DOCUMENT CONTROL 

DIN 50286-154-1 PAGES 668 

COPY NO. 624 OF 67 COPIES 

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DATE: 12 February 1968 

ENGINEERING ANALYSIS REPORT 

ACQUISITION SUBSYSTEM 

CDRL ITEM NO. 278 

SECRET SPECIAL HANDLING 

kEB 
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TABLE OF CONTENTS 

Section Page 

1.0 REQUIREMENTS AND SUBSYSTEM DESCRIPTION 

1. 1 Requirements 

1.2 Comments Concerning GE Compliance with Specifications 

1-1 

1-1 

1-13 

1. 2. 1 Pointing Accuracy 1-13 

1. 2. 2 MCS Pointing Corrections 1-15 

1. 2. 3 Jitter Performance 1-15 

1. 2. 4 ATS Resolution 1-15 

1. 2. 5 Cue Retrieval Time 1-16 

1. 2. 6 Protective Shroud 1-17 

1. 2. 7 Weight 1-17 

1. 2. 8 Decoupling 1-17 

1.3 Subsystem Description 1-19 

1.3. 1 Purpose of the Acquisition Subsystem 1-19 

1. 3. 2 Operational Modes 1-23 

1. 3. 3 Description of Subsystem Hardware 1-27 

1. 3. 3. 1 Acquisition Optics 1-29 

1. 3. 3. 2 Drive K Electronics 1-39 

1. 3. 3. 3 Visual Display Projector Design and 

Description 1-40 

1. 3. 3. 4 Cue Module Storage Rack 1-47 

1. 3. 3. 5 Magnification Control Stick 1-47 

1. 3. 3. 6 Head Rests 1-49 

1. 3. 4 ATS Controls Description 1-53 

1. 3. 4. 1 General Functional Description 1-53 

1. 3. 4. 2 Detailed Functional Description 1-53 

iii 

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TABLE OF CONTENTS (Cont) 

Section Page 

2.0 ATS ANALYSIS 2-1 

2.1 ATS Optics Analysis 2-3 

2.1.1 Design Description 2-3 

2.1.1.1 Optical Prescription 2-3 

2.1.1.2 Optical Design-Low Power 2-6 

2.1.1.3 Optical Design-High Power 2-6 

2.1.1.4 Eyepiece Assembly 2-6 

2.1.2 Analysis of Proposed ATS Optical Design 2-15 

2.1.2.1 Introduction 2-15 

2.1.2.2 Conclusions 2-16 

2.1.2.3 Recommendations 2-17 

2.1.2.4 Discussion of Results 2-17 

2.1.2.5 Discussion of Analysis Techniques 2-36 

2.1.3 GE Analysis Optical Performance Under 

Dynamic Conditions 2-44 

2.1.3.1 Target Characteristics 2-47 

2.1.3.2 Summary of Static Performance Analysis . 2-47 

2.1.3.3 Summary of Dynamic Performance 

Analysis 2-49 

2.1.3.4 Static Performance Analysis 2-52 

2.1.3.5 Dynamic Performance Analysis . . ..... 2-59 

2.1.4 Secondary Design Performance Information 2-68 

2.1.4.1 Vignetting 2-68 

2.1.4.2 Gimbal Angles Versus Scan Field 

Position 2-70 

Exit Pupil Shift 2-70 

Total Transmittance Summary 2-70 

Focus 2-74 

2-79 

Controls 2-81 

Requirements Interpretation 2-81 

Discussion of Approaches 2-84 

Rate Control Servo Configuration 2-85 

Position Servo Configuration 2-103 

Position and Rate Servo Configuration • • 2-119 

Baseline Configuration-Description, 

Requirements and Design Margins 2-128 

Command Generation 2-156 

SECRET SPECIAL HANDLING 

2.1.4.3 

2.1.4.4 

2.1.4.5 

2.2 ATS Controls 

2.2.1 Scanner 

2.2.1.1 

2.2.1.2 

2.2.1.3 

2.2.1.4 

2.2.1.5 

2.2.1.6 

2.2.1.7 

iv 
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TABLE OF CONTENTS (Cont) 

Section 

2.2.1.8 Slaving Mode 

2.2.1.9 ATS Performance with the Man 

in the Loop 

2.2.1.10 Initial Tracking Error 

Page 

2-164 

2-167 

2-169 

2.2.2 Image Orientation (Derotation) Control System. . . .. 2-172 

2.2.2.1 Requirements-System Level 2-172 

2.2.2.2 Servo Configuration 2-172 

2.2.3 Dynamic Component Requirements 2-178 

2.2.3.1 Command Filter 2-178 

2.2.3.2 Amplifier 2-178 

2.2.3.3 Motor 2-179 

2.2.3.4 Position Transducer 2-179 

2.2.4 Zoom Control Servo 2-180 

2.2.5 Blanking Control Servo 2-184 

2.2.6 Power Change Servo 2-184 

2.3 ATS Pointing Error Analysis 2-187 

2.3.1 Discussion of Results 2-187 

2.3.2 Description of Error Sources and Errors Removed 

by Bores ighting 2-195 

2.3.3 Procedure for Determining Pointing Errors 2-202 

2.3.4 Statistical Treatment of Error Sources and the 

Total Pointing Error 2-202 

2.3.5 Results of Monte Carlo Computer Programs 2-205 

2.3.6 Procedure Required to Determine Allowable 

Misalignments 2-215 

2.4 ATS Thermal Analysis 2-223 

2.4.1 External Optics Assembly 2-223 

2.4.1.1 General External Component Design 

Philosophy 2-223 

2.4.1.2 Temperature Constraints 2-229 

2.4.1.3 Environment 2-229 

2.4.1.4 External System Temperature Behavior . 2-231 

2.4.1.5 Coating Selection 2-254 

2.4.1.6 Insulation Materials 2-256 

2.4.1.7 Area Requiring Additional Analysis 2-257 

2.4.2 Telescope Assembly 2-259 

2.4.2.1 Design Philosophy 2-259 

2.4.2.2 Temperature Constraints 2-259 

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TABLE OF CONTENTS (Cont) 

Section Page 

2. 4. 2. 3 Environment Description 2-259 

2. 4. 2. 4 Temperature Profile 2-260 

2. 4.2. 5 Conclusions 2-274 

2. 4. 2. 6 Additional Analysis Required 2-274 

2. 4. 3 Thermal Development Test Plans 2-275 

2. 4. 3. 1 General 2 -275 

2. 4. 3. 2 Externally Mounted Equipment 2 -275 

2. 4. 3. 3 Internal Mounted Equipment 2-280 

2.5 ATS Structures Analysis 2 -283 

2. 5. 1 Structural Analysis of Telescope Under Powered 

Flight Loads 2 -285 

2. 5. 2 Structural Analysis of Tracking Pedestal Under 

Powered Flight Loads 2 -307 

2. 5. 3 Structural Analysis of the Protective Shroud Under 

Powered Flight Loads and On-Orbit Operating Loads . . 2-319 

2. 5.4 Deflection Analysis of Telescope and Tracking 

Pedestal Due to 1G Load Relief 2-337 

2. 5. 5 Determination of Optical Path Misalignments Due to 

Mechanical Joint Slippage 2-345 

2. 5. 6 Determination of Misalignments Due to Shell 

Pressurization and "Hot-Dogging" 2-351 

2. 5. 7 Dynamic Analysis and Loads Determination 2 -355 

2. 5. 8 ATS Weight Analysis 2 -381 

2.6 Bore s ighting . 

 2-389 

2.6. 1 Normal Mode Equations 2-389 

2. 6. 2 Failed Pellicle Mode 2-395 

2. 6. 2. 1 Nomenclature 2-396 

2. 6. 2. 2 Analysis 2 -396 

2. 7 Image Derotation and Manual Control Stick Recoupling 2-401 

2. 7. 1 Introduction 2-403 

2. 7.2 Notation and Definition 2-403 

2. 7. 3 Coordinate Systems 2-405 

2. 7.4 Tracking Mirror Matrix 2 -409 

2. 7. 5 Image Derotation 2-411 

2. 7. 6 Image Motion Decoupling 2-415 

2. 7. 7 Approximations to the Derotation Equations 2 -421 

2. 7. 8 Standardized Equations 2-423 

2. 7. 9 Image Derotation 2-425 

2. 7. 10 Image Motion Decoupling 2 -427 

vi 

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TABLE OF CONTENTS (Cont) 

Section Page 

3.0 CUE SYSTEM ANALYSIS 3-1 

3. 1 Introduction 3-3 

3.2 VDP Optical Analysis 3-5 

3.3 Subcontractor-Compliance Comments 3-17 

3. 3. 1 Access Time (Single Frame Operation) 3-17 

3. 3. 2 Operating Times 3-17 

3. 3. 3 Display Screen Brightness 3 -18 

3. 3. 4 Display Resolution (Projector) 3-19 

3.3. 5 Alpha-Numeric Display Time 3-19 

3. 3. 6 Search and Fine Positioning Power 3-19 

3. 3. 7 Screen Illumination Power 3 -20 

3. 3. 8 Alpha Numeric Display Power 3 -20 

3. 3. 9 Standby Power 3 -20 

3.3. 10 Power Interrupt 3 -21 

3.4 Controls and Structure 3-21 

4. 0 Subsystem Design Information 4-1 

4. 1 Drive K Electronics 4-3 

4. 1. 1 Introduction 4-3 

4. 1.2 K-0 Electronics 4-5 

4. 1. 2. 1 Functions (K-0) 4-5 

4. 1. 2. 2 Performance Characteristics (K-0) 4-7 

4. 1. 2. 3 Additional Buffering Requirements for K-0 4-10 

4. 1. 2. 4 Digital Integrator 4-11 

4. 1. 2. 5 Feedback Processor 4-12 

4. 1. 2. 6 Position Error Adder 4-13 

4. 1. 2. 7 Timing and Control Circuitry 4-14 ,- 

4. 1.3 K-1 Electronics 4-15 

4. 1. 3. 1 Description 4-15 

4. 1. 3. 2 Requirements Summary 4-18 

4. 1. 3. 3 Performance Characteristics (K-1) 4-21 

4. 1. 3. 4 Stability Analysis (K-1) 4-23 

4. 1.4 K-2 Electronics 4-27 

4. 1. 4. 1 Description 4-27 

4. 1. 4. 2 Performance Characteristics (K-2) 4-27 

4. 1. 4. 3 Stability Analysis (K-2) 4-30 

4. 1. 4. 4 Test Results on the K-2 Electronics 4-30 

4. 1. 5 K-3 and K-4 Electronics 4-37 

4. 1.6 K-5, K-6, and K-7 Electronics 4-41 

vii 

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TABLE OF CONTENTS (Cont) 

Section 

4.2 

4.3 

4.4 

Magnification Controller 

Vendor Selection Trade-Offs and Criteria for the ATS 

Rate Gyro 

Alpha Subsystem Power Requirements 

4.4.1 Alpha Weight Status 

Page 

4-45 

4-47 

4-49 

4-53 

5.0 CONSOLES AND DISPLAYS 5-1 

5.1 Consoles and Displays Requirements 5-3 

5.2 Panel Layout 5-3 

5.3 B-Panel Controls and Displays 5-9 

6.0 RELIABILITY, CONTAMINATION CONTROL, AND 

MAINTAINABILITY 6-1 

6.1 Reliability Requirements and Evaluation 6-1 

6.1.1 Summary 6-1 

6.1.2 Systems Reliability Requirements 6-3 

6.1.3 Mission Reliability Design Criteria 6-5 

6.1.4 Design Exceptions 6-7 

6.1.5 Acquisition System Reliability Estimate 6-7 

6.1.6 System Failure Mode and Effects 6-9 

6.2 Contamination Control 6-49 

6.2.1 Philosophy 6-49 

6.2.2 Visual Display Projector (VDP) 6-49 

6.2.3 Acquisition Tracking Scope (ATS) 6-51 

6.3 Maintainability 6-53 

6.3.1 Concept 6-53 

6.3.2 Analysis 6-53 

6.3.3 Effects of Maintainability on Other Subsystems 6-54 

6.3.4 Summary 6-55 

6.4 Eye Hazard Analysis 6-57 

6.4.1 Summary of Conclusions 6-59 

6.4.2 Discussion 6-59 

6.4.3 Calculations 6-62 

vi" 

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INTRODUCTION 

It is the purpose of this report to support the Acquisition Subsystem PDR by documenting 

analysis, and descriptive material, to demonstrate the capability of subsystem design concepts to satisfy the proposed CEI requirements of ECP No. 17R5*. Most of the analysis 

presented represents work performed by GE. In a few cases, material has been extracted 

from subcontractor PDR reports to demonstrate compliance with principal requirements. 

In general, however, the PDR reports of Itek and Lear Siegler are considered part of the 

GE PDR documentation, and detailed analysis presented in these reports is not repeated 

here. 

Section 1.0 of the report contains a table listing the requirements negotiated with the Air 

Force for the Acquisition Subsystem, briefly describing the design solution proposed for 

the requirement, and referencing the sections of the report which relate to the requirement. 

A general statement of the state of compliance with the requirement is included in the 

table; further discussion of compliance with specific requirements is presented in Section 

1. 2, where required. 

Section 1.3 presents a description of the Acquisition Subsystem functions and hardware. 

*Issued 20 Nov 1967 

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SECTION 1.0 

REQUIREMENTS AND SUBSYSTEM DESCRIPTION 

The material in this section is subdivided as follows: 

1.1 Requirements and Compliance Summary 

1.2 Discussion of Compliance 

1.3 Subsystem Description 

1.1 REQUIREMENTS 

Table 1.1-1 presents a summary of the subsystem functional requirements, briefly describes 

the design solution for each requirement, and references sections of the EAR report which 

pertain to each requirement. A yes/no statement of compliance is also included. Where 

further discussion is required concerning compliance, it is included in Section 1.2. 

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TABLE 1.1-1. 

1-2 

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SECTION 

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DESIGNSOLUTION 

Separatemanual sticks,and VOand AO eyepiece; are located at each crew 

console topermit targetsto besimultaneouslytracked and centered by 

twocrewmen,oneusingthe VO and theother the ATS. 

Whenatargetis 

centered Ineach eyepiece,thecomputercan bedirected toaccept 

positioninformation fromeach trackingmirror and calculatealignment 

correctionparameters.It will benecessary to boresight on at least twotargets. 

a. 

AnATS and itsassociated eyepieceareavailableateach crew 

console.Head restsareprovided toaidinviewingtheexitpupil. 

Separatesetaof crewcontrolsareprovided ateach console. 

Separatecuescreens are located topermit vie

wing fromeach head 

rest. 

b. 

Manual controlsarelocated oneach projectorwhich permit frame call-up withoutcomputerparticipation. 

Theoperationof themagnificationstepchangemechanismandthegyro 

heatersareinhibited during ATS stews toconservepeak power. 

Thecomputerwill monitorthepositionreadoutsfromthegimbals, andthe 

Pechanprism,and executeadecouplingroutinetogeneratedecoupled 

commandsderived fromthe MCS inputs. The decoupled MCS commandswill 

besummedwith thecomputergenerated trackingcommandsandappliedto 

thegimbal controls. 

APechanprismwill berotated inthe ATS optical train byaclosed loop control systemascommanded bycomputer Inputs. 

Theprismwill be 

positioned duringslewand remainstationaryduringtrack. Boresightingroutines,control loopoffsets,and alignment errorapportionments have been developed topermit thisrequirement to benearlymet, 

Therequired alignment budget istight,however. 

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REQUIREMENT 

Providemeans to boresight ATS toMO, 

LOS and toupdatetargetand/or 

ephemerisinformation 

a. 

ProvideAO display,cuedisplay, 

and associated controlsateach crew console. 

b. 

Providecontrols forpre-pass cue 

review 

GE-AVEpeak powerallotmenttoAcquisitionSubsystemis480 watts 

a. 

Manual Control Stick shall causescene 

displacementsinthe ATS eyepiecewhich areparallel tothestick motiontowithin +.20,exclusiveof vehiclealtitude 

errors. 

b. 

Orient thetargetimagedisplayed in the 

ATS eyepiecesothatthetrack vector of theimageiswithin+5° of vertical in 

theeyepieceat theend of slew. 

a.TheLOS shallpointtoa target within 

+8 minutesof arcinautomaticoperationassuming; 

a. 

Boresightingcomplete 

b. 

Perfect ephemerisand target locationinformation 

c. 

LMstructural deflectiondoesnot 

exceed 0. 5arcminute 

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DESIGN SOLUTION 

Theresolutionof thetelescopewill permitthisrequirementtobemet. 

The formulationof anMCStransfer functionand a demonstrationof the 

crew's abilitytooperateinthecontrol loopwill beaccomplished in 

futuresimulationwork. 

Considerationof serviceoffsetsto therate(roll)servoloop,dynamic errors,and ephemeris uncertainties indicatethatthisrequirementcan be 

readilymet. 

A positionplusrateservoisutilized inpitch,and aratesysteminroll. 

Gyros areused on both axes to increasenoiserejection.Both loops can be 

operated in backupmodesinevent of gyrofailure.Theeffectsof bearing 

and encodernoisearerelevant to thisrequirementand havebeenanalyzed 

extensively. 

Thedesignsatisfiesthisrequirementinpitch only. Theselectionof a rateservo loop fortheroll gimbal minimizes theeffects 

of vehicle disturbances CVO ATS rateerrors. 

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b.TheLOS shall pointwithin& 2minutes 

(2o)aftercrewcorrectionsvia the stick. 

a. TheATS LOSshall track such that 

target driftisless than540 A rad/sec 

attheend of slew. 

b.Inauto

maticoperation,co

mbined drift 

and jittererrorsshall be lessthan 

described inFigure2. 1.1 exclusiveof: 

1. 

Effectsof OVmotionduetonon- GE equipment,ACTS operation, crewmotions, maintracking mirror, and slidingmask. 

2. 

Effectsof ATS gimbal rates 

below0. 01 deg/sec. 

li 

c.WhenMainTrackingMirrorratesare 

12 deg/sec(each axis) with associated 

slidingmask rates,theATS LOS errorshall not bemorethan130μrad/ sec over that of Figure2. 1.1. 

d.With OV rateIncrementsduetoACTS 

thrusteroperationof 0. 1, Oftip and 

0..13 deg/secbzroll,pitch, ilia yaw, respectively,theadditional LOS rate errorshall not exceed 75011 rad/sec. 

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Pointing Accuracy 

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SECRET SPECIAL HANDLING 
TABLE1. 1-1. SUMMARY OF REQ

UIREME

NTS AND DESIGN SOLUTIONS (Cont) 

1-4 

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DESIGN SOLUTION 

A backup loop isprovided topermit operationof both pitch and roll loops as 

positionservosystemswithout gyrosupport.It isassumed thatgimbal 

rates to0. 05 deg/secneed be tolerated in theroll looponly. 

Thecrewwill boresight theATS lineof sight tothe MO LOS by trackingand centering thesametargetIn both FOV's simultaneously,and initiatinga 

softwareroutinetocalculatealignmentcorrectionparameters to beapplied 

tosubsequent targets. Thesoftwareroutineassumesthreeorthogonal 

parameterswill define themisalignmentwhich exists. 

Separatesoftwareand control capabilitiesareprovided foreach ATS and themainoptics toper

mit simultaneousobservationsof threeseparate 

targets.Twoseparatemanual controlsticksprovidethecapabilityto 

manuallycontrol both ATS's, orif theIVS isoff,tocontrol oneATS and the MO. 

ThecrewcancausethemainopticsLOS to beslavedtoeither ATS LOS by actuatingapanel switch tocommand thecomputertooperatethemain 

opticstotheslavemode.InthismodetheATSstewsandtracks thesame 

target astheMO.Manual stickcorrectionsappliedbytheMCS forthe ATSaresimultaneouslyappliedtotheMO. Thefocal length of the telescope,and thesizeoftheentrancepupil tothe 

eyepieceassembly,aresuch astoprovidethedesired fieldsof view. 

Thescannerprovidessufficient gimbal freedomtosatisfythescan field requirement. Theconfigurationof theshroudand external mirrors 

satisfythevignettingrequirements. 

. 

`-) 

8 ., REQUIREMENT In theevent of asinglegyro failure,co

mbined LOS drift, and Jittershall beless 

than fivetimesthat shownin Figure 2. 1. 1 

(with theexclusioninb)aboveand excludingcaseswhere ATS gimbal ratesare less than0. 05 deg/sec. a.Provide for boreeighting both ATS LOS's 

totheMO LOS.After boresighting, the 

ATS and MO LOS's shall remainaligned 

towithin16minuteswhen tracking the 

same target(excludeserrors due to 

LM structural deflectionsof more than 

*0.5 arcminuteafter boresighting). 

Providecapabilitytosimultaneously track 

separatetargetswith the two ATS's and the MO and toaccept manual control signals foranytwoof thethreeL

OS's. 

ProvidecapabilitytoslaveeitherATS to the 

MO so that ATS position,and rate corrections, aresimultaneouslyapplied to the MO 

a. 

TheFOV shall beat least 4. 0 degrees atminimummagnificationand at 

least 1 degreeatone-halfof maximummagnification. 

b. 

Thescan field shall permitpointing 

between+70 degreesand -40 degrees 

inpitch and between+45 degrees in 

roll (ref tovehicleco-ordinates). Thereshall benovignetting between +35 degreesinroll and+10 to+40' in 

pitch. 

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T

ABLE 1. 1-1. SUM

MARY OF REQUIREM

ENTS AND DESIGN SOLUTIONS (Cont) 

EAR 

SECTION 

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DESIGN SOLUTION 

Amechanicallycompensated zoomassemblyisprovided topermitcon -

tinuouslyvariablemagn ificationovera 2:1 range. Powerchangelensesare 

inserted into theoptica l path toprovidea4:1 stepreductioninmagnification. Thesizeof deelectrical motorsrequ ired tomeetthetimeresponserequire- ment has been determined.Thedrivemotorsarecontrolled byelectronics 

intheDriveK electron icsoncommands fromtheMagn ificationControl Stick. 

W 

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C 4'1 

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r L. > 4 

E -, 

Theminimummagnificationshall be16\ 

1X. Themaximumpowershall be125X 

+5,-3. A zoomshallbeprovidedbetween 

minimumandtwiceminimummagnification 

(lowpower)and betweenmaximummagn ificationandone-half ofmaximummagnification(high power). Thezoomshall respond 

(end toend)in less than0. 5second.The optics shall respondfromhigh tolowpower 

in1 second. 

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SECRET SPECIAL HANDLING 
TABLE 1. 1-1. SUMMARY OF REQUIREMENTS AND DESIGN SOL

UTIONS (Cont) 

1-6 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 -SEM+ SPECIAL HANDLING EAR SECTION DESIGN SOLUTION A10-inch objectiverefractivetelescopewith a high powerof 127X,will beprovided tosatisfythisrequire

ment. ModulationTransfer function 

analysis byItek isprovided to demonstrateconformancewith this 

requirement. BackupGE analysisisalsopresented which indicates 

thatanextremeworst-case buildupof tolerancescouldresultina slightout-of-speccondition. 

Analytical predictionsof off-axisperformanceareincluded inthis 

report.Thesepredictionshave beenmade by GE. Thesubcontractor has notyetsubmitted hisestimatesof off-axisperformance. 

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3. 

TheLOS betweentargetand 

ATSisstatic(nojitter). 

4. 

The1. 65 feetwide(pro- jected)barsareatthecenter 

of theapparent field of view. 

Within10 degreesof thecenterof the apparent field-of-view, theobserver 

shallbeabletoresolve lines of TBD feetwidth and spacing(projected)under 

theabove(except4)conditions. 

Attheedgeof therealfield-of-viewin 

theareasnotobscured bytheperipheral display,theobservershall beableto 

resolvelinesof TBD feetwidth and spac- ing(projected)undertheabove(except 

4)condition. 

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Thevignettingand transmissionof thetelescopearecontrolled 

tosatisfythisrequirement. 

Thetelescopewill beslightly defocused ontheground sothatitwill beproperly focused inanorbit,5 psi,environment.Thefocal 

length isshort enough,(73 inchesathigh power)topermitground scenes to beaccommodated withoutrefocusingas afunctionof slantrange.Theeyepiecewill per

mitmanual focusing toacco

m- modateeye dissimilarities. 

35 lightsaremounted attheperipheryof theeyepieceFOV. The 

lamps arecontrolled bycomputercommandsviatheconsole 

controller. Pinlightsareutilized toconserve powerand space. 

The display has beenmocked upforcrewreviewof itseffectiveness.Priorityand timinginformationtothecrew.Thedisplay 

conformswith theobscurationrequirements. Theeyerelief is0. 77 inch.Eyerelief varies less than0. 02 inch duringzoomand magnificationstepchanges. 

Anoptical flat islocated intheoptical trainattheentrancepupil ofthezoomeyepiece. Thiselementisinscribed with acircle 

equal in diametertotheFOVof theeyepieceathigh zoom. 

A VisualDisplayProjectorwith a6. 5 inch circular,rearprojectionscreenwill belocated ateach crewstationinaposition 

topermitviewing from theATS headrest. Thealphanumeric 

displaywill beplaced overthescreen. 

A singlestoragerack isplanned forall16modulesonconsole 

IA. 

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REQUIREMENT 

Theimage brightnessattheedgeof the 

FOVshallbeatleastone-half of thatat 

thecenterwhennovignettingoccurs and 

fortargetsof uniformbrightness. Meansshall beprovided toinsurethat ground scenesareproperlyfocused when theMOLaltitudeisbetween 75 and 230nm. 

At anygivenaltitude,thesceneshall remaininfocus throughoutthescan field 

withoutmanualfocusing.Theeyepiece 

shall per

mit focusingfrom-2.5 to+3. 5 

diopters. 

a. 

TheATSshallprovideamonocular 

eyepiece.Theeyepieceshall incorporateaperipheral display. ThePDshall notobscurethetargetscenewithin25 degreesof the 

in-trackvector. 

b. 

Theeyerelief shallbeatleast 0.4 

inch and shall notvarybymorethan 0. 06inch. 

c. 

A singlecirclereticlewhich indicatesthereal FOVatmaximumzoom 

positionshall beprovided.The 

reticleshallincludecross hairs for 

centering. 

A pictorial screenprojected cueand a sixcharacteralphanu

mericshallbedisplayed ateach crewstation. 

Storage for16filmmodulesshallbe provided. 

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1-7 

SECRET SPECIAL HANDLING 

TABLE 1. 1-1. SUMMARY OFREQ

UIREMENTS AND DESIGN SOL

UTIONS (Cont) 
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TABLE 1. 1-1. 

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DESIGNSOLUTION 

Thecomputerwill output a binary, serial callupcommand, which will be held intheconsolecontrollerand presented totheVDP oracon- tinuous, parallel signal.Thecomputerwill change thecueafter K 

seconds bysendinganewcallupco

mmand. 

Thecrewcan hold acue beyond K seconds bytwistingthehandleof 

themagnificationstick tosignal thecomputertohold thecue. 

Thecueprojectorprovides atwo-speed film transport systemtoslew the film.Each film frameisread duringretrieval bysiliconphoto transistorswhich detectbinary descriptor bitsonthe film.Avelocity 

trackonthefilmisalsoopticallymonitored duringslewtocontrol filmslewspeed.Propermotorsizingand break designper

mitcornpliancewith therequirement. 

Thealphanumerics shalldisplayinformationreceived fromthecomputerwithin0.3 seconds.Theinformationis decoded for displayin theVDP.TheANdisplaywill not have thecapabilityof displaying 

informationfromthecue film. Manual selectiondialscalibrated indecimal numberswillbelocated oneach VDP.TheVDP will be designed toacceptoctalnumbers from these dials and call upthecommanded frame. 

Headrests havenot beendesigned at this time. Trial headrests have beeninstalled intheconsolemockupforcrewreview. 

A hinged protectiveshroud willcoverand protect all external componentswhen theATS isnotinuse. Theshroud doorwill becontrolled byredundantmotorsco

mmanded byeitherthecomputerora 

panelswitch.Theshroud providespassivethermal control forex- ternal ATS components,protectsagainst Orbit Adjust thruster 

impingement,and protectsagainstcontaminantswith asealing gasket.A final shroud designconcept hasnot beenaccepted. 

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a. 

Thecueframeshall bedisplayed 

and held automatically for K secondsafter beginningof slewto 

target. Thetargetshall beautomaticallyreleased whenatarget 

votebuttonisactuated. 

b. Acuehold control shall beprovided 

topermitlongerviewingof thecue desired. 

c. 

Thecueframeforatargetshall be displayed within0. 45 + N 

100 

secondsafterthereleaseof the 

previous cue. 

d. 

Thealphanumericsshallbedisplayed within0.3 secondsafterthestartof 

ATS slew. 

Controlsshallbeprovided topermitthe 

crewtomanuallycommand the display 

of anycueframewithoutcomputerpro- cessing. 

Adjustable headrestsshallbeprovided to 

facilitateviewingof theATSdisplay. Protectionshallbeprovided forexter- nal ATS componentsagainstACTS and 

wasteproduct contamination. Protectionshall beprovided asnecessary to 

adequatethermal stability. 

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TABLE1.1-1. S

UMMARY OF REQ

UIREME

NTS AND DESIGN S

OLUTIONS (Co

nt) 

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DESIGN SOLUTION 

Themagnificationcontrol stick hasnot been designed.Itwillbe 

designed topermitthecrewtoeffecttherequired controls with the 

left hand.A "slew" signal will bederived fromthecomputergimbal commandsintheDriveKelectronicsand used toinhibitstepmagnificationchanges duringATSslew. 

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a 

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41 41 

41 41 

a 

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a. 

Amagnificationcontrol stick shall 

beprovided tocontrol ATS magnification,VO magnification,target 

voting,and cuehold. 

b. 

Thestick shall beinhibited from 

commandingstepchanges during 

ATSslew. 

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TABLE 1. 1-1. 

SUM

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1-10 

NRO APPROVED FOR 

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EAR 

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DESIGN SOLUTION 

c. 

Themagnificationstick will providesignalsto thecomputertoindicatewhich of fourmagnificationrangesareineffect. Thecomputer 

will applysuitablecorrections tothemanual control stick transfer 

functiontoreduceapparentrateand sensitivitychangeswhenthe ATS 

magnificationischanged. Themagnificationstick will generatesig- nalsusedbythecomputertovarythegainof theMCS. 

d. 

Themagnificationstickwill actuateamicroswitch to indicatetothe 

computerwhetherthecrewwantsthe MCS tocontrol the ATS ortheVO. 

Thecrewmanwill beallowed tocontrol themainoptics with thestick onlyifthe IVS isoff and the LEFT-RIGHT AUTO switchis inthe 

propernon-AUTO position,orif theMO targetwas assignedtohis 

ATS. Tentativedesignwill permitthe MCS tocontrol the VOonly when: 

1. 

Themagnificationstick isintheVOposition. 

2. 

The IVS isoff. 

3. 

Themainopticstargetwas assignedtothecrewconsolerequestingMCScontrol. 

Thecomponentswill benominallyaligned asdescribed tominimizethe 

occurrenceofzeroroll rates.Thetoleranceswill beconsiderably 

tighterthanthe0. 5 degreesspecifiedinordertomeetpointingrequire- ments. 

Anautomaticallytriggered blankingshutterwill beprovided whose responseand sunsensor FOV aresuchthatthesuncannotenterthe 

FOVatmaximumslewspeeds. It isplanned thattheshutterwill be 

manuallycockedthoughthisfeature hasnotnegotiated with thesub 

contractor. 

Twotechniques have beenconsidered forslaving.Onecauses themain optics and theATStotrackthesame target In lapeedantlywithMCS 

correctionsapplied toboth subsystemssimultaneously.Theothercauses 

thepositioninformation fromthetwosubsyste

m's encoders to becompared 

and the differencenulled. 

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Thestick shall generatesignalswhich vary 

thegainsof themanual control stick. Whenthemagn ificationstick isinthe ATS position,themanual control stick shall con- trol only ATS pointing. Wheninthe VO 

position,themanual control stick will con- trolonlymainopticspointing. 

Theroll axisofeach ATS shall becanted outboard (fromthe+X axisof theOV)9e0. 5 degrees. Theoptical axisbetweenthecen - tersofthefixedandtrackingmirrorsshall 

beparallel tothe X-axisof theOV within 

0. 5 degree. Theroll axisof each ATS shallbeparallel tothe X-Y planeof theOV 

within0. 5 degree. 

Auto

maticblankingprovisionsshall bein- corporated topreventcrewinjuryresulting 

from inadvertentsunviewingthrough the 

acquisitionoptics. Theblankingmechanismshall bereset manually. 

Themainoptics LOS shall track within 

100 /A rad/secof thedesired ratewhen 

operatingintheslavemode in the track 

period. 

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The ATS gimbal drivemotorsaresized and controlledtopermitcompliance. The dynamicrangeof theservo loopamplifiergainsaredictated bytheslew 

requirements. 

Energyabsorbingstops areprovided,topermitcompliance. A "wheel" containingfourfilterswill beinsertedinthe ATS optical train 

intheeyepiece. Thecrewwill selectfiltersmanually. 

Theoptical systemof the VDP employs acook triplet projectionsystem 

anda high-resolution high-gainscreencapableof providingtherequired 

resolution. GE and Learanalysisispresented to demonstratecompliance. 

ral i z 

8 0. 

co to, 

4 4 

REQUIREMENTS 

Theslewcapabilityof theacquisitionscopes shall satisfythe followingrequirements: 

Slew 

j- * tseconds, where A 9 

istheLOS angle(degrees) betweenthe 

currenttargetand thesubsequenttarget 

atthestartof slew. 

Mechanical stops shall beincorporated on 

each ATS scannertoprevent damagetothe 

ATS trackingmirror intheeventof a worst-caseslewsignal failure. 

Meansshall beprovided foracrewmember toselectand insertanyoneof fourspectral 

orneutral density filters intotheimage 

path intheacquisitionoptics within 5 se- conds. 

Thevisual displayprojector(VDP)with the associated VDP filmmodule,shall project a6. 5 inch (minimum) diameterimageonto a backlighted screen froma0. 500t0. 005 

inch diametertargetimageonthecue film inthe VDP filmmodule. Theprojector, 

whenprojectingamediumcontrastratio 

(6:1) Standard Resolving Target,per MIL-STD-150A,within2 inches of the 

centerofthescreen,shall permitaviewer, with visioncorrespondingtothe flightcrew 

standardsper AFM 160-1 Flying Class I, 

with a10X (maximum) magnifiertoresolve 

uptoandincluding228 linepairs/mm 

(Group 7-6)asreferenced tothesource 

material. 

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TABLE1. 1-1. 

S

UMMARY OF REQ

UIRE

M

ENTS AND DESIGN S

OLUTIONS (Cont) 
TABLE 1. 1-1. 

1-12 

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RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

cc 

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DESIGN SOLUTION 

(Sameaspreviouspage) 

Theprojectorwill permitviewing fromthe ATS headrests. A knobonthe VDP will permitadjustment from0 to25 foot-lamberts. An incandescent projection lampwill providethenecessaryillumination. Projection lamps 

will be interchangeableonorbit. Theprojection lampwill be turned off 

during frameretrieval (slew). 

A bluefilterwill permitthe lampintensityto bevaried with aminimum 

shift inapparentcoloronthescreen. 

a. 

Cuemoduleswill have capacity for4094 useful frames. Sufficient 

leaderfilmoneitherend will beprovided topermitanyof the4094 

framesto beretrieved and displayed. 

b. 

Theprojectorwill permit modulechangeinaboutfourseconds. Provisionswill bemadetostoremodulesneareach crewstationtoassure compliancewith theoverall tensecond requirement. 

c. 

The filmreadercircuitry hasnotbeenfinalized.GE belivesthata designcanbe developed which will permitcompliance. 

We ightproblemsexist inthe Drive K electronics. A weight figure forthe 

shroud hasnot beennegotiated with Itek. 

4) 

I C.) 

Z 

P. 

m 

L'' 4 

PPi 

4) Z 

CD 

CO 

REQUIREMENT 

With thesametargetprojected within1/4 inch of theedgeof thescreen,theviewerwith a 10 

10X (max)magnifiershall becapableof re- solvinguptoand including204. 5 linepairs/mm 

(Group 7-5)asreferencedtothesourcemater- ial. Themodulationtransfer functionof the 

VDP atone linepair/mm,asreferenced to thesourcematerial,shall beno lessthan 

0. 83. 

Thedisplayshall beoptimized forviewing fromthecrew head positiondictated bythe 

acquisitionopticseyepiece. The brightness 

ofthecuescreenshall beadjustable,contin- uously, from0 to15foot-lambertsasmea- sured within 1 inch of thecenterof the screen. Atthemaximumscreen brightness 

setting,the brightnesswithin 1/4 inch of the edgeof thescreenshall beno less than half of thatatthecenter. 

a. 

The VDP shall becapableof automatically displayinganyof4094 cue frames 

perfilmmodule. 

b. 

Thecrewshall becapableof exchanging filmmodules within 10 seconds. 

c. 

The VDP shall satisfyretrieval requirementswith 6:1 minimumcontrast 

ratiobetweenl's and 0's onthe film. 

Weights havenot beennegotiated forthe 

aquisitionoptics hardware. 

Thisweight is based uponaweightof 329 lb forthesubcontracted optical asse

mbly(Spec. 

No. EC331), notincludinggyrosand gyro electronics. The GE-AVE weightwill beadjusted at PDR toreflecttheprimecontractor/ 

Air Forceweightagreementforthesubcon- tracted optical assembly. 

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1.2 COMMENTS CONCERNING GE COMPLIANCE WITH SPECIFICATIONS 

Table 1.1-1 indicates that most of the negotiated requirements are being complied with at 

the preliminary design stage. This section presents a brief discussion of those areas 

where a flat statement of compliance can not be made in the table. 

1.2.1 POINTING ACCURACY - Refer to entry 0 in Table 1.1-1 

A detailed error analysis is presented in Section 2.3 of this report which demonstrates that 

the eight-minute pointing accuracy is generally met. A worst case set of alignment errors 

and error buildup is shown to produce 8. 2 minutes of pointing error, however. 

The analysis shows that pointing error is a function of over twenty independent error parameters, most of which are alignment uncertainties prior to boresighting and time fluctuating 

errors after boresighting. It is shown that static misalignments between the various ATS 

components do influence pointing accuracy despite boresighting. 

The determination of an attainable set of on-orbit alignment tolerance requirements depends 

to a large extent on the degree to which the structure preserves ground established alignment in the presence of pressurization changes, launch loads, and gravity field relief. 

Analysis is planned to further investigate the capability of the structure and to establish 

realistic alignment error apportionments based on this capability. A precise estimate of 

status relative to the 8-minute pointing requirement can not be made until this work is performed. In the meantime, it is felt that the analysis presented herein indicates that 

the pointing accuracy requirement will be difficult to comply with. 

