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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LOCKHEED-HUNTSVILLE
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
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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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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
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2-1
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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
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Section
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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
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Section
3-83
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4-1
6-1
6-2
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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
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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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LOCKHEED- HUNTSVILLE
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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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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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
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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.
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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.
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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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Figure 2-5. LOS Pointing Errors Introduce Errors into Wind Velocity Vector Measurements
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„ 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 ).
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LOCKHEED- HUNTSVILLE
Maximized Signal to Noise
Ratio Factors
Figure 2-7. Contributing Factors for Maximized Signal-to-Noise Ratio
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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
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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.
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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.
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X
IVCIOCITTI
Figure 3-2. LAWS Baseline Design Flight Configuration
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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.
3-57
LOCKHEED-HUNTSVILLE
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
3-58
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
3-59
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
. 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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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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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
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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
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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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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.
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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
3-77
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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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&
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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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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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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LOCKHEED-HUNTSVILLE
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
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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
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Command Command Command
LMSC-HSV TR F320789-II
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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.
3-87
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LMSC-HSV TR F320789-II
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.
3-88
LOCKHEED-HUNTSVILLE
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
LMSC-HSV TR F320789-II
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
^ gg FOLDl»wE
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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
LMSC-HSV TR F320789-II
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
LMSC-HSV TR F320789-II
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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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
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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
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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
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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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4-2
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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
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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.
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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
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HeNe ALIGNMENT
V LASER
LMSC-HSV TR F320789-II
6-2
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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
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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
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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
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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
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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
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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
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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
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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
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T/dlv 1 jje
Figure 6-13. Chirp Measurement from Fast-Fourier Transform, Example Measurement B
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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
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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
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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
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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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LMSC-HSV TR F320789-II
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)
6-17
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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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LMSC-HSV TR F320789-II
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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LMSC-HSV TR F320789-II
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.
7-1
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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
7-2
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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
7-3
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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
1A
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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%
7-5
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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
7-6
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
• 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
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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.
7-8
LOCKH EED-HU NTSVILLE
LMSC-HSV TR F320789-II
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
7-9
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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
7-10
LOCKHEED-HUNTSVILLE
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
7-11
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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
7-12
LOCKHEED-HUNTSVILLE
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.
7-13
LOCKHEED-HUNTSVILLE
Table 7-6. Analytical Tools for Contamination Analysis
LMSC-HSV TR F320789-II
7-14
LOCKHEED-HUNTSVILLE
and redistribute by vibroacoustic excitation
LMSC-HSV TR F320789-II
Appendix A
Functional Specification
LAWS Laser Breadboard
A-1
LOCKHEED-HUNTSVILLE
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
A-2
LOCKHEED-HUNTSVILLE
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
A-3
LOCKHEED-HUNTSVILLE
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
1116P
ii
A-4
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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
A-5
LOCKHEED-HUNTSVILLE
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.
2
A-6
LOCKHEED-HUNTSVILLE
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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LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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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LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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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LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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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LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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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LMSC-HSV TRF320789-II
LMSC-HSV SPEC F312362
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 TR F320789-II
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 TR F320789-II
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.
1136P
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LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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.
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LMSC-HSV TR F320789-II
LKSC-HSV SPEC F312362
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 TR F320789-II
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.
1136P
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LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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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LOCKH EED-HU NTSVILLE
LMSC-HSV TR F320789-II
Appendix B
Enhanced LAWS Laser Test Results
From ID Program
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LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
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.
B-2
LOCKHEED-HUNTSVILLE
LMSC-HSV TR F320789-II
CD
CN
S3inor ni A0d3N3 mdino
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CD
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Q_
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