas Mixture . The laser breadboard working gas mixture will be 50 +25 

percent He, 25 ±10 percent C0 2 , remainder with residual gasses Less 
than 0.1 percent. 


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LMSC-HSV SPEC F312362 

Rev. 6 

1 April 1991 

10.7.2 Gay Pressure . The laser breadboard working gas pressure will be 
greater than 0.2 and less than 0.5 atmospheres (atm). 

10.7.3 Cavity Flush Factor . The laser breadboard power laser cavity flow 
flush factor will be greater than 2.5. 

10.7.4 Cavity Acoustic Transits . Acoustic pulse transits within the 
discharge cavity of the laser breadboard will be greater than 50 across the 
acoustic mufflers. 

10.7.5 Catalyst . The laser catalyst will be an in-the-f low-loop monolith or 
honeycomb structure. 


*_4 


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Appendix B 

Enhanced LAWS Laser Test Results 
From ID Program 


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The LAWS laser hardware has been u t ; depicts operadons at 330 and 380 
Program Z773 to obtain enhanced I**™ ' ^ No(e ^ increaS e in measured efficiency 

Torr at specific energy load.ngs tom 60 to 90 j t code predic tions and 

as pressure and energy load.ng ts uicreased . ^ ^ „ pwards of 10 J output for the 

test data. The figure depicts a demons fflciencies of 8 to 10 percent The top two figures 

LAWS laser. Note the measured disc g ^ ^ ditferenl opetat ing pressures. The 

validate the model (o). with actual test d < ) J/L . atm/475 Torr design point from the 

bottom figure extrapolates to ou pu achievable through a minor modificadon of the 

validated model. Arc-free design load.ng will ^ ^.ion of side-by-side 

current electrode dielectric configuration ^ electric material prior to machining to 

electrodes by approximately 1 mm, or by test of the 
eliminate minor voids in strategic regions. 


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CD 


CN 



S3inor ni A0d3N3 mdino 


n 

CD 

eg 

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