GE is presently studying several approaches to improving pointing performance, including; 

a. Increasing boresighting capability to desensitize ATS pointing to misalignments between the folding mirror and the scanner. 

b. Analytically predicting on-orbit alignment, and pre-biasing the ground alignment accordingly (either in the hardware or the software). 

c. Stiffening the structure between the folding mirror and the scanner to minimize 

alignment changes. 

The effectiveness of increasing boresighting capability will be available prior to P. D.R. 

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1.2.2 MCS POINTING CORRECTIONS (Refer to entry in Table 1.1-1) 

It is required that the crew be able to correct ATS pointing accuracy to within two minutes 

of arc with the Manual Control Stick. The accuracy and speed with which a man can control 

the pointing of the ATS will depend on, among other things, the transfer function of the 

Manual Control Stick and the characteristic response of the man as a servo loop component. 

The necessary simulation has not been performed to determine either the MCS transfer 

functions or the characteristics of the man at this time,and it is thus not possible to present relevant analysis to demonstrate compliance with this requirement. 

Preliminary simulation and servo analysis, however, has not revealed any difficulties 

which will prevent compliance with this requirement. 

1.2.3 JITTER PERFORMANCE (See entry UTable 1.1-1) 

Present analysis shows that the design will meet the jitter requirement as specified in roll. 

The design will not meet the specified performance with pitch gimbal rates as low as 0.01 

deg/sec. Since these rates will not occur during the mission it is assumed that intent of 

this requirement applies only to the roll gimbal at gimbal rates of 0.01 deg/sec. The jitter 

requirement will be met in pitch for the gimbal rates to be experienced during the mission. 

1.2.4 ATS RESOLUTION (Refer to Item 19 in Table 1. 1-1) 

The optical resolution predicted for the telescope is dependent on the magnitude of the 

various manufacturing errors assumed and, to some extent, the analytical methods utilized. 

Itek has presented analysis as part of their PDR report which shows that the required resolution of 3.3 feet can be achieved with the present design if the combination of manufacturing 

error sources does not produce more than 0.152 X of wavefront deformation. They feel that 

this represents a tight, but achievable tolerance budget for the system, and that they will 

meet the requirement. 

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In view of the critical nature of this requirement GE has performed extensive independent 

analysis of the system using the optical prescription curve supplied by Itek. The results 

of this analysis are comparable to, but not identical with the Itek analysis. GE analysis of 

the prescription optical system, with no allowance made for errors, yields somewhat 

better resolution than the Itek analysis. With a worst case tolerance buildup, however, 

GE has produced one case where the predicted on-axis resolution was 3.4 feet. 

On the basis of the analysis performed to date it is concluded that the specified resolution 

can be met, but that considerable effort will have to be expended in establishing and maintaining a realistic error budget for the system. 

Visual Threshold Curve 

Uncertainties concerning the exact nature of the eye's modulation threshold characteristic 

introduces a significant element of uncertainty in the prediction of ATS performance. The 

curve determined by Otto Shade has been used in all Itek and GE analyses. Itek is preparing an experiment designed to gain some verification of the Shade curve and to determine 

the effects of dynamic jitter on resolution. 

Dynamic Resolution 

In order to better understand the degrading effects of jitter, GE has derived an analytical 

method of predicting the effects of monotonic, high frequency jitter (above 6 cps) on resolution. This work is presented in Section 2.1.3 and yields results which are similar to those 

obtained using a method applied by Itek. The analysis indicates that a dynamic resolution of 

no worse than 3.6 ft. will be achieved. 

1.2.5 CUE RETRIEVAL TIME (Refer to entry 0 in Table 1.1-1) 

Lear presented an analysis and demonstrated on a breadboard that the required retrieval 

time can be met if at least three consecutively numbered frames precede the commanded 

frame. In the event that three such frames do not exist, the projector may require an 

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additional 0.1 second for retrieval. GE interprets the negotiated requirement to apply 

only when the commanded frame is preceded by the three consecutively numbered frames. 

1.2.6 PROTECTIVE SHROUD (Refer to entry in Table 1.1-1) 

The design concept proposed by the subcontractor at their PDR was not accepted by GE 

due to excessive weight, unacceptable mechanism design, and an excessive reliance on 

pyrotechnic Caging devices. At the time of this publication both Itek and GE are independently establishing new design concepts. A review of the structural aspects of the GE 

design is included in Section 2. 5. The Itek new design has not been presented to GE 

at this time. 

1. 2. 7 WEIGHT 

Weight figures have been negotiated for all elements of the subsystem except the Acquisition 

Optics. At time of this publication GE has not yet established an acceptable weight bogey for 

the A. 0. subcontractor (Itek), and the weight bogey established for the prime K electronics 

is being exceeded. 

1. 2. 8. DECOUPLING - (See Entry ® in Table 1. 1-1. ) 

While the two degree decoupling requirement can be met, GE proposes increasing the permissable decoupling to 5 degrees to relieve hardware and equipment alignment constraints. 

At present a 1 degree error allotment is required by the position readout potentiometer on 

the pechan prism, and a 1/2 degree allotment is required by the A/D converter. Thus less 

than 1/2 degree remains for errors in the alignment of the Manual Control Stick and the 

ATS. Considering the relative motion between the consoles and LM structure during powered 

flight, this would dictate the need for on-orbit alignment of the stick to the ATS. 

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1.2.9 JITTER PERFORMANCE - (See Table 1-1, entry 0 ). 

The best data available from the gimbal encoder vendor (Wayne George) indicates that encoder noise from the 18 bit pitch encoder will exceed the level required to meet the jitter 

specification in the CEI specification. A 20 bit pitch encoder will permit the specification 

to be met, but will exceed the space envelope available for the encoder with the current 

scanner design. GE is continuing to evaluate the use of a 19 or 20 bit encoder to meet the 

requirement. Consideration is also being given to switching to a straight rate loop in pitch 

which does not depend on the encoders in the track mode. 

GE intends to comply with the present CEI jitter requirement and will discuss the alternate 

design changes which are necessary to meet the requirement with the Airforce. 

G. E. 's interpretation of the jitter requirement is discussed in section 2.2.1.1 of this report. It is believed that this interpretation is in agreement with Airforce/Aerospace intent. 

1-18 

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1.3 SUBSYSTEM DESCRIPTION 

This section presents a brief description of the subsystem hardware and functions. The 

material is subdivided as follows: 

1. 3. 1 Purpose of the Acquisition Subsystem 

1. 3. 2 Operational Modes 

1. 3. 3 Description of Subsystem Hardware 

1. 3. 4 Description of ATS Controls 

1. 3. 1 PURPOSE OF THE ACQUISITION SUBSYSTEM 

The Acquisition Subsystem provides the crew with high-resolution telescopes whose 

LOS's are computer-controlled in the same manner as the main tracking mirror. With 

these devices the crew can observe weather and activity at the selected target areas and 

can select the most desirable targets for photographic operations. By manually correcting computer commands, the crew can minimize pointing and tracking errors due to 

target location information, ephemeris prediction, attitude reference, computation, and 

servo errors. Therefore, the crew and an effective Acquisition Subsystem will eliminate 

most of those factors which limit performance in the Automatic Mode of operation, and 

increase the quantity and quality of intelligence data acquired in the manned-automatic 

mode. 

Specifically, the Acquisition Subsystem - crew combination permits the following: 

a. Target appraisal for activity 

b. Target weather evaluation 

c. Selection of optimum targets for photography 

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d. Target centering and rate nulling 

e. Stored target position parameters to be updated by crew observations. 

A schematic showing the relationships between the various elements of the Acquisition 

Subsystem and related Subsystems is shown in Figure 1.3-1. Reference to panel controls 

shown on this Figure can be found in Section 5. 

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1. 3. 2 OPERATIONAL MODES 

Detailed operational modes are still being refined. The modes described here are intended 

to summarize, in general terms, some of the ways the ATS's may be utilized. 

A. Activity Mode - The activity mode represents the primary mode of ATS operation. 

In this mode, each crewman observes pre-programmed targets for activity. The selected 

targets will be identified in target "groups", where each group will contain one primary 

target and up to six secondary targets. One target in each group will be selected by the 

computer for photography by the main optics at a predetermined time designated "decision 

time". The main optics will photograph selected targets without direct crew support. In 

selecting the target to be photographed, the computer will consider stored information concerning the targets importance and the crews inputs concerning current target activity and 

weather conditions. 

The crew will be able to indicate three levels of target visibility and activity by depressing 

one of three "voting" buttons. The ACTIVITY button will tell the computer that the target 

is visible and that an unusual level of activity exists (such as the loading of a missile in a 

silo, etc.). The INACTIVITY button will tell the computer that the target is visible but 

that no unusual level of activity exists. The REJECT button will indicate that the target is 

cloud covered or too hazy to be photographed. The activation of any of these voting buttons 

will cause the ATS to slew to the next programmed ATS target. The computer will select the 

primary target if a reject vote is not entered for this target, and if no active votes are entered for the secondary targets in the group. 

Each target will be assigned a weighted number in the computer to provide the basis for the 

computer to select the most desirable target when one or more ACTIVE votes are entered 

for secondary targets or when the primary target is rejected. 

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In some cases the computer logic might be such that an INACTIVE primary target will take 

precedence even over certain ACTIVE secondary targets. If the crew observes a level of 

activity at a given target which he feels makes photography mandatory, he can override the 

computer selection logic by actuating the PRIMARY OPTICS OVERRIDE button to cause the 

primary optics to immediately slew to the target under observation by the ATS. 

In addition to appraising activity and weather conditions, the crew will have the capability 

of updating the computer stored target position parameters. When the crew observes a 

primary or an active secondary target which is not positioned near the center of the FOV, 

he may center the target with the Manual Control Stick and actuate the POSITION UPDATE 

switch to cause the actual target position as recorded by the ATS gimbal encoders to update 

the target location in the computer. 

B. Single Operator - In this mode, one crewman performs a complete target operation, 

using first the ATS to acquire the target for weather assessment and to null rates and 

center the target (if required). The crewman then switches his attention to the visual (main) 

optics and nulls rates (if required). If the IVS is operating within spec. , the crew involvement with the visual optics would only be occasional to check on IVS performance. The 

crewman would then revert to the ATS and repeat the above process with another target. 

C. Specialist - In this mode, one crewman acquires all targets with the ATS, performing 

the same ATS operations as described for the single operator mode. The second crewman 

works only with the visual (main) optics, taking over control of the target after the main 

optics slew to the target. If the IVS is operating, visual optics operations would not be 

necessary. 

D. Leap-Frog - This mode consists of two crewmen, each operating in the single operator 

mode. Since only one crewman can use the visual optics at a time, the operations of the 

two crewmen are programmed such that while one crewman is using an ATS, the other is 

using the visual optics. If the IVS is operating, visual optics operations would not be 

necessary. 

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Although the Acquisition Subsystem can be operated in any of the four primary modes as 

described above, it is most likely to be used in the Activity Mode. 

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1.3.3 LIST OF PRINCIPAL COMPONENTS COMPRISING THE ACQUISITION 

SUBSYSTEM 

The acquisition subsystem is comprised of the following elements; 

a. Acquisition Optics (on ATS) - Includes; 

• Telescope 

• Window 

• Scanner 

• Folding Mirror 

• Peripheral Display 

• Shroud and Shroud Door 

b. Cue Projector and Alpha Numerics Display 

c. Drive K Electronics 

d. Cue Module Storage Rack 

e. Magnification Control Stick 

f. Head Rest 

g. Gyroscopes 

The manual control stick is considered part of the N&C subsystem. 

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1.3.3 DESCRIPTION OF SUBSYSTEM HARDWARE 

1.3.3. 1 Acquisition Optics 

The Acquisition Optics (or ATS) consists of the optical elements of the subsystem and 

their supporting hardware. All elements of the Acquisition Optics are supplied by Itek 

under subcontract to GE, and are shown schematically in Figure 1.3-2. 

1.3.3.1.1 Telescope Assembly and Optical Performance Description 

The heart of the acquisition subsystem is the Acquisition Telescope whose size and 

elements are shown schematically in Figure 1.3-3. This component consists of a 

10-inch objective, refractive, high power, telescope mounted within the pressurized 

Lab Module of the spacecraft. It presents a magnified ground scene to the crew 

through an eye piece with a 60-degree apparent FOV and a 2 mm (minimum) exit pupil. 

Two such telescopes are included in each spacecraft as shown in Figure 1.3-4. 

The peak magnification of the telescope is 127 power. At this power, the telescope will 

have a field of view of 0.5 degree and will view approximately 0.7 nautical miles on the 

earth at a slant range of 80 nautical miles (nadir). 

The on-axis resolution of the telescope from 80 mu is predicted to be about 3.6 feet when 

operating in orbit in the presence of a low level of predictable "jitter" caused by servo 

noise. The static resolution of the telescope is about 3.3 feet. 

The scanner (to be discussed later) is designed to permit the LOS of the telescope to be 

directed through a scan field from 70 degrees forward to 40 degrees aft in pitch, and from 

-45 degrees to +45 degrees in roll. In operation it will be possible to lock on and track a 

target at something less than a full 70 degrees forward in pitch. 

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There will be no vignetting in the optical system of the acquisition optics from +10 degrees 

to +60 degrees in pitch and from -35 degrees to +35 degrees in roll. Outside of this region, 

vignetting will occur in the external optical elements. 

The magnification of the telescope can be continuously varied from 16 to 32 power, and 

from 63 to 127 power, by means of two independent mechanisms as described below. 

The magnification of the telescope can be reduced by a factor of four by inserting a set of 

low power lenses into the optical train. These lenses reduce the effective entrance aperture to 2.5 inches; they do not change the exit pupil diameter. The lenses are inserted and 

removed by an electric motor on command from the flight crew via the Magnification Control Stick and the Drive K Electronics. A manual override exists to permit operation of the 

power lenses in the event of control system failure. 

The magnification of the telescope can vary continuously over a two-to-one range by means 

of a zoom mechanism. This 2:1 change can be obtained with the low-power lenses either 

in or out of the optical train. The zoom is part of the eye piece optical train and its operation affects the exit pupil size, causing it to increase to 4 mm at one-half maximum power. 

The zoom magnification changes are effected by two sets of lenses which are axially positioned in the telescope by an electric motor. The relative movement of the two sets of 

lenses is controlled by a cylindrical cam which in turn is driven by an electric motor. The 

commands for the zoom mechanism are originated by the crew at the Magnification Control 

Stick. Electronic controls for the zoom are contained in the Drive K electronics. A manual 

backup is provided to permit control of the zoom mechanism in the event of automatic 

control failure. 

A pechan prism is contained in the optical train to permit the target scene to be rotated to 

present the crew with a scene whose in-track component is always vertical in eyepiece 

FOV at the end of slew. The prism is prevented from rotating during the tracking of a 

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target so as not to introduce jitter into the scene. The scene will thus rotate slightly in 

the eyepiece as the target is tracked. Commands generated by the manual control stick 

will be decoupled in the software to remove the effects of this scene rotation such that 

vertical displacements of the stick always cause vertical scene displacements. 

The pechan prism is driven by a control system in the Drive K electronics in accordance 

with commands generated in the computer. A manual backup will permit setting the prism 

to a zero reference position in the event of mechanism failure. 

A solar blanking shutter and associated sun sensor are included to protect the crew from 

inadvertant viewing of the sun. This mechanism will be controlled by the Drive K Electronics in such a manner as to prevent any portion of the solar disk from entering the FOV of 

the ATS when the gimbals are slewing at any speed up to the maximum gimbal rate. After 

closing, the shutter will be re-cocked (opened) manually by the crew. 

A filter wheel is included in the telescope to permit the crew to insert any one of four 

filters into the optical train to improve optical performance. 

The telescope will not require any refocusing when the target range lies anywhere within the 

scan field for spacecraft altitudes from 75 to 230 nm. The eyepiece of the telescope will be 

focusable over a range from -2.5 to +3.5 diopters to permit individual crewmen to adapt 

the telescope to their eyes. 

The telescope eyepiece includes a reticle with cross hair references to locate the center of 

the FOV, and a circle to reference the size of the real field of view attainable at maximum 

zoom magnification. 

The telescope eyepiece also includes a peripheral display to be described later in this 

section. 

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The telescopes are mounted in the spacecraft by a rigid mount to the pressure shell at the 

objective end. The objective lens serves as a backup for the window which provides the 

primary optical feedthrough in the pressure shell. A shear mount near the elbow in the 

telescope provides additional mounting restraint. The eyepieces for the two telescopes 

are brought through consoles 2 and 8 for crew observations. The telescope can be broken 

down into three pieces at "field breaks" to facilitate installation in the Lab Module. Precise alignment of the telescope will not be required; the external folding mirror will provide the necessary alignment adjustment. 

Peripheral Display 

The Acquisition Telescope contains a set of peripheral display lights arranged in the eyepiece as shown in Figure 1.3-5. The brightness of the lights in the display can be controlled 

from a single panel control. 

The operation and function of the display groups is presently planned as described below. 

a. Group I - Time Reference Display - This group of 25 lights will convey information to the crew concerning: 

1. The number of seconds remaining to the next decision time (up to 25 seconds). 

2. The number of targets remaining to be evaluated in the group. 

3. The ground estimate of the point in time at which the crew should have released 

each target (by voting). 

At the beginning of a new target group, a light will be lit in the display at a position 

indicative of the time remaining to decision time (that is if 24 seconds remain, the 

24th light will be activated). The position of the activated light ("bouncing ball") 

will be advanced one position in the display each second. In addition the first light 

in the display will remain lit during the countdown to serve as a zero reference 

point. 

In addition to the time reference countdown described above, one light will be 

activated in Group 1 for each target which remains to be evaluated in the group. 

The position of each "marker" light in the group will be indicative of the time by 

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APPARENT FIELD 

OF VIEW 

GROUP 2 

5 LIGHTS 

GROUP 1 

25 LIGHTS 

GROUP 3 

5 LIGHTS 

Figure 1. 3-5. Peripheral Display Pattern 

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which ground personnel recommend that a vote be entered to release the target. 

As each target is released, its marker light in the display will be extinguished, 

while the time wipeout countdown continues. 

b. Group II and DI - These two 5-light groups will be controlled by the computer to 

convey information on target priorities, voting instructions, etc. Final determination of the precise information to be conveyed will be made later. The switching 

control for these lights will permit them to be individually controlled. 

The lamp drivers for the peripheral display lights are located in the console controller. 

1.3.3. 1.2 Window - The window provides the optical feedthrough through the Lab 

Module pressure shell. It is mounted in the pressure shell in a DAC-supplied 

penetration fitting which controls the temperature of the circumference of the 

window within ±2 degrees. 

1.3.3. 1.3 Folding Mirror (Fixed) - The folding mirror is located external to the spacecraft and functions to fold the telescope LOS parallel with the roll axis. The 

mirror provides alignment adjustments in its mount to permit accurate alignment of the LOS between the folding mirror and the scanner. The folding 

mirror is mounted to the same penetration fitting which supports the window 

and the telescope to minimize relative alignment changes between these components. 

1. 3. 3. 1.4 Scanner - The scanner provides the gimballed flat mirror which permits the 

telescope line of sight to be directed within the prescribed scan field. It consists of the scanner mirror, supported on two orthogonal gimbals, and associated gyros, bearings, encoders, and alignment references. 

The outer axis of the gimbal is the roll axis of the scanner and is mounted 

such that it is canted 9 degrees from the spacecraft roll axis in the horizontal 

(X-Y) plane. This gimbal cant minimizes the occurrence of zero roll rate 

conditions during target tracking and thus minimizes jitter due to bearing 

stiction. 

Incremental shaft encoders are mounted on both gimbal shafts. The incremental pulses are accumulated in the Drive K to generate shaft position 

signals for the computer and the Drive K servo loop. 

The gimbals are driven by brushless dc torquers which are controlled by servo 

electronics in the Drive K in response to computer-generated commands. 

Gyros, mounted on each gimbal, are an integral part of the rate loops in each 

servo control system. Energy-absorbing stops limit the gimbals' movement. 

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The bearings supporting the gimbals are the object of a continuing development program to insure smooth jitter-free tracking by the gimbals. 

The scanner is mounted to the pressure shell on a three-point mount. Alignment references will be provided to permit accurate installation. 

1.3.3. 1. 5 Shroud and Shroud Door - The external components of the ATS are enclosed 

in a hinged shroud which provides thermal control and protection from contaminants when the ATS is not in use. When closed, the shroud will compress 

a seal to insure that contaminants from the thrusters and the waste management system do not degrade the optical quality of the mirrors. 

Thermal control of the shroud and its internal components is achieved passively. The shroud is opened and closed by redundant electric motors, controlled via the Drive K electronics in accordance with commands generated 

either by the computer or the flight crew. The shroud door drive will be 

interlocked to prevent closing when the scanner is not in its stow position. 

1.3.3.2 Drive K Electronics 

Most electronics associated with AO operation will be in a single package called the Drive 

K Electronics. The functions of the Drive K Electronics are categorized in ten groups, 

designated K)hrough K10, and are described briefly here. The K designations are 

utilized in non-classified documents to refer to the functions listed. A more detailed 

description of the controls is presented in paragraph 1.3. 4. 

KO - Digital Processor - Processes digital signals from the MDAU for the Drive K. 

K1 - Gimbal Control - The K1 electronics receive inputs from the MDAU and the 

MCSA (in backup mode) and provides drive signals to gimbal gyro torquers. 

K2 - Gimbal Control - The K2 electronics receive gyro rate signals from the 

gimbal gyros and process them for application to the gimbal torquers. 

K3 - Derotation - The K3 electronics receive derotation signals from the MDAU/ 

computer, and the analog prism position pickoff, and generates the drive 

signals for the derotation prism in the AO. 

K4 - Zoom - The K4 electronics receive inputs from the Magnification Control 

Stick and the analog zoom position transducers and generate the drive signals 

for the zoom motor in the AO. 

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K5 - Power Change - The K5 electronics receive signals from the Magnification 

Control Stick and generate signals to control the power change lens in the ATS. 

K6 - Solar Blanking - The K6 electronics receive analog signals from the AO sun 

sensor, provide threshold detection, and generate the necessary signals to 

actuate the solar blanking mechanism. 

K7 - Door Control - The K7 electronics receive signals from the control console 

door switch and the door limit switches, and generate a signal to drive the external AO door. 

K8 - Power Supply - The K8 electronics receive unregulated power from the OV 

power system, and generates necessary regulated DC voltage. 

K9 - Telemetry - The K9 electronics perform signal conditioning necessary for 

telemetry. 

K10 - Monitor and Failure Switching - 

1. 3. 3. 3 Visual Display Projector - Design and Description 

The Visual Display Projector is primarily intended to aid the crewman in locating targets 

in the ATS field of view by cueing him with high-resolution pictures of the scene. It is 

intended that the crew review the cues for the targets assigned before each target pass 

(prepass briefing), and that he be briefly presented a cue for each target during the pass 

while the ATS is slewing to the target. A secondary function of the VDP is to provide 

maintenance or instructional material to the crew. 

In order to satisfy the mission requirement the VDP will be capable of rapidly retrieving 

and displaying cue frames from 16mm strip film. The retrieval time will be less than 

0.45 + N/100 seconds where N is the number of frames traversed in retrieving the cue 

(provided that at least three consecutively numbered frames precede the desired cue). 

The resolution of the VDP will permit 228 1p/mm material on the film, with a 6:1 contrast 

ratio, to be visible on the screen with the aid of a 10-power loupe. 

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During target pass operations, cues will be called up by computer commands which are 

held in the console controller and presented to the VDP as parallel binary words. Each 

cue frame so retrieved will be held for a short fixed duration and then released to permit 

the projector to retrieve the next cue. The operation will be planned so that the cue is being retrieved during the periods when the astronaut is viewing an ATS scene, and displayed 

during the slew period (perhaps with some overlap into the track period). The cue will be 

released automatically if the crew votes on the target while it is still being displayed. The 

crew can retain a cue frame beyond the planned release time by actuating a cue hold control 

on the magnification control stick. 

The VDP is located on the console in a position to permit the crew to view the projected 

scene from the ATS head rest position to maximize usefulness during the target pass. 

The 6 1/2-inch rear projection screen has a high gain to provide minimum screen resolution 

degradation and maximum illumination efficiency. 

During prepass briefing periods, it is intended that the computer automatically command 

each frame retrieval as in the target pass operation. The cue would be released and the 

next cue displayed, however, by the crew by operating a manual control (probably the 

REJECT button on the magnification control stick). 

Cue frames can be called up manually by dialing the desired frame number, in octal form, 

on the manual selector control on the face of the VDP panel. 

Film for the VDP is stored in special modules which can be quickly inserted or removed in 

the VDP front panel. Schematics of the projector and film module are shown in Figures 

1. 3-6 and 1.3-7. 

Alpha-Numeric Display 

A six-character alpha-numeric display is included on the VDP to project cue words relating 

to each ATS target. The information for these frames is stored in the computer, not on 

the film. 

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Figure 1.3-6. Visual Display Projector Layout 

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Operational Description 

The Visual Display Projector has the capability of displaying any cue frame in a library of 

32,752 frames. These cue frames will be stored in 8 film modules, each containing 4,094 

frames. Each frame is uniquely identified by a binary code which is printed on the film. 

When the frame called for is not on the film, the servo will recognize this and stop to prevent damage to the film. 

The beginning of the film reel is identified by a series of frames identified by all zeros, 

and the end of the film is identified by all ones. In the event that the frame called for 

appears to be off the end of the reel, the servo will recognize the end of the reel and stop 

to prevent damage to the film. 

Drawing 47D405106 (Figure 1.3-8) is the schematic arrangement for the Visual Display 

Projector. 

When the Off-Auto-On switch is in the Auto position, the projector is in the automatic mode. 

In this mode, the projector will accept binary commands from the computer via the console 

controller, and in the absence of commands from the console controller it will accept octal 

commands from the manual frame selector. 

When the Off-Auto-On switch is in the On position, the projector is in the manual mode and 

will accept commands from the manual frame selector only. The console controller input 

is inhibited. 

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In the Automatic mode, inputs from the console controller will command a particular cue to 

be displayed. Upon receipt of this command, the projector will read the code of the cue 

frame in the film gate and compare it to the new input code. As a result of comparison, the 

logic will determine the direction the film should be moved and the speed. If the command 

frame is more than three frames away from the frame in the film gate, the servo drives the 

film at 100 frames per second, until the frame which is three frames prior to the commanded frame is reached. At this time, the servo dynamically brakes until a speed of approximately 10 frames per second is reached. The film travels at this speed until the binary 

code of the frame in the film gate corresponds to the commanded binary code. When this 

occurs, the servo switches to fine position and centers the frame in the film gate. The 

motors are de-energized, the platen presses the film against the prism face to insure flatness, and the projection lamp is turned on. If the commanded frame is three frames or 

less away from the frame in the film gate, the servo drives the film at the low speed until 

the commanded frame is in the film gate. The sequence is then the same as above. 

In the manual mode, the thumbwheel switch is set to the octal number that describes the 

frame to be displayed. The octal-to-binary converter converts this signal to binary code. 

The execute button injects the binary code into the signal control and the servo handles this 

input in the same manner as an input from the console controller. 

In either the manual mode or the automatic mode, the astronaut may select the frame 

immediately ahead or behind the projected frame by depressing the forward-reverse switch 

in the proper direction. This command is overridden by a new input from either the manual 

frame selector or the console controller, depending on the mode of operation. 

When the VDP is energized, the alpha-numeric circuitry will accept binary commands. 

Each readout segment is controlled by a 5-bit binary code and is capable of displaying any 

number, 1 through 9, or any letter, A through Y, with the following letters missing: 0 I S Q. 

The display remains illuminated as long as the command is present. 

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The outline of the Visual Display Projector is shown in Drawing 47D405104 (Figure 1.3-9). 

The screen size and location, the controls provided, and their description are shown on the 

drawing and need no further description. The test switch, the 2X/1X switch, frame missing 

light, and end of reel light have been deleted since the preparation of this drawing. In addition, the location of the Visual Display Projector is being changed in the console, and the 

change in relocation will require a complete repackaging of the Visual Display Projector. 

The outline drawing presented here represents the previously defined baseline design with 

the Visual Display Projector subcontractor. 

1.3.3.4 Cue Module Storage Rack 

16-cue modules will be stored in a single-cue module storage rack in the Lab Module. 

Detailed design of this rack is not complete. It is presently intended that this rack be 

located in console 8A. During the launch phase, the cues will be stored in the Gemini. 

1.3.3.5 Magnification Control Stick 

The Magnification Control Stick will be designed for operation by each crewman's left hand. 

Its functions will be to: 

a. Provide the means to electrically select any of four incremental magnification 

levels for the visual optics. 

b. Provide a signal to switch the manual control stick between the Main Optics and 

the Acquisition Subsystem. 

c. Provide two separate continuous control ranges for controlling ATS zoom magnification in the high-and the low-magnification ranges, respectively (63 to 127X, 16 

to 32X). 

d. Provide a switch for commanding magnification step changes to the ATS. 

e. Provide three finger-operated voting switches for target REJECT, ACTIVE, and 

INACTIVE votes on the stick handle. 

f. Provide a control for commanding CUE HOLD by twisting the stick handle. 

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The Magnification Control Stick will move in a two-range track on the console as shown in 

Figure 1.3-10. One range will provide four discrete positions corresponding to four discrete VO magnification levels. The other range shall provide two continuous ATS zoom concontrol ranges separated by a detent corresponding to the position at which a magnification 

step change is commanded. The two ranges (ATS and VO) are separated by a switch to 

signal the computer which subsystem the manual control stick controls (Acquisition or Main 

Optics). 

Visual Optics Control 

It will be possible for the crew to select any of the four visual optic magnification levels 

without commanding intermediate power levels (i. e. , it will be possible to switch from 

1000X to 125X without going through the intermediate steps of 250 and 500 power). Detents 

or other means shall be provided to permit the crew to position the stick by "feel". 

A functional block diagram of the magnification stick is shown in Figure 1.3-11. A detailed 

design of the stick has not been completed. 

1.3.3.6 Head Rests 

Head rests will be provided at each console to aid in stabilizing the observer's head in the 

head rests. Detailed design of the head rests has not begun and will ultimately be constrained by simulation results and flight crew preference. The following general guidelines are considered relevant: 

a. The head rest will provide maximum crew comfort over the duration of full target 

pass. 

b. It will provide stable, comfortable positioning of the crewman's head to allow 

viewing through a 2-mm AO exit pupil in a zero-g environment. 

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-HOW- SPECIAL HANDLING 

MAGNIFICATION 

CONTROL 

STICK 

PROVIDES 

1) V.O. MAG. SELECTION 

2) ATS ZOOM CONTROL 

3) ATS MAG. CONTROL 

4) CUE HOLD 

5) TARGET VOTING 

6) DIRECTION OF MCSA 

TO ATS OR M.O. 

 TO TIM 

• CUE HOLD STATUS 

• MCS STATUS 

TO DRIVE K 

• DISCRETE MAG. 

STEP COMMAND 

• ZOOM CONTROL 

SIGNAL 

TO CONSOLE 

CONTROLLER 

• V.O. MAG. LEVEL 

SIGNAL 

• TARGET VOTING 

SIGNALS 

• ATS MAGNIFICATION 

STATUS SIGNAL 

CREW 

INPUT 

Figure 1.3-11. Functional Block Diagram, 

Magnification Control Stick 

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V. 0. CONTROL 

RANGE 

127X 

63X 

32X 

16X 

Figure 1.3-10. Magnification Control Stick Slot Configuration 

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ATS CONTROL 

RANGE 

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NRO APPROVED FOR 

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c. It will maximize the crew's ability to scan the FOV at all magnification and pupil 

sizes produced by the AO. 

d. It will provide sufficient adjustment to allow either eye to be used and to allow 

each crewman to position his head in accordance with the AO eye relief. It shall 

be adjustable with a gloved hand. 

e. It will prevent contact of crew head with ATS eyepiece. 

f. It will not interfere with the crew's ability to move quickly from the ATS to the VO 

eyepiece. 

1-52 

SECRET SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

1.3.4 ATS CONTROLS DESCRIPTION 

1.3.4.1 General Functional Description 

The functions of S/S Alpha Controls will be described with reference to Figure 1.3.4-1. 

The primary functions of S/S Alpha controls are to provide: 

a. Proportional control of the ATS two axis gimballed LOS scanner mirror in 

response to computer generated digital steering commands (Functions KO, Ki, 

and K2) 

b. Bi-valued position control of an environmental protective shroud door in response 

to computer controlled (primary) or manually initiated commands originating on 

the control console (Function K7) 

c. Proportional position control of an image orientation prism in the ATS telescope in response to a computer digital command (Function K3) 

d. Proportional position control of a zoom lens assembly on the ATS telescope 

in response to signals originating from a manually operated zoom/magnifi- 

-- cation controller. (Function K4) 

e. Bi-valued position control of a magnification lens assembly on the ATS telescope to select a high or low range of magnification in response to a command 

from the zoom/magnification controller (Function K5) 

f. An ATS telescope solar blanking mechanism which will protect the astronaut 

against viewing the sun through the telescope (Function K6a) 

1.3.4.2 Detailed Functional Description 

The detailed functional description of the S/S Alpha controls is presented in the following 

paragraphs: 

1.3.4.2.1 Scanner Control 

The functional block diagram of the scanner controller is presented in the top portion of 

Figure 1.3.4-1 and, in greater detail, in Figure 1.3.4-2. The scanner control operates 

on separate inner and outer axis digital rate command signals received from either MDAU 

No. 1 or No. 2. 

1-53/1-54 

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1. 3. 4. 2. 2 Inner Axis (IA) 

For the Inner axis, control is provided by either of two configurations. In both modes the 

IA rate command is integrated at a 19.5 KC rate and stored in the command register. 

The contents of this register is compared to the "whole-word" gimbal feedback information 

formed in the aui feedback register from the incremental data generated by the IA gimbal 

encoder. The error signal, Ea_ is sent to the K1 D/A converters. ulA 

In the primary mode the error signal is used as the input to a rate minor loop formed 

around the IA Rate Gyro Package (RGP) mounted on the IA gimbal. Functional block K1aIA 

provides the D/A conversion, compensation and gyro command torquer drive. 

Functional block K2 IA operates on the gyro output signal and provides the compensation 

and the torquer motor drive capability. 

In the back-up configuration, the rate minor loop is not required and the digital error signal 

is converted and compensation is supplied by functional block K1bIA. The compensated 

signal directly drives the IA torque motor power amplifier. In both primary and back-up 

configurations, the position feedback data is read out to the computer (via the MDAU) from 

the 6IA readout buffer. 

1.3. 4. 2. 3 OUTER AXIS (OA) 

Two control loops are also available in the OA. The primary configuration operates directly 

on the rate command word from the 9

0A register. This rate command is converted to an 

analog signal, compensated and conditioned to drive the OA rate gyro torquer in functional 

block K1a0A. 

The gyro output signal is then compensated and applied to the OA torquer motor driver in 

functional block K2OA. In the back-up mode the integrated rate command in the 00A register is compared with 90A feedback register to produce a position error, ceoA, as in 

1-59 

1- SPECIAL HANDLING 
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RELEASE 1 JULY 2015 IfeRET-- SPECIAL HANDLING 

the IA loops. As in the back-up IA loop the OA position error is converted, compensated 

(in K1bOA), and applied directly to its torque motor amplifier (via K2OA). 

Also as in the IA loop, the 610A feedback data is read out to the computer via the MDAU's, 

in both primary and back-up modes, from the 8/A & OAreadout buffer. 

1. 3. 4. 2. 4 Operational Control 

On orbit, during normal operation, certain scanner functions are automatically carried. 

These include power turn on and off, initialization and synchronization of digital registers, 

in KO and stowing of the scanner. In addition, an interlock with the shroud door (K7) is 

provided to prevent either moving the scanner from its stowed position while the shroud 

door is closed (by interrupting unregulated power to K2 until door is fully open), or closing 

the shroud door with the scanner not in its stowed position (by K2 interlock signal to K7 

input logic). The signal flow for these functions can be seen in Figure 1.3.4-1. 

Launch lock release is manually actuated from the Operational Panel. In the present baseline design the scanner and environmental launch lock employ electro-explosive devices 

(EED's) to supply the force to remove locking pins. (Studies are under way to eliminate 

EED's.) 

Both scanner and environmental door are released by operation of switches on the Operational Panel, employing the primary and backup actuation circuits in sequence. An EED 

circuit enable switch (also on the Operational Panel) must be actuated in conjunction with 

the release switches, as a safety precaution. 

Contingency control, by redundant command signal paths, can be initiated by manually 

actuated switches on the Overrides and Status Panel to perform the following functions: 

a. Stow the scanner 

b. Synchronize the position feedback registers 

1-60 

---SEGRET- SPECIAL HANDLING 
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c. Remove power manually from torque motor amplifiers (to hold scanner position 

while switching configuration) 

d. Perform an operational check of the scanner 

e. Select the backup loop configuration 

f. Select backup encoder electronics 

A description of the Initialization and Synchronization, Stow Procedure and the Scanner 

Operational Check is given in the following paragraphs: 

Initialization and Synchronization (See Figure 1.3.4-2) 

These functions serve to minimize the "turn-on" transient (initialization) and force agreement (synchronization) between the feedback registers and the gimbal positions (IA and 

OA) initially, and periodically throughout the active period. 

Initialization begins immediately after power turn-on or when commanded manually from 

the control console and results in blocking the normal rate commands; and setting the rate 

command registers (euREG and OOA REG) to zero; the OA position register (90AREG) 

to zero; the IA position register (611AREG) to approximately 11.25 degrees; and the application of a constant rate command (via the IA and OA D/A converters) to drive the IA and OA 

axes toward their respective "synch" positions, (0° for OA, and 11.25° for IA). As each 

gimbal reaches its "synch" position the constant rate drive in that axis is removed, and 

when both gimbals have reached their respective "synch" positions the normal rate commands are again admitted. 

Synchronization takes place in each axis independently, and occurs not only as part of the 

initialization process but whenever the gimbals pass through their respective synch positions during the periodic slews. Synch is inhibited from occurring during the "rate" mode. 

1-61 

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Synchronization is accomplished by resetting the position feedback registers (e0Aand 

e

IA FDBK REG) to 0° and 11.25°, respectively, as the associated gimbal passes through 

its synch position. 

For manual initizalization, which is employed in a contingency mode, the synch pulse in 

each axis is stretched out to approximately one second duration and used to momentarily 

light indicators on the Overrides and Status Panel as an indication that initialization was 

successful. The initialization and synchronization function described above are implemented 

by the various gate, pulse, drive and word generators and register input logic shown in 

Figure 1.3.4-2, operating on the sector signals generated by the OA and IA shit position 

encoders. 

1.3.4.2.5 Stow Procedure 

The stow operation consists of commanding the scanner gimbals to a position which will 

permit the environmental shroud door to be closed. The scanner will be maintained in this 

position, with power off, until the shroud door is opened at the commencement of the next 

operating period. A non-power consuming latching device will be supplied, if necessary, to 

hold the scanner in the stowed position. As shown in Figure 1.3.4-2, the present baseline 

selection for this holding device is a permanent magnet. 

Stowing is initiated by the computer, via the MDAU's, at the end of an active period, or by 

the manual actuation of a switch on the Override and Status Panel. On the receipt of a stow 

command the normal rate commands are blocked and a 5°/sec rate is introduced in both 

rate command registers (80A and 61A REG's). As each gimbal reaches its stow position, 

the stow rate command to that axis is terminated. The normal rate commands, however, 

remain blocked, and effectively zero, rates are commanded, as long as the stow command is 

continued. 

Stowing is implemented by employing the stow commands to the rate register input logic to 

block the normal commands and to set up to IA and OA "ST" Gates. These gates are reset 

1-62 

SECRET SPECIAL HANDLING 
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by the stow position signals (CW Limits) from the encoders. The "ST" gates go to their 

respective rate command registers to gate in the stow rate command until the stow positions are reached. An indication of stow for each gimbal is provided on the Override and 

Status Panel. 

1.3.4.2.6 Scanner Operational Check 

An operational check of the scanner can be quickly performed with the controls and displays 

available on the Override and Status Panel, and a stop watch. The first step is to manually 

initiate synch by means of the manual synch switch. Completion of manual initialization 

is signaled by the synch indicators. 

The next step is to manually initiate the stow procedure and time the period elapsed before 

the stow lamps are lit in each axis. Knowledge of the stow rate and the gimbal displacement 

between the synch and stow locations in each axis provides the astronaut with the nominal 

reference times, for a properly operating system, with which to compare his measured 

values. 

1.3.4.2.7 Image Orientation Control (Function K3) 

The functional features of the Image Orientation Control is shown in Figure 1.3.4-3. The 

control consists of a proportional position servo loop operating on a digital step position 

command received from the computer through the MDAU's. The input is converted to an 

analogue command signal and lagged by the shaper transfer function before being compared 

to analogue position feedback signal. The shaper function limits rate of build-up of a step 

command to reduce the tendency to saturate the following amplifier stages. (Saturation 

would reduce the loop response time.) The Summer and Power amplifier generates the 

error signal, provides compensation and drives a pechan prism to orient the image. 

Image orientation is made to some angle predetermined by the computer for each slew, and 

executed during the slew period. In the tracking mode the control loop is opened to immobilize the prism. This function is performed by solid state switching in the power amplifier 

1-63/1-64 

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stage responding to the position/rate synch from KO. The position feedback signal is read 

out to the computer, to be used in conditioning the MCSA transfer function with respect to 

prism position. 

The pechan prism can be manually positioned in case of failure of the electrical control 

loop. The electrical drive can be disabled by a switch provided on the Override and Status 

Panel if it becomes necessary to use the manual back-up. The electrical disable control 

operates through the LMMC to cut-off unregulated power to the output stage. 

1.3.4.2.8 Zoom Control (Function K4) 

The Zoom control is functionally depicted in Figure 1.3.4-4. The zoom function employs 

a proportional control loop with an analogue command signal derived from the zoom/mag 

stick on the Operational Panel and an analogue position feedback signal. Shaping is provided on the command signal to minimize saturation and resulting loss in response time, 

in the event of a steep-sloped input. 

A signal indicating high or low zoom magnification setting is generated from the position 

feedback signal to be used by the computer in adjusting MCSA sensitivity. This signal is 

transmitted through the MDAU's. 

Manual zoom positioning is provided as a back-up to the servo loop control. 

A manually operated switch, located on the Override and Status Panel is used to disconnect 

the unregulated power to the output stage to disable the motor drive during manual operation. 

1.3.4.2.9 Magnification Control (Function K5) 

The K5 functional block diagram is shown as part of Figure 1.3.4-1. In operation, a constant voltage is applied to the load actuator, to drive the lens assembly to the commanded 

position. A limit switch removes the drive when the attained position agrees with the 

command. The command is bi-valued, corresponding to lens assembly "out" (for high 

1-67/1-68 

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magnification power) or "in" (for low power) of the optical path. Position accuracy in 

the optical path is established by a cam configuration that assures the lens is firmly positioned against its precisely located stops before a limit switch signal is generated, and the 

incorporation of a brake that is applied when•.power is removed. The state of the limit 

switches is also sent to the computer (via the MDAU's) for adjusting sensitivity of the MCSA. 

Commands to the lens assembly are inhibited during scanner slew. Manual positioning of 

the lens assembly is employed as a backup. Disconnection of power to the power switch 

during manual operation is accomplished by actuating a switch on the Override and Status 

Panel. 

1.3.4.2.10 Solar Blanking (Function K6a) 

The solar blanking control functional block diagram is shown as part of Figure 1.3.4-1. 

The solar blanking mechanism is a lightweight shutter that is removed from the optical 

path when an actuation spring is wound up and latched. The shutter will be manually cocked 

(reset) by the crew. A "sun-presence" will light an alarm indicator on the Monitor Alarm 

Panel 

1.3.4.2.11 Cue Insertion (Function K6b) 

Space provisions are included in the telescope for a cue insertion mechanism in the event 

such a feature is desired. The cue insertion mechanism would drive a mirror into the 

optical path to allow cue information to be presented to the astronaut in place of the ground 

image. The functional block diagram of K6b shown in Figure 1.3.4-1, represents a possible configuration for this control if it is added to the subsystem. This function is not a 

part of the baseline design. 

1-71 

SPECIAL HANDLING 
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1.3. 4.2. 12 Shroud Door Drive and Launch Lock Release Actuation (Function K7) 

The functional characteristics of the shroud door control is shown as part of Figure 1.3.4-1. 

a. Shroud Door Control - The Shroud door is driven by two motors, one placed on 

each end of the door (forward and aft). In the primary mode of operation the 

command to open or close the door is received from the LMMC (shown as the 

K7 command) and is applied to the drive logic of both aft and forward motors. 

A separate set of limit switches is associated with each end of the door and is 

used to cut off the associated motor drive when its end is positioned as commanded. 

The input logic also incorporates an interlock signal which prevents the door from 

being closed unless the scanner is stowed (K2 interlock). The scanner stowed 

signal is derived from the scanner encoder position limit and sector signal. 

(Shroud door launch locks are also provided and are released as described in 

paragraph 1. 3. 4. 2 . 4. ) 

For back-up to the primary mode of operation, the following functions are provided: 

1. A manually initiated switch signal from the Override and Status Panel can 

operate the doors in place of the normal command from the LMMC. 

2. Power can be removed from either motor by switches on the Override and 

Status Panel in case of an electrical control loop failure, permitting control through the remaining motor and control loop. 

3. Either motor can be mechanically disconnected from the drive shaft by 

manually actuating a switch on the Override and Status Panel, in case of 

binding of one of the motor shafts. 

4. A spring incorporated on the door drive mechanism can be released to open 

the door after mechanically disconnecting both motors. (Note that in the 

current baseline configuration the disconnects and door release function of 

(3) and (4) are accomplished by EED actuated pins, and for safety reasons 

an EED circuit enable switch is also provided. Currently a study is being 

made to reduce the number of/or eliminate completely, these EED's.) 

5. A continuous position indicator is also provided on the control console to aid 

in determining which motor must be mechanically disconnected in case of a 

binding shaft. The position signal is derived from one end of the door so 

that the response, or lack of response of the position indicator to a door 

drive command determines which side is binding. 

SECRET SPECIAL HANDLING 

1-72 
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1.3.4.2.13 Power Conditioning and Distribution 

A functional diagram of the power conditioning and distribution is shown in Figure 1.3.4-5. 

The general features of the arrangement shown are dictated by the following considerations: 

a. Power turn on will be accomplished on computer commands to the LMMC via the 

MDAU's. 

b. flexibility to turn on gyros separately for peak power management. 

c. Two separate power paths. Primary and Secondary Power Generation and Routing. 

1. Primary power for primary scanner, shroud door and launch lock control 

2. Secondary power for all remaining functions. 

This separation will prevent propagation of electrical failures through a common 

power path, from the functions in path (2) to function in path (1). 

This is advantageous since all functions in (2) can either be performed without 

electrical power, in a back-up mode, or are part of back-up provisions for the 

path (1) functions. 

d. Separate control of unregulated power to individual functions to allow overload, 

isolation, failure mode switching, and interlocking on an individual function 

basis. 

An additional feature which is shown in Figure 1.3.4-5, but is still tentative, is 

the inclusion of fuses on the regulated power inputs to the functions of path (2). 

A possible reliability penalty may be tolerable for these functions since they have 

function manual back-up. 

The advantage lies in the ability to operate in a normal mode with the remaining 

functions after one function fails in a manner which overloads its conditioned 

power lines. 

The functions of path (1) will not be fused, since reliability of these functions 

cannot be compromised. 

1-73/1-74 

SPECIAL HANDLING 
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SECTION 2.0 

ATS ANALYSIS 

This section contains analysis relative to the design and performance of the acquisition 

optics components. The material is subdivided as follows: 

2.1 ATS Optics Analysis 

2.2 ATS Controls Analysis 

2.3 ATS Performance Error Analysis 

2.4 ATS Temperature Control Analysis 

2.5 ATS Structure Analysis 

2 . 6 Bores ighting 

2. 7 Image Rerotation and Decoupling 

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2.1 ATS OPTICS ANALYSIS 

This section presents material related to the optical design and performance of the ATS. 

The material is subdivided as follows: 

2.1.1 Design Description 

2.1.2 GE Performance Analysis 

2.1.3 GE Analysis of Optical Performance Under Dynamic Conditions 

2.1.4 Secondary Design and Performance Information 

2.1.1 DESIGN DESCRIPTION 

This section defines the optical design as it exists for preliminary design review. 

2.1.1.1 Optical Prescription 

The optical prescription for the ATS is given in Table 2.1-1. The prescription is not 

expected to change further. Minor adjustments will be made based on melt and manufacturing 

tolerances. Column headings are as follows: 

1. Surface Number - The outside surface of the telescope objective is Surface No. 1 

The exit pupil is surface No. 53. Surface Nos. 5 through 15 inclusive are contained only in the power changer mechanism. 

2. Radius - Radius of curvature at each surface 

3. Radius tolerance 

4. Thickness 

5. Thickness tolerances 

6. Air Space 

7. Air Space tolerance 

8. Clear Aperture 

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TABLE 2. 1-1. OPTICAL PRESCRIPTION FOR ATS 

2 -4 

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RADIUS THICKNESS AIR SPACE CLEAR GLASS 

SURFACE 

RADIUS TOL. THICKNESS TOL. SPACE TOL. APRTR TYPE AND 

NO, (INCHES) (INCHES) (INCHES) (INCHES) (INCHES) (INCHES) (INCHES) n* v** MELT 29 - 1.6161 * 0.002 0.3996 *0.005 0.3342 *0.005 0.66 1.62004 36.37 F-2 

High 

30 - 0.8979 * 0.005 1.1384 Mid. 0. 67 

Low 

1 31 1.7353 * 0.015 0. 3246 *0.005 1.3933 0. 75 1.51680 64. 17 BK-7 

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1 38 5.4668 * 0.069 0. 1500 *0.005 1.18 

6 

39 . Filter 0.1700 *0.005 1.18 1.51680 64. 17 BK-7 

L 40 0. 1500 *0.005 1.18 

1 41 -1.9993 * 0.009 0.2001 *0.005 1.18 1.51680 64. 17 BK-7 

42 - 4.1623 * 0.035 1.0426 *0. 005 1.24 1 43 m 0. 2000 *0. 005 1.50 ImagePlane 

44 - 7.4978 * 0.084 0. 1670 *0. 005 1.54 1.68893 31. 18 SF-8 

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50 0. 9348 * 0.002 0. 500 *0. 005 1.44 1.62041 60. 33 SK-16 

I 51 - 11.9989 * 0.300 0. 1490 *0.005 1.29 1.68893 31. 18 SF-8 1 

52 1.3461 * 0.006 0. 7716 *0. 005 1.02 

53 Eye Point 

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SURFACE RADIUS TOL, THICKNESS TOL. SPACE TOL. 

APRTR TYPE AND 

NO. (INCHES) ( INCHES) (INCHES) (INCHES) (INCHES) (INCHES) (INCHES) n* v** MELT 

1 40. 1302 * 0.080 1.3544 +0. 000 10.00 1.62041 60. 33 SK-16 

-0.005 

2 - 18.4179 + 0.005 0. 2656 +0.002 9. 98 

3 - 17. 8183 * 0.005 0. 6000 +0. 000 9.84 1.61340 44. 30 KzFSN-9 I 005 

4 273. 6105 + 3.000 28. 8180 +0.030 9.73 

5 - 8.9940 +0.036 0. 3613 +0.005 2. 79 1.62041 60. 33 SK-16 

6 10. 7028 + 0.050 0. 0200 +0. 005 2.79 

7 7. 6130 + 0.026 0. 2934 +0. 005 2. 80 1.62004 36.37 F-2 8 486. 4318 *105. 000 13.9069 *0. 010 2. 79 

9 m 10. 0873 *0. 010 2. 06 Aperture I 

10 - 5.4627 * 0.010 0.4468 +0. 005 3.28 1.62041 60. 33 SK-16 11 - 3.1057 * 0.003 0. 1337 *0.005 3. 35 12 - 3.0157 * 0.003 0. 4516 *0.005 3.32 1.62004 36. 37 F-2 

13 - 6. 7970 * 0.013 0. 0220 *0. 005 3.55 

14 11.9795 * 0.038 0. 5711 *0. 005 3.63 1.62041 60.33 SK-16 15 - 18.5349 * 0.090 12. 0696 *0.010 3.63 16 m 9. 1391 Unfolded 2. 22 1.51680 64. 17 BK-7 

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18 2. 2310 * 0.008 0. 2391 *0. 005 1.53 1.51680 64. 17 BK-7 

19 2. 2002 0. 008 0. 7916 +0. 005 1.46 

20 m 0.200 *0. 005 1.36 1.51680 64. 17 BK-7 21 m 0.3229 *0.005 1.37 Reticle 

22 2. 5493 0. 011 0. 3959 *0.005 1.39 1.62041 60. 33 SK-16 1 23 - 1. 8689 0. 006 0. 0837 *0. 005 1.36 

24 - 1. 3669 0.004 0. 2310 +0. 005 1.33 1.68893 31.18 SF-8 

25 1.3808 0. 004 0.4500 *0. 005 0. 2274 *0.005 1.35 1.62041 60. 33 SK-16 

{High id 26 -2.2523 * 0.009 0. 5215 Mow 1.37 

27 2. 9031 * 0.022 0.3200 *0. 005 0. 9993 1.09 1.51680 64. 17 BK-7 28 - 1.4431 * 0.006 0. 3750 1.05 

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TABLE 2. 1-1. OPTICAL PRESC

RIPTION FOR ATS (Cont) 

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9. Refractive Index measured at 5876 Angstroms 

10. Dispersion 

11. Glass Type 

2.1.1.2 Optical Design - Low Power 

The optical design of the telescope design at low magnification is shown in Figure 2.1-1. 

Optical elements 5 through 16 are part of a power change mechanism inserted into a basic 

telescope design to reduce the magnification and to increase the field of view. Optical elements 1 through 4 make up the objective lens assembly. Optical elements 20 through 52 

are part of the eyepiece assembly. These parts will be described in more detail in paragraph 2.1.1.3. 

2.1.1.3 Optical Design - High Power 

The optical schematic of the telescope design at a magnification of 127 times is shown in 

Figure 2.1-2. Optical elements 5 through 15 have been removed to increase the magnification of the telescope. The optical element between surface 16 and 17 is the Pechan Prism 

used for image rotation. The optical element between surface 18 and 19 is a field flatner. 

The reticle is contained on surface 21 of an optical flat represented by surface 20 and 21. 

Optical elements between surface 22 to 46 are part of a zoom mechanism. The peripheral 

display will be located between surface 46 and surface 47. 

2.1.1.4 Eyepiece Assembly 

The eyepiece assembly consists of the following subassemblies: 

a. Eyepiece 

b. Peripheral Display 

c. Filter Assembly 

d. Zoom 

e. Reticle 

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In addition to supporting the optical elements, the eyepiece mechanical design must allow 

a focusing adjustment to be made. This adjustment is to compensate for the eye differences 

between various observers. 

The eyepiece of the AO Telescope is specially designed to compensate for residual aberrations of the zoom relay assembly. Because of this, the alignment tolerances are considerably tighter than the standard eyepieces found in military instruments. These tighter 

tolerances impose a restraint on the design of the focusing mechanism. 

The individual lenses are kept in optical alignment via accurate fabrication of the glass, 

tube bore and spacers. Decenter and tilt errors are kept small by manufacturing tolerances 

and can only be built up by the operational clearances between the tube and the outer housing. 

The final scheme will reflect the final allowable optical design tolerance on tilt and decenter. 

Weight is minimized by using the smallest elements optically allowable thus producing a 

minimum eyepiece envelope. A weight reduction study on the eyepiece housing and tube 

will be performed as soon as final tolerances are established. 

2. 1. 1. 4. 1 Peripheral Display 

Another design restriction imposed upon the eyepiece focusing mechanism is the space 

required for incorporation of the peripheral display. 

Three types of light sources for the peripheral display have been considered: solid state, 

electro-illuminance, and incandescent. 

For any type of light source, it may be viewed directly or the light transmitted to the image 

plane by an optical means. While direct viewing is more efficient and possibly lighter 

weight, it does present packaging problems. 

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During the course of display evolution the pattern as seen in the apparent field has gone 

from two complete annular rings of possible sources to the present arrangement of three 

groups, each being a single row. 

The lamps will be epoxied into recesses machined in the circuit board. Multi-colored lamps 

will be visible from the front side through clear viewing ports. The wiring side of the board 

will be epoxied over to provide both structural integrity and electrical protection. 

The present approach using the printed circuit board is geared toward quickly getting a 

workable system to prove out the optical design problems through simulation. Full attention 

will be given to developing a minimum size, reliable, minimum weight system that will provide optimum display information while allowing minimum degradation of the optical system. 

It is intended that this approach be developed further through evaluations of such techniques 

as flexible printed circuits and complete encapsulation. 

The design will minimize package size to allow for the - 2.5 to + 3.5 eyepiece dioptric 

adjustment. The optical design of the peripheral display will provide patterns visible in 

the eyepiece field of view while at the same time providing for minimal degradation of system performance. Optical design objectives include minimizing random reflections, minimizing obscuration of the apparent field of view, and minimizing the undesirable effects 

cuased by moving the eyepiece through its diopter range. The display luminosity level is 

vet to be determined. The display does not encroach upon the eyepiece field within ± 30 

degrees of the in-track vector thus meeting the requirement of ± 27 degrees. 

2.1.1.4.2 Filter Assembly 

Various methods of inserting a filter into position on the optical axis were studied. One 

method evaluated has four individual filters mounted in arms that pivoted on a common shaft 

but were inserted into position separately. This design used at a minimum four to five 

tiller thickness of space on the axis. As the optical design became firmer the space available 

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......•■■.... 

for filter insertion became smaller until only one filter and its mount could fit into position. This requirement led to the bar approach whereby all four filters in a rectangular 

plate were racked into and out of position. This scheme made for a large bulky envelope 

requirement. The wheel or disc filter was studied in order to minimize space. Compacting 

the filters as close together as possible led to the final filter design. 

Several drive mechanisms for the filter wheel were studied and refined until a minimum 

space requirement was achieved. 

The filter wheel drive mechanism consists of a single pass, 4 : 1 ratio gear drive. This 

system provides filter-to-filter indexing with only a 22-1/2 degree motion of the filter control knob. Total knob motion is 90 degrees. Sealed ball bearings are used to support the 

control knob shaft. A sector gear transmits knob motion to the pinion gear which also 

serves as a shaft bearing for the filter wheel. A stub shaft extending from the zoom housing 

supports and retains the filter wheel pinion gear assembly. The filter wheel is firmly 

positioned by a spring loaded roller into a local groove detenting on the wheel periphery. 

There are four grooves, one corresponding to each filter position. 

The filter wheel design was limited by the spacer envelope remaining after adjacent glass 

elements were mounted in their respective positions. The most compact approach of 

mounting the lenses was used so as to provide maximum space for the filter wheel. Even 

so, the filter glass must be bonded into place because of lack of room for a retaining ring. 

However, optical tolerances on the filter glass are looser than on other elements; therefore bonding and staking the filter into place is satisfactory for all load conditions. 

The filter wheel assembly, as now designed, permits the use of four selectable spectral or 

neutral density filters. Additional study is required before the characteristics of these 

filters can be specified. Based upon current knowledge of atmosphere haze, human eye 

response and chromatic characteristics of the AO Assembly, it is expected that two yellow 

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filters, plus one neutral density filter will be specified, plus a clear optical window in the 

fourth filter slot. The two yellow filters will be a deep yellow filter for very hazy viewing 

conditions and a light yellow filter for normal viewing. The neutral density filter will permit more comfortable viewing of very bright scenes; and the clear window may be used for 

viewing targets at night and for viewing poorly illuminated targets (e.g. high latitude targets). 

2. 1. 1.4.3 Zoom 

The optical design tradeoffs between mechanically compensated zoom configurations are 

discussed in this section. 

The three major zoom types are: 

a. Mechanical Compensation - Two or more lens groups move with respect to the 

focal plane and each other. 

b. Linear Compensation - A mechanically compensated lens with the added 

restriction that the motions of the moving groups are linearly related. 

c. Optical Compensation - A linearly compensated lens with the added restriction 

that the moving groups move as a unit with one stationary group between each 

pair of moving groups. 

Optically compensated zooms are a small subset of mechanically compensated zoom lenses. 

['hey are used based on mechanical considerations, provided satisfactory optical performance 

can be achieved. The mechanical tradeoff is not straight forward. In the mechanically compensated AO Zoom, only two groups move with no stationary elements in between. In 

optically compensated zooms, two or more groups move with mechanical means required 

to keep a lens group stationary, between each pair of moving groups. In addition, the 

defocusing as a function of zoom position is exactly zero for mechanically compensated zoom 

lenses but a non-zero value for optically compensated zoom lenses, reaching a small value 

as the number of moving groups is increased. 

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The practical question arises of whether two conjugate pairs (in this case the objective 

image pair and the entrance/exit pupil pair) can be kept stationary in an optically compensated zoom. This has been shown to be possible paraxially by Silvertooth (Wooters 

and Silvertooth, Josa 55, 347-1965) and is thus clearly achievable with mechanically compensated zooms. However, the ability to control the real exit pupil position to ±0.02 inch 

over a 2 : 1 zoom range and 60-degree apparent field is not as easily proven and in fact is 

a property of the higher order aspects of the design. No optically or mechanically compensated zoom in the open literature has been found with these properties. For this and 

other reasons, the AO Telescope zoom was designed to achieve these specifications without 

the added restriction of optical compensation. The resulting zoom relay eyepiece has less 

than 0.2 wave spherical aberration and chromatic aberration throughout the zoom range on 

axis. The variations of field aberrations with zoom position is very small and the actual 

field correction is limited by the eyepiece portion rather than the zoom portion of the system. 

2. 1. 1. 4. 4 Reticle 

,..•••••••■., 

A single circle reticle which indicates the real field of view at maximum zoom position is 

provided. The reticle includes cross hairs for centering. 

The reticle material has been selected as BK-7 glass with a clear aperture of 1.36 inches 

and is 0.200 inch thick. The reticle pattern described above will be deposited on the glass 

by a method yet to be determined. 

Prime consideration has been given to the reticle mounting requirements because of the 

need for adjustment to set focus and because of the critical cleanliness problems associated 

with reticles. The reticle is mounted as an integral part of the zoom lens assembly. This 

allows us to fix the position of the reticle at the front focal point of the zoom lens while 

still isolating it from the mechanism, thus maintaining its tilt and decenter requirements. 

The compensating adjustments, at assembly, are accomplished by shimming the entire zoom 

reticle assembly at the field break behind the Pechan Prism, after the reticle zoom relationship is established. 

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To control the cleanliness of the reticle surface the optical design has been modified to 

place the reticle surface inside the zoom lens (i. e. , second surface of the reticle lens). 

If the telescope tube is separated at this field break the critical surface will not be contaminated or damaged. 

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2.1.2 ANALYSIS OF PROPOSED ATS OPTICAL DESIGN 

2.1.2. 1 Introduction 

In the analysis that follows, the static physical frequency response of the ATS optical design has been computed. Physical frequency response curves-have been generated at six 

magnification powers for the proposed design subject to manufacturing tolerance and for 

the nominal (pristine) design. The resolution generated for the design subject to manufacturing tolerance represents a realistic performance level while the resolution level of 

the nominal design is over-optimistic since every manufacturing and alignment tolerance 

is zero. The estimate achievable ground resolutions have been obtained by superimposing 

the Schade Visual Resolution Threshold curve on the physical frequency response curves 

and determining the points of intersection. 

The effect of element curvature and air space/glass thickness on total system performance 

was also investigated. To accomplish this, the partial derivative of the total system performance relative to each curvature and air space/glass thickness was determined. These 

partial derivatives, given the same sensitivity coefficients, revealed that the first two elements of the objective have a more significant relative effect on the total system performance than the relative effect of the remaining elements combined. 

The following conditions were used in the analysis: 

a. Target Definition - A white bar, on a black background, separated by an equal 

distance from a similar bar. 

b. Orbital Altitude - 80 nautical miles 

c. Apparent Contrast - 2:1 at aperture 

d. Illumination at target - 530 foot lambarts 

e. Vibration Level - Static 

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2. 1. 2. 2 Conclusions 

a. The zero degree (in-axis) field physical frequency response of the optical prescription computed for the high power 127X indicates a ground resolution capability of 2.4 feet. The value determined from Itek data is 2.8 feet. Both values, 

although considerably better than the 3.3-feet requirement, are very optimistic 

since manufacturing tolerances have not been considered. 

b. The estimated worst case achievable ground resolution, determined by applying 

a tolerance budget to the nominal design prescription is 3.4 feet. The value submitted by Itek is 3.3 feet. The static requirement is 3.3 feet. The GE estimate 

is the worst case resolution found after applying eight different tolerance budgets. 

Six of the eight tolerance budgets resulted in an estimated ground resolution of 

2.7 feet or better. 

c. A level of confidence that the proposed design can be improved upon if tolerance 

budgets bring the system performance below an acceptable level has been established. Of eight different tolerance budgets applied five immediately resulted in 

an acceptable performance level; three did not. A further analysis demonstrated 

that the performance level of the three could be significantly improved by tweaking the doublet airspace and applying an appropriate reticle refocusing. 

d. The analysis demonstrated that the doublet is the most sensitive portion of the design. Assigned manufacturing tolerances to the doublet have a more significant 

effect on total performance than the effect of manufacturing tolerances assigned 

over the remainder of the design. This result has been predicted by the sensitivity 

coefficients which had the doublet account for 89.64% of the total sensitivity coefficients. 

e. The analysis demonstrated that assigned manufacturing tolerances to the doublet 

radii of curvatures which alternate in sign have a significant degrading effect on 

ground resolution. Such tolerances even after a doublet airspace adjustment is 

made, result in the worst case estimated ground resolution of 3.4 feet. Tolerances in the doublet radii of curvature which are of one sign also have a degrading 

but less significant effect to the resolution capabilities. Such tolerances, when 

accompanied by a doublet airspace adjustment, result in an estimated 2.7 feet. 

f. The analysis demonstrated that the resolution is considerably degraded as the 

apparent field is transversed from 0° to 300. For the high power design - 127X, 

the estimated achievable ground resolution at 0° is 3.4 feet; at 10° it is 5.5 feet; 

at 21° it is 8.4 feet; at 27° it is 8.8; and at 30° it is 9.3 feet. A resolution degradation across the field for the other zoom powers is also apparent. For the lower 

power design - 15. 88X, the estimated achievable ground resolution at 0° is 18.8 

feet; at 10° it is 35.0 feet; at 21° it is 171.9.feet; at 27° it is 244.8 feet; at 30° it 

is 266.1 feet. (Note: The system resolution figure for off axis performance are 

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the result of very preliminary General Electric analysis and have neither been 

substantiated or disproved by Subcontractor analyses. Off-axis performance is 

an area where additional analysis is required to predict system capability). 

2. 1. 2. 3 Recommendations 

a. Manufacturing tolerances of the doublet radii of curvatures should be reassigned 

to values which do not alternate in sign. Two alternate procedures would be to 

either minimize the tolerances or fabricate duplicate elements which could then 

be subject to selection. 

b. The use of a triplet design in place of a doublet design should be re-evaluated to 

determine if the sensitivity coefficients of the objectives lens can be reduced. 

Itek has completed an optical prescription using a triplet objective lens, but discarded this design based on weight considerations when the doublet objective lens 

met performance requirements. The triplet objective lens weighs 2 to 3 pounds 

more than the doublet design; however, the use of a triplet design may result in 

a total system weight saving by reducing environmental control requirements in 

the area of the objective lens. The use of a triplet design will also permit a 

further reduction in chromatic aberration and may result in improved ground 

resolution better than 3. 4 feet. 

2. 1. 2. 4 Discussion of Results 

The primary objective of this analysis was to determine if the proposed design can meet 

or surpass the specified performance level and if so, with what level of confidence. To 

carry out this result the physical frequency response of the nominal design was determined, 

the level of confidence of the design was investigated, a realistic performance estimate 

was established and the effect of tolerances was determined. 

A. Nominal Design Capabilities - The physical frequency response of the nominal 

prescription was computed for six magnification powers. This was done for the on-axis 

- 0° field since the specified CEI requirements are given in terms of on-axis resolution. 

The response curves generated are given in Figures 2.1-3 to 2.1-8. Also included in each 

figure is the corresponding Itek supplied curve and the Schade Visual Resolution Threshold 

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1.0 

0.9 

0.8 

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0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 

Figure 2.1-3. Nominal Design - Physical 

Frequency Response - 127X 

SCHADE 

THRESHOLD 

SCHADE 

THRESHOLD 

GE 

1.56 C/MRAD (3.7 FT 

0.3 

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SCHADE 

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SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

Figure 2.1-4. Nominal Design - Physical 

Frequency Response - 84.67X 

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0.2 

0.1 

0 

0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 

o° 

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I I I I I I 1 I 

0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

Figure 2.1-5. Nominal Design - Physical 

Frequency Response - 63.5X 

1.0 

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THRESHOLD 

1 1 I I I I 1 I 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

Figure 2.1-6. Nominal Design - Physical 

Frequency Responses - 31.75X 

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1 

00 

1.52 C/MRAD (10. FT) 

1.57 C/M RAD (9.8 FT) 

GE 

0.7 

>-■ U 

Z 41 • 0.6 

a w 

w E0 

g z 

w o a, 0.5 

K2, 

2-19 
Figure 2.1-7. Nominal Design - Physical 

Frequency Response - 21.17X 1.0 

0. 9 

0. 8 

1.76 C/MRAD (17.4 FT) 

1.62 C/MRAD (18.9 FT) 

SCHADE 

THRESHOLD 

PHYSICAL FREQUENCYRESPONSE 

0.3 

0.5 

0.4 

0.7 

0.6 

GE 

0. 2 

0 . 1 

0 

0.2 0.4 0.6 0. 8 1.0 1.2 1.4 1.6 

NRO APPROVED FOR 

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1.0 

0.9 

0. 8 

0. 7 

>-■ 

• 0.6

z 

a w 

• cn 

° 0.5 

•-] con 

C4 

4=7 

cn 0. 4 0 

1.67 C/MRAD (13.7 FT) 

0.3 

0.2 SCHADE 

THRESHOLD 

0.1 

0 

0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

SPATIAL FREQUENCY (cycLEs/mILLIRADIAN) 

Figure 2.1-8. Nominal Design - Physical 

Frequency Response - 15.88X 

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Curve. The intersections of the physical frequency response curves and the Threshold 

curve are denoted by an arrow with the frequency of intersection and the corresponding 

ground resolved distance written above. 

In general, the GE computed curve and the Itek Curve are in good agreement. Any differences between the curves may have resulted from the fact that Itek used nine spectral 

wavelengths to compute their curves while the computer program used to generate the GE 

Curves is capable of incorporating only three spectral wavelengths. 

At the high power design - 127X, the GE curve indicates a theoretical ground resolution of 

2.4 feet and the Itek curve indicates a theoretical ground resolution of 2.8 feet. Both are 

better than the static requirement of 3.3 feet. However, they are overoptimistic, since 

manufacturing tolerances have not been considered in the computations. (Unless otherwise specified, system performance curves have been calculated by tracing 75 rays 

through the system. An explanation of the computation techniques is included in Section 

2.1.4.) 

B. Design Level of Confidence - After the nominal prescription performance as a function of magnification was established, the level of confidence of the design was investigated. 

The purpose of this level of confidence investigation was to determine if tolerance budgets 

would bring the system performance below an acceptable level. If so, could anything be 

done to improve the performance level? 

In this section each computation was made for the 127 power at 0° field. The manufacturing 

tolerance budgets specified in Table 2.1-2 were applied to the doublet, zoom relay, eyepiece, and over the complete design. In each case, the physical frequency response of 

the high power 127X system was computed. The curves are plotted in Figures 2.1-9 

through 2.1-13. In Figure 2.1-9, curve 1 is the system physical frequency response 

when case "e" tolerances are given to the zoom relay; curve 2, the response when case "Fr 

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TABLE 2.1-2. TOLERANCE BUDGETS 

TYPE 

AND 

NUMBER a b c d e f aT dT 

R1 +0.08 -0.08 +0.08 -0.08 A +0.08 -0.08 

T1 0 -0.005 -0.005 -0.005 0 -0.005 

R2 -0.005 +0.005 +0.005 -0.005 -0.005 -0.005 

T2 -0.002 +0.002 +0.002 +0.002 -0.002 +0.002 

R3 +0.005 -0.005 +0.005 -0.005 +0.005 -0.005 

T3 0 -0.005 -0.005 -0.005 0 -0.005 

R4 -3.00 +3.00 +3.00 -3.00 -3.00 -3.00 

T4 -0.03 +0.03 +0.03 +0.03 -0.03 +0.03 

R16 A A A 

2 W 

T16 A 

1-1 

NOMINAL DESIGN • 

R17 ;!3 Poi 

T17 0 

R18 Z -0.008 -0.008 

T18 c/3 szl 'cn I rn

ul l -0.005 -0.005 

PA 41 W W 

R19 A A A A -0.008 -0.008 

1-4 14 1-4 1-4 

T19 -,, (4 -,, -..„ <4 -0.005 -0.005 

01. ic1 PO A 

R20 Z 0 

T20 

0 Z 0 Z 0 Z Z 

R21 

T21 V c/3 cn 

R22 -0.011 W C.) W C.) 

Z Z 

T22 -0.005 

A 

R23 -0.006 W 0-1 W 1-4 

0 0 

T23 -0.005 E l 

a) 

El 

a) 

R24 -0.004 

T24 -0.005 

R25 -0.004 

T25 y -0.005 

2-22 

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TABLE 2.1-2. TOLERANCE BUDGETS (CONT) 

TYPE 

AND 

NUMBER a b c d e f aT dT 

R26 -0.009 A 

T26 -0.005 

R27 -0.022 

T27 -0.005 

R28 -0.006 

T28 -0.005 

R29 -0.002 

T29 -0.005 

R30 -0.005 

T30 -0.005 

co 

R31 

NOMI

NAL DESIC 

NOMINAL DESIC 

- NOMINAL DESIC 

[OMINAL DESI( 

-0.015 

IOMINAL DESI( 

e 

T

OLE

RANCE 

W C.) 

T31 -0.005 Z 

g 

R32 -0.003 4.1 1-4 

T32 -0.005 0 E-1 

a> 

R33 -0.003 

T33 -0.005 

R34 -0.041 

T34 -0.005 

R35 -0.009 

T35 -0.005 

R36 -0.004 

T36 -0.005 

R37 -0.004 

T37 -0.005 

R38 -0.069 

T38 -0.005 

R39 

T39 -0.005 

2-23 

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TABLE 2.1-2. TOLERANCE BUDGETS (CONT) 

TYPE 

AND 

NUMBER a b c d e f aT dT 

R40 

f TOLE

RANCES --10- 

f TOLE

RANCES 

T40 -0.005 

R41 -0.009 

T41 -0.005 

R42 -0.035 

T42 

R43 

T43 

R44 -0.084 

T44 

gOMINAL DESIGN - NOMI

NAL DESI

GN 

NOMINAL DESIGN - 

-0.005 

1145 

sTOMLNAL DESIC 

-0.002 

T45 -0.005 

1146 -0.004 

T46 -0.005 

Z 

R47 -1 

co 

-0.006 

T47 

rzl 

A -0.005 

1-1 

R48 .,4 -0.002 

T48 0 

', -0.005 

1149 Z -1.253 

T49 -0.005 

1150 -0.002 

T50 -0.005 

1151 -0.300 

T51 -0.005 

R52 -0.006 

T52 

2-24 

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0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 

CURVE 1 - ZOOM TOLERANCES 

CURVE 2 - EYEPIECE TOLERANCES 

NOMINAL 

DESIGN 

CURVE 2 

NOMINAL 

DESIGN 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

Figure 2.1-9. Physical Frequency Response - 127X 

• 0.1 

0 

TOLERANCES - AFTER ADJUSTMENT 

TOLERANCES - BEFORE ADJUSTMENT 

TOLERANCES - AFTER ADJUSTMENT 

TOLERANCES - BEFORE ADJUSTMENT 

0.9 

1.0 

d

T 

d

T 

aT 

aT 

\ 

2 

\ 3 

• • • • 1.14 C/MRAD (3.35 FT) 

•• Zs 

1.55 C/MRAD (2.5 FT) 

SCHADE 

THRESHOLD 

0.8 

›, 0.7 

z 

0.6 

CY41 

GLI Z 

° 0.5 

w 

M 

• 

0.4 

• 0.3 

0.2 

\ •

, •

• • .• • • 

0.2 0.4 0.6 0.8 1.0 

•••••••• 

1.2 1.4 1.6 

• 1.082 C/M D (3.5 FT) 

1 

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1..0 

0.9 

0.8 

0.7 

U 

z fa) 0.6 

C4 Z 

0 4.. a, 0.5 

C.) 

Fd 0.4 

0.3 

0.2 

0. 1 

0 

Figure 2.1-10. Physical Frequency Response - 127X 

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0.2 0. 4 0. 6 0.8 1.0 1.2 1.4 1.6 

SPATIAL FREQUENCY (CYCLES PER MILLIRADIAN) 

Figure 2.1-11. Comparison of a, b, c, d Tolerance 

Budget-Physical Frequency Response-127X 1.0 

0.9 

0. 8 

0.7 

0.6

z 

0' 44 

44 [14 Z 

• 0.5 

cn 

.4 44 

• G4 

0.4 

z 0. 

0.3 

0.2 

0 . 1 

2 

1 1 

SCHADE THRESHOLD 

1.0 

0.9 

0.8 

.4 44cis 0. 5 

aw 

U r:4 

0.4 

• 0.3 

0. 2 

0. 1 

0 

• 

0.7 

0.6

z 

44 Eon 

• Z 

0 I 1-- I I I 1 

0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 

SPATIAL FREQUENCY (CYCLES PER MILLIRADIAN) 

Figure 2. 1-12. Effect of Tolerance Budget a 

Before and After Adjustment-127X 

2 -26 

ET SPECIAL HANDLING 

1.6 
GE WORST CASE 

PERFORMANCE ESTIMATE 

ITEK 

PERFORMANCE 

ESTIMATE 

1.12 C/MRAD (3.4 TT) 

SCHADE 

THRESHOLD 

1.16 C/MRAD (3.3 FT) 

1.0 

0.9 

0.8 

0.7 

0.6 

0.5 

0.4 

0.3 

0.2 

0.1 

0 

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0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 

SPATIAL FREQUENCY (CYCLES PER MILLIRADIAN) 

Figure 2.1-13. Comparison of Performance 

Estimate - Zero Degree Field 127X 

tolerances are given to the eyepiece; curve 3, the response of the prescription design. Included in this figure is the threshold curve. For an estimated ground resolution of 3.3 feet, 

the physical frequency response curve and the threshold curve must interset at 1,160 cycles/ 

radian. Each case intersects at a frequency better than the limiting frequency, 1,160 

cycles/radian. By inspection, it is apparent that the zoom relay or eyepiece tolerances 

have little effect on the physical frequency response. 

Figure 2.1-10 contains the physical frequency responses generated for two different tolerance budgets distributed over the complete design before an element adjustment is made 

(curves 1 and 3). These responses are significantely improved by applying an adjustment to 

the doublet airspace (curves 2 and 4). The fact that the performance level was significantly 

improved by an element adjustment increases the level of confidence of the design. 

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The level of confidence was further investigated by applying tolerances only to the doublet. 

Four combinations of doublet tolerances were formulated and added to the nominal design. 

The physical frequency response curve was computed for each combination. (A more detailed account of this treatment is given in Section 2.1.4.4 of the analysis). The results 

are given in Figure 2.1-11 along with the threshold curve. Curves 1 and 2 indicate a degraded performance while 3 and 4 prove acceptable. An element adjustment was given to 

case "a". The response was vastly improved. The improved response is shown compared 

to the original response in Figure 2.1-12. 

From this discussion, it is apparent that for each tolerance grouping applied, the system 

performance either immediately surpassed the minimum required or was significantly improved to an acceptable level by an element adjustment. 

C. Performance Estimate and Estimated Ground Resolution - To arrive at a realistic 

performance estimate, tolerance budgets were applied to the nominal design. A performance estimate, to be realistic, would have to be generated by distributing tolerances over 

the complete design. The physical frequency response curve selected would then have to 

represent a worst case level, and yet be generated after an element adjustment is made to 

simulate a worst case in the laboratory. Such a worst case tolerance performance estimate is represented by curve 1, of Figure 2.1-13. This curve was calculated by tracing 

750 rays through the system to obtain a more accurate prediction of system performance. 

Comparing this curve to the Itek performance estimate curve (curve 2), curve 1 runs about 

30% lower than the Itek curve at the lower frequencies; however, the curves tend to approach 

each other in the frequency range 1000-1200 cycles/radian. The Itek curve intersects at 

1,160 cycles/radian, which results in an estimated achievable ground resolution of 3.3 feet. 

The GE performance estimate curve intersects at 1,082 cycles/radian, which results in an 

estimated achievable ground resolution of 3.4 feet. The difference in the curves can not be 

explained since the Itek tolerance budget was not available. The tolerance budget used to 

generate the GE worst-case performance estimate curve has then been used to compute the 

2-28 

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...■.... 

0° field physical frequency response performance estimates for each zoom power. Furthermore, with the same tolerance budget applied, it was then possible to compute the off-axis 

resolutions as well, and be confident that they correspond to a realistic performance level. 

The off-axis physical frequency response curves for each zoom power was thus computed. 

The generated curves are given in Figures 2.1-14 through 2.1-19. Included in these figures 

is the Itek-supplied 127X Performance estimate curve, which is for on-axis only, and the 

Schade Visual Resolution Threshold curve. All GE-calculated curves in these six figures 

were produced by tracing 75 rays through the system due to the large amount of computer 

computation time required to generate each curve using 750 rays. However, it is felt 

that the error introduced by using the smaller number of rays is quite small since the difference in performance between the 750-ray case (Figure 2.1-13) and the 75-ray case 

( Figure 2. 1-14) is only 0.1 foot. 

The intersections of the physical frequency response curves and the Threshold curves are 

the frequencies from which the estimated achievable ground resolutions are obtained. Estimated achievable ground resolutions at five points in the apparent field of view have been 

tabulated for each zoom power and is given in Table 2.1-3. The results show a significant 

resolution degradation that the viewer will experience as he transverses the field: 3.4 feet 

ground resolution at 0° compared to 9.3 feet at 30° for the 127 power (worst-case). 

Another result demonstrated by the on-axis performance estimate curves is that the 127X 

curves are inferior to the curves obtained for the other magnifications. Thus, the 127 

power should be used as the worst case power in the tolerance analysis. 

2-29 

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,vefort 

TABLE 2.1-3. TABULATION OF ESTIMATED WORST CASE 

ACHIEVABLE GROUND RESOLUTIONS (FEET) 

FIELD 

ANGLE 

MAGNIFICATIONS 

127X 84.67X 63.5X 31.75X 21.17X 15.88X 

0° 3.5 4.74 5.7 11.5 15.3 18.8 

10° 5.2 7.3 9.6 18.9 26.8 33.5 

21° 7.8 7.6 41.4 18.6 104 185 

270 8.1 27.9 61.2 53.4 163 306 

30° 8.5 54.0 69.5 72.9 201 340 

D. Sensitivity Coefficients and Tolerances - The sensitivity coefficients indicate the 

relative effect of each curvature and airspace/glass thickness on total system performance. 

Parameters with larger sensitivity coefficients have a more pronounced effect on total system performance than those with smaller sensitivity coefficients. 

The sensitivity coefficients obtained for the high-power case (127X) have been tabulated 

and are given in Table 2.1-4. These coefficients represented by an algebraic sign and 

magnitude were determined using the General Electric Auto-Spryte computer program. 

In order to express the sensitivity in a numerical manner, the sensitivity coefficients were 

all added together and each sensitivity coefficient was expressed as a percentage of the 

total. These Relative sensitivity coefficients are also included in Table 2.1-4. In this 

table, surface number pertains to the surface place in the total collection of surfaces that 

comprise the design (i. e. , Surfaces 1 through 4 comprise the four surfaces of the doublet). 

The letter "C" designates the sensitivity coefficient related to curvature (equal to reciprocal 

of radius of curvature). The letter "T" designates the sensitivity coefficients with respect 

to airspace or glass thickness, whichever applies. 

2 -3 0 

SECRET SPECIAL HANDLING 
1.0 

0.9 

ITEK PERFORMANCE 

ESTIMATE - 0° 

SCHADE 

THRESHOLD 

0.2 

0.1 

1.082 C/MRAD (3.5 FT) 

1.160 C/MRAD (3.3 FT) 

00\ 

10° 

0.8 

0.7 

0.3 

1.0 

0.9 

0.8 

z 

0.7 

a Ea Ex) Z

En 

4. 00.6 a. a m 

ew c.) 0.5 

0.4 

0.3 

0.2 

0.1 

0.140 C/MRAD 

0 

0 0.2 0.4 0.6 0.8 1.0 1.2 1. 4 1.6 

1.220 C/MRAD (4.74 FT) 

SCHADE 

THRESHOLD 

0.760 C/MRAD (7.6 FT) 

I I 

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0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

Figure 2.1-14. System Performance Estimates-127X 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

Figure 2.1-15. System Performance Estimates - 84.67X 

SECRET SPECIAL HANDLING 

2-31 
PHYSICAL FREQUENCY RESPONSE 

0.2 

0. 3 

0. 5 

0. 4 

0. 6 

0.8 

0.7 

0. 9 

1.0 

1. 330 C/MRD (11.5 FT) 

1.0 

0.9 

I I 

0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

0 0.2 

0.8 

'HYSICAL FREQUENCYRESPONSE 

0.7 

0. 6 

0.5 

0.4 

0.3 

0.2 

0.1 

0 

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SCHADE 0.1 THRESHOLD 0. 110 (C/M RAD) 

0 

0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

Figure 2.1-16. System Performance Estimates - 63.5X 

Figure 2. 1-17. System Performance Estimates - 31. 75X 

2-32 

SECRET SPECIAL HANDLING 
1.505 C/MRAD (15.3 FT) 

SCHADE 

THRESHOLD 

0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

PHYSICAL FREQUENCY RESPONSE 

1.0 

0.9 

0.8 

0.7 

0.6 

0.5 

0.4 

0.3 

0.2 

0.1 

1.0 

0.9 

0.8 

c.) 0.7 

a 

O' a) 

z 

44 

0.6 

o 

wa 

1.625 C/MRAD (18.8 FT) 

0.2 0.4 0.6 0.8 1.0 1.2 1.4 

0.2 

0.1 

SCHADE 

THRESHOLD 

0 1.6 

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Figure 2.1-18. System Performance Estimates - 21.17X 

1.8 

SPATIAL FREQUENCY (CYCLES/MILLIRADIAN) 

Figure 2.1-19. System Performance Estimates - 15.88X 

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TABLE 2.1-4. ABSOLUTE AND RELATIVE SENSITIVITY COEFFICIENTS 

SURFACE ABSOLUTE RELATIVE SURFACE ABSOLUTE RELATIVE SURFACE ABSOLUTE RELATIVE 

1,C -0.461489 0.18411 26,C -0.007289 0.002908 40,C - - - 

1,T -0.040055 0.015950 26,T -0.001998 0.000797 40,T 0.002495 0.000995 

2,C -0.518593 0.206892 27,C -0.006490 0.002589 41,C 0.002194 0.000875 

2,T -0.212488 0.084772 27,T -0.001031 0.000411 41,T 0.002555 0.001019 

3,C -0.545727 0.217718 28,C -0.007377 0.003182 42,C -0.002615 0.001043 

3,T -0.038295 0.015278 28,T - - 42,T 0.002653 0.001058 

4,C -0.397200 0.158463 29,C 0.010264 0.004095 43,C - - 

4,T -0.033166 0.015232 29,T -0.002962 0.001142 43,T 0.003221 0.001285 

16,C - - 30,C -0.011479 0.004580 44,C 0.007204 0.002874 

16,T - - 30,T -0.000244 0.000097 44,T 0.003696 0.001475 

17,C - - 31,C 0.007790 0.003108 45,C -0.004139 0.001651 

17,T - - 31,T 0.000858 0.000342 45,T 0.006027 0.002404 

18,C -0.002499 0.000997 32,C -0.009322 0.003799 46,C -0.009110 0.003634 

18,T -0.003686 0.001471 32,T - - 46,T -0.003203 0.001278 

19,C 0.001572 0.000627 33,C 0.009853 0.003931 47,C 0.010930 0.004361 

19,T -0.002727 0.001128 33,T - - 47,T 0.00873 0.000348 

20,C - - 34,C 0.008877 0.003541 48,C -0.000908 0.000386 

20,T -0.002322 0.000926 34,T 0.000644 0.000257 48,T -0.000318 0.000127 

21,C - 35,C -0.002966 0.000824 49,C -0.009112 0.003635 

21,T -0.002411 0.000962 35,T 0.002428 0.000985 49,T 0.000576 0.000230 

22,C 0.004436 0.001790 36,C 0.001940 0.000774 50,C 0.012859 0.005130 

22,T -0.002249 0.000897 36,T 0.002874 0.001147 50,T 0.008373 0.003340 

23,C -0.005251 0.002095 37,C 0.003010 0.001201 51,C 0.001362 0.000543 

23,T -0.001290 0.000515 37,T 0.002255 0.000891 51,T 0.007016 0.002798 

24,C 0.002217 0.002480 38,C -0.002127 0.000849 52,C -0.008217 0.003278 

24,T 0.002667 0.001064 38,T 0.002463 0.000979 52,T - - 

25, C 0.000791 0.000316 39,C - 

25,T 0.002859 0.001141 39,T 0.002375 0.000948 

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The sensitivity coefficients with respect to curvature and airspace/glass thickness of the 

first four surfaces which comprise the doublet are overwhelmingly greater than the coefficients of the remaining surfaces. The relative sensitivity coefficient of the first four surfaces is 89.64 percent of the total. It can be stated that departures from nominal of these 

parameters affects total performance more significantly than the departures from nominal 

in the zoom relay or eyepiece. 

This is apparent when the curves in Figure 2.1-9 are compared to the curves in Figure 

2.1-11. The curves computed for tolerance budgets applied to the zoom relay and eyepiece 

(shown in Figure 2. 1-9) run consistently better than the curves generated for doublet tolerance budgets shown in Figure 2.1-11. The percentage of the total sensitivity coefficients 

contained by the zoom relay is 5.96 percent. That of the eyepiece is 3.87 percent. That 

Case 3 is betten than Case 2 confirms the prediction of the sensitivity coefficients. 

Since the doublet tolerances have such a strong bearing on the total performance, it is 

highly probable that a poor performance level of the adjusted system will be caused by 

misalignment, poor adjustment, or excessive manufacturing tolerance in the doublet portion 

of the design. Furthermore, any adjustment of the doublet elements will have to be accompanied by an appropriate shift of the reticle plane. 

Another result is apparent when the on-axis performance estimate curves (shown in 

Figures 2.1-14 through 2.1-19) are compared to the physical frequency response of the 

nominal designs (shown in Figures 2.1-3 through 2. 1-8). The high power magnifications 

(127X - 63.5X) curves show a considerable difference between the nominal design and the 

performance estimate curves, but in the low power magnifications (31.75X - 15.88X) the 

curves are in close agreement. In the analysis, it will be pointed out that the entrance 

pupil diameter of the high-power magnifications is 10.0 inches compared to 2.54 inches in 

the low-power magnifications. With more lens surface being used, tolerances are seen to 

have much greater effect on the system performance, as is expected. 

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2.1.2.5 Discussion of Analysis Techniques 

A. Method of Determination of Physical Frequency Response - Image evaluation computer 

programs have been used to determine the physical frequency responses of the proposed 

design. The following procedure was taken to obtain the final curves: 

B. Determination of Entrance Pupil and Relative Aperture (F/Number) - The entrance 

pupil is obtained by determining the limiting aperture of the design and finding the Gaussian 

image and magnification of this aperture formed by the elements which precede it. In the 

high-power mode, the limiting aperture is the first surface of the doublet. The entrance 

pupil thus coincides with this surface and has a radius of 5 inches. In the low-power mode, 

the aperture stop is surface 9. The image of this stop, formed by the forward elements, 

has a radius of 1.27 inches. This is, then, the entrance pupil radius for the low-power 

mode. 

The relative aperture of the design is the effective focal length divided by the entrance 

pupil diameter. The effective focal length, which is the distance from the second principle 

plane of the objective grouping to the image at the reticle of an infinitely located point 

object, has been determined. Focal lengths are shown below with the corresponding relative apertures: 

Eff. Focal Length (Inches) Relative Aperture 

High Power 

(127X - 63. 5X) 111.87 11.18 

Low Power 

(31.75X - 15.88X) 28.00 11.0 

C. Computation of Geometric Frequency Response - The IBM Program for Optical System 

Design has been used to compute the geometrical frequency response of the proposed design. 

In making the computations, a selected number of rays from a single object point at infinity 

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are traced through the entire system and emerge through the exit pupil having close degree 

of parallelism. The total number of rays traced are divided into three portions, each portion for a selected wavelength of the visible spectrum. The number of rays in each wavelength portion is governed by the standard eye luminosity curve published in the MIL-141 

Handbookl. In this way, the spectral distribution of the ATS optical system is given the 

eye spectral response. The wavelengths and number of rays in each portion is given in the 

table below: 

Spectral Wavelength Number of Rays 

(Microns) 

0.56 57 

0.47 12 

0.65 6 

The Program for Optical Systems Design (POSD) computes the geometrical frequency 

response by doing a computer analysis of the spot diagram traced from a point object. The 

computation technique is standard and is described in several up-to-date references2. 

The response at a given frequency is obtained from the summations: 

A(x) cos2ir v X. A X 

A

C (v ) - 

A(x) x 

J 

A(x) sin2 77' v X. X 

AS(v ) - j 

EA(x)6x 

1. "Handbook of Optical Design, " MIL-HDBK-141. 

2. Smith, "Modern Optical Engineering, " McGraw Hill, 1966 

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Geometrical Frequency 

Response C

(v )1 2 + [As(v)] 2 

where: 

v = Spatial frequency of the image 

x. = Jth coordinate of ray intercept in image plane. 

i 

A(x) = Line spread function in image space 

The POSD computes the geometrical frequency response by performing the above summations 

over the spot diagram that is formed by tracing rays through the system from a point object. 

Since the emerging rays of the proposed system are nearly parallel, a special technique 

pertaining to afocal systems was required to form a spot diagram. This was done without 

introducing any additional aberrations. Such a technique is described in the MIL-1411 

Handbook, paragraph 8.8. 

Once the spot diagram is formed, the geometrical frequency response is computed by 

making the summations that have been written above. Off-axis as well as on-axis spot diagrams may be formed from which off-axis and on-axis geometrical frequency responses are 

obtained. 

D. Determination of Physical Frequency Response - The POSD forms the spot diagram 

which is then analyzed using a ray trace that is governed by a strictly geomatrical theory. 

To account for diffraction effects, the geometrical response is multiplied by the physical 

response of a diffraction-limited optical system of the same relative aperture. This product tends to err toward the low side, thus resulting in a conservative or safe performance 

estimate. 

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The physical frequency response of diffraction-limited systems have been computed for the 

low-and high-power relative apertures of the design. This was done by using a computer to 

evaluate the physical frequency response expression: 

PFR (v) = (0 - cos cp sin 0) (cos e)

K 

where: 

v = Spatial frequency 

ca = cos

-1 (X vF ) 

0 = half-field angle 

X = Spectral wavelength (taken to be the mean wavelength), 0. 56 microns 

k = 1 and F = relative aperture # 

The product of this evaluated expression and the geometrical frequency response has been 

obtained using computer methods on a frequency-by-frequency basis. The products are the 

physical frequency response which are shown plotted in the results section. 

It should be pointed out that the diffraction effects result in a limiting frequency beyond 

which the system can not resolve. This limiting frequency is obtained by setting the physical frequency response (PFR) expression to zero: 

PFR (v) =-,27 (0 - cos 0 sin 0) (cos A) = 0 

The solution of this equation is 

cD = 0 

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or 

= cos

-1 (X v oFf) 

This occurs when 

X v oF#= 1 

or 

1 

V 0 = X F# 

Assuming A to equal 0.22 x 10-4inches and the F#value taken for the high-power design 

11.18, 4150 cycles/radian is then the cutoff spatial frequency. 

E. Formation of Sensitivity Coefficients - Sensitivity coefficients are computed but usually 

stored internally by optical automatic design programs, such as the General Electric AutoSpryte Program. In solving an automatic design problem, the program seeks to find a 

solution or something approaching a solution of an equation of the form. 

FkP1,P2, P.) = Sk 

The P1, P

2 P. are the j-designated variables of the system. In this case, they are 

the curvatures and air spaces/glass thickness of the system. The Skis some designated 

target value, representing the desired system performance. Fk(P1, P2, Pj

) = Sk 

says that the target value Skis a function, Fk, of the system parameters, Pi, P2 P.. 

To solve this equation, the program employs an iteration technique which, as a method of 

solution, computes the partial derivarives. 

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F

k F

k 

F

k 

P1 ' aP2 ' aPj 

These partial derivatives are the sensitivity coefficients. 

For this determination, the target value was the slope angle at the exit pupil of the 30-degree 

chief ray. The chief ray was chosen since it is the mean ray of the bundle, and ideally the 

other emerging rays of the bundle should have an identical slope angle. The sensitivity 

coefficients resulting from this analysis were tabulated in Table 2.1-4. 

F, Treatment of Manufacturing Tolerances - Manufacturing tolerances have been assigned 

by Itek as shown in Section 2.1.1.1. In assigning these tolerances, Itek has been careful to 

emphasize that the assigned tolerances are tentative, and that the problem of tolerances will 

be given more serious consideration prior to the critical design review. The resolution 

level obtained for nominal designs is, in general, overoptimistic since all the tolerances 

are zero and the elements are assumed to be perfectly aligned. The performance level obtained when manufacturing tolerances are included in a design is the realistic level and was 

the one used to determine the estimated ground resolutions reported in Section 2.1.2.4.C. 

The difficulty in making a tolerance analysis is knowing how to combine the tolerances to 

compute the physical frequency response. There is a multitude of combinations over the 53 

surfaces of this design which are possible. What is required to simplify the problem is to 

determine if there is a grouping of elements in the design which most significantly affects 

the total system performance. If so, the emphasis of the analysis should be placed on them. 

It was pointed out in Section 2.1.2.4.D that the first four surfaces which comprise the 

doublet contain 89.64 percent of the relative sensitivity coefficients, compared to 5.96 

percent contained by the zoom relay, and 3.87 percent contained by the eyepiece. This says 

that the effect of the doublet tolerances should be the most significant on the total system 

performance level. To substantiate this prediction, the total system physical frequency 

response of the high power 127K design was computed for the following three cases. 

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a. Tolerance budgets a, b, c, d applied to the doublet portion of the design. 

b. Tolerance budget "e" applied to the zoom relay portion of the design. 

c. Tolerance budget "f" applied to the eyepiece. 

In applying these tolerances, it became necessary to adjust the three focal distances in the 

design to obtain optimum imagery. The adjustments are given in the following table (refer 

to Table 2.1-1 for surface notation): 

Tolerance Budget Adjusted Airspace New Value 

a AS-8 1.237 

b AS-8 0.354 

c AS-8 0.788 

d AS-8 0.701 

e AS-31 1.256 

f AS-32 0.216 

The generated physical frequency responses are shown plotted in Figure 2.1-9 and Figure 

2.1-11. A comparison of the curves demonstrates that doublet tolerances have a more 

significant degrading effect on system performance, as predicted by the sensitivity coefficients. It then follows that the performance level of the proposed design can be investigated 

with a high degree of confidence by applying the assigned manufacturing tolerance just to 

the doublet portion of the design. This is a significant simplification to the tolerance 

analysis. 

The assigned manufacturing tolerances of the first four surfaces which comprise the doublet 

are given in Table 2.1-1. 

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The four tolerance budgets, a, b, c, d, have been arbitrarily formulated as typical groupings of tolerances which could be distributed over the doublet. Two combinations (a, b) have 

the radii of curvature tolerances alternating in sign while two (c, d) have tolerances all of 

the same sign. The physical frequency responses for the four cases were computed; for the 

high power design 127X and plotted in Figure 2.1-11. Budgets c and d which do not alternate 

in sign (shown as curves 3 and 4) have a less degrading effect than budgets a and b (shown as 

curves 1 and 2) which do alternate in sign. 

By inspection, tolerance budget "a" is the worst case. It then follows that a worst-case 

tolerance budget for the complete system would have to include tolerance budget "a". To 

determine this worst-case performance level, tolerances were applied over the entire design. The physical response function was computed, after focal adjustments were made for 

optimum imagery, for the a, e and f budgets combined with tolerances of the remaining 

surfaces of the objective assembly. This combined budget is denoted a .The result is 

shown in Figure 2.1-10 as curve 1. If such degraded resolution ocurred in the laboratory, 

it would be normal procedure to tweak the doublet airspace. To simulate this procedure, 

tolerance budget "a" was treated again. A significant improvement was found for a doublet 

airspace of 0.2755 inch and reticle separation of 0.725 inch. The improved response, 

compared to the original response, is given in Figure 2.1-12. The zoom-relay and eyepiece tolerances were then combined to this modified budget. The resulting response is 

curve 1 of Figure 2.1-13. This curve represents the worst case found and was the budget 

used to generate the performance estimate curves for each magnification. 

■■•••■,, 

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2.1.3 GE ANALYSIS OF OPTICAL PERFORMANCE UNDER DYNAMIC CONDITIONS 

Introduction 

This section documents the results of a GE study to analytically predict ATS optical resolution under dynamic (jitter) conditions. The study presents a method of treating the degrading 

effects of jitter on image modulation curves presented, demonstrating ground resolution as 

a function of jitter amplitude. A curve is also included showing the maximum jitter amplitude 

as a function of frequency which will permit 3. 6-foot ground resolution to be achieved. 

The method employed has been derived by GE and is not the same as that previously 

employed by Itek. The results are similar, however. 

An interesting secondary conclusion of the analysis is that the static resolution can not be 

rigorously obtained from the intersection of the MTF curve and the eye threshold curve. 

It is shown that the actual resolution is somewhat better than predicted by the intersection 

of the curves thus providing a safety margin in the analysis presented in the previous section. 

The Itek MTF curve is used as the analytical model for the telescope. 

The specific objectives of this performance study were: 

a. Analytically determine the dynamic performance of the acquisition telescope 

and relate the dynamic performance to that obtained under static conditions. 

b. To specify the magnitude of half-amplitude jitter, as a function of jitter 

frequency, below which resolution of a three bar, 3. 6 feet per cycle, 2 to 1 

contrast target (viewed from 80 nautical miles, at nadir) is possible. (Note: 

Unless otherwise specified, all jitter amplitudes referred to in this report 

will be half-amplitude values). 

c. To verify the methods and results from the Acquisition Subsystem Subcontractor, 

Itek Corporation, in the areas of static and dynamic performance analysis. 

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1.0 

0.2 

0 

0 0.75 0.90 1.05 1.20 0.15 0.30 0.45 0.60 

ANALYTICAL MTF - e 

ACTUAL SYSTEM MTF - A(w) 

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The analysis was conducted using the final lens design Modulation Transfer Function (MTF) 

provided by Itek Corporation in their Acquisition Optics Subsystem PDR Report dated 25 

September 1967. (See Figure 2.1-20.) 

The Schade visual resolution threshold characteristic specified by Itek in the same report 

was also adopted for use in this analysis. (See Figure 2.1-21.) A literature search was 

conducted to determine if experimental data was available to establish a more reliable eye 

resolution threshold characteristic; however, no data was found which appeared to be more 

reliable. It is imperative that a reliable eye resolution threshold characteristic be generated 

in order to accurately predict both static and dynamic performance of the Acquisition Subsystem. Itek Corporation is presently engaged in the design of an experiment that will yield 

the necessary data on eye modulation and contrast threshold. 

SPATIAL FREQUENCY (CYCLES/FT GND) 

Figure 2.1-20. Acquisition Telescope Transfer Function 

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50 

PERCENT MODULATION 

40 

30 

2 

10 

9.05 0.10 0.15 0.20 0.25 0.30 

SPATIAL FREQUENCY (CYCLES/FT - GND) 

0.35 0.40 

Figure 2.1-21. Low Contrast Visual Threshold (Schade 1964; MDM = 0.015) 

It should be clearly understood that the intent of this study is only to verify that the Acquisition Subsystem meets the system performance requirements as contained in the GE-AVE 

CEI Specification. No attempt was made to predict system performance under actual onorbit operating conditions. 

For example, in reality, jitter imposed on the system will have a random spectrum (i. e. , 

at any given time will be composed of many frequencies with definite phase and amplitude 

relationships) and will be two dimensional in nature. 

For purposes of this analysis, only the degradation in resolution due to discrete frequencies 

was considered. Also, motion of the line-of-sight (LOS) was reduced to motion in a plane 

perpendicular to the axis of the bars being viewed. These assumptions are valid for purposes of verifying that the system design meets the specification requirements, but they are 

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not valid when predicting system performance on-orbit. Consideration of these additional 

factors will further reduce system resolution. It is felt that work in the area of system 

performance prediction should be continued to further define system resolution under actual 

operating conditions. However, a more realistic eye threshold characteristic must be 

established before this can be done. 

2.1.3.1 Target Characteristics 

The analytical results obtained for the Acquisition Telescope static and dynamic performance 

were derived for a target with the characteristics specified below: 

a. Three light bars on a dark background 

b. Contrast ratio of 2 to 1 at the input to the optical system 

c. Apparent average brightness of 530 ft. -lamberts at the input to the optical 

system 

For purposes of this analysis, it was assumed that the target is being viewed from 80 

nautical miles (at nadir) with a magnification of 127X. Figure 2.1-22 is a graphic representation of a target with a spatial period of 3.6 feet per line pair on the ground (T = 0.9 feet). 

2.1.3.2 Summary of Static Performance Analysis 

Results obtained from analysis of the Acquisition Telescope performance under static conditions show that for a tri bar target input (square wave brightness distribution) with a 

spatial period between 3.0 and 4.2 feet per cycle, the image formed by the telescope has a 

sinusoidal brightness distribution. For a given ground spatial period in this region the output image contrast and modulation may be determined fromFigure 2.1-23. This is based on 

three bar target input with characteristics as described above. Using the Schade threshold 

specified in Figure 2.1-22, the static resolution of the Acquisition Telescope is 3.12 feet 

per cycle on the ground. This result is considerably lower than that calculated by Itek. 

They obtained a minimum resolution of 3.3 feet per cycle from the direct intersection of 

the telescope Modulation Transfer Function and the Schade Threshold multiplied by a 

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2 

0 

I I I 

2T 4T 6T 8T 10T X 

0 

T = 0.9 FT 

Figure 2.1-22. Graphic Representation 

3.0 3.2 3.4 3.0 3.8 

TARGET SPATIAL PERIOD (FT/CYCLE) 

4.0 4.2 

Figure 2.1-23. Acquisition Telescope Static Performance 

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factor of three (multiplication by three is necessary since the MTF is normalized to 1.0, 

thus representing a high contrast input). A direct intersection of these two curves is not 

possible unless the input and output brightness waveforms are considered. 

The increase in system resolution can be most easily understood by analyzing the static 

system performance using a target with an infinite number of bars. In this case, a 2 to 1 

contrast square wave input to the Acquisition Telescope produces an output waveform equivalent to that obtained for a 2.5 to 1 contrast sinusoidal input. In other words, the first harmonic of the square wave brightness distribution has a greater contrast ratio than 2 to 1. The 

result is a higher output modulation at the specified threshold frequency than computed in 

the Itek analysis. 

2.1.3.3 Summary of Dynamic Performance Analysis 

ECP Number 17 to the GE-AVE CEI Specification contains a curve of acceptable half amplitudes of oscillation of the LOS as a function of frequency. This is reproduced in Figure 

2.1-24 (solid curve). Analysis in this report shows that a system designed to meet this 

allowable jitter characteristic will have the capability of resolving a 3.38 feet per cycle 

three bar target on the ground. 

Also included in Figure 2.1-24 (dashed curve) is the half amplitude jitter as a function of 

frequency that will produce a minimum resolution of 3.6 feet per cycle on the ground. The 

allowable jitter specified by the broken curve between 0.001 and 1.0 Hz is conservative 

since it is based on the value of jitter which begins to produce resolution degradation. The 

shape of the characteristic between 1.0 and 6.0 Hz has been estimated. Degradation in 

resolution due to jitter in this region is a function of both amplitude and frequency. Therefore, it is not possible at this time to analytically predict system performance. 

Figure 2.1-25 (solid curve) shows the minimum system resolution as a function of jitter 

half amplitude (above 6.0 Hz). Due to the numerical integration techniques used to generate 

this curve, accuracy below five-foot minimum resolution is approximately three percent. 

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100,000 

10,000 

rn 

1,000 

H 

P. 100 

H 

H 

10 

0 z 1 

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\\ 

_ 

ii k 

I 

1 

1 

1 

1 

I 1 

I 

1 

1 

G.0 

Figure 2.1-24. Acceptable Half Amplitudes of Oscillation 

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10 100 

45,000 

20,000 

560 

0.35 

0.25 

0.1 

0.001 0.01 0.032 0.1 1 

FREQUENCY - Hz 

ALLOWABLE JITTER. FROM ECP 17 TO GE-AVE CEI SPEC. 

— — — ALLOWABLE JITTER FOR 3.6 FT/CYCLE RESOLUTION 

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N 

0 °0f c°0 

0 0 0 

CD Cr, CO t••• CD 

0 

UD mt. 

(J.Naouad) Nouvavuosa mounlosali 

0 01 CO C• 4D O CO 

0 0 0 0 0 0 0 

CD CO CO CO III 

O O 

N 

O co co t- t0 o N 

9-4 

III 1 1 I 1 I 111 1 I I 1 I MANUFACTURING 

TOLERANCE= 0. 152 X 

(STATICRES =3. 12 FT) 

7 

SYSTEM PERFORMANCE USING 

ITEK METHOD OFANALYSIS 

......•••••(STATIC RESOLUTION = 3. 3 FEET) R 

MANUFACTURING TOLERANCE 

EQUALS 0 X. 

(STATIC RES =2. 64 FT) 

1 l I 1 I1 1 1 1 I 1 1 1 1 1 1 

RESOLUTION 

1----.........................< DEGRADATION 

(PERCENT) 

111 1 1 1 1 1 111 1 1 1 

I ... 

1 

Eli I 

(3'I3AO/J,3) NoilrYlOsau aisznotiO AinpuxiNi 

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Above five-foot, accuracy, is approximately seven percent. From this curve it may be 

seen that as the maximum allowable jitter half amplitude increases above the 0.35-to 0.45-

arc second region, a sharp decrease in system resolution occurs. 

The System MTF used to generate the solid curve contains a manufacturing tolerance of 

0.152 l . The broken curve in Figure 2.1-25 shows the minimum system resolution with a 

manufacturing tolerance of 0 A. 

A comparison of the two curves shows that under static conditions, the minimum resolution 

is degraded 19 percent by a manufacturing error equivalent to 0.152 A . However, as the 

maximum allowable jitter is increased, the effect of the manufacturing error becomes less. 

The system degradation in percent as a function of maximum allowable jitter amplitude is 

also shown in Figure 2.1-25. For example, a system with a maximum allowable jitter of 

0.6 arc seconds will show a 2.3 percent decrease in resolution as the manufacturing and 

alignment errors increase from zero to 0.152 A. 

Also shown in Figure 2.1-25 (dash curve), is the minimum system resolution as a function 

of jitter obtained using the Itek Corporation method of analysis. 

2.1.3.4 Static Performance Analysis 

E. L. O'Neill(1) demonstrates an analogy between linear time filters and optical spatial filters 

(conventional optical systems with lenses, prisms, etc.). Using this analogy, it is possible 

to specify an optical system in terms of its spatial frequency characteristic or transfer 

function G(w). The output of such a system at any frequency womay be represented as 

G(wo) times the input. The output wave form may then be found by use of the inverse 

Fourier Transform. 

(1) O'N

eill, E. L. , Introduction to Statistical Optics; Addison-Wesley Publishing Co. , 

Inc., 1963 

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F

o

(w ) =G(w) F. (w) 1(

) 

R0(w)e icPo(W) = A(w) eje(w) Ri(w) ej(41(w) 

fo(X) = F

-1 

[F

o(w)] 

F

-1 ei [e(w) + coical = [A(w) . R.(w) . f(X) 

The three bar target of Figure 2.1-22 has a continuous frequency spectrum described by: 

Fi(w) = 

= 

2AT 

2AT 

SI W) T [ -j WT 

+ 

e 

-jWT] re 

e-j5wT 

+ e 

-RLIT 

+ e 

j - 5wT 

-j4wT 

e 

-j9w 

+ 

T 

e 

-j9wT1 

e 

-j6WT 

(N 

WT 

[ 

e

jWT 

tor 

A 

[

2j 

[ e-jOWT

e 

-j2wr 

ice jw 

e 

-j8wr 

jw 

e

-jlOwT 

jw 

ice jw 

An attempt was made to find an analytical function to represent the magnitude A(w) of the 

Acquisition Telescope transfer function specified by Itek Corporation with no regard for the 

phase characteristic. Due to the nature of the analyses, neglecting the phase of G(w) will 

have a negligible effect on the results. A Fourier Spectrum with a magnitude equal to an 

exponential decay e-a/f/ and o = 6 provides a good fit for A(w) over the spatial frequency 

region defined in the PDR report. A comparison of e-04/f/ and A(w) is shown in Figure 2.1-20. 

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Now 

Fo(w) = G(w) Fi(w) 

a 

IPI [-OPT 

21r e e 

-2PT 

= A e - P P + 

- PT 4 - PT 6

e e 

P P 

+ 

e 

-8PT e 

-10PT 1 

P P 

where 

P = jw 

The output function fo(X) can be obtained by taking the inverse transform of the above equation. The analytic l expression for the resultant is an extremely complex equation and is 

difficult to interpret. A graphical interpretation is much more meaningful and is obtained 

as follows: 

g(X) = 

fi(X) = 

fo (X) = 

The graphic interpretation of the convolution of g(X) and fi(X) is shown in Figures 2. 1-26a 

through e. 

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X 

a. 

1 

IT a 

g(X, - X) - a (XiIT [(X, - X) 2+a0 X1 X 

b. 

2.0 

fi(X) 

A l X 2 X3 X 4 X 5 

c. 

Figure 2.1-26. Graphic Interpretations 

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The product of. g(X, -X) and fi (X) is: 

2 

ira fi(X) g(X, - 

1 

X 

A O X1 

X

2 X1 

Figure 2. 1-26d. Graphic Interpretations 

The ordinate of the output waveform fo(X) at X = X1 is the product of the two functions with 

= X. For an input square wave with a period of 3.6 feet per line pair (7 = 0.9 feet) 

the resultant output waveform is: 

1.510 1.538 

1.510 

1.307 

] 1.307 

fo(")2 

1.083 

\\Ik 

1.5 

1 0 

1.083 

-2T 2T 4T 6T 8T 10T 12T 

X1 

Figure 2. 1-26e. Graphic Interpretations 

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Figure 2.1-26e shows that under static conditions, the acquisition telescope degrades a 3.6 

feet per cycle three bar target input with 2 to 1 contrast and 33 percent modulation to a 

low contrast three cycle sinusoidal waveform with an average contrast of 1.17 to 1 and 

average modulation of 7.6 percent. 

Figure 2.1-23 is a curve which shows output waveform modulation and contrast, for a 2 to 1 

contrast three-bar target input, as a function of spatial frequency on the ground. When the 

Schade modulation Threshold characteristic is superimposed, the two curves intersect at 

3.12 feet which means that under static conditions the smallest three-bar target that can be 

resolved is 3.12 feet per cycle. This contradicts the 3.3-foot result presented by Itek 

Corporation in their PDR report. The reason for this discrepancy is most easily explained 

by performing a similar image waveform analysis for a bar target consisting of an infinite 

number of light and dark bars (instead of three). This is a much simpler problem since a 

periodic function consisting of an infinite number of pulses may be represented by the following series: 

1 2 fi(X) = 2Aav 7 ( + 77 COS y - 37r COS 3y i- 51T COS 5y 

) 

where 

2 TT X 

y - 

T 

and A

av 

is a function of average brightness. 

This function has a discrete Fourier spectrum of the form: 

SIN Ceir Fi(co) = A cor 6 (co - co g) 

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• 

..1•■■ 

• 

5/T 

\ I I. 

1/T %..1./ 3/7 

; 

• 

• 

-3/7 I -1/T -w -5/r 

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where 

cog = 21rnfi 

n = 1, 2, 3, 4, 

f. = Frequency components of the Fourier series representation 

A plot of this Fi(w) is shown in Figure 2.1-27. 

Operating on this discrete frequency spectrum with the transfer function A(w) of the telescope shows that each frequency component is multiplied by the ordinate of the transfer 

function at that particular value of spatial frequency. The resultant is a brigthness distribution represented by the equation: 

Fi(W) 

Figure 2.1-27. Plot 

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%(X) = 2 A

av 

(0.190) 

2 

IT 

COS y 

- (0.006) 

2 COS 3y + 3 IT 

The amplitude of the fundamental frequency term is a factor of 80 times the amplitude of 

the third harmonic term. Therefore, all higher order terms may be neglected and 

fo(X) = 2 Aav [ (1) -j-- + (0.190) cos 

2 IT X 

2 ff 

This shows that a 2 to 1 contrast infinite bar target input produces an output waveform equal 

to the output produced by infinite sinusoidal target input with 2.5 to 1 contrast. Therefore, 

the direct superposition of a threshold curve on a modulation transfer function to yield 

system resolution is not valid unless the type of input and output waveforms are considered. 

2.1.3.5 Dynamic Performance Analysis 

Oscillation of the acquisition telescope LOS with respect to a point in the object plane produces a point in the image plane which appears to oscillate at the same frequency as the 

LOS oscillation but with an angular amplitude of 127 times that of the LOS oscillation amplitude. The same image motion result is obtained with a static LOS and an oscillating target. 

Therefore, for purposes of analysis, all jitter will be introduced as oscillation of the target 

with respect to a static LOS. 

To analyze the effects of jitter on system resolution, jitter frequencies in the range 0.001 

to 100 Hz were considered. ECP number 17 to the GE-AVE CEI Specification contains a 

curve of acceptable amplitudes of oscillation of the LOS as a function of frequency. This 

characteristic was shown in Figure 2.1-24 (solid curve). This spectrum was broken into 

four regions (0.001 to 0.032, 0.032 to 1.0, 1.0 to 6.0, and 6.0 to 100 Hz) and the allowable 

jitter amplitude determined in each region. 

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Low frequency, large amplitude jitter (in the 0.001 to 0.032 Hz range) makes the image 

appear to move across the field-of-view at a constant rate. Therefore, the factor affecting 

resolution due to jitter in the 0.001 to 0.032 Hz range is the linear velocity at which the 

eye must tract the image. The maximum linear image velocity with respect to the eye for 

the allowable jitter specified of Figure 2.1-24 was calculated and found to be 4.5 degrees 

per second. Data presented in the Bioastronautics Data Book (1964) indicates that very 

little degradation in resolution is introduced for velocity up to 10 degrees per second. This 

corresponds to a jitter half amplitude of approximately 45,000 arc-seconds at 0.001 Hz and 

1,400 arc-seconds at 0.032 Hz. 

Jitter in the frequency range 0.032 to 1.0 Hz will produce small sinusoidal oscillations of 

the image with respect to the eye. Again, this motion will not produce a significant resolution degradation around 0.032 Hz if the angular rate of image motion is less than 10 degrees 

per second. Experimentation done by Jones & Draxin and Guignard & Irving show that for 

frequencies around 1 Hz, the eye can track sinusoidal motion with half amplitudes up to 0.5 

degree with respect to the eye. This corresponds to a jitter amplitude of approximately 

14 arc-seconds. 

The next range of jitter frequencies to be considered is the 1.0 Hz to 6.0 Hz range. Figure 

2.1-28 is a plot showing the ratio between actual motion of the eye when viewing an oscillating target and the motion of the target with respect to the eye. This characteristic shows 

that in the 1 to 6 Hz region, degradation of the image resolution is a function of both jitter 

amplitude and frequency. Because the degradation is a function of both variables, an 

analytical performance prediction cannot be obtained. 

When the frequency of oscillation of a point in the object plane is above 6 Hz, the point in 

the image plane viewed by an observer appears to be a continuous line. Therefore, image 

degradation of a point oscillating above 6 Hz is independent of the frequency of oscillation. 

The primary factors that contribute to the degradation are oscillation amplitude and waveshape (i. e. , sinusoidal, linear, random, etc.). For this analysis, the degradation effects 

of both linear and sinusoidal jitter were investigated. 

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1.0 

0.8 

0.6 

0.4 

0.2 

0 

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RELATIVE AMPLITU

DE 

1 2 3 4 5 6 7 

FREQUENCY - Hz 

Figure 2.1-28. Ratio Plot 

The brightness distribution characteristic for a three-bar target (the object) as a function 

of half amplitude linear and sinusoidal jitter is shown in Figures 2.1-29 and 2.1-30, 

respectively. The object waveform contrast ratio as a function of jitter amplitude is shown 

for linear and sinusoidal jitter in Figure 2.1-31. This shows that the contrast of the object 

remains constant up to and including amplitudes of jitter which produce peak-to-peak LOS 

excursions on the ground equal to half the spatial period of one target cycle. The object 

contrast then decreases to 1 to 1 as the p-p LOS excursion increases above this value. The 

contrast of the image formed at the output of the telescope begins to degrade immediately 

as jitter is introduced. This is due to the fact that the Fourier components of the object 

brightness characteristic (shown in Figures 2.1-29 and 2. 1-30) change as target motion is 

introduced. For example, when the linear half amplitude jitter is such that the brightness 

distribution of the object is a triangular wave (half-amplitude jitter (a) equals r), the 

Fourier expansion of this brightness distribution (assuming an infinite target) may be represented as: 

fi(X) = 1.5 4 1 

COS y +-L 25 COS 3y + cos 5y + 

9 

IT 

2 

where 

2 ir X 

Y - T 

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14— 2Tt 2T -el 

7 -a 

2(7+a)-01 

2a 7-a 

a = 7- 

le-- 2 T 

a-7 

7 /a 

a > 7 

1- /a 

a- 7 

Figure 2.1-29. Target Brightness Characteristics as a Function 

of Linear Jitter Amplitude 

2-62 

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a= 

a < T 

A 

/ 

1 

a > T 

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2-63 

l

os- 2 T le— 2T -11 

a= 

—NH I.— a -T 

a - T 

Figure 2.1-30. Target Brightness Characteristics as a Function 

of Sinusoidal Jitter Amplitude 

-S-EREPf SPECIAL HANDLING 
1.6:1 

SINUSOIDAL JITTER 

T/2 T 3T/2 2T a 

-4 

2:1 

1.8:1 

LINEAR JITTER 

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HALF-AMPLITUDE JITTER 

Figure 2.1-31. Object Waveform Contrast as a 

Function of Jitter Amplitude 

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UNDEMENTALFREQ. DEGRADATION 

1. 144: 0.8 

0 

1,108:1 0.6 

Cd 

0 1.072:1 

0 

0.4 

1.036:1 0.2 

_ 1:1- 

1.180:1 

z 

4.56 

z 0 

E. 

3.01 

A 

O 

7.60 

6.08 

1.52 

0 

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■ •••••■.... 

As was the case for the square wave brightness distribution (zero jitter), harmonics above 

the fundamental frequency may be neglected. Therefore, with linear jitter of a = 7, the 

image at the output of the telescope (for a spatial period of 3.6 feet per cycle) has a sinusoidal brightness distribution with contrast and modulation reduced from 1.18 to 1 and 8.0 

percent (at zero jitter) to 1.10 to 1 and 5.1 percent. 

Figure 2.1-32 shows the effect of half amplitude jitter (both sinusoidal and linear) on a 2 to 

1 contrast 3.6 feet per cycle three-bar target. This figure demonstrates that 0.35 arcsecond half amplitude sinusoidal jitter will yield a sinusoidal brightness distribution with 

1.10 to 1 contrast and 4.3 percent modulation. This corresponds to visual threshold as 

defined by the Schade Threshold characteristic. 

This analysis was extended to include spatial periods from 3.0 to 4.2 feet per cycles on the 

ground. The result of this analysis for the sinusoidal jitter is shown in Figure 2.1-33. 

0 0.125 0.25 0.375 0.50 0.625 0.75 

HALF-AMPLITUDE JITTER - (ARC-SEC) 

Figure 2.1-32. Image Modulation and Contrast as a Function of Jitter 

(3.6 Ft/Cycle Tri-Bar Target) 

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10.0 

MODULATION (PERCENT) 

8.0 

6.0 

4. 0 

2.0 

0 1:1 

3.0 3.2 3.4 3.6 3.8 4.0 4.2 

TARGET SPATIAL PERIOD (FT/CYCLE) 

JITTER 

0 ARC-SEC 1.223:1 

STATIC 

PERFORMANCE 

1.179:1 

- O 1.135:1 

a 

r- 1.091:1 

_ 1.046:1 

SCHADE 

THRESHOLD 

LOWRY & DEPALMA 

THRESHOLD 

0.25 ARC-SEC 

0.35 ARC-SEC 

0.45 ARC-SEC 

0.55 ARC-SEC 

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Figure 2.1-33. Acquisition Telescope Dynamic Performance 

Curves for Sinusoidal Jitter 

The curves depicted were calculated using a numerical integration technique to convolve 

the three-bar brightness distribution waveforms shown in Figure 2.1-30 with the transfer 

function of the Acquisition Telescope in the spatial frequency domain. Comparing the 

results obtained using the numerical integration techniques with an analytical solution of 

zero jitter and for an infinite target indicated an accuracy of three percent. 

Figure 2.1-33 shows the degradation in system performance from the static performance 

(both image contrast and modulation) as a function of target period on the ground for discrete 

half amplitudes of jitter (0.25, 0.35, 0.45, and 0.55 arc-seconds). Also drawn on this 

graph are two eye threshold curves, the Schade Threshold previously discussed and a more 

optimistic curve by Lowry and DePalma(2). Using the Schade curve, a three-bar target 

(2)DePalma, J. J. and Lowry, E. M. ; Journal of the Optical Society of America, 

Volume 52, March 1962, Page 328. 

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with a spatial period of 3.38 feet per cycle is the minimum size that can be resolved if the 

system is acted upon by half-amplitude jitter of 0.25 arc-second. The maximum jitter that 

can be experienced to allow a resolution of a 3.6 feet per cycle target is 0.35 arc-second. 

It is felt that the data obtained by Lowry and DePalma to generate their threshold curve is 

optimistic. However, if this curve is used, 0.53 arc-second of jitter can be tolerated to 

obtain a minimum resolution of 3. 6 feet per cycle. 

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2.1.4 SECONDARY DESIGN PERFORMANCE INFORMATION 

2.1.4.1 Vignetting 

This subsection describes analysis performed to determine the amount of vignetting present 

in the design. The description is presented in two parts. 

A. Internal Telescope Vignetting - To determine if there is any internal vignetting in the 

telescope, the following analysis was performed by Itek. 

The vignetting in the telescope was calculated at the peak wavelength (0. 56 microns) and 

also throughout the visible (0.47 to 0.65 micron) spectral region. At the peak wavelength, 

253 rays were traced at normalized field positions of 0.0, 0.5, and 1.0, since the transmission specifications are given at these field positions. To calculate the vignetting over 

the visible spectral region, 70 rays were traced at each of the following wavelength, 0.65, 

0.605, 0.56, 0.515, and 0.47 microns. In both the monochromatic and polychromatic cases, 

the rays were traced at six magnifying powers; namely, 127X, 84.67X, 63.5X, 31.75X, 

21.17X, and 15.88X. 

The percent vignetting for all these different cases is given in Table 2.1-4 where percent 

Rays in - Rays out vignetting - x 100. Rays in 

This shows that only at the condition of low zoom power is there any internal vignetting and 

this is not expected to significantly effect resolution. 

B. External Component Vignetting - A complete study of vignetting (including the effects 

of external components) is in progress. 

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RELEASE 1 JULY 2015 -S-EGRET- SPECIAL HANDLING PERCENT VIGNETTING = O. 5 =1.0 

0 

0 

0 

. 

0 

0 

0 

0 

0 

CV 

CS; 

0 

0 

0 

c; 

0 

0 

0 

c: 

0 

0 

C 

c; 

0 

0 

POLYCHROMATIC 

= 0. 0 

o 

. 

0 

0 

0 

0 

0 

0 

0 

0 

. 

0 

0 

PERCENT 

VIGNETTING 

= 1.0 

0 

0 

0 

0 

CV 

0

. 

CV 

0 

 0 

0 

0 

0 

cl; 

I-1 

MONOCHROMATIC 

= 0. 0 = 0. 5 

0 

0 

o 

. 

0 

. 

o 

d 

0 

. 

o 

d 

0 

0 

0 

d 

0 

0 

0 

0 

0 

0 

0 

0 

MAGNIFYING 

POWER 

X t••■• C. q 

,--I 

X C.- CO 

't 00 

X C) LO 

CO 

0 

X In N 

1-4 

ea 

X C-• r•-i 

1-.1 

CV 

X CO 00 

ld 

,--I 

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.......... 

TABLE 2. 1-4.PER

CENT VIGNETTING 
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2.1.4.2 Gimbal Angles Versus Scan Field Position 

Line of sight and gimbal angles required for the scan field coverage were derived by Itek 

for the nine-degree total cant configuration and are presented in tabular form. 

The results are listed in Tables 2.1-5 and 2.1-6. The tables give either the LOS or gimbal 

angles for the nine-degree cant angle. The top chart shows the extreme or worst case 

angles alone with the corresponding LOS angles with respect to vehicle axes. This gives 

the pitch and roll ranges required to cover the entire + 70 degrees to - 40 degrees in pitch 

and ±45 degrees in roll. The lower chart shows the angles corresponding to various points 

of interest within the total range. Reference to Dwg. 905842 shows that the required (and 

computed) gimbal angles were used in the design of the scanner gimbal assembly. However, 

further analysis is needed to determine if the LOS intersects the Lab Module contour. 

2.1.4.3 Exit Pupil Shift 

The exit pupil shift as a function of zoom position, or magnification, is shown in Figure 

2.1-34. The specification limits are also shown on this figure. The curve shown in Figure 

2.1-34 clearly indicates that the specification requirements on exit pupil shift have been 

met. A more detailed discussion of this topic can be found in the Itek PDR Report. 

2.1.4.4 Total Transmittance Summary 

The total transmittance of the complete AO Telescope system has been calculated by Itek 

(see Table 2. 1-7) and found to be 35.4 percent for the 127X (high power) mode and 31.2 

percent for the 31.75X (low power) mode. The following assumptions were made: single 

layer magnesium fluoride anti-reflection coatings on all refracting surfaces, both external 

mirrors are aluminized and have a silver oxide protective coating (87 percent reflection 

each). The internal folding mirror aluminized with a two-layer enhancement coating (93 

percent reflectance), and multilayer coatings (98.5 percent reflection) on two surfaces of 

the Pechan Prism. The transmittance calculation takes into account the expected reflection 

2-70 

SECRETSPECIAL HANDLING 
+55.85 

o 

co 

W 

A- 

+68.14 

co 

cu 

ri 

cu 

+ 

+68.25 

co 

- 

a; 

N I 

+24.13 

1-4 

N 

GO 

7 

+39.41 

co 

..4' 

L-: + 

+36.81 

co M 

'Wc 

+0.71 

co 

't. 

iti 

.14 

+ 

+9.88 

co 0 

+ 

+13.41 

m m 

V 

14 

-6.35 

0 

co 

;... 

±0 

$ 

+6.35 

0 

co 

4 

-36.81 

co 0 

O 

+ 

-39.41 

co 4 

` 3" 

-24.13 

cu 

cd 4 I 

I 

+70 

+40 

PITCH 

(LOS Angles 

with respect 

to vehicle 

axes) 

+10 

0 

-40 

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Ittle SPECIAL HANDLING 

TABLE 2.1-5. LOS ANGLES FOR PLUS NINE—DEGREE CANT 

EXTREME ANGLES 

CORRESPONDING 

LOS ANGLES (WITH 

RESPECT TO 

VEHICLE AXES) 

ROLL PITCH ROLL PITCH 

-48.21 

+54.80 

-40.83 

+70.26 

-45 

+18 

-26 

+45 

-40 

-40 

+70 

+70 

i O -1' 

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TABLE 2.1-6. GIMBAL ANGLES FOR PLUS NINE—DEGREE CANT 

EXTREME ANGLES 

CORRESPONDING 

LOS ANGLES (WITH 

RESPECT TO 

VEHICLE AXES) 

ROLL PITCH ROLL PITCH 

-40.75 

+62.26 

+22.25 

+80.39 

-45 

+29 

-45 

+45 

-40 

-40 

+70 

+70 

+70 

+40 

PITCH 

(LOS Angles 

with respect 

to vehicle 

axes) 

+10 

0 

-40 

+69.05 

ua N 

oi 

co 

+ 

+76.58 

co 

uo 

+.7 

1 

+80.39 

CO ol 

1; 

+53.47 

o 

co 

0 

+ 

+63.53 

to U) 

cO 1-1 

-I- 

+65.93 

co 

On 

e V); 

I 

+51.08 1" 

+49.17 

co to 

o .-1 

+ 

+54.45 

..t. o 

0 VI I 

+38.33 

c; 0 

+ 

+44.30 

.:0 to 

+50.95 

t- .0. 

C. V) I 

+23.12 

op 

o 

0 

+ 

+24.67 

t- 

+35.79 

ID 

0 

1 

0 

2 —72 

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130 OR 65 

120 OR 60 

PA 

1100R 1 55 

100 OR 50 

90 OR 45 

80 OR 40 

70 OR 35 

tiv 

o• 

EXIT PUPIL SHIFT INCHES 

60 OR 30 

-0.03 -0.02 -0.01 0 0.01 0.02 

0.7516 0.7616 0.7716 0.7816 0.7916 

lie 0.040 INCH 

SPEC 

0.7466 

Figure 2.1-34. Exit Pupil Shift as a Function of 200 M Position 

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TABLE 2.1-7. TRANSMISSION OF COMPLETE SYSTEM 

HIGH POWER MODE 

0.56 12 

(PERCENT) 

LOW POWER MODE 

0.56 1.1 

(PERCENT) 

Transmission due to reflection 

losses alone 

39.2 34.8 

Transmission due to absorption 

losses alone 

90.2 89.6 

Transmission due to combined 

effects 

34.4 31.2 

and absorption losses of each element, rather than the theoretically predicted values. 

The performance assumed for the single layer magnesium fluoride coatings, for example, 

is based on actual measurements made on coatings on different index glasses. Although 

the variation of reflection with glass index and angle of incidence of the light are considered, 

the effects of vignetting and filters (aside from surface reflections) has been ignored. 

0 

All values are given for 5600 A light. The theoretical data is obtained from a CDC 3300 

computer using the MULTIFILM program. The practical data are measured values for the 

anti-reflecting films and estimates for the reflectors. The estimated values are based 

upon prior knowledge of typical performance. 

2.1.4.5 Focus 

Means have been provided to insure that ground scenes are properly focused when the MOL 

altitude is between 75 and 230 nautical miles. At any given altitude the scene remains in 

focus throughout the scan field without manual refocussing. The eyepiece has a focusing 

range (manual) from - 2-1/2 to + 3-1/2 diopters to compensate for individual eye differences 

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between various observers. With linear defocus (moving the image plane off the reticle) the 

reticle may appear to lose some of its definition, causing the observer's eye to have to refocus as he concentrates on either the reticle, or the scene. This will be a cause for concern only if the reticle to scene defocus exceeds the depth of field of the eye. The eye has a 

depth allowance of one-half diopter so that any lesser depth will be within the focusing tolerance. 

Defocus errors were calculated by Itek for altitudes of 70 to 230 nautical miles, pitch 

angles from zero degrees to 70 degrees and roll angles from zero degrees to 45 degrees. 

The greatest error of any altitude will be for the maximum pitch and roll. 

ALTITUDE WORST CASE (INCHES) 

70 0.0008 

80 0.0007 

100 0.00056 

150 0.00038 

230 0.00025 

At 80 nautical miles (at high power), the worst case slant range defocus is 0.0007 inch 

at the objective lens focal plane, corresponding to 0.003 inch at the eyepiece object plane. 

At 230 nautical miles (at high power), the defocus is 0.00025 inch at the objective or 

0.00107 inch at the object plane. 

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The effects of the curvature of the earth has been calculated on the slant range with the 

following equation: 

d h + R (1 - cos y) - 

cos f3 

h = altitude 

R = Radius of earth 

y = Geocentric angle to intercept pt. 

g = Vehicle centered angle 

Beta ($) includes obliquity and stereo factors as well as roll and pitch angles and is 

measured on a plane through vehicle, earth center and intercept point. Gamma (y) is 

on the same plane and is measured as follows: 

y = cos 

-1 A

2

B ±11A

4

B

2 - (A2 + 1) (A2

B

2 - 1) 

A

2 + 1 

where 

A = Tang 

B = — +1 

The slant range at 70 degrees goes from 215 for a flat earth to 238 on a curved earth, a 

10 percent increase. The slant range at 230 nautical miles is 706.5 nautical miles for a 

flat earth and 708 nautical miles on a curved earth, less than one percent. The 10 percent 

increase will increase the defocus error by approximately 10 percent and the defocus will 

then amount to 0.0033. 

The defocus is less at higher altitudes, so we need only consider the 80 nautical miles 

worst case, i. e. , 0.0033 at the eyepiece object (reticle) plane. 

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The defocus effect on the eye of a 0.0033 difference between reticle and image corresponds 

to a 0.14 diopter error. The depth of field of the eye is 0.5 diopter which is four times as 

large as the image distance, so no visible effect is predicted. 

The requirement for a manual focussing range from - 2-1/2 to + 3-1/2 diopters is met by 

the mechanical eyepiece design as follows: 

The eyepiece of the AO Telescope is specially designed to compensate for residual aberrations of the zoom relay assembly. Because of this, the alignment tolerances are considerably tighter than the standard eyepieces found in military instruments. These tighter 

tolerances impose a restraint on the design of the focusing mechanism. 

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SECTION 2.2 

ATS CONTROLS 

This section presents servo-analysis pertinent to the ATS controls. A description of the 

operation of the controls is included in Section 1.3.4. The material in this section is subdivided as follows; 

2.2. 1 Scanner Controls 

2. 2. 2 Image Orientation Controls 

2.2.3 Zoom Control 

2. 2. 4 Blanking Servo 

2. 2. 5 Power Change Servo 

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2.2.1 SCANNER CONTROLS 

2.2.1.1 Requirements Interpretation 

The LOS of the acquisition optics shall be driven such that the nominal rate error (drift) is 

less than 540 microradians/sec at the end of slew. The LOS displacement (jitter) requirement about the nominal LOS is defined in Figure (2.2-1). This requirement must be met 

when the roll tracking rate is 0.1 degrees/sec or greater and when no vehicle transients 

are being induced into the vehicle by the Main Tracking Mirror or the ACTS operation. 

The interpretation of the jitter requirement is that gimbal displacements at frequencies 

below 0.032 cps are considered D.C. drifts and above 0.032 cps the gimbal displacements 

are weighted by the inverse of curve shown in Figure 2.2-1. 

The gimbal displacements are determined by relating the noise input (where the noise is 

defined in terms of its PSD signature) to gimbal displacements. This relationship makes 

use of a weighting function F (w) which normalizes the specified jitter amplitude over the 

frequency spectrum above 0.2 radians per second and the control loop transfer function 

0 (w) where 8(w) is the gimbal displacement and N(w) is the noise input. N(c4.9 

The Power Spectral Density is defined as 

cp (w) = lim 1— T T 

f 2 

f(t) e-jc4.)t dt 2 

where f(t) is the noise to be analyzed. 

SECRET SPECIAL HANDLING 

2-81 
O 

O 0 

0 r-1 

0 0 0 

O 0 

O 1-1 

awl C.0 

FREQ

UENCY CYCLE

S/SEC 

Fig

ure2. 2-1. JitterPerformanceEnvelope 

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W Lt 

9 

44 

z 

z

0 U 

C:4 

E-1 

0 1-1 

0 0 

o 

(vwois g) 

(oas - auv) 

aCIIIIrldIAIV - drIVH 

so'i 

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The weighting function F(w) is defined as 

0 F(w) -.027S (S/37 + 1) 

The closed loop transfer function for the gimbal displacement as a function of noise at a 

specific input point is defined as 

G(w) - e(W) 

The RMS gimbal displacement weighted with respect to the curve of Figure 2.2-1 is 

determined using the following relationship 

co 

eRms = f 2(43(w) . F(w) 2 . ) 2d 

0.2 

where 

e = radians of displacement 

(ft lbs) 

c13(w) Radian per second 

Radians G(W)(W) - ft lb 

F(w) = is dimensionless 

In determining the optimum servo configuration a "hand off" position was assumed for the 

crewman. This assumes no noise rejection or D. C. offset removal by the crewman. 

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2.2.1.2 Discussion of Approaches 

In attempting to design the scanner control loop to meet the stated requirements, three 

servo configurations are synthesized and evaluated. The three configurations are: 

a. Rate control loop 

b. Position control loop 

c. Combination position and rate control loop 

The rate loop configuration utilizes a gyro on the gimbal to sum the command and gimbal 

rates. The input to this loop is a rate command. 

The position loop which is the least complex of the three considered, utilizes position feedback via a position encoder. The same encoder is used in the slew configuration. The 

gimbal rate commands are integrated by digital techniques and the input to this loop is a 

position command. 

The position and rate configuration consists of a position loop with feedback via a position 

encoder on the gimbal. This configuration has a rate minor loop with feedback via a rate 

gyro on the gimbal. This configuration is the most complex of the three configurations considered, but offers the greatest flexibility with respect to minimizing extraneous inputs. 

The bandwidths of the configurations considered allow only one of the tracking configurations to be made part of the slew configuration during the slew mode of operation. 

The following three sections discuss the performance of each of the configurations in detail with respect to the scanner tracking mode requirements. 

2 -84 

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As a result of the studies of the three servo configurations it is concluded that a position 

plus rate servo configuration for the pitch axis and a rate servo configuration for the roll 

axis allows the following: 

a. The jitter requirement to be met (see Table 2.2-1). 

b. Provides flexibility for changing gains and bandwidths to minimize noise 

inputs as these become defined. 

c. Low sensitivity to errors in the command signal resulting from vehicle 

disturbances. 

The following three sections discuss the performance of each of the configurations in detail with respect to scanner tracking mode requirements. 

2. 2.1. 3 Rate Control Servo Configuration 

The rate control configuration is shown in Figure 2.2-2 and ultilizes a rate gyro to make 

the summation of the command rate (ei) and the gimbal rate (6G) to develop the rate 

error signal (E; e) where de = bi-b G. 

The open loop GH has a crossover frequency of 150 radians per second and allows the command signal to be followed with a dynamic error of less than 5 microradians per second. 

The crossover frequency was selected using the curves of Figures 2.2-2 and 2.2-3. 

The following analysis was used to determine the dynamic error. 

2. 2. 1. 3. 1 Rate Loop Dynamic Error 

The dynamic servo error will be less than 10 μ radians per second during the rate mode. 

This is determined from an analysis of the input function generated during the rate mode. 

The equivalent inputs are determined considering the derivatives of a sinusoidal input. 

-SURE+ SPECIAL HANDLING 

2-85 
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SECRET SPECIAL HANDLING NRO APPROVED FOR 

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1:4 

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Figure 2. 2-2. RateServo Config

uration 

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SECRET SPECIAL HANDLING 
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2 3 4 

a & Figure 2.2. -3. Plot of --2- , , 

2 2 

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2a 3 

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2-87 
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If 

• • • 

e 

M ... 

e sin co t 

M ... 

e 

co e cos cot 

e 

M ... 

8 

2 

w 

sin cot 

M ... 

e cos cot _ 

W 3 

Therefore the equivalent position amplitudes of the derivatives are 

M ... 

e 19 

s 

_ 

2 

e A 

M .. 

e 

2 

w 

M . 

e 

(43 

M - M .. M .. 

Where e, e, e

. 

 are the half amplitudes of the sinusoidal portions of those 

derivatives and w = 2ff — where T is the period of the equivalent sinusoid. If the fundamental

T 

position signal and the derivatives occur at the same time the equivalent sinusoid 

must be considered to be impressed on the system at the same time and hence the error 

are additive also. The split should be made so that the total error at that time, t, does 

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not exceed the allowable error, but the error split should be balanced to a slope that is 

realizable (i. e. , 0 db/dec, -20 db/dec, -40 db/dec, etc.). If E/R for the system is 

known the required value of — can be determined. 

C 

G 

R - C = E 

R - E = C 

For most systems R >> E, therefore R = C, and 

R = M sin cot 

at max value R = M 

and M C 

E E 

If the input is of such a form that the rate, acceleration, and jerk are not sinusoidal 

derivates of the position input, the input can be represented by equivalent position inputs 

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which have the characteristics of the several components. The middle range of frequencies 

of C/E can then be determined so that the error will be within the constraints dictated. 

When the requirements are written in terms of rate of change of output and input the 

several components of the actual input can be converted to equivalent rate inputs. 

R 

and SE is known, the value of K can be determined by the allowable 

value of Se and the macimum value of SR. 

Since the input is a sum of sinusoidal inputs, 

E _ E1 E

2 E

3 

R R1R2 R

3 

For sinusoidal inputs of rate are considered and the allowable input is considered to be 

1/4 of each component. Figure 2.2-3 shows the curves of — — 

a a a and — and 2 a 

a

2 

which were used to determine the following relationships: 

Position: 

= 0.0412 sin 0.0412 t (equivalent rate input) 

= 0.013 sin 0.1372t (from a/a) 

= 0.021 sin 0.412t (from ix/a) 

SECRET SPECIAL HANDLING 

2a3 

A

w

l 

Rate: 

A

w

2 

A 3 

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Acceleration: 

A

w

6 

A

co

7 

= 0.013 sin 0.1372t (from ti/a2) 

• • 

= 0.443 sin 0.0485t (from a /2a3 ) 

Jerk: 

A 4 = 0.00176 sin 0.275t (from '&72a3) 

A

w

5 = 0.000935 sin 0. 1372t (from .tY/2a3) 

These relations are used to develop the error locus shown in Figure 2.2-4. 

The locus indicates a maximum error of 1.6g rad at 0.137 radians for an unconditionally 

stable loop with an open loop gain of 10,000 at 0.137 radians and LOS rate inputs. Allowing a line-of-sight error of 5g radians, the necessary gain at 0.137 radians can be determined. Since the scanner error is equal to 2 x LOS error, the gain (K) required at 0.137 

radians is 

10 10,000 

1.6 

K = 625 

Therefore, designing the scanner pitch and roll servos such that the gain at 0.137 radians 

is a minimum of 1000 insures that a servo dynamic error requirement of 45 jj, rad/sec 

will be met. 

2-91 

SECRET SPECIAL HANDLING 
1 1 I I 1 I I I 

0 

00 

0 

CV 

0 

0 

Figure2. 2-4.ErrorFocus 

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•:}4CV 0 OD •:t,CV CV '410 00 0 CV 0 00 

0 C; C; 0 0 

Das/crvu 

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2.2.1.3.2 Jitter Performance 

The performance with respect to the requirements of Figure 2.2-1 is evaluated by determining the LOS jitter resulting from the following extraneous inputs. 

Bearing Noise (Tn) 

Electronic Noise (An) 

Gyro Noise (Gn) 

Input Noise (D/An) 

The contributions of these sources to the total line-of-sight (LOS) jitter error are calculated using their respective Power Spectral Density (cD(w)) characteristics and closed loop 

N(w) transfer function G(w) = weighted by the function F(w). 

N(w) 

For the rate configurations of Figure 2.2-2 the error resulting from gyro noise (Gn) is 

EGn = 0.206 x 106 f 

2000 

2 (9(w) )2 d(.1.) sec (1)( 0) • (FP)) • Gn(w) 

0.2 

2000 2 6/2 

-162 db S -31 db S -56 db (S/100+1) 

= 0.206 x 10 

0.2 

((S/100+1)(S/300+1))( (S/37+1) 

)

2

( 

S(S/150+1)(S/100+1) d w 

EGn = 0.023 second of arc 

The transfer function of the PSD of the gyro noise was taken from a PSD measurement 

made by Honeywell on a gyro similar to the one selected for the ATS scanner control 

system. 

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The noise at the input due to electronic noise generated in the D/A convertor and buffer 

amplifier is not known. To make the necessary configuration tradeoffs it was assumed 

this noise source would not cause an RMS gimbal error of greater than 0.016 arc seconds. 

Using this error allotment the allowable RMS value of D/A converter and buffer amplifier 

noise can be determined as follows: 

2000 

ED/A = 2f C'D/A(W)• 

O. 2 

assuming ED/A = 0.016 sec 

0.016 sec = 0.0775 x 10-6radians 

(-31 db S 1 

(S/37+1) ((S/150+1)(S/1000+1)3) 

2 2 

dW 

E

D/A 

/0.2 2000 

cD D/A(w) dw (2.95) 

2000 

(0.0262 x 10-6)2= D/A(W) dw 

0.2 

with gyro scale factor of 0.00436 rad/sec/ma and torque resistance of 100 ohms 

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2000 dw 

6.8 x 10-16 

= f

(I'D/A(w) 

0.2 

2 (1 

6.8 x 10-16 rad2 /sec2 

 x 0.0436 volt

2

/rad

2

/sec

2 = 1.56 x 10-14 volts2 

) 

2000 

1.56 x 10-14 volts

2 

= f D/A(w) dw 

0.2 

If the noise is assumed to be bandwidth limited white noise, where the bandwidth is equal 

to 2000 radians per second, then the level of the noise is equal to 

1.56 x 10-14 volts2 

- 0.78 x 10-17 0.078 x 10-16 volts2 

2000 rad/sec rad/sec 

This is then the level of noise that will produce a position error of 0.016 arc seconds 

RMS gimbal error. It should be noted that a roll-off of 30 radians per second at the buffer 

amplifier could be implemented without affecting loop dynamics in track or slew modes. 

This will increase the allowable noise at the D/A convertor output and make it much easier 

to implement. 

The allowable noise at the input of the compensating amplifer in the Main Tracking Mirror 

control loop was assumed to be representative of that which would be present in a rate 

servo configuration. This noise in terms of bandwidth limited white noise (4.5 x 10-8 

volts2

/rad/sec to 2000 rad/sec) was used to determine the position error resulting from 

electronic noise within the control loop. 

2-95 

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(co) -86 db (1) (ft - lb)2 

rad/sec 

(S/0. 1+1)

2 

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The position error resulting from this noise source is: 

2000 

( -31 db s ( -51 db (S/100+1) ) 

EAn = 0.206 x 106 2 f (-147 db) (S/37+1) 

S(S/150+1)(S/1000+1)2 

2 

d ir.) 

0.2 

E = 0.022 sec M 

This error can be decreased by decreasing the allowable amplifier noise. It should be 

noted that the allowable noise requirements are much more severe on the buffer amplifier 

electronics than on the amplifier within the control loop. This is due to the noise rejection, 

provided by the high gain gyro, to amplifier noise within the control loop. 

The bearing noise information on the tracking mirror drive indicates the bearing noise 

PSD for the Main Tracking Mirror is approximately equal to 

-6 db (ft - 1b)2 

(S/0.1+1)2 rad/sec 

Assuming the pre-load on the ATS Scanner gimbal bearings is approximately equal to 

50 times less than that of the tracking mirror bearings and the bearings are 1/3 as large, 

the ATS scanner bearing noise PSD signature is assumed to be 

For the roll servo configuration the error resulting from this bearing noise (ETn) is 

equal to 

2-96 

SPECIAL HANDLING 
2000 2 -86 db -31 db S -75 (S/0.4+1) 

(S/0.1+1)2 (S/37+1) S(S/40+1)(S/150+1) 

du) 

)2 

NRO APPROVED FOR 

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Roll ET = 0.206 x 106 2 

0.2 

= 0.014 sec 

Pitch ETn = 0.206 x 106 /2 f 

0.2 

0.042 sec 

SPECIAL HANDLING 

2000 

-86 db (-31 db S)2 (-65 db (S/0.4+1) )2 dw 

(S/0.1+1)2 (S/37+1) (S/40+1)(S/150+1) 

Figures 2.2-2 and 2.2-3 show the pitch and roll RMS gimbal errors as a function of the 

bandwidth for lag-lead ratios of 10 to 1 and 100 to 1. It should be noted that the error 

resulting from the assumed bearing noise PSD can be reduced by a factor of approximately 

2 to 1 by increasing the gain and using a 100 to 1 lag-lead network rather than a 10 to 1 

lag-lead network. 

The error values shown in Table 2.2-1 are based on a 100 to 1 lag-lead ratio. The electronic noise error shown in Figures 2.2-5 and 2.2-6 is based on a white noise input to the 

electronics in the loop and the bandwidth is changed by changing the amplifier gain. 

Figure 2.2-6a shows the open loop frequency response and Figure 2.2-6b shows the slew 

loop frequency response with the rate loop incorporated. 

The rate loop configuration has the advantage of being able to control the gimbal rates down 

to the structure level of the drive. It also is the least sensitive, of the three configurations 

considered, to errors in the rate commands induced by vehicle transient disturbances. 

The rate loop configuration has the disadvantage of providing less bearing noise rejection 

than either of the other configurations studied. It also requires that the D/A convertor 

noise be less than the configurations with the D/A convertor inside the servo loop. 

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TABLE 2.2-1. RATE SERVO CONFIGURATION ERROR SUMMARY 

(JITTER PERFORMANCE) 

ERROR SOURCE PITCH AXES ROLL AXIS 

Bearing Noise 

Gyro Noise 

Loop Electronic Noise 

Input Noise (D/A) 

0.043 sec 

0.023 sec 

0.016 sec 

0.022 sec 

0.0143 sec 

0.023 sec 

0.016 sec 

0.022 sec 

RMS Gimbal Error 0.054 sec 0.0375 lel 

RMS LOS 0.1105 sec 

2cr LOS 0.221 sec 

I = 0.12, 0.36, 0.08 

100:1 lag-lead 

150 rad/sec We 

SECRET SPECIAL HANDLING 

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CV 

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NRO APPROVED FOR 

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OPEN LOOP FREQ

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THE SLEW L

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INCORPORATED AS A MINORLOOP 

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( S2 

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RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

saailoaa - aMillti

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Roll (ETn) = 0.206 x 1062 -86 db -31 dbs2 -41 db 

(S /0 . 1+1)2 (S/37+1) (S/10+1) (S/30+1) 

2000 

0.2 

) 2 

dw 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

2.2.1.4 Position Servo Configuration 

The position servo configurations are shown in Figure 2.2-7. A bandwidth of 60 radians 

per second in pitch and 30 radians per second in roll was selected for the position servo 

configurations. 

The selected bandwidths allow the scanner tracking control loop to be incorporated into the 

slew servo configuration as a minor servo loop. This allows the transistion from slew to 

track mode to be made with a minimum servo transient. 

Figures 2.2-8 and 2.2-9 show the RMS gimbal error in pitch and roll as a function of the 

position configuration bandwidth. The open loop frequency response of the servo loops is 

shown in Figures 2.2-10 and 2.2-11. The loops have an approximate gain of 50,000 and 

13,000 at 0.15 radians per second in pitch and roll respectively 

Figures 2.2-12 and 2.2-13 show the open loop frequency response of pitch and roll slew 

servo loops with the position servo configuration of the tracking loops incorporated as 

minor loops. 

For the position configurations of Figure 2.2-7, the error resulting from the assumed 

bearing noise PSD is: 

2000 

Pitch (ETn) = 0.206 x 106 2 f -86 db ( -31 dbs -43 db 2 

2(S/37+1) • (S/20+1)(S/60+1) 

dw 

0.2 

2-103 

4E-RET- SPECIAL HANDLING 
SPECIAL HANDLING 

C 

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rn 

a 

Figure2. 2-7.Position Servo Configurations 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

DIGITAL PROCESSOR 

0 

8 

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0 0 

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t 

0 0 

ti 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

--SftR-ET-- SPECIAL HANDLING 

2-105 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 -S-E-CRET-- SPECIAL HANDLING 

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2-106 

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Figure2.2-9. Roll Po

sition Servo 
Figure2.2-10. OpenLoop Frequency Resp

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—ERR SPECIAL HANDLING 

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onseof SlewLoopwith Po

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r-1 

0 

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2-109 

SECRET SPECIAL HANDLING 
— I — — — — — 

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— 

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S 

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Figure2. 2-13. Roll Op

enLoop FrequencyResp

onseof SlewLoopwith PositionMinorLoop 

O 

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SaH2IDaa - HSVIld 

CD 

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SECRET SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

Assuming electronic noise in the loop is represented as bandwidth limited white noise at 

the input to the position buffer amplifier and the allowable position error resulting from 

this noise is 0.016 seconds of arc. The value of the input noise is 

2 

Roll (An) = 0.16 x 4.85 x 10-6 = 2 f (1) An( w S/37+1) 2• (5/30+1)(S/100+1) dw 

-31 dbs -60 db 

0.2 

2000 

0.0775 x 10-6 

:0.200 

ilf 

• I 2000 

= f il 

0.2 

c13An(w) dw • (0.00506) 

1. 52 x 10-5 

(10 AN (w) dw 

2.30 x 10-10 2 

An - = 1.15 x 10-13 Volts 

2000 Radian/Sec 

2000 2 

Pitch (An) = 0.0775 x 10-6 = 2 f (D An(w)

( -66 db dw (S/37+1) (S/60+1)(S/200+1)) 

2000 

0.0775 x 10-6= / f (13 co ) dw An ( (0.00454) 

0.2 

2000 

17.6 x 10-6 = f cI3An((to) dw 

0.2 

3.1 x 10-10 -13 Volt2 

c)An = 1.51 x 10 2000 Radian/Sec 

2-111 

SECRET SPECIAL HANDLING 

0.2 
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It should be noted that the position configuration provides a higher rejection to D/A buffer 

amplifier noise than does the rate configuration; this permits the design of the buffer 

amplifier in the position configuration to be less complex than in the rate configuration. 

An additional source of jitter in the position configuration that is not present in the rate 

configuration is the jitter generated by digital summation of the encoder output and position 

command signal. (See Figures 2.2-13 and 2.2-14.) 

The jitter resulting from the generation of the quantized error signal can be calculated by 

determining the amplitude of the fundamental of the square wave and operating on this with 

the response of the servo at the error signal quantizing frequency. 

For the case of the pitch axis with a 

220th encoder quantum level on the gimbal the quantizing is equal to: 

G 

2 

r

A sin co (D) [G] 

where G is the attenuation of the control loop at the quantizing frequency. 

Assuming (D) = 0.5 (worst case) and the gimbal rate is 0.3 degrees per second, the loop 

roll off will attenuate the quantizing error by a factor of approximately -100 dg 

6 = 2 x 1'25x 1 x 0.00001 = 0.00008 sec ' GE 3.14 

This error can further be reduced if the input command has a quantum level which is s 1/4 

of the position encoder quantum level. This allows the quantized error signal to approach 

a triangular wave, the fundamental of which is equal to 0.5 times the fundamental of a 

square wave. 

2-112 

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-1E-R-E-T- SPECIAL HANDLING 

Figure2. 2-14.Generationof QuantizedErrorSig

n

al 

141H 

■•■•• ■ •■• 

e•M •••■ MIM■ ••••••• 

•NROMO •■■•• MIME. 

(CD 

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2-113 

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20 BIT 

INPUT 

22 BIT 

INPUTS _FL/Lit_ 

An additional source of error results because the encoder quantum level varies about the 

nominal value. A digital simulation was run to determine the effects of this bit to bit 

variation. Figure 2.2-15 shows the servo configuration used in the simulation and the encoder bit distribution which was simulated. 

The error resulting from the bit to bit variations can be controlled by controlling the allowable bit variations of the encoder with respect to the nominal quantum level. Using an 

allowable variation of 10.1 arc seconds per quantum level with the restriction that the maximum accumulative error never exceeds 0.2 arc second at any time nor 0.1 arc second 

over any 5 bit interval it is possible to restrict the gimbal jitter to less than 0.0125 

second of arc at a gimbal rate of 0.3 degree per second with a 30 radian per second-servo 

bandwidth. 

Figure 2.2-16 is a plot of quantizing error as a function of gimbal rate for a 220th encoder 

and a 30 radian per second bandwidth. Figure 2.2-17 shows the same error information for 

a 60 radian per second bandwidth. 

2-114 

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•11 

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2-115 

SPECIAL HANDLING 

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t =0. 0055 SEC 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 41-410-- SPECIAL HANDLING 

C> 

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2-116 

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0 

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0 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 __SterRit- SPECIAL HANDLING 

ucaraa avawm 

2-117 

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The position servo configuration is the least complex of the three configurations considered 

and does not require a rate gyro as a sensor. Thus it also uses less power and weighs 

less than the other two configurations. Position feedback is via the gimbal encoder which 

is also used in the slew configuration, thus an additional sensor is not required. 

The bearing noise rejection characteristic of the position configuration is approximately 

equal to that of the rate configuration for the bearing noise profile assumed for the tradeoff studies: 

( ct, _-86 db volts2 

(S/0.1+1)2 rad/sec )

where 0.1 = 20 log db 

The sensitivity of the position control loops to errors in the rate commands induced by 

vehicle transient disturbance is the highest of the three configurations considered. Table 

2.2-2 shows the tabulated error inputs for the position configurations. 

TABLE 2.2-2. POSITION SERVO CONFIGURATION ERROR SUMMARY 

ERROR SOURCE PITCH ROLL 

Bearing Noise 

Electronic Noise 

Quantizing Error 

0.18 sic 

0.016 sec 

0.025 sec 

RMS Gimbal = 0.182 sec 

RMS = 0.403 sec 

2 a = 0.806 sec 

0.165 sec 

0.016 sec 

0.062 lee 

0.172 sec 

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2.2.1.5 Position and Rate Servo Configuration 

This configuration is shown in Figure 2.2-18. The inner loop is a rate loop with feedback 

via the gyro and the outer loop is a position loop with feedback from a position encoder. 

The dynamic error of this configuration with respect to the tracking input command is less 

than one microradian per second (see Section 2.2.1.3.1). 

The performance with respect to the requirements of Figure 2.2-1 is evaluated by considering the following noise error sources: 

Bearing Noise (Tn) 

Electronic Noise (An) 

Gyro Noise (Gn) 

Quantizing Noise (Q) 

Using the same noise profiles as in the evaluation of the rate and position servo configurations, performance of the rate plus position configuration was determined as follows: 

6 

200 

f 0 ( -162 db s ( -31 db s 

 2 

( 

-105 db (S/1+1)(S/100+1) 2 

EGn = 0.206 x 101i 2 dw 

(S/100+1)(S/300+1) (S/37+1) (S/10+1)(S/40+1)(S/150+1)(S/1000+1)2 0.2 

= 0.022 sec 

2000 

-86 db -31 dbs -96 db (S/4+1)(S/1+1) 2 

Pitch ETn = 0.206 x 10;1 2 f ((s/0.1+1)2) (s/37 +1))2( 0 (S/1 +1)(S/37+1)(S/30+1)(S/40+1)(S/150+1) dw 

0.2 

= 0.02 sec 

2000 

( -86 db ( -31 d(s) ( 2 -106 db (S/4+1)(S/1+1) )2 Roll ETn = 0.206 x 106/2 dw 

(S/0.1+1)2 (S/37+1) (S/10+1)(S/37+1)(S/30+1)(S/40+1)(S/150+1) 

0.2 

2-119 

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t 

wa 

oa 

yy

d 

a d 

PITCH SERVO 

4 4, 

Aw 

T 

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rlVi 

Figure2. 2-18. Po

sitionand Rate Servo Configuration 

r - - 1 

LO 

l 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 --SH-41-ET- SPECIAL HANDLING 

2 —120 

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Quantizing Noise: 

2 

Roll = ( .1 ) 

2.28 + (0.004) 0 2 8

8 = 0.0635 sec at 0.05o

/sec with a 2

20th 

Pitch =

11( 0.18 

2.828 x 5 

)2 + (0.00008) = 0.0125 sec at 0.3°/sec with a 220th 

The error due to electronic noise has been assumed to be less than 0.016 seconds of arc 

for the configurations shown in Figure 2.2-18. This is consistent with the assumed 

electronic noise error used in the evaluation of the rate and position configurations. To 

meet this error allotment the electronic noise also is budgeted such that the D/A buffer 

noise contributes 15/16 of the allowable noise error. This is a reasonable allotment in 

that the gimbal position is approximately 30 times more sensitive to this noise source than 

to the noise within the rate minor loop. 

Figures 2.2-19 through 2.2-21 show the RMS gimbal error for the pitch and roll control 

loop with the position plus rate configurations. Table 2.2-3 gives an error tabulation for 

this servo configuration. 

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encoder 

encoder , 
I I I I 

O co co +t■ N 

ftft 0 0 0 0 

0 0 d d 

0 N 

0 

R

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a. 0 0 

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x 0 

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.■ . ..■......■ - --.------ --, 

SPECIFICATION LIMIT RMS GIMBAL ERROR 

RATE =0.05 DEGREES/SECOND 

ENCODER 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRETSPECIAL HANDLING 

11011H3 rIV8W10 '1'1011 SIADI 

2 -122 

-SEC-1114- SPECIAL HANDLING 
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0 

CD 

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O 

(.0 

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0 

O

4-4 

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z 

0 

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a, 

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0 z 

RATE = 0. 30 DEGREES/SECOND 

ENCODER= 2

20 BITS 

SPE

CIFICATIONLIMIT RMS GIMBAL ER

ROR 

RMS Q

UANTIZING ERROR 

GH OUTER LOOP 

Figure2. 2-

20.Po

sitionand RateCo

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2-123 

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0 

aC 

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00 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 -SE-41-ET--- SPECIAL HANDLING 

0 0 

O 

r-1 

C/1 

CD 

8 

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0

0 

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U) 

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1 

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1E-R-E-T--- SPECIAL HANDLING 

Figure2. 2-21.Positionand RateConfiguration(Jitterequals 0. 12slug ft2) 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 -1{411-E-T- SPECIAL HANDLING 

TABLE 2.2-3. POSITION PLUS RATE SERVO CONFIGURATION ERROR SUMMARY 

ERROR SOURCE PITCH ROLL 

Bearing Noise 0.02 sec 0.0066 sec 

Electronic Noise 0.016 sec 0.016 tree 

Gyro Noise 0.023 sec .... 0.023 sec 

Quantizing Error 0.0125 sec 0.068 sec 

RMS Gimbal O. 037 sec 

la LOS 0.105 sec 

0.074 sec 

2a LOS 0.21 sec 

Figure 2.2-22 shows the open loop frequency response of the position plus rate configuration. 

Figure 2.2-23 shows the open loop frequency response of the slew loop with the position 

and rate configuration incorporated as minor loops. 

The 30-radian position loop allows the tracking requirements to be met by the servo while 

allowing it also to be incorporated into the slew configuration. 

The 150-radian per second minor loop is used to develop a high rejection of bearing noise 

(Tn). The 150-radian bandwidth requires the scanner structural frequency be greater than 

150 radians and the allowable frequency ratio and damping are functions of the exact 

structural frequency. 

An advantage of the rate and position configuration is that it has flexibility such that the 

minor loop bandwidth and gain can be increased for maximum rejection of bearing noise or 

gyro and amplifier noise depending on which becomes the more significant as the noise 

characteristics of the individual components become better defined. 

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saatioaa - ova HSVHcl 

C> O

rl 

0 0 0 

CO 

0 

CO 1-4 

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2-126 

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ADIANS/SECOND 

Figure 2. 2-22. OpenLoop FrequencyResp

onseof Pitch SlewServo 

-11C-R-ET- SPECIAL HANDLING 
O 

0 

O 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

gamma' - DV'l 3SVHd 

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N 

0 0 0 Co Co

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Figure2.2-23.Closed LoopFreque

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2-127 

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2.2.1.6 Baseline Configuration - Description, Requirements and Design Margins 

2.2.1.6.1 General 

The baseline configuration incorporates the three servo configurations previously discussed. 

It consists of a primary mode and secondary mode of operation. The primary mode 

utilizes a position plus rate servo configuration for pitch axis control and a rate servo configuration for roll axis control. The bandwidths of these servos are as discussed in Paragraphs 2.1.3.1 and 2.1.2.5. These are a 30 radian per second position loop with a 150-

radian per second rate minor loop in pitch. 

The primary configuration has the following advantages: 

a. Increased bearing noise rejection in the pitch axis. This is required due to the 

higher sensitivity of the LOS to pitch jitter and the lower inertia of the pitch 

gimbal. 

b. Increased flexibility, in that the bandwidths and gains can be varied to minimize 

specific unwanted disturbances. 

c. Bandwidths can be reduced to accommodate lower structural natural frequencies 

if this should become necessary while still being capable of meeting the jitter 

performance requirements. 

d. Inherent capability of tracking at the minimum rates of 0.3 degree per second in 

pitch and 0.01 degree per second in roll (requirement). The use of a rate loop in 

roll allows the lower tracking rate, while the pitch encoder quantum level limits 

the minimum pitch tracking rate to 0.3 degree/second. 

The primary pitch axis servo configuration provides increased bearing noise rejection and 

reduced noise requirements on the D/A converter at the expense of added complexity and 

encoder uncertainty. The increased complexity also allows a backup mode to be incorporated. 

2-128 

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The primary roll axis servo configuration provides sufficient bearing noise rejection with 

relative simplicity while decreasing the ATS error sensitivity to vehicle transient disturbances. It requires tighter control on the D/A converter and buffer amplifier noise and 

does not provide as large a design margin with respect to bearing noise increases as does 

the pitch servo configuration. 

The backup configuration utilizes a 60-radian per second position loop in pitch and a 30-

radian per second position loop in roll. The characteristics of these servos are the same 

as those discussed in Paragraph 2.2.1.4. 

The backup configuration provides redundancy should a gyro or compensating amplifier fail 

in the primary servos at the expense of a decrease in pitch performance. 

The baseline servo configurations are shown in Figure 2.2-24 and the tabulated error date 

is listed in Table 2.2-4. 

The baseline configuration has the advantage of being able to be converted to any one of the 

three configurations discussed by the use of a few switching points in the primary and 

backup servo loops. 

2.2.1.6.2 Dynamic Requirements for the Baseline Scanner Control System. 

The scanner pitch and roll configurations contain the primary mode (switch number 1 in the 

one position) and the backup configuration (switch number 1 in the two position). 

In operation all components would be energized and the Digital Processor would always have 

the current position stored. This would allow the switch to be closed and the change made 

from the primary mode to backup mode with a minimum transient in the loop. 

2-129 

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o4 

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Figure 2. 2-

24. 

SPECIAL HANDLING 

2 —130 

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TABLE 2.2-4. PRIMARY MODE ERROR SUMMARY 

PITCH ROLL 

Bearing Noise 0.02 sec 0.026 sec 

Electronic Noise 0.016 sec 0.022 sec 

Gyro Noise 0.022 sec 0.023 sec 

Input Noise 0.016 se--c 

Quantizing Noise 0.0125 sec 

RMS Gimbal error - 0.037 sec 0.0434 sec 

RMS LOS - 0.086 sec 

2 cr LOS - 0.172 sec 

BACKUP MODE ERROR SUMMARY 

PITCH ROLL 

Bearing Noise 0.180 sec 0.165 sec 

Electronic Noise 0.016 sec 0.016 sec 

Quantizing Noise 0.0125 sec 0.062 sec 

RMS Gimbal - O. 182 sec 0.172 sec 

RMS LOS - 0.403 sec 

2o-LOS - 0.806 sec 

The dynamic requirements for the D/A converter and buffer amplifier have been combined 

and are given as requirements for each combination. This will allow the electronic 

designer to determine the optimum gain distribution between the two components with 

respect to noise and other design parameters. The gain distribution selected must allow 

the dynamic requirements specified for the combination of the D/A and buffer to be met. 

2-131 

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A. Scanner Pitch Loop - Digital Processor (Pitch) - The Digital Processor shall receive 

the pitch command (9 ) from MDAU and integrate this command (e ) by digital summation 

Pc Pc 

techniques. It shall receive the pitch incremental encoder output and convert this to a 

position reference ). The integrated rate command signal and the position reference 

shall be summed and the resultant shall be the output of the Digital Processor. 

(6)

E =f epc - P

) 

• 

The integrated rate command ( 

pc

) shall have a maximum quantum level of 0.32 sec and 

the position reference (El ) a maximum quantum level of 1.25 se-c. 

Timing reference inaccuracies in the Digital Processor shall not cause an error at the 

output of greater than 1 A radian for any one second period. This error shall not vary at 

a frequency between 0.2 radian per second and 1000 radians per second. 

B. D/A Converter and Gyro Buffer (Pitch) - The D/A converter in pitch receives its inputs 

from the Digital Processor in both the primary and backup modes of operation. The D/A 

converter and Gyro Buffer combination shall process the pitch position command signal 

from the Digital Processor and develop a signal to command the pitch rate gyro. 

The dynamic requirements of the D/A converter-Gyro Buffer combination are: 

Gain - 72.5 ± 10 percent amperes per radian into a 100-ohm load ± 10 percent with a 

stability of er_ 1 percent. The gain shall be linear for inputs from -0.0105 to +0.0105 

radian. 

Frequency Response - The frequency response of the Buffer amplifier shall have the 

following characteristics: 

where 

T1= 0.1 sec 

7

2 

= 1 sec 

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Additional breaks shall not contribute more than 6 degrees of phase shift below 1000 

radians per second. 

Offset - The dc offset and long term drift shall not exceed 1 ma. Long term drift shall 

include all noise components below 0.2 radian per second. 

Noise - The noise at the output of the D/A Converter - Gyro Buffer shall not cause an 

rms position error of greater than 0.016 units when the pitch gimbal position with 

respect to noise at the Buffer output is expressed as: 

eG 4.1 x 10 (S/0,2+1) (S/1+1) sec 

NGB (S/10+1)(S/30+1)(S/37+1)(S/150+1)(S/1000+1)

3 rad/sec 

for frequencies above 0.2 radian per second. 

C. D/A Converter and Position Buffer (Pitch) - The D/A Converter and Position Buffer 

combination is used in the backup mode. It shall process the position error signal from the 

digital processor for input to the power amplifier. The dynamic requirements of the D/A 

converter and Gyro Buffer combination are: 

Gain - 13,600 volts per radian f 5 percent 

The gain shall be linear for inputs from -0.003 to +0.003 radian. 

Frequency Response - The frequency response of the D/A Converter-Position Buffer 

combination shall have the following characteristics. 

1 S+1) 7

1 = 0.05 sec ± 10 percent 

(1-2S+1) ( T3 +1) where 

Additional breaks shall contribute no more than six degrees of phase shift below 1000 

radians per second. 

Offset - The dc offset and long term drift shall not exceed 40 millivolts. Long-term 

drift shall include all noise components below 0.2 radian per second. 

2-133 

SECRET SPECIAL HANDLING 

T2 = 0.005 sec 10 percent 

2 

T

3 

= 0.001 sect 10 percent 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

Noise - The noise at the output of the Pitch Position Buffer shall not cause an RMS 

pitch gimbal error of greater than 0.016 unit when the pitch gimbal position with 

respect to noise at the Position Buffer output is expressed as: 

eG O. 6 x 10-1 (S/0.2+1) sec 

N

PB (S/20+1)(S/37+1)(S/60+1)(S/1000+1)

2 volt 

for frequencies above 0.2 rad/sec. 

D. Gyro Pitch - Gyro shall be as presently specified in TR-1717. 

E. Compensating Amplifier (Pitch) - The pitch compensating amplifier receives the output 

from the gyro and processes this signal for input to the power amplifier. The dynamic 

requirements of the compensating amplifier are: 

Gain - 28 volts/volt t 10 percent 

The gain shall be linear for inputs from -4 to +4 volts. 

Frequency Response - The frequency response of the compensating amplifier shall 

have the following characteristics. 

( T

1 

(T3 S+1) 7

1 

 = 0.025 sec ± 10 percent 

where 

( 72S+1) (T4S+1) T2 = 0.250 sec ± 10 percent 

T

3 

= 0.010 sec ± 10 percent 

7

4 = 0.001 sec f 10 percent 

Additional breaks shall contribute no more than 6 degrees of phase shift below 2000 

radians per second. 

Offset - The dc offset and long term drift shall not exceed 100 millivolts. Long term 

drift shall include all noise components below 0.2 radian per second. 

2-134 

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RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

Noise - The noise at the output of the compensating amplifier shall not cause an rms 

gimbal position error of greater than 0.004 unit when the pitch gimbal position with 

respect to noise at the output of the compensating amplifier is expressed as: 

1.50 x 10-4 (S/2+1)(S/1+1)(S/4+1) sec 

NCA (S/10+1)(S/30+1)(S/37+1)(S/40+1)(S/510+1)(S/1000+1) volt 

for frequencies above 0.2 radian per second. 

F. Power Amplifier (Pitch) - The power amplifier receives the output of the compensating 

amplifier in the primary mode and the output of the position buffer in the backup mode. 

The dynamic requirements of the pitch power amplifier are: 

Gain - In the current feedback configuration the gain shall be 0.0224 amp per volt 

± 5 percent into a load of 64 ohms and a back emf of 0.31 emf of 0.31 volt max. 

The gain shall be linear to ± 0.4 amp output. 

Frequency Response - The frequency response of the Pitch Power amplifier shall have 

the following characteristics 

K 

( T1 S+1) where T

1 

 = 0. 001 sec max. 

K = gain 

Additional breaks shall contribute no more than 6 degrees of phase shift below 2000 

radians per second. 

Offset - The dc offset and long term drift shall not exceed 2 ma. Long term drift 

shall include all noise components below 0.2 radian per second. 

Noise - The noise at the output of the power amplifier shall not cause an rms gimbal 

position error of greater than 0.004 unit when the relationship between noise at the 

power amplifier output and the gimbal position is expressed as 

eG 3.34 x 10-3 (S/0.2+1)(S/1+1)(S/4+1) sec 

NPA (S/10+1)(S/30+1)(S/37+1)(S/40+1)(S/150+1) volt 

frequencies above 0.2 radian per second. 

2-135 

SECRET SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 -SECR-E-T--- SPECIAL HANDLING 

G. Torquer (Pitch) - The following characteristics were given by Subcontractor "A" for 

the pitch axis torquer: 

Rated Torque = 37.6 oz - in with 16 volts across both windings 

Resistance of each winding = 155 ohms 

Time Constant = 0.001 second (both windings) 

KV = 0.31 volt/radian/second 

These characteristics were used to determine the pitch loop component dynamic requirements. 

Tolerances which should be applied to the torquer parameters are 

Torque = Level given should be the minimum allowable level 

Resistance of windings ± 10 percent 

Time constant ± 10 percent 

Back Emf constant (KN

) ± 5 percent 

H. Encoder (Pitch) - The pitch gimbal encoder is a 2

18 incremental encoder. 

Accuracy - Each quantum of encoder output shall be equal to 1.25 ± 0.1 sec of shaft 

position angle. 

Noise - The Power Spectral Density of the components of the encoder position readout 

error resulting from bit variations with respect to the nominal quantum level shall be 

within the envelope of Figure 2.2-25 when the encoder shaft is rotated at a constant 

rate of 0.065 degree per second. 

Response - The accuracy requirements shall be met for all shaft rotation rates from 

0.3 to 4.5 degrees per second. 

2-136 

R-E- SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

Note: The test information presently available from the encoder vendor indicates that 

they can not meet the noise specification detailed above with an 18 bit encloder and that 

the CEI jitter requirement could not thus be met. Studies are in progress evaluating 

the expected noise from a 219and a 220 bit encoder. 

Analysis of the latter, (page 2-112), demonstrates that a 220 bit encoder adequately meets the 

design objective. This device, however, is larger than the 218bit encoder space envelope 

designed into the scanner. Recent information from the Wayne-George vendor, indicates 

that a219bit encoder can be manufactured in the same disk size, and consequently same 

space envelope, as the 218bit encoder. Studies are continuing to evaluate its noise 

characteristics and packaging requirements. 

2-136A/B 

SPECIAL HANDLING 

..,...... 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

4-E&RET" SPECIAL HANDLING 

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2-137 

SECRET SPECIAL HANDLING 
SPECIAL HANDLING 

I. Mass Properties (Pitch) - The pitch axis inertia shall be no less than 0.12 slug foot 

squared. 

Structural Dynamics (Pitch) - The Structural dynamics of the pitch gimbal assembly 

gyro assembly and mount location shall be such that the frequency response of the 

gyro output signal to motor excitation normalized to the response at one radian per 

second and compensated for the motor electrical time constant and the gyro transfer 

function shall have no amplitude above the profile of Figure 2.2-26. 

J. Mechanical Noise (Pitch) - Power Spectral Density of the torque variations resulting 

from mechanical noise input such as Bearing noise and cable noise shall be within the 

envelope shown in Figure 2.2-25 for all gimbal rates from 0.1 to 1.8 degrees per second. 

K. Friction (Pitch) - The mean value of running friction torque for the pitch gimbal shall 

be no greater than 2 in. -oz. Starting friction torque shall be no greater than 1.5 times the 

mean value of running friction torque. 

SCANNER ROLL LOOP 

L. Digital Processor (Roll) - The Digital Processor shall receive the roll rate command 

signal from the MDAU. It shall integrate this command signal by digital summation techniques. It shall receive the roll incremental encoder output and convert this to a position 

reference, (OR). The integrated rate command signal and the position reference shall be 

summed and the resultant shall be the output of the Digital Processor. 

The integrated rate command shall have a maximum quantum level of 0.32 unit. 

The position reference (OR) shall have maximum quantum level of 1.25 units. 

2-138 

SECRET SPECIAL HANDLING 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 -4E-C-Rff SPECIAL HANDLING 

r-- 

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2-139 

SECRET SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

Timing reference inaccuracies in the digital process shall not cause an error at the output 

of greater than 1 un radian for any one second period. This error shall not vary at a 

frequency greater than 0.2 radian per second or less than 1000 radians per second. 

M. D/A Converter and Gyro Buffer (Roll) - The D/A Converter in roll receives its inputs 

from the MDAU in the primary mode of operation and from the Digital Processor in the 

backup mode. The D/A Converter and Gyro Buffer combination shall process the roll rate 

command signal from the MDAU and develop a command to the roll rate gyro. 

The dynamic requirements of the D/A Converter and Gyro Buffer combination are: 

Gain = 0.25 ma/deg/sec into a 100 ohm load f 10 percent with a stability oft 1 percent. 

The gain shall be linear for inputs from f 0.005 deg/sec tot 45 degrees/sec. 

The D/A gain shall be equal to 0.1 times the buffer gain in the position mode and 

equal to 10 times the buffer gain in the rate mode. 

Offset - The dc offset and long term drift shall not exceed 10 amperes. Long term 

drift shall include all noise components below 0.2 radian per second. 

Frequency Response - The frequency response of the D/A Converter and Gyro Buffer 

shall have the following characteristics: 

K where T

1 

 = 0.033 sec ± 10 percent 

K = gain 

( T1 S+1) 

Noise - The noise at the output of the D/A Converter - Gyro Buffer shall not cause 

an rms gimbal position error of greater than 0.016 unit when the roll gimbal position 

with respect to noise at the Gyro Buffer output is expressed as 

0 5.55 x 10

+3 

sec 

N

GS (5/37+1)(S/150+1) (S/1000+1)3 volt 

for frequencies above 0.2 radian per second. 

2-140 

-S-E-H-ET- SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

SPECIAL HANDLING 

N. D/A Converter and Position Buffer (Roll) - The D/A Converter and Position Buffer 

combination is used in the backup mode. It shall process the position error signal from 

the digital processor, and the output shall go to the power amplifier. The dynamic requirements of the D/A Converter and Gyro Buffer combination are 

Gain (K) - 1976 volts per radian ± 5 percent 

The gain shall be linear for inputs from -0.005 to +0.005 radian 

Frequency Response - The frequency response of the Buffer amplifier shall have the 

following characteristics 

T 1 S+1) 

Additional breaks shall contribute no more than 6 degrees of phase shift below 1000 

radians/sec. 

Offset - The dc offset and long term drift shall not exceed 20 millivolts. Long term 

drift shall include all noise components below 0.2 rad/sec. 

Noise - The noise at the output of the output of the Roll Position Buffer shall not 

cause an RMS roll gimbal error of greater than 0.016 units when the roll gimbal 

position with respect to noise at the Position Buffer output is expressed as 

0 6.5 x 10 (S/0.2+1) sec 

N

PB (S/10+1)(S/30+1)(S/37+1)(S/1000+1) volt 

frequencies above 0.2 radian per second. 

Gyro (Roll) - Gyro shall be as presently specified in TR - 1717. 

0. Compensating Amplifier (Roll) - The roll compensating amplifier receives the output 

from the gyro and processes this signal for input to the power amplifier. 

2-141 

SECRET SPECIAL HANDLING 

(r2S+1) (73 S+1) where 

T

1 

 = 0.1 sec ± 10 percent 

T

2 

= 0.01 sec ± 10 percent 

T

3 

= 0.001 sec f 10 percent 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

-S-EDRET-SPECIAL HANDLING 

The dynamic requirements of the compensating amplifier are: 

Gain (K) - 16 volts/volt ± 10 percent 

The gain shall be linear for inputs from -4 to +4 volts. 

Frequency Response - The frequency response of the compensating amplifier shall 

have the following characteristics: 

(71 S+1) (73 S+1) 

( 72 S+1) (T4 S+1) where 

T

1 

 = 0.025 sec ± 10 percent 

T2 = 0.250 sec ± 10 percent 

T3 = 0.010 sec ± 10 percent 

T

4 

= 0.001 sec ± 10 percent 

Additional breaks shall contribute no more than 6 degrees of phase shift below 2000 

radians/sec. 

Offset - The de offset and long term drift shall not exceed 20 my rms. Long term 

drift shall include all noise components below 0.2 rad/sec. 

Noise - The noise at the output of the compensating amplifier shall not cause an rms 

gimbal position error of greater than 0.004 unit when the relationship between the 

noise at the compensating output and the gimbal position is expressed as 

6.5 x 10--1 (S/4+1) sec 

N

CA (S/37+1)(S/40+1)(S/150+1)(S/1000+1) volts 

frequencies above 0.2 rad/sec. 

P. Power Amplifier (Roll) - The power amplifier receives the output of the compensating 

amplifier in the primary mode and the output of the position buffer in the backup mode. 

The dynamic requirements of the roll power amplifier are: 

Gain - In the current feedback configuration the gain shall be 0.133 amp per volt ± 

5 percent into a load of 6.5 ohms and a back-emf of 0.87 volt max. The gain shall 

be linear to ± 2.5 amps output. 

2-142 

SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

---SERE-T- SPECIAL HANDLING 

Frequency Response - The frequency response of the Roll Power amplifier shall have 

the following characteristics 

K 

( T1 S+1) where T

1 

 = 0.001 sec max. 

K = gain 

Additional breaks shall contribute no more than 6 degrees of phase shift below 2000 

radians per second. 

Offset - The dc offset and long term drift shall not exceed 4.0 milliamps. Long term 

drift shall include all noise components below 0.2 rad/sec. 

Noise - The noise at the output of the power amplifier shall not cause an rms gimbal 

position error of greater than 0.004 unit when the relationship between the noise at 

the power amplifier output gimbal position is expressed as 

6 4.5 x 10 (S/4+1) sec 

N

PA (S/40+1)(S/150+1)(S/37+1) volt 

above 0.2 radian/sec. 

Q. Torquer (Roll) - The following characteristics were given by subcontractor "A" for 

the roll axis torquer: 

Rated Torque = 210 - with 16 volts across both windings 

Resistance of each winding = 13 ohms 

Time constant = 0.028 second (both windings) 

K

v 

= 0.87 volt/radian/second 

These characteristics were used to determine the roll loop component dynamic requirements. Tolerances which should be applied to the torquer parameters are: 

2-143 

-SE-R-ET-- SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

Torque - Level given should be the minimum allowable level 

Resistance of windings ± 10 percent 

Time constant ± 10 percent 

Back Emf constant (K) ± 5 percent 

R. Encoder (Roll) - The Roll gimbal encoder shall be a 2

20 incremental encoder. 

S. Accuracy - Each quantum of encoder output shall be equal to 1.25 ± 0.1 sec of shaft 

position angle. 

T. Noise - The Power Spectral Density of the components of encoder position readout 

error resulting from bit variations with respect to the nominal quantum level shall be 

within the envelope of Figure 2.2-25 when the encoder shaft is rated at a constant rate of 

0.065 degree per second. 

U. Response - The accuracy requirements shall be met for all shaft rotation rates from 

0.01 to 45 degrees per second. 

V. Mass Properties - The roll axis inertial shall be no less than 0.36 slug foot squared. 

W. Structural Dynamics (Roll) - The structural dynamics of the roll gimbal assembly, 

gyro assembly and mount location shall be such that the frequency response of the gyro 

output signal to motor excitation, normalized to the response at one radian per second 

and compensated for the motor electrical time constant and the gyro transfer function shall 

have no amplitude excursions above the profile of Figure 2.2-26. 

X. Mechanical Noise (Roll) - Power Spectral Density of the torque variations resulting 

from mechanical noise inputs such as bearing noise and cable noise shall be within the 

2-144 

IHRET--SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

--S-EMET- SPECIAL HANDLING 

envelope shown in Figure 2.2-25 for all gimbal rates from 0.01 to 0.04 degree per 

second. 

Y. Friction (Roll) - The mean value of running friction torque for the roll gimbal shall be 

no greater than 2 in. -oz. Starting friction torque shall be no greater than 1.5 times the 

mean value of running friction torque. 

DESIGN MARGINS 

The selected configuration for the primary mode of operation has the highest design margin with respect to meeting the jitter requirement based on the present assumed noise 

input. The Design Margins for the possible servo configurations are: 

Pitch Roll Jitter Amplitude Design Margin 

Rate Rate 0.221 1.14 

Position Position 0. 806 

Position and Rate Position and Rate 0.21 1.19 

Rate Position 0.406 

Position Rate 0.76 

Position and Rate Rate 0.166 1.51 

Rate Position and Rate 0.23 1.09 

All of the servo configurations considered have sufficient phase and gain margins for 

operation in both the slew and track modes of operation. 

2-145 

SECRET SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 -41411-ET-- SPECIAL HANDLING 

SLEW MODE 

Requirements 

The slew maneuver shall be completed in (-4-9-- +1 and 15 +1 seconds for the 

30 

roll and pitch axes respectively. The end of the slew maneuver is defined as that time at 

which the gimbal position error is less than 1 arc minute and the rate error is less than 

0.01 degree per second per axis. 

Servo Configuration 

Figure 2.2-27 shows the pitch and roll slew loop configurations. Advantages of the slew 

servo configuration are: 

a. Software is similar to that required for the Main Tracking Mirror. 

b. It allows the tracking servo configuration to be an integral part of the slew servo 

(minimum transient at completion of slew). 

c. Generates shaped torque pulses (minimum structural excitation) 

d. Allows dead beat operation 

e. There are no saturations - linear operation 

f. Limits command rates 

The pitch and roll baseline slew loop servo configurations are position loops with a bandwidth of 18 radians per second. Figures 2.2-28 and 2.2-29 show the open loop frequency 

response of the slew loops with the baseline tracking mode servos incorporated as minor 

loops. 

2-146 

l'ECRET- SPECIAL HANDLING 
O 

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SECRET SPECIAL HANDLING 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 -S-EC-R-ET-- SPECIAL HANDLING 
0 3,1 O O O 03 O O O O O C•1 

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Slew Servo 

2-149 

0 0 Co 

1MT SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

Digital simulations were made using both 12-radian and 18-radian position loops. The 

results indicate that the slew time requirements can be met with the 12-radian slew loops, 

however the baseline configuration of 18 radians provides some margin in meeting the 

time requirements. Figures 2.2-30 and 2.2-31 show the slew performance in pitch and 

roll as a function of A e . Figure 2.2-32 shows the shape of the slew command signal for 

two different A 9's. Figures 2.2-33 and 2.2-34 show analog representations of the slew 

command generation equations. 

2-150 

SPECIAL HANDLING 
a, 

8 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 -4(410— SPECIAL HANDLING 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 —S-EGRET-- SPECIAL HANDLING 

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2 -152 

SECRETSPECIAL HANDLING 
11 

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111111111_1111 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.:3 1.8 2.0 2.2 2.4 2.6 

TIME SECONDS 

Figure 2.2-32. Pitch Slew Command 

2-153 

l'ECRE-T- SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 -S-Efilff- SPECIAL HANDLING 

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2 —154 

1%1E4- SPECIAL HANDLING 
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NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

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2-155 

SECRET SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

2.2.1.7 Command Generation 

The slew command is generated in such a manner as to command the servo to have a deadbeat response, generate shaped torque pulses, and limit the gimbal rates to ± 45 degrees 

per second. 

The slew command is shaped to command the servo to the desired response characteristics. 

The command signal (0c) generation equations can be represented by the analog configuration of Figures 2.2-33 and 2.2-34. 

For values of Pet greater than 8.8 in roll and 17.6 in pitch the inner loop of the slew command generator will go to zero and the output of the generator will become a ramp with a 

slope of 45. This limits the maximum gimbal rates to 45 degrees per second for any /le 

input. 

Figures 2.2-35 through 2.2-39 show the variations of torque, gimbal rate, gimbal position, 

and gimbal position and rate errors as a function of time for a 10 degree roll slew. 

Digital simulations have been made of the slew manuever using the backup servo configurations. A slew performance requirements are met with this servo configuration. 

D/A Gain Switching - To reduce the noise present at the output of the position electronics, 

A

np

, the gain of the D/A convertor can be increased. This increase is accomplished by 

changing the inputs to the D/A ladder network and does not cause a corresponding increase 

in noise at the D/A output. Thus the signal to noise ratio of Anpcan be increased if the 

D/A gain is increased. An increase of D/A gain must be accomplished by a reduction in the 

dynamic range of the D/A. 

2-156 

SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

0 

0 

• 

TIME - SE

CONDS 

Figure2. 2-

35. TenDegree SlewRoll 

saaupaa NIOLIISOcI 

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2-157 

SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

t-- 

cD 

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T-14 

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TIME-SECONDS 

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O 

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SHRIIDHCE 

2-158 

SECRET SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SIG„" SPECIAL HANDLING 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 —isigkET- SPECIAL HANDLING 

10 DEGREE SLEW ROLL 

TORQUE APPLIED 

2 —160 swi-JA - anbuoz. 

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39. Ten Degree 

Slew

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sitionand RateErrors 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

C`l 00 Cr) 

2-161 

SPECIAL HANDLING 
D/A INPUT BUFFER 

AMP 

I 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

(Anp) 

To optimize the gain-dynamic range relationship it is necessary to change the gain and range 

when switching from the slew mode to the track mode. Thus, if the noise requirements of 

the tracking mode configuration require a high D/A gain it is possible to accomplish this by 

incorporating gain switch logic in the electronics. 

This condition can be simulated on a digital computer. The effects of a gain change equal 

to 30 are shown in Figure 2.2-40. This change is made in the roll channel at 2.8 seconds 

after the start of a 60 degree slew. The buffer amplifier gain is reduced by a factor of 30 

and then 0.1 second later the D/A gain is increased by a factor of 30. The results indicate 

that dis gain change can be made just prior to the end of a slew and the effects on the acceleration and rate error will be minimum when the gain change in the buffer amplifier precedes 

the change in the convertor. The 0.1 second delay used in this simulation is a severe case 

and it is assumed that this delay will be much smaller in most cases. 

2 -162 

SECRET SPECIAL HANDLING 
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CHANGE 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

2. 7 2. 8 2.9 3.0 3. 1 

TIME, SECONDS 

Figure 2.2-40. Effects of a A — gain change at End of Slew 

2-163 

SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

-MK+ SPECIAL HANDLING 

2.2.1.8 Slaving Mode 

The requirement for slaving the Main Tracking Mirror to the ATS exists, and methods of 

accomplishing this were investigated. 

The configuration shown in Figure 2.2-41 appears to be the optimum in that the closed loop 

configuration causes the gimbals of both systems to be driven to the same position. The 

rate input to both systems is then updated by the crewman in the ATS control loop. 

Both loops would be slewed to the same target and switch no. 1 would be closed. The crewman would reduce the rate error in the ATS. The change in rate needed to accomplish this 

would also be summed into the Tracking Mirror control loop. 

When the rate was reduced switch no. 1 would be opened and the value present would be 

clamped to prevent the generation of a transient due to rate biases in the tracking mirror 

control loop. 

Figure 2.2-42 shows that the two drives could be slaved together to a rate accuracy of less 

than 1 microradian per second and a position accuracy of less than 30 seconds of are. This 

could be accomplished in less than three seconds starting from the normal end of slew position 

and rate error values. 

The method referred to above involves the actual comparison of the ATS and main optics gimbal positions as well as the application of the man's rate corrections to both subsystems. A 

method of slaving which does not provide position coupling between the two subsystems, but 

simply provides the man's rate corrections to both subsystems is also under consideration. 

This approach would simplify the software needed to implement slaving. 

The table below shows the error allocation assumed to determine the rate error in the slave 

mode. The numbers are considered conservative and demonstrate that the 100 prad/sec 

requirement is met. 

Tracking Mirror LOS Error - Slaving 

Structural Vibration 

TM Servo Noise 

ATS Servo Noise 

Crew Error 

Slaving Error 

Misalignment Error 

26 p rad/sec 

26 fi. rad/sec 

45 /./ rad/sec 

17 p rad/sec 

20 p rad/sec 

50 p rad/sec 

RSS Total 81 p rad/sec 

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z 

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0 

z 0 

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0 

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2 -166 

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nt Profile 
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,•••■■•, 

2.2.1.9 ATS Performance with the Man in the Loop 

Only very preliminary analysis has been performed relative to the performance of the ATS 

with the man in the loop via the MCS. A few comments are included here describing the 

general philosophy being applied to man's role and performance in the control loop. Man's 

performance in the loop will be optimized by simulation experiments to determine optimum 

manual stick transfer functions. 

It has been shown that operators in manual control systems exhibit a type of behavior 

analogous to the behavior of compensating elements which have been inserted to improve 

the overall dynamic performance. When tested, it was found that the human control performance characteristics depended on the dynamic nature of the controlled element as well 

as the forcing function. This characteristic can be obscured in testing; but for averaged 

data taken with trained subjects, it is possible to show (McKruer and Kreudel, "The Human 

Operator as a Servo System Element, " Journal of the Franklin Institute vol 267 (1959), 

pp 381-403 and 511-536) that an analytical describing function form is adequate to describe 

human behavior for a given controlled element and forcing function. When these describing 

functions are generalized, a servo model of human operation and adaptation (in compensatory tasks with random-appearing forcing functions) is obtained. This model is a dynamic 

description of human capabilities in such situations. The model consists of two elements: 

a. A generalized describing function form 

b. A series of "adjustment rules" which specify how to "set" the parameters in the 

generalized describing function. 

The generalized describing function form for one- and two-dimensional compensatory control tasks is 

K e-tS (TLS + 1) 

N P (TNS + 1) (TeS + 1) 

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where 

t is the reaction time delay. 

T

N is the neuromuscular time constant. 

T

L and T

e 

are the compensation characteristic time constants. 

The gain K and the compensation time constants are the only parameters subject to control by the operator. 

When in a task, the human operator adjusts these parameters to achieve the following 

objectives: 

a. To achieve stable control 

b. To achieve good low-frequency closed-loop response to the forcing function, where 

"good" means something like minimum rms error. 

c. To achieve a phase margin of between 60 and 110 degrees based on the operatorgenerated criteria of b, above. 

This describing function, with the parameters adjusted by the rules, applies quite well to 

all situations where the system crossover frequency, estimated by the application of the 

describing function model, is greater than the cutoff frequency of the system forcing function; that is, when the operator can make the controlled element follow the fastest fluctuations of the input. 

The ATS design has incorporated the philosophy described above, including the time 

response requirements of the mission function, in the following fashion: 

a. The describing function used in the synthesis of the tracking loop is the generalized 

describing function described above, but does not depend on the compensation 

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characteristic. This approach has been recommended by Kreudel and McRuer 

in their WADC report as "The best performance is achieved when the describing 

function required of the operator is as simple as possible. " 

b. The bandwidth (equivalent) of the tracking loop, including the human operator, 

has been set by the response time requirements. (The response of the main 

optics has been used as a requirement in the ATS because of the similarity of the 

two tasks. The concept of limited slew or search has not been included as it is 

not part of the SPDR requirements.) This approach far exceeds the bandwidth 

criteria discussed earlier in this section for the validity of the describing function. 

The philosophy was also tested in the Interim Simulation, which was a compensatory 

tracking problem with the AVE forcing function (less noise) and the baseline controlled 

element. The results showed that no major discrepancies existed in this area. One of the 

aspects of AVE characteristic that has not been verified (except analytically) are the effects 

of sampling, quantizing, and extrapolation on human performance in this problem. 

2.2.1.10 Initial Tracking Error 

At the end of slew the LOS error must be less than 540 u rad/sec (See Table 1.1-1 entry 8). 

In addition 2 minutes of position error are alloted in the pointing error analysis (See 

Section 2.3). It is the purpose of this section to briefly summarize the calculations which 

demonstrate compliance with these requirements. 

Rate errors are introduced by off sets in the roll servo loop, a dynamic rate error of 0.01 

deg/sec which can exist at the end of slew, and ephemeris errors which are principally 

due to altitude uncertainties. 

The ephemeris error is taken here as 130 g rad/sec. This value is undoubtably high in 

that it reflects an altitude error of 0.2 mn, considerably more than is now expected with 

or without the low "g" accelerometer. The off set error in roll is calculated in Table 2.2-5, 

to be 52 g rad/sec. The position loop in pitch does not contribute to the rate "off-set" 

error. The 0.01 deg/sec (1741.trad/sec) dynamic error exists for both axes. The total 

error is thus, 

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1/(52)2+ (130)2+ 5(174)2= 400 μ rad/sec 

Thus the 540 μ rad/sec requirement is satisfied. 

The two minute position error allotment is used in the pointing error analysis in section 

2, 3. The pitch axes contribution to this error is derived and compiled in Table 2.2-6 and 

is found to be virtually negligible. 

The effect of the 9° cant has not been considered in this analysis. The effect of ephemeris 

errors is being revised downward. The results of this re-analysis will be documented in the 

next issue of CDRL 100. 

TABLE 2.2-5. 

COMPONENT OFFSET LOOP ERROR 

Roll D/A Convertor 

Roll Gyro 

Roll Compensating Amp 

Roll Power Amp 

Roll Gimbal Friction 

10-6Amp 

4 X 10-5 Rad/Sec 

2 X 10-2 Volts 

4 X 10-3 Amps 

1 X 10-2 ft-lbs 

43.6 it rad/sec 

0. 7 /./ rad/sec 

22.0 u rad/sec 

33.0 p rad/sec 

18.5 it rad/sec 

Roll Off-Set Error = 52.5 it rad/sec 

LOS Rate Error = 52.5 p rad/sec 

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TABLE 2.2-6. 

COMPONENT OFFSET LOOP ERROR 

Pitch D/A Convertor 

Pitch Gyro 

Compensating Amp 

Power Amp 

Friction 

0.001 Amp 

4 X 10-5 Rad/Sec 

1 X 10-2 Volts 

4 X 10-3 Amps 

1 X 10-2 ft-lbs 

0.043 Min 

0.0000 Min 

0.0012 Min 

0.0018 Min 

0.0001 Mil 

Pitch Off-Set Error = 0.043 Min 

LOS Position Error = 0.096 Min 

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2.2.2 IMAGE ORIENTATION (DEROTATION) CONTROL SYSTEM 

The image orientation control system is used to orient the target image that is displayed to 

the crewman. The target image is oriented in such a way that the in-track component of 

the target coordinate vector is aligned vertically at the completion of a scanner slew. Subsequent to the slew, the system is not driven to "track" the angular alignment. In the 

absolute worst case situation the initial alignment of the in-track vector will deviate by 

16 degrees. 

2. 2. 2.1 Requirements - System Level 

a. Slew time - The prism slew shall be accomplished in —30 + 1 seconds where 

AA is the LOS angle in degrees between the correct pointing angle and the target 

when the slew command is applied. 

b. Position accuracy - The position accuracy of the control system shall be within 

(A6 +2 .0 degree of the command at 30 — + 1 seconds after command initiation. 

c. Position Feedback to MDAU - The decoupling computation requires that the 

prism position be known to within ± 1.0 degree and be available to the MDAU 

in digital form. 

2. 2. 2. 2 Servo Configuration 

The servo was configured to meet the dynamic requirements while minimizing software, 

power, and weight. 

The software is minimized by requiring that the input command be a step output from the 

computer. This type of input command avoids updating as a function of time until the start 

of the next slew. 

The motor was sized for this application to minimize weight and power. The rated torque 

available from the motor is 4 inch-ounces for a 16 volt input. 

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The servo configuration for the Image Orientation drive is shown in Figure 2.2-43. This 

configuration allows the control loop to meet the performance requirements with a minimum of software, power, and weight. 

A step input is applied to the command filter. The filter shapes the input to the servo loop 

such that the overshoot which results due to amplifier saturation is minimized. Amplifier 

saturation is a result of the high gain necessary to reduce the offset, caused by motor and 

load friction, to less than 1 degree. The open loop frequency response of the control loop 

is shown in Figure 2.2-44. Figure 2.2-45 shows the image rate and position as a function 

of time for step commands of 0.5 and 1.0 radian. Figure 2.2-46 illustrates the torque 

applied to the load as a function of time for input steps of 0.5 and 1.0 radian. 

2.2.2.3 Error Analysis 

The following error apportionment has been established 

Servo Static Error ± 1.0 degree 

Position Transducer Error ± 1.0 degree 

A/O Conversion Error f 0.5 degree 

Command Equation Error* t 0.5 degree 

The RMS prism error is thus 1.6 degrees, considerably better than the allowed ±5 degrees. 

This apportionment does not consider misalignments between the ARS and the OV. It is 

possible that approximations may be made in the drive equations which will increase the 

command error above 0.5 degree. 

*Conservative Estimate 

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I I 

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2.2.3 DYNAMIC COMPONENT REQUIREMENTS 

2.2.3.1 Command Filter 

The command filter receives a step command from the MDAU and processes this signal 

for input to the K - 3 amplifier. The gain of the command filter is equal to 1 volt per volt ± 

2 percent. The gain shall be linear for inputs from +4 to -4 volts. 

The frequency response of the command filter shall have the following characteristics: 

K 

(TiS + 1) 

where 

ri= 0.166 second ± 5 percent 

K = gain 

Additional breaks shall not contribute more than 6 degrees of phase shift below 100 radians 

per second. 

2. 2. 3. 2 Amplifier 

The K - 3 amplifier shall sum the input signal from the command filter and the feedback 

signal from the position transducer. It shall process this signal and output it to the motor 

winding. The gain of the K - 3 amplifier is equal to 60 volts per volt ± 10 percent. The 

output shall be linear from + 16 to - 16 volts. 

The frequency response of the K - 3 amplifier shall have the following characteristics: 

(T1 S+ 1) 

(72 

S + 1) 

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where 

T

1 = 0.111 second ± 10 percent 

7

2 --- 0.0055 second ± 10 percent 

Additional breaks shall not contribute more than 6 degrees of phase shift below 1000 radians 

per second. 

2.2.3.3 Motor 

The following characteristics were submitted by Subcontractor "A" for the K - 3 motor. 

a. Rated Torque = 4 inch-ounces 

b. Resistance of winding = 26 ohms 

c. Time constant = 0.0008 second 

d. Kv= 0.06 volt/radian/second 

The characteristics were used to determine the K - 3 servo loop dynamic requirements. 

2.2.3.4 Position Transducer 

The position transducer converts the motor shaft angular position to an analog voltage 

for feedback to the K - 3 amplifier. The gain of the position transducer is equal to 3.8 

volts per radian ± 2 percent. The gain shall be linear for inputs from +1 to -1 radian. 

The position transducer shall contribute no more than 6 degrees of phase shift below 1000 

radians per second. 

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2.2.4 ZOOM CONTROL SERVO 

The zoom control system functions identically in the two zoom magnification ranges: 15 to 

30 power and 63-1/2 to 127 power. The difference being the insertion of the power change 

lenses into the system when the zoom control is operating in the high range. 

a. Requirements - The servo response to a step input shall be such that the position 

error is less than 5 percent of the step amplitude in 0.5 second. 

The zoom assembly position shall be monitored and fed to the MDAU to modify 

the gain of the scanner control loops with the crewman in the loop. 

b. Servo Configuration - The servo was configured to meet the dynamic requirements 

while minimizing weight and power. 

Figure 2.2-47 shows the servo configuration of the zoom servo assembly and 

Figure 2.2-48 shows the open loop frequency response. The servo was designed to 

get as much gain as possible in front of the point of motor and load friction while 

still minimizing the servo overshoot to transient inputs. 

Figure 2.2-49 shows zoom assembly rate and position as a function of time for 

a step input equal to the total zoom range. 

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a 0 

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(.1 Z a> 

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Fig

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mbly 

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RATE AND POSITION OF THE 

ZOOM ASSEMBLY AS A 

FUNCTION OF TIME FOR A 

STEP INPUT COMMAND 

2. 5 

2.0 

1.5 

1. 0 

0. 5 

AMPLITUDE OF COMMAND 

IS EQUAL TO FULL RANGE 

OF ASSEMBLY TRAVEL 

0.1 0.2 0.3 0.4 0.5 0.6 

TIME - SECONDS 

Figure 2.2-49. Rate and Position of the Zoom Assembly as a 

Function of Time for a Step Input Command 

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2.2.5 BLANKING CONTROL SERVO 

This blanking control servo serves to protect the ATS optics and the flight crew from a 

direct view of the sun. The blanking control unit drives the shutter into the optical path 

when the sun sensor detects illumination above a preset level. The shutter has two discretes 

positions. A signal from the sun sensor to the control electronics will allow the shutter to 

be driven to the closed position by a mechanical spring. 

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POSITION 

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2.2.6 POWER CHANGE SERVO 

The magnification controller provides the command signal to the power change control. 

Either of two signals are provided, depending upon whether the magnification controller 

is in the low or high zoom ranges. The power change control system drives a set of two 

lenses into or out of the optical path. Functionally, the power change lenses are inserted 

to provide the lower range of magnification as the magnification control stick passes from 

the high to low range. The lenses shall be able to move from one position to the other in 

1.0 second. The power change drive shall be provided with a manual backup in the event 

of drive failure. 

The power change servo is an open loop control which drives the lenses into position when 

the magnification controller passes from low to high power range. The lenses have two 

discrete positions. Figure 2.2-50 illustrates the control servo. It shows a discrete signal 

feedback from the lens assembly to indicate its arrival at either extreme of the travel with 

this signal being used to terminate the drive excitation. 

COMMAND 

FROM 

MAGNIFICATION 

CONTROL 

STICK 

Figure 2.2-50. Control Servo 

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2.3 ATS POINTING ERROR ANALYSIS 

This section presents an error analysis which identifies all major contributors to ATS 

pointing error in the automatic and slave modes and demonstrates a general capability 

to meet the CEI requirements. The alignment error apportionment considered does permit the pointing requirements to be slightly exceeded in the event of a worst case tolerance 

buildup. 

The material is subdivided as follows: 

2.3.1 Discussion of Results 

2.3.2 Description of Error Sources and Errors Removed by Boresighting 

2.3.3 Procedure for Determining Errors 

2.3.4 Statistical Treatment of Error Sources and the Total Pointing Error 

2.3.5 Results of Monte Carlo Computer Programs 

2.3.6 Procedure Required to Determine Allowable Misalignments 

2.3.1 DISCUSSION OF RESULTS 

Table 2.3-1 presents a list of the sources of pointing error for the main optics and ATS, 

a brief description of each, the current allocated value, and describes the type of each 

error for any given vehicle (i.e., random, time varying, or constant error.) It is expected that all of the alignment tolerances listed in Table 2.3-1 can not be met during onorbit operation. As a result, more realistic misalignments were assumed in some of the 

error sources and are shown in Table 2.3-2. 

Table 2.3-3 summarizes the pointing errors that result from the error values in Tables 

2.3-1 and 2.3.2. The errors in Table 2.3-3 correspond to the targets "2 sigma" pointing 

errors that occur over the scan field of: (Continued on page 2-194.) 

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TABLE 2.3-1 

ATS/MAIN OPTICS ERROR SOURCES 

VALUE 

ARC MIN. (2 o) 

(ON-ORBIT) 

TYPE OF ERROR 

(GIVEN A PARTICULAR 

VEHICLE) 

o bya Optical Axis 1.2 Constant 

o bz 

a 

Errors 1.2 Constant 

o A la Roll Gimbal Direction 0.5 Constant 

o kb 

la 

Errors 0.5 Constant 

o 0 2a Mirror Mounting Error 0. 5 Constant 

A TS ( " 3a 

Gimbal Non-Orthogonality 0.5 Constant 

ae 

a

m Gimbal Angle 0.333 Random 

' 8 0 

a

m Measurement Error 0.333 Random 

o (4

a

cs Control System 0.894 Random 

`' o 0 

a

cs Pointing Error 0.894 Random 

o by Optical Axis Measurement 0.667 Random 

6 bz Errors 0.667 Random 

o 8

1 Roll Gimbal Direction 0.5 Constant 

6 kb 1 Errors 0.5 Constant 

Prime A Constant 

Optics - ' 2 Mirror Mounting 0.5 

6 0 3 Gimbal Non-Orthogonality 1.5 Constant 

aem Gimbal Angle 0.333 Random 

o Dm Measurement Error 0.333 Random 

oecs Control System 0.447 Random 

6 0 cs Pointing Error 0.447 Random 

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Table 2.3-1 (Cont) 

ATS/MAIN OPTICS ERROR SOURCES 

VALUE 

ARC MIN. (2a) 

(ON-ORBIT) 

TYPE OF ERROR 

(GIVEN A PARTICULAR 

VEHICLE) 

Vehicle 

Attitude 

6 g x 

6 g , 

Y 

6 g

z 

P xv 

p -- 

yz 

6 byT 

a 

6 bzT 

a 

6 byTa 

6 bz

T 

6 a 

x 

6 ry 

Y 

6 a 

z 

bail 

6a y1 

6 a z 1 

6 Rx 

6 R 

Y 

6 R

z 

P xz 

Vehicle Attitude 

Error (Referenced to 

TM Reference) 

Correlation in Attitude 

Error 

ATS Tracking 

Error 

Prime Optics 

a 

Tracking Error 

. 

' ATS/Prime Optics 

Alignment 

Increase in Attitude 

Error Due to DACO Rate 

 Gyro Error 

, 

Target Location Error 

2.311 

1.477 

1.477 

* 0.873 

*-0.8395 

*-0.7467 

0.667 

0.667 

0.667 

0.667 

1.0 degree 

(total) 

1.36 **2 

1.025 **2 

500 ft 

500 ft 

100 ft 

Random + 

time variant 

component 

N.A. 

N.A. 

N.A. 

Random 

Random 

Random 

Random 

Increases with 

time from last 

star tracker 

update 

Random 

Random 

Random 

* Correlation Coefficients are dimensionless. 

**2This value corresponds to 44 seconds after the last star tracker update. 

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TABLE 2.3-1 (Cont) 

ATS/MAIN OPTICS ERROR SOURCES 

VALUE 

ARC MIN (2cr) 

(ON-ORBIT) 

TYPE OF ERROR 

(GIVEN A PARTICULAR 

VEHICLE) 

6 r

x 

3200 ft 

Vehicle position error Random with 6 r

y without low "g" accel. 370 ft bias component 

6 r

z 210 ft 

Orx 1300 ft 

Vehicle position error Random with 6 r Y with low "g" accel. 145 ft bias component 

6 rZ 210 ft 

TABLE 2.3-2 

ERROR SOURCE VALUE AND TYPE OF ERROR GIVEN A PARTICULAR VEHICLE 

(ARC MIN.) "2 SIGMA" 

ATS 

6 by 

6 bz

a 

619 la 

6°la 

42a 

64)3a 

"aill 

6 f) m 

a 

ae

a

cs 

op cs 

a 

(2.4) 

(2.4) 

(1.224) 

(1.224) 

0.5 

0.5 

0.333 

0.333 

0.894 

0.894 

Constant 

Constant 

Constant 

Constant 

Constant 

Constant 

Random 

Random 

Random 

Random 

(0. 707) Time Varying 

(0. 707) Time Varying 

(0.52) Time Varying 

(0.52) Time Varying 

( )----Numbers in parentheses indicate changes from Table 3.2-1. 

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TABLE 2.3-2 (Cont) 

ERROR SOURCE VALUE AND TYPE OF ERROR GIVEN A PARTICULAR VEHICLE 

(ARC MIN.) "2 SIGMA" 

6 by 0.667 Random 

6 bz 0.667 Random 

5 e l 0.5 Constant 

'501 0.5 Constant 

Prime 6 0 2 0.5 Constant 

Optics 

s3 1.5 Constant 

sem 0.333 Random 

6 nin 0.333 Random 

s ecs 0.447 Random 

6 0 cs 0.447 Random 

15 g

x 

2.311 

g

y 

1.477 

Vehicle 15 gz 

1.477 

{.6 

Attitude p 

xy 

0.873 * 

0 yZ 

-0. 8397 * 

P xz -0.7467 * 

6 byT 

a 

0.667 Random 

6 bT

za 0.667 Random 

8 bTy 0.667 Random 

a b Tz 

s« 

x 

0.667 Random 

Less than 40 min. each axis constant error plus 0.707 arc 6 

«y 

s« 

z 

) 

min. each axis time varying 

*Correlation Coefficients are Dimensionless. 

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TABLE 2.3-2. (Cont) 

ERROR SOURCE VALUE AND TYPE OF ERROR GIVEN A PARTICULAR VEHICLE 

(ARC MIN.) "2 SIGMA" 

8ax1 1.36 ** Time Varying 

8 a yl 1.025 ** Time Varying 

6 a zl 1.025 ** Time Varying 

**Corresponds to 44 seconds after the last star tracker update. 

TABLE 2.3-3. SUMMARY OF POINTING ERRORS "2 SIGMA" 

BASED UPON ERRORS 

IN TABLE 2.3-1 

BASED UPON ERRORS 

IN TABLE 2.3-2 

Primary Optics Pointing Error 

(Automatic Mode) 

Including all errors, after 2.5 rev. 

(without low "g" accelerometer) 

23.7 arc min. 23.7 arc min. 

Including all errors, after 2.5 rev. 

(with low "g" accelerometer) 

11.3 arc min. 11.3 arc min. 

Neglecting ephemeris and target 

location errors--2 axis alignment monitor on star tracker 

and at time of star tracker 

update 

4.2 arc min. 4.2 arc min. 

Neglecting ephemeris and target 

location errors, and with 

perfect knowledge of attitude 

3.2 arc min. 3.2 arc min. 

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TABLE 2.3-3. SUMMARY OF POINTING ERRORS "2 SIGMA" (Cont) 

BASED UPON ERRORS 

IN TABLE 2.3-1 

BASED UPON ERRORS 

IN TABLE 2.3-2 

Primary Optics Pointing Error 

(Pointing angles derived from 

the ATS, or Slave Mode) 

Boresighting Angles 4.0 arc min. 5.3 arc min. 

0= 20; r = -20° 1 1 

0

2 = -209; E2 = -209 

ATS Pointing Error (Automatic Mode 

and After ATS Boresighting) 

Neglecting ephemeris and target 

location errors but including 2 

axis alignment monitor on 

star tracker and at time of 

star tracker update 

4.58 arc min. 5.65 arc min. 

Same as above, but 44 seconds 

after star tracker update 

5.0 arc min. 6.0 arc min. 

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-30 < E s 15° for the primary optics 

I I < 40° 

-300< E s 600 for the ATS 

I (7< 450 

(In the discussion section of this report, it can be seen that the pointing error varies with 

the pointing angles.) 

These "2 sigma" error values mean that given a large number of vehicles, 95.4 percent of 

them will always have pointing errors less than the value shown in Table 2.3-3. (For the remaining 4.6 percent of the vehicles, the pointing error can be always greater than the 

values shown in Table 2.3-3.) 

It is required that, given any vehicle, the ATS and primary optic pointing error be respectively less than 8 arc min. and 6 arc min. for 95.4 percent of the times that these systems 

are automatically pointed at a target. (This is neglecting target location and vehicle 

ephemeric errors.) In addition, in a slave mode, the primary optics must point to within 

6 arc min. of the ATS line of sight for 95.4 percent of the time. 

Sections 2.3.3 and 2.3.5 of the discussion show that even though the pointing errors summarized in Table 2.3-3 are less than 6 arc min. -- "two sigma" -- the overall pointing requirements are not met for any given vehicle. 

This analysis assumes time varying errors from the PAC structure which exceed 0.5 

degrees of LOS error, however, and are thus in violation of our CEI spec guarantees. 

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2. 3. 2 DESCRIPTION OF ERROR SOURCES AND ERRORS REMOVED BY BORESIGHTING. * 

Figure 2.3-1 and 2.3-2 show sketches of the ATS and primary optical systems. To physically describe each error source, it is helpful to consider that only one error exists at any 

one time. To define optical axis errors, assume that the image of a spot on the center of 

the tracking mirror could be observed on the display screen, (see Figure 2. 3-3) and that 

only optical axis errors, 6 by

, 

6 bz

a

, exist. The angle between the actual and the desired 

optical axes (i. e. , the angle between the lines ao and do in Figure 2.3-3) is defined as 

the optical axis error. 6 by

a 

and 6 bz

a 

are the pitch and yaw components of this error. 

The roll gimbal direction errors, S e laand 6 0 laare defined in Figure 2.3-4 with respect 

to a set of ATS references axes, x

a

, y z . 

a

, 

a 

The mirror mounting error, 6 12a' is defined as the acute angle between a line on the 

mission surface and the mirror pitch axis. The gimbal orthogonality error, 6 0 3a, is 

defined as the angular deviation from 90° between the gimbal pitch and roll axes. A 

similar set of errors is defined for the primary optical system. Next, the errors removed 

by ATS to primary optics boresighting will be discussed. 

When the ATS and primary optical systems are installed within the vehicle, they are 

aligned to a reference system located at the Laboratory Module bulkhead, xb, yb, zb. 

The boresighting procedure assumes that each element of the ATS and primary optics is 

perfectly aligned to coordinate systems, x

a' ya' Zs, 

and xp, yp, z . 

P 

procedure is then used to find the angular rotation between the 3Z

a 

and 

systems. (Or equivalently, the angular rotation between the -2-Z

a 

and 3E'b 

The boresighting 

coordinate 

p 

and rCb, xpcoordinate systems.) 

Now, if any ATS or primary optic components shift in alignment with respect to the 

-x

a 

or

p 

coordinate system (see Figure 2.3-4) the boresighting procedure will equate the 

effect of these misalignments to motions of the entire ATS with respect to the primary 

■••••■•••..._ 

*See also Section 2.6, Boresighting. 

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Figure2. 3-1. ATS and Primary Optics Co

nfiguration(Side View) 

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.............. 

Figure2.3-2. ATS and Main Optics Co

nfiguration(Top View) 

.....-...., 

1 1 

1 

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DESIRED OPTICAL AXIS 

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ACTUAL OPTICAL AXIS 

Figure 2.3-3. ATS Configuration 

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X

a 

6 bz

a 

6 91a 

LOS a 

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OPTICAL AXIS 

6 bya, 6 bza- OPTICAL AXIS ERRORS 

6 0 I A ' 6tG is - ROLL GIMBAL DIRECTION ERRORS 

Y

a 

1 a 

OPTICAL AXIS X

a 

6 bYa 

Figure 2.3-4. ATS Errors 

—SECTREF SPECIAL HANDLING 

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and; 

a (Los Error) 

by a 

- 1.06 without boresighting. 

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optics. This causes an error in the computed value of the alignment or , , a and also 

x y z 

in the primary optics pointing when in the slave mode or in ATS pointing for the automatic 

mode. 

Even though the boresighting procedure does not determine the misalignments of all of the 

ATS and primary optic elements, it does tend to reduce the effects of the static and random 

error. Table 2.3-4 shows the absolute value of the sensitivity of the primary optic and 

ATS pointing error (at nadir) to some of the sources of error for the case with no boresighting and the case with boresighting. It is important to note that even though the boresighting at P 1 = 20°, E = -20° and C22= -20°, E 2= -20° decreased the sensitivity of 

the pointing error to most of the error sources for nadir pointing, the error sensitivities 

can increase for other pointing angles. 

For example, at pointing angles of E = -30, S2 = 45 the sensitivity of the main optic or 

ATS pointing error to the ATS optical axis error in the pitch direction is: 

(Los Error) = 1.0 without boresighting 

by a 

In conclusion, the boresighting procedure tends to reduce the sensitivity of the pointing 

error to the static or random type errors. (Static errors are misalignments that do not 

change during the time interval from the start of a boresighting operation until the next 

boresighting operation.) Also the best reduction in sensitivity of the pointing error can 

be expected to occur for pointing angles close to the boresighting angles. 

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TABLE 2.3-4. SENSITIVITY OF POINTING ERROR AT NADIR 

PRIME OPTICS (SLAVED) ATS AUTOMATIC 

ERROR 

SOURCE 

WITH 

BORESIGHTING 

WITHOUT 

BORESIGHTING 

WITH 

BORESIGHTING 

WITHOUT 

BORESIGHTING 

6 81A 0 2.0 0 2.0 

6 6 LA 0.28 1.0 0.28 1.0 

8 th 2a 0.19 

. 

1.414 0.19 1.414 

6 0 0.3 1.0 0.3 1.0 3a 

6 bya 0 1.0 0 1.0 

(5 bz

a 

0.34 1.0 0.34 1.0 

6 8 1p 0 2.0 2.0 0 

6 0 0.34 1.0 0.657 0 1p 

6 0 2p 0.392 1.414 1.81 0 

6 0 3p 0.43 1.0 1.43 0 

a 

x 

0 1.0 0 1.0 

a 0 1.0 0 1.0 

Y 

a

z 

0 0 0 0 

Boresighting performed for targets at: 

P 1 = 20°, E 1 = -20° 

= -20°, E 2 = -20° 

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2.3.3 PROCEDURE FOR DETERMINING POINTING ERRORS 

To obtain the pointing errors that result from the sources of error in Tables 2.3-1 and 

2.3-2, a computer program using Monte Carlo techniques was written. (The Monte Carlo 

technique was used to allow appropriate probability distributions to be used for each 

error source and to correctly combine the effects of the error sources to obtain the total 

pointing error probability distribution.) Figure 2.3-5 shows the procedure for obtaining 

pointing errors after ATS boresighting. (To obtain pointing errors for the automatic mode 

without boresighting, the same program is used but the part enclosed by the dashed line is 

omitted.) In this technique, values of all errors are initially selected from appropriate 

probability distributions, normal, uniform, etc. -- and gimbal angles are computed for the 

first boresighting target. New error values are then found for randomly varying errors, 

and the ATS and primary optic gimbal angles are computed for the second boresighting 

target. The error in computed ATS/Primary Optics alignment is then found. Finally, 

new error values are found for the random and time varying errors and the ATS or primary optic pointing errors are found for several different pointing angles. This process 

is then repeated several hundred times to gain confidence in the probability distributions 

obtained for the pointing error. The sources of error that are considered to be constant, 

randomly varying or time varying given any particular vehicle were shown in Tables 2.3-1 

and 2.3-2. 

The method used in obtaining the ATS/TM alignment involves a least squares solution. 

(There are three unknowns and four independent equation.) Computer results have indicated 

that the accuracy of the boresignting procedure is degraded if one of the independent equations is dropped and ATS/TM alignment found by standard matrix inversion. 

2.3.4 STATISTICAL TREATMENT OF ERROR SOURCES AND THE TOTAL POINTING 

ERROR. 

Tables 2.3-1 and 2.3-2 show whether each error sources is constant, randomly varying, 

or time varying given a particular vehicle, and show the two sigma values for each error 

source. For the present, it is assumed that the value of each error is normally distributed 

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INPUT ERROR CHARACTERISTICS 

AND BORESIGHTING ANGLES 

E., a; E 2' 2 

INITIALIZE 

PROBABILITY 

DENSITY FUNCTIONS 

SELECT VALUES 

FOR ALL ERRORS 

RANDOM NO, 

GENERATOR 

COMPUTE GIMBAL ANGLES 

FOR FIRST BORESIGHTING 

TARGET E1, C1 

SELECT NEW VALUES 

FOR RANDOM ERRORS 

COMPUTE GIMBAL ANGLES 

FOR SECOND BORESIGHTING 

TARGET E2 , 

OBTAIN SOLUTION FOR 

ERROR IN COMPUTED 

ATS/PRIME OPTICS 

ALIGNMENT, ax, ate, az 

PLACE a x, or , a IN 

y z 

PROBABILITY DENSITY 

REPEAT FOR 

1000 TRIALS 

SELECT NEW VALUES FOR 

RANDOM AND TIME 

VARYING ERRORS 

 NO, 

GENERATOR 

COMPUTE LOS POINTING 

ERROR FOR POINTING 

ANGLES -300 < E < 15° 

I rt < 40° 

PLACE EACH LOS 

POINTING ERROR 

IN PROBABILITY 

IN DENSITY 

PRINT OUT 

PROBABILITY 

DISTRIBUTIONS 

Figure 2.3-5. Boresighting Error Analysis Computer Program 

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and has a zero mean value. Then, for example, if 1,000 vehicles were constructed, the 

"2 sigma" value of 0.5 arc min. for allowable mirror mounting error requires that 954 of 

the vehicles must have mirror mounting errors less than 0.5 arc min. and 46 of the vehicles 

can have mirror mounting errors greater than 0.5 arc min. Also, since the mirror mounting error is considered to be constant, its value must remain constant during the time interval between boresighting. Random and time varying errors are also considered to vary 

randomly from one vehicle to the next. However, given a particular vehicle, it is assumed 

that the time varying errors remain constant only during the boresighting operation, while 

the randomly varying errors can change even during the boresighting procedure. Then to 

properly interpret the total pointing error obtained, one must consider a large number of 

vehicles. For example, the "2 sigma" pointing error of 4 arc min. for the main optics in 

the automatic mode at the viewing angles of E = -30° and ri = 40° means that 954 out of 

1000 vehicles will have pointing errors less than 4 arc min. when pointing to a target located 

at = -30° and SZ = 40°. Now, if a vehicle is constructed with the alignment tolerances in 

Table 2.3-1 or 2.3-2, the pointing error observed with this particular vehicle may always be 

greater than 6 arc min. for some particular stereo and obliquity angle. This indicates that 

it was one of the 46 out of 1000 "bad vehicles". 

Now if the requirement is that for any vehicle the pointing error must be less than 6 arc 

min. at least 95.4 percent of the time ("2 sigma"), then the alignment tolerances in 

Tables 2.3-1 or 2.3-3 are not acceptable and additional work is required to define an 

acceptable set of alignment tolerances. (This is discussed in Section 2.3.) If the requirements mean that pointing errors must be less than 6 arc min. for 95.4 percent of the 

vehicles, then the tolerances in Table 2.3-3 are acceptable. 

Finally, some discussion of the use of the words "2 sigma" is warranted. "Two sigma" as 

used in describing the pointing error, means that given a large number of vehicles, 95.4 

percent of the vehicles will have pointing errors less than the value shown, and does not 

refer to twice the standard deviation of the pointing error. Twice the standard deviation 

of an error only gives the value that is greater than the actual error 95.4 percent of the 

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time if that error is distributed normally and has a zero mean value. From Figure 2.3-6, 

showing the distribution of pointing error, it can be seen that the error is not distributed 

normally -- in fact it closely resembles a Rayleigh distribution. Thus, "2 sigma" value 

used in describing the pointing error refers only to the 95.4 percent level and not to twice 

the value of the standard deviation of the pointing error. 

If all errors are RSS'd to give the pointing error, the difference between the 95.4 percent 

confidence level pointing errors for an assumed normal distribution and the actual distribution for the pointing error becomes even more noticeable. 

If we assume for simplicity that only optical axis errors exist, the primary optics pointing 

error, P 9 , for the automatic mode can be written as; 

e = Sbz2+ 6 by2 

If RSS techniques are assumed, and the "2 sigma" value of optical axis errors, 6 by and 

6bz are 0.667 arc min. , Figure 2.3-7 shows the assumed distribution for the total pointing 

error. For this simple case an exact combination of errors shows the pointing error 

distribution to be a Rayleigh distribution. Comparing the exact and the RSS'd summation 

of errors in Figure 2.3-7, one can see that overly large errors are indicated if only the 

RSS'd technique is used. 

2.3.5 RESULTS OF MONTE CARLO COMPUTER PROGRAMS. 

Figures 2.3-8 through 2.3-12 show the cumulative probability distributions of ATS and 

Primary Optics Pointing errors based upon the error values in Tables 2.3-1 and 2.3-3 

and for selected stereo and obliquity angles. Figure 2.3-13 and 2.3-14 show the error in 

computed value of AT S/Primary Optics alignment. (However, in Section 2.3.1 it was noted 

that the error in the computed value of ATS/TM alignment does not necessarily increase 

pointing error. Constant errors, such as gimbal non-orthogonality, will cause error in 

--lferla SPECIAL HANDLING 

2-205 
•75 

as 

05 

C75 

01 

O 

• 

% PROBABILITY POINTINGERROR IS LESS T

HAN VALUE OFO

RDI

NATE 

Figure2.3-6. Comp

aris

onof Pointing Error Distributions 

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Oi co L CO 1-1 

( • MIN DHV) 1101111a ONLLNIOd SIN 

2-206 SECRET SPECIAL HANDLING 
O 0 

g g 

N 

• ci) 

>/ 

cn 

as 

cn 

O 0 

g 

w a 

a C) 

-tC 

II II 

N .0 .0 

.0 .0 0.667 ARCMIN. 0. 667 ARC MIN. 

1 

1,4 

0 

0 

co 

a) 

a) 

Cr) 

CA

0 

a) 

0 to 

co 

0 

40 

O 

0 

O 

O 

LO 

cs3 

% PROBABILITY POINTING ERROR IS LESS THAN VALUE OF ORDINATE 

Figure2.3-7. Comp

aris

onof RSS'd and Exa

ct Summationof Errors 

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a) OD O 

,—; 

Craw Dui() uoutia ONI.LNIOd SOLI,c10 AA/Wilk' 

iffriaLSPECIAL HANDLING 

2-207 
P 

r-1 

O 

O 

cD eM O CO O CV O 

••••1 

CV 

0 0 .11 

V 

a 

0 

0 

C1') 

11 

r..) 

z o gi 

-,4 C) 44 

.3 B `

El f.-4 44 E.-, 

E. 

El El 

F 4, 44 

-,C g 

4 f1,4 a 

41 En 

w 

41 

,24 ..4 44 w 

E- P 

WC4'- .4 0 3 0 14 0 44 144 Z 44 Z 

""4 41 

0H0 

g 

•tC 

8 F 

az 

Ei 4 

rcS1 

-14 

E. ..4 A Z 4 <4 

4 c- „

M r, 

;1 B 

w 4 

w E-I 

.4 ao 00 14 

44 f:4 g Z 44 0 

•( rn 

E-4 

cn 

t:4 0 

1-1 

1-4 

44 

4C ) 

BASEDUPON ERROR VALUES IN TABLES 1 OR3 

O 

C) C) 

co 

0 

C) 

O oo 

O

tO 

0 

0 

•14 

O 

O N 

% PROB POINTING ERROR IS LESS T

HAN VALUE OF THE ORDINATE 

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2-208 

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Figure2. 3-8. Primary Optics Pointing Error(Automatic Mode) 
% PROBABILITY POINTING ERROR IS LE

SS THAN VALUEOF ORDINATE 

Figure2.3-9. Primary Optics Pointing Error(Slave Mode) 

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CI) N co in ..t. VD CNI 

........., 

•1411141 DIIV 

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2-209 
% PROBABILITYPOINTINGERROR IS LESS T

HAN VALUEOFORDINATE 

Figure2.3-10. Primary Optics Pointing Error(

Slave Mode) 

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cm . , w .., .1, Co) CN) 

•NITAI 311V 

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2-211 ET SPECIAL HANDLING 

Oi OD N C) •zr 

C) 

.t0 

rn 

ci 

U) 

I0 

Figure2.3-11. ATS Pointing Error(Arc Min.)(Automatic Mode- After Boresighting) 
0 

0 

1 cr5 

FA --- a; 

r. 

2 

co 

• c) 0 0 

C c., 0 W = 

if II 

cr.) 

rn 

0 

G 0 CD 

II 

0 

110 .44 

II 

14 a 

0 0 

co 0 

eq 

N N 

CV 

W 

0 

0 CV 

CV 

N N ■-•1 CNI 

c a 

00 CD d c,0 CV BORESIGHTING ANGLES BASED UPON ERROR VALUES IN TABLE 3 

ti 

O 

O 

O 

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° NIW 31IV 

2 —212 

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Figure2.3-12. ATS Pointing Error(After Boresighting) 
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E-4 

• 

1-4 

0 

• 

• • 

• • N 

• • • 

• 

"2 SIGMA" 

4 

• • 

• 

• 

• 

• 

• 

• 

• 

• • • 

• • 

OD t— CD O C■1 1-1 

2-213 

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• 

BORESIGHTING ANGLES 

O 

0 0 0 

N I C■I 

;z1 

1-4 

PI 

aO

o

z

GL 

• 

• 

• 

• 

• 

• 

•-• 

0 

a) 

co a) 

C) 

O 

O a) 

O 

oo 

O

to 

co 

O 

0 

0 M 

O

N 

0 v-1 

C) 

t—I 

% PROBABILITY ERROR LESST 

VALUE OF ORDI

NATE 

Figure 2.3-13. Error In Computed Valueof ATS/Primary Optics Alignme

nt(Arc Min.

• 

) 

• 

• 

• 

I I I I I I I I 
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C3 

C) 

rn 

C) 

0 

0 

GO 

0 ti 

0 

0 

0 

0 

Co3 

O 

N 

O 

ti 

L) 

I-1 

% PROBABILITYERROR IS LESS T

HAN VALUE OF THEORDINATE 

Figure2. 3-14. Error In Computed Valueof ATS/Primary Optics Alignme

nt(Arc Min.) 

2-214 

OD Ls. cp d4 

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the computed value of ATS/TM alignment, but since the computed alignment does contain 

information on the constant error, the total effect is to improve the ATS or primary optics 

pointing when using the ATS/TM alignment. 

2.3.6 PROCEDURE REQUIRED TO DETERMINE ALLOWABLE MISALIGNMENTS. 

In Section 2.3.1, it was stated that if the pointing requirements mean that for any MOL vehicle, 

the on-orbit pointing errors must be less than 6 arc min. (8 arc min. for the ATS) at least 

95.4 percent of the time, then the alignment tolerances in Tables 2.3-1 and 2.3-3 are not 

acceptable. Assuming this to be the case, additional work is required to determine correct 

alignment tolerances. 

One approach is to select tolerances from Table 2.3-1 or 2.3-2 and accept the risk of 

having any particular vehicle always show excessive pointing errors for some particular 

pointing angles. The magnitude of this risk will depend upon the probability distribution of 

the pointing angles as well as the values of allocated errors. Based upon the curves shown 

in Figures 2.3-8 to 2.3-12, Table 2.3-5 shows the limiting values of the risk involved in 

accepting either Table 2.3-1 or 2.3-3 as being acceptable error tolerances. The actual 

value of the risk involved can be obtained by using the Monte Carlo computer programs with 

appropriate probability distributions of ATS and primary optics pointing angles. 

Another approach is to: use the worst possible combination for the constance errors, allow 

random and time varying errors to vary, and use target pointing angle probability distributions to obtain the probability of having pointing errors be less than 6 or 8 arc min. 95.4 

percent of the time. 

If the targets are assumed to be uniformly situated on the ground, and the ATS or Primary 

Optics is assumed to randomly point at a target, the probability density distributions of the 

pointing obliquity and stereo angles are: 

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1 

f

o 

- 2 2 tan O

ce 

cos 0 

1 

f E ( E) - E 2 - Ell 

0 

E

1 <Es E2 

and 

where 

Ct

o 

is the maximum obliquity angle. 

E 1 is the maximum aft stereo angle. 

E 2 is the maximum forward stereo angle. 

Using these probability distributions, and assuming that all constant errors have the worst 

case and sign, the Monte Carlo program was used to obtain the preliminary on-orbit error 

apportionment shown in Table 2.3-6. The pointing errors obtained from this error 

apportionment are shown in Figure 2.3-13. 

From Figure 2.3-15, the primary optics pointing error is less than the required 6 arc min. 

for the 95.4 percent of the time that any particular vehicle points at a target. The ATS 

pointing error is 8.2 arc min. as compared to the required 8 arc min. However, since the 

probability is very small that all constant errors would have the worst case magnitude and 

sign the probability of having the actual pointing error observed for any ATS exceed 8.0 arc 

min. 95.4 percent of the time is small. 

Future efforts will refine the apportionment of allowable system errors. 

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TABLE 2.3-5 

PROBABLE RISK USING ERROR 

VALUES FROM: 

TABLE 2.3-1 TABLE 2.3-3 

Prob. Prime Optics Pointing Error is 

always greater than 6 arc min. 

< 0.002 < 0.002 

• Automatic mode* > 0.0001 > 0.0001 

• Slave mode < 0.003 < 0.02 

(after boresighting) > 0 > 0.0001 

Prob. ATS Pointing Error is always 

greater than 8 arc min. 

• Automatic mode* < 0.001 < 0.004 

(after boresighting) > 0 > 0.0001 

*Neglecting vehicle ephemeris and target location errors 

TABLE 2.3-6. APPORTIONMENT OF SYSTEM ERROR 

MAXIMUM VALUE REMARKS 

Constant Errors 

1.2 arc min. 

1.2 arc min. 

0.5 arc min. 

0.5 arc min. 

0.5 arc min. 

0.5 arc min. 

15 arc min. 

(total) 

apportionment 

apportionment 

apportionment 

apportionment 

ITEK Spec. EC-331B 

ITEK Spec. EC-331B 

apportionment 

6 by a 

6 bz 

a 

691a 

641)la 

6 /P2A 

6 kb 3A 

6°1 x 

8« 

Y 

6a 

z 

ATS Optical Axis 

ATS Roll Gimbal Direction 

ATS Mirror Mounting 

ATS Gimbal Orthogonality 

ATS/Prime Optics Align 

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TABLE 2.3-6. APPORTIONMENT OF SYSTEM ERROR (Cont) 

MAXIMUM VALUE REMARKS 

Constant Errors (Cont) 

0.5 are min. 

0.5 arc min. 

0.5 arc min. 

1.5 arc min. . 

2 Sigma Value 

(arc min.) 

0.707 

0.707 

0.52 

0.52 

0.707 

0. 707 

0.707 

0.3535 

0.3535 

2 Sigma Value 

(arc min.) 

f. 

±. 33 *1 

2.0 (total) 

Apportionment 

Apportionment 

EK Spec 101.2 

EC-1204 

Apportionment 

Apportionment 

±. 75 min. total due to 

pressure variations 

Estimate--includes 

pressure and thermal 

variations 

0.5 min. total due to 

thermal effects 

16 bit shaft encoder 

Estimated error after 

settling 

6 0 1 Prime TM Roll Gimbal Dir. 6 th 1 

6 4 2 TM Mirror Mounting 

80 3 Gimbal Orthogonality 

Time Varying Errors 

8 by 

ATS Optical Axis 6 bz 

a 

6 0 la) ATS Roll Gimbal Dir. 8 0 la 

8

a 

x 

6a ATS/Prime Optics Align 

Y 

8 a 

z 

6 e 1 ) Prime TM Roll Gimbal Dir. 6 1 4 

Random Errors 

6 w 

6 0 am 

a 

o 0cs 

a 

8 0 cs 

a 

33 *1 m 

ATS Gimbal Angle Meas. 

Control System 

Pointing of TM 

*1 This error has a uniform distribution. f. 33 arc min is the absolute maximum value. 

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TABLE 2.3-6. APPORTIONMENT OF SYSTEM ERROR (Cont) 

MAXIMUM VALUE REMARKS 

Random Errors (Cont) 

0.667 

0.667 

f0.33 *1 

±0.33 *1 

1.0 (total) 

0.667 

0.667 

Estimate of alignment 

monitor error 

16 bit shaft encoder 

Estimated error after 

settling 

Corresponds to 95 ft. 

each axis at nadir 

8 by Primary Optical Axis 

0 bz Alignment and Monitor Error 

) 

8 ("2m 

Gimbal Angle Meas. 

8 e m 

8 CI Control System 

cs Pointing of TM 8 e 

8 byT ATS or Prime Optics 

0 bzT Tracking Error 

*1 This error has a uniform distribution. ±0.33 arc min is the absolute maximum value. 

-SECRET SPECIAL HANDLING 

2-219 
% PROBABILITY POIN

TINGER

ROR IS LESS T

HAN VALUE OF THE ORDINATE 

Figure2. 3-15. Lineof Sight Pointing Error 

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2.4 ATS THERMAL ANALYSIS 

The material in this section pertains to the thermal design of the Acquisition Optics Components. The material is subdivided as follows: 

2.4.1 External Optics Assembly 

2.4.2 Telescope Assembly 

2.4.3 Thermal Development Test Plans. 

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2.4 ATS THERMAL ANALYSIS 

2.4.1 EXTERNAL OPTICS ASSEMBLY 

2.4.1.1 General External Component Design Philosophy 

2.4.1.1.1 Passive Versus Active Thermal Control 

Both active and passive thermal control systems have been investigated in developing a 

design concept. An active control system was discarded very early in the analysis on the 

basis of the additional weight and reliability penalty associated with such a system. Two of 

the active control systems considered were: 

a. Thermostatically actuated heaters 

b. Movable louvers 

A passive system is one employing coatings, component thermal inertia, internal conductance, and insulation to maintain a system within a given temperature range. 

2. 4.1.1. 2 Design Concepts and Configurations 

The basic configuration consists of a shroud with a hinged panel which totally encloses the 

tracking mirror and fixed fold assembly. During a target pass, the hinged panel opens up 

to provide a viewing slot approximately 24 in. x 44 in. This shroud serves four purposes: 

a. Provides an accurately controlled temperature enclosure. 

b. Protects the optical system from direct insolation. 

c. Protects the external optics from contamination by the orbital maintenance 

thrusters and the output of the waste management system. 

d. Protects the external optics from the heating effects resulting from firing of the 

orbital maintenance thrusters. 

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2-223 
ALUMINIZED MYLAR OR 

KAPTON 

Al FOIL/TISSUGLAS OR 

S/S FOIL OR Ni FOIL 

INSULATION 

Be, Ti OR S/S SHIELD 

HIGH TEMPERATURE 

PLASTIC TEMPERATURE 

CONTROL SURFACE 

RADIATOR SHELL 

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Very early in the program, the concept of a single shroud was proposed. This shroud 

would serve two functions: 

a. Serve as an aerodynamic fairing during ascent. 

b. Function as an orbital protective shroud once orbit has been achieved. 

This concept was discarded by Aerospace as representing potentially too large a weight 

penalty plus an extremely cpmplex interface. In lieu of this, a dual shroud concept has 

evolved. The outer shroud serves solely for protection against ascent heating and pressure 

loading effects. The outer shroud is ejected prior to orbit injection. The inner shroud or 

on-orbit shroud functions as a protective cover during all orbital operations. The following 

design concepts for the protective shroud have been identified along with their relative advantages and disadvantages. 

a. Single wall shroud 

2-224 

-HUH- SPECIAL HANDLING 
Ti, Be OR S/S SHIELD 

(GOLD PLATED INSIDE) ALUMINIZED MYLAR OR 

KAPTON INSULATION 

INSULATING STANDOFF 

RADIATOR SHELL 

AL OR Be INNER SHELL 

HIGH TEMPERATURE PLASTIC 

SUCH AS IMIDITE 1850 OR 

TEXTOLITE GRADE 11617B 

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Advantages: 

1. Composite insulation layup optimizes insulation weight. 

2. High temperature plastic temperature control surface helps isolate 

control surface from hot surfaces. 

Disadvantages: 

1. Design fairly sensitive to accuracy with which one can analytically 

predict transient characteristics of composite insulation. 

2. S/S or Ni foil, if required from temperature standpoint, represent a 

significant weight penalty. 

b. Double wall shroud 

Advantages: 

1. Somewhat easier to analytically predict characteristics than configuration a. 

2. High temperature plastic surface tends to help isloate temperature control 

surface from the higher temperature structure. This should help to lessen 

impact of heating effects from the on-orbit thruster plume. 

----11C-REL SPECIAL HANDLING 

2-225 
RADIATOR SHELL 

HEAT OF FUSION SALT OR WAX 

T., Be OR S/S SHIELD 

ALUMINIZED KAPTON OR 

MYLAR INSULATION 

HIGH TEMPERATURE PLASTIC 

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3. The shield or shields could be optimized to cover just the most critical 

are as . 

Disadvantages: 

1. Possibly higher weight than configuration a. 

c. Heat of fusion shroud 

Advantages: 

1. Maintains constant temperature during plume impingement. 

2. Highly predictable performance analytically. 

Disadvantages: 

1. High weight penalty. Most potential heat of fusion materials have latent 

heat of fusions on the order of 100 - 150 BTU/lb. This system might 

represent as much as a 30 lb weight penalty per system. 

2. Significant development work required. 

2-226 

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(GOLD PLATED BACK FACE) 

•••• 

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CONFIGURATION 

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d. Shadow shield 

ACTS 

Advantages: 

1. Minimizes direct heating problem. 

2. Minimizes import of plume impingement problem on system design. 

Disadvantages: 

1. Creates very difficult interface problems on account of weight penalty 

plus impact on DACO radiator design. 

The proposed approach is a combination of concepts a and b. This concept would consists of 

a high-temperature shroud with a composite insulation blanket. Covering a portion of this 

shroud would be a partial shield which would provide a degree of protection from the 

peak heating effects of the orbit adjust thruster firing. This configuration is shown in 

Figure 2.4-1. 

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COMPOSITE INSULATION 

OUTER LAYERS - S/S OR 

AL FOIL 

INNER LAYERS - ALUMINIZED 

MYLAR 

i1 

INSULATION 

STAND-OF F 

Figure 2.4-1. Proposed Shroud Configuration 

All of the interior surfaces of the shroud are insulated with the exception of the lower earthfacing surface. This lower earth-facing surface constitutes the basic temperature control 

surface for the externally mounted system. The coating pattern on this control surface is 

adjusted to minimize the temperature variation over a given orbit and to limit the temperature variation over the entire orbit mission. The lower earth-facing surface was chosen 

for this purpose since it is exposed to the smallest solar heating input. Hence, the variation this surface is subject to in heat input is minimized. Also, this surface is relatively 

isolated from the effects of the orbit adjust thruster firing. 

2-228 

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2.4.1.2 Temperature Constraints 

Table 2.4-1 gives a summary of allowable temperature variations for the external components, both for operational and non-operational modes. 

TABLE 2.4-1. ALLOWABLE TEMPERATURE VARIATIONS FOR 

ACQUISITION TRACKING SYSTEM 

MINIMUM 

TEMPERATURE 

MAXIMUM 

TEMPERATURE 

Roll and Pitch Gyro 0°F +100°F 

(Mounting surface) 

Pitch and Roll Motor -65°F +160°F 

Pitch and Roll Encoders - 5°F +160°F 

Roll and Pitch Amplifiers -65°F +160°F 

Mirrors -10°F +140°F 

2.4.1.3 Environment 

2.4.1.3.1 Orbit Envelope 

Figure 2.4-2 shows the mission orbit envelope. Five cases have been analyzed to bracket 

the expected range of external heat inputs. 

2.4.1.3.2 Orbital Heat Rates 

Table 2.4-2 summarizes the range of orbital heating rates predicted for the mission as 

presently defined. 

'MKT SPECIAL HANDLING 

2-229 
80 100 

MAX. APOGEE (ha) 

220 

200 

100 nm 

-60 + 60 

ALTITUDE NM 

MIN. PERIGEE (h ) 

z 

75 

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INCLINATION 

Figure 2.4-2. Mission Orbit Envelope 

TABLE 2.4-2. TABULATED VALUE OF NOMINAL VALUES OF SOLAR (S), ALBEDO (A), 

EARTH (E), AND MOLECULAR HEAT FLUXES (QmoL) ON THE ON-ORBIT SHROUD 

CASE NO. $ DAY PERIGEE APOGEE 

E 

Btu/hr-ft2 

§-F1 

Btu/hr-ft2 

1,:' MO L 

Btu/hr-ft2 

1 -60° 2 75 nm 125 nm 27.5 114.0 71.4 

2 0° 2 75 nm 125 nm 26.6 88.7 71.1 

3 0° 82 80 nm 180 nm 26.6 83.9 28.6 

4 +60° 182 220 nm 220 nm 24.0 51.2 0 

5 -60° 2 100 nm 100 nm 27.5 114.0 35.1 

Note: (1) The average molecular heat flux cited is for the forward face only since 

the flux level on the remaining surfaces is negligible. 

(2) /3 = Angle between earth-sun line and orbital plane. 

2-230 

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2.4.1.3.3 Free Molecular Heating 

The predicted molecular heating curves for the above cases are shown in Figure 2.4-3. 

2.4.1.4 External System Temperature Behavior 

2.4.1.4.1 Ascent Heating 

The on-orbit shroud is covered at liftoff, as previously described, by an external shroud, the 

function of which is to protect the on-orbit shroud from excessive temperatures and pressure loading during ascent. It also serves the auxiliary function of limiting the system 

temperature during the prelaunch period when it is exposed to an extended period of solar 

heating. Figure 2.4-4 shows the ejectable shroud peak temperature time behavior. The 

integrated average temperature-time behavior is considerably less than this. The period 

between ejection of the ascent fairing and orbit injection has not been evaluated in detail 

since the time of fairing ejection has just recently been defined. Figure 2.4-5 shows the 

expected on-orbit shroud temperature-time behavior during ascent, based on the assumption that ejection of the aerodynamic fairing will not occur before 290 seconds after launch. 

2.4.1.4.2 Normal Orbit Characteristics 

The first 90-minute portion of Figures 2.4-6 and 2.4-7 shows the shroud temperature-time 

behavior for the hot orbit condition. Both the tracking mirror and the fixed fold assembly 

are maintained at virtually a constant 70°F over the entire orbit during a non-operational 

mode. The tracking mirror varies less than 2°F over the orbit. Since in an operational 

mode the mounting surface for the roll gyro is approximately 30°F hotter than the temperature of the shroud control surface, the choice of coatings was dictated by the requirement of maintaining the average temperature of the control surface on the shroud to within 

0-70°F overall normal orbit conditions. The actual estimated operating range for the 

tracking mirror is 55 1 15°F, and for the folding mirror, the predicted variation is 

53 ± 17°F. 

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Face of On-Orbit Shroud 

2-232 1-EtRET-- SPECIAL HANDLING 
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Figure 2.4-6. Shroud Temperature History During Thruster Firing 

-IMO- SPECIAL HANDLING 

2-235 
280 

240 

200 

120 

40 80 120 160 200 240 

-40 

0 

2 

V 

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TIME, MIN. 

Figure 2.4-7. Shroud Temperature History During Thruster Firing 

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Table 2.4-3 gives a summary of the assumed coating pattern for the on-orbit shroud. 

TABLE 2.4-3. SUMMARY OF COATING PROPERTIES 

ASSUMED FOR ON-ORBIT SHROUD 

SURFACE EMITTANCE SOLAR ABSORPTANCE 

1 0.75 0.3 

2 0.75 0.3 

3 0.10 0.35 

4 0.16 0.14 

5 0.75 0.3 

6 0.2 0.12 

2.4.1.4.3 Operational Mode - Gyro Warmup 

The effect of the warmup period for the gyro on the tracking mirror has been analyzed on 

the following basis: 

a. Cold case (sink temperature = 0°F) 

b. 60-watt coarse heater energized in gyro when gyro temperature is in the temperature range 159°F. 

c. A 10-watt fine control heater is utilized to maintain the gyro at 160 f 5°F. 

d. Vehicle pressure shell maintained at 70 ± 5°F. 

The following temperatures have been predicted for the end of a twenty minute operational 

period: 

Roll motor = 55°F Bezel = 17°F 

Yoke = 30°F Tracking Mirror = 11°F 

Belly band = 40°F 

2-237 

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2.4.1.4.4 Orbit Thruster Heating 

A. Orbit Adjust Maneuvers - The following is a description of the expected modes of 

operation: 

Orbit adjustments are made every three days to reinitialize the orbit starting with revolution No. 44 and every 48 revolutions thereafter. 

Mode: Non-intersecting initial and decayed orbits 

a. Two-burn procedure 180° apart 

b. Both thrusts forward 

c. Firings occur at equator 50 

d. First burn of 4 minutes' duration occurs when vehicle is proceeding from north to 

south (sunlit side of earth) 

e. Second burn is of 0.4 minute duration occurring when vehicle is proceeding from 

south to north (earth's umbra) 

Mode: Intersecting initial and decayed orbits. Two procedures are currently being considered to reinitialize the orbit. 

a. Two-burn procedure 

1. First thrust vector forward 

2. Second thrust aft 

3. First burn occurs when vehicle is proceeding from north to south and is of 

4 minutes' duration. 

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4. Second burn is of 0.4 minute duration occurring when vehicle is proceeding 

from south to north 

5. Firings occur at equator ±5° 

b. Single-burn procedure 

1. Thrust vector forward 

2. Single burn would occur at the equator 15° in the descending mode, i.e. , 

vehicle is proceeding north to south 

3. The firing cycle would be four 4-minute burns each separated by a 3-day period 

followed by a 3-day interval and then a 10-minute burn 

The two-burn technique is currently the baseline design, although the single-burn technique 

is under consideration. The single-burn technique would not fully correct for perigee drift 

so that the one-burn technique might be alternated with a two-burn technique. In the twoburn technique described for an intersecting degraded and initial orbit, the vehicle is 

turned around the yaw axis. This maneuver takes approximately one minute. 

B. Plume Characteristics - The plume characteristics supplied to us by DACO have been 

reproduced in the accompanying figures (Figures 2.4-8 and 2.4-9) and in Tables 2.4-4 

and 2.4-5. 

Figure 2.4-10 shows the relative position of one of the Y-axis thrusters and the on-orbit 

shroud. For purposes of analysis, the shroud has been divided into a number of nodes 

for which the heat flux distribution has been calculated. Figure 2.4-11 shows the shroud 

surface identification. 

The plume characteristics supplied to us by DACO essentially assume that the flow before 

impinging on the shroud is that of a free jet. This neglects the interaction of the boundary 

SECRET SPECIAL HANDLING 

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O 1-1 O to 

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2-240 

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NRO APPROVED FOR 

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0., 

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layer and the shock on the cylindrical lab module with the shroud. The oblique shock 

(Figure 2.4-12) is estimated to lie about 2° above the radiator skin; thus the majority of the 

shroud extends above such viscous effects. Also neglected is the effect on the plume properties of the constriction of the plume by the radiator. The effect of the radiator on the 

inviscid flow field near the radiator is to increase the density and hence the heat flux. On 

the other hand, the viscous effects on the lab module radiator tends to reduce heat transfer 

to the body since the gas in the boundary layer is cooled. Also the adverse pressure 

gradient caused by the shroud in the boundary layer might cause the boundary layer to 

TABLE 2.4-4. MOL FRACTION OF CHEMICAL 

SPECIE IN THRUSTER PLUME 

SPECIE MOLECULAR FRACTION 

CO 0.1287 

CO2 0.0401 

H 0.0179 

H9 0.1576 

H2O 0.3343 

N2 0.3075 

NO 0.0013 

O 0.0008 

OH 0.0111 

02 0.0007 

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TABLE 2.4-5. FLOW PROPERTIES FOR ORBIT ADJUST PLUME 

MACH 

NUMBER 

DENSITY 

LB/FT3 

TEMP 

°R 

VELOCITY 

FT/SEC 

GAMMA 

5.0 0.1423 x 10-3 1259 9,445 1.327 

6.0 0.5586 x 10-4 916 9,757 1.349 

7.0 0.2544 x 10-4 692 9,942 1.362 

7.5 0.1791 x 10-4 609 10,009 1.366 

8.0 0.1290 x 10-4 540 10,064 1.370 

8.5 0.9486 x 10-5 482 10,110 1.373 

9.0 0.7096 x 10-5 432 10,148 1.375 

9.5 0.5392 x 10-5 390 10,181 1.377 

10.0 0.4161 x 10-5 353 10,209 1.379 

10.5 0.3244 x 10-5 322 10,233 1.381 

11.0 0.2559 x 10-5 294 10,254 1.382 

11.5 0.2042 x 10-5 269 10,272 1.384 

11.75 0.1830 x 10-5 258 10,280 1.385 

12.0 0.1645 x 10-5 248 10,289 1.385 

12.5 0.1338 x 10-5 229 10,303 1.386 

13.0 0.1097 x 10-5 212 10,315 1.383 

13.5 0.9062 x 10-6 197 10,326 1.389 

14.0 0.7543 x 10-6 183 10,336 1.389 

14.5 0.6327 x 10-6 171 10,345 1.389 

15.0 0.5337 x 10-6 160 10,353 1.389 

15.5 0.4527 x 10-6 150 10,361 1.389 

16.0 0.3859 x 10-6 141 10,367 1.389 

17.0 0.2842 x 10-6 125 10,379 1.389 

18.0 0.2129 x 10-6 112 10,389 1.389 

19.0 0.1620 x 10-6 101 10,397 1.389 

20.0 0.1249 x 10-6 91 10,404 1.389 

27.5 0.2471 x 10-6 48 10,435 1.389 

2-243 

-lteRg- SPECIAL HANDLING 
Y 

ON-ORBIT 

SHROUD 

STA 

698.4 

STA STA 

658.5 652.5 

STA 

565 

ON-ORBIT SHROUD 

ORBIT ADJUST THRUSTER 

SURFACE 2 

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SURFACE 3 

FORWARD FACE 

SURFACE 4 

z 

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Figure 2.4-10. Relative Position of Orbit Adjust Thrusters 

and the On-Orbit Shroud 

NOTE: AXES SHOWN DO NOT CORRESPOND TO VEHICLE AXES 

Figure 2.4-11. Velocity Coordinates and Shroud Surface Identification 

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separate from the lab module somewhere in front of the on-orbit shroud. As a result, 

there would be a dead region at the base of the shroud with a resultant reduction in heat 

transfer. As a conservative analysis, we have evaluated the increased density po, in 

the inviscid flow to evaluate the heat transfer. 

Several other simplifying assumptions have been made: 

a. There is no X component of the flow (see Figure 2.4-11). The flow vector is 

radial from the nozzle and impinges with angles of a = -1.7° and y = 13.75°. 

b. Boundary layer theory can predict the heat transfer even though Reynold's 

number is low. 

c. Radiation from the gas behind the normal shock is negligible compared with the 

convective flux. 

The heat transfer to the stagnation point of the cylindrical corner between surfaces 1 and 2 

is given by: 

q

o 

= 0.57 Pr2

-0.6 /3 p2 μ2 Cp (Taw- T)(1) 

where: 

Subscript 2 denotes conditions behind the normal shock formed in front of the cylinder 

Taw= adiabatic wall temperature = Tstagnation 

$ = velocity gradient at the stagnation point 

= 0.3 U /R 

co o 

Pr 

(1)Truitt, R. W. , "Fundamentals of Aerodynamic Heating, " New York, Ronald Press Co. , 

1960. 

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P2 = 6 Pco 

p = density 

The calculated values of these constants are: 

M = 19 

p = 0.162 x 10-6lb/ft3 

U = 10,379 ft/sec co 

T

stagnation = 7550°R 

C

p 

 = 0.285 Btu/lb, °R (= 1/2 U2,0/Tstag notion) 

Pr = 0.7 

= 5 x 10-5lb/ft, sec 

Hence for those parts of the shroud which protrude above the oblique shock (X coordinate 

7 in) the heat transfer is: 

0.779 x 10-4 

clo 

(6310 Tv") Btu 

R

o 

ft, sec 

The first order effect of the interaction between the plume and the cylinder will be a 

region (inviscid) of increased density along the surface of the cylinder. This flow, impinging on a normal surface, will result in higher heat transfer than that due to impingement 

of the jet without interaction. The increased density will lie behind an oblique shock which 

lies close to the cylinder. Thus, only a part of the on-orbit shroud will be affected. 

The flow along a line to the on-orbit shroud is considered. Local mach number and flow 

angle determines a local shock angle a shown in Figure 2.4-12. The abscissa is the 

2-246 

SECRET SPECIAL HANDLING 
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2-247 

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station number; thus, at station 568, the local conditions are evaluated at 26° radial position from the nozzle. The shock shape is plotted to scale in Figure 2.4-13 using local a 

to determine the intersection of the shock with the shroud. The ramp at station 558.5 

would produce a shock and an expansion fan, but these have been neglected since the ramp 

width (6 in.) is small. About 7 in. of the on-orbit shroud are below the shock. Figure 

2.4-14 shows the flow conditions behind the oblique shock evaluated using local angles and 

mach numbers. Note that the density is increased by a factor of 1.8 over po:. . The flow 

between the cylinder and the shock is supersonic and produces a normal shock in front of 

the shroud with a further increase in density by a factor of about 6. The analysis is valid 

with p replaced by 1.8 p. 

Those parts of the shroud which do not protrude above the oblique shock (1. e. , X coordinate 

less than 7 in.) have heat transfer increased by a factor of AIT..T ; i, e. , 

1.035 x 10-4 

q - (6310 - T ) Btu 

o R

o 

w ft , sec 

There are two other significant errors. One of these is due to the unknown composition and 

the thermodynamic properties behind the shock. The composition in the free jet, approximately 

CO = 0.1387 

112 = 0.167 

H2O = 0.360 

H

2 = 0.3343 

2-248 

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by mole fraction, gives (based on R. H. Kohrs Advanced Aerospace Phys. Tech. Memo 

No. 30) 

Pr = 0.815 

= 5.26 x 10-5 

0.99 x 10-4 

OR q - (6810 - Tw) 

approximately 4 percent lower than the previous formula. 

Another inaccuracy is due to the assumption of two-dimensional flow around the cylinder. 

R o 

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tween 1 and 2). For > 1, the curved nature of the cylinder would make the flow tend 

toward that of an axisymmetric stagnation point which would have heat transfer given by 

\r2- times the qogiven previously. 

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function of distance from the stagnation point. The heat flux to surface 3 is the same as 

surface 2 along their common edge and the same as surface 1 along their common edge. 

The two-dimensional distribution within surface 3 is assumed to be the product of the two 

distributions in perpendicular directions. The other surfaces receive no convective flux. 

2-251 

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C. Shroud Temperature Behavior During Thruster Firing - One condition has been analyzed 

to date. For the purpose of evaluating the average temperature behavior of the shroud, a 

3-inch radius was assumed for the edge common to the forward face of the shroud and the 

top sector of the shroud. The maximum temperature on the forward face of the on-orbit 

shroud is estimated to be about 800°F based on (1) neglecting condction effects; and (2) using 

the relationship 

(Tmax T 

avg 4) Qmax) Qavg 

where 

Q = heating rate 

T = absolute temperature 

The average temperature peaks out at about 600°F. Figures 2.4-6 and 2.4-7 shows the 

impact of the thruster firing for the one typical case cited. The first 90-minute period is 

typical of a:hot-orbit case as previously cited. The thruster firing is assumed to occur at 

the equator:on the sunlit side of the earth. 

The temperature of the external optics rose approximately 2°F during the thruster firing 

transient. Figures 2.4-6 and 2.4-7 also indicate that the system will essentially recover 

from the thruster firing transient within two orbits. The insulation was represented, for 

the purposes of this analysis, as a material with an effective emittance of 0.03. Most of 

the data accumulated to date on typical carefully installed multi-layer insulation blankets 

shows a range of 0.01 - 0.03 for the effective emittance at temperatures closer to room 

temperature. As the average blanket temperature increases, the effective emittance 

actually decreases so that a value of 0.03 is a highly conservative value. As discussed 

in section 2.4.1.6, the insulation can, under normal orbital conditions, have a greater 

effective thermal conductivity by a factor of almost 10 without deleterious effects. 

2-253 

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2.4.1.4.5 Gyro Response Characteristics 

In order to maximize the gyro reliability, the gyro is activated just prior to an activity 

mode and then deengerized following the activity mode. Temperature control is achieved 

by means of two thermostatically regulated heaters. These heaters consist of a 50-watt 

coarse control heater for rapid heat-up plus a 10-watt heater for fine control. The 50-watt 

heater is activated when the gyro temperature is more than 5°F below the set point. The 

gyro set point is 160 ± 5°F. The fine control heater is activated when the gyro temperature 

is less then the set point. Figure 2.4-16 shows a typical gyro temperature-time curve 

for a 0°F sink temperature. In order to analyze the heatup and cooldown transient, it was 

assumed that the gyro case itself was maintained at 0°F. This assumption results in predicting a longer heatup period than will actually occur. Also, this tends to predict a lower 

temperature after cooldown than will actually occur during a series of active orbits. 

2. 4. 1. 5 Coating Selection 

The function of the coatings employed on the on-orbit shroud is twofold: 

a. Prevent spurious light reflections from entering the optical system 

b. Control the shroud temperature and, concurrently, the temperature of the 

optical elements within the defined limits. 

The first objective is accomplished by coating all of the non-optical surfaces internal to 

the on-orbit shroud with a thermodynamically black coating. The second objective is 

achieved by means of an array of coatings. The method for selecting the coating array 

is as follows: 

a. The shroud is divided into several discrete zones. 

b. The variation in the orbital average environmental heat fluxes can be calculated 

for each zone. 

c. For those zones which are subject to any significant heating from the orbit adjust 

thruster firing, a coating with a low solar absorptance and high emittance will 

2-254 

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be used to limit the peak temperature which these surfaces will be subject to 

during the aforementioned short-term transient. 

d. For the remaining zones, a coating pattern will be chosen with two characteristics: (1) Lowest emittance coating available compatible with the required solar 

absorptance. This tends to minimize the temperature variation of the skin over 

the course of an orbit; (2) Choose coatings which, under the condition of peak 

environmental heating (exclusive of the thruster firing temperature) and peak 

internal heat generation, will limit the average skin temperature to less than 70°F. 

This is compatible with maintaining the internal equipment within the temperature 

range of 0 - 100°F. Coatings have not yet been specified in detail. The above 

philosophy will be followed in selecting specific materials. 

2.4.1.6 Insulation Materials 

Two insulation blanket configurations are being proposed. These two assemblies are shown 

in Figure 2.4-17. Table 2.4-6 shows some of the materials currently being reviewed for 

this application. 

/ VEHICLE 

SKIN 

...%"./.4,/ 

VEHICLE 

SKIN 

•1/1/Aev-v 4

HIGH TEMPERATURE 

BLANKET 

LOW TEMPERATURE 

BLANKET 

(a) HIGH TEMPERATURE (b) COMPOSITE BLANKET 

BLANKET 

Figure 2.4-17. Insulation Assemblies 

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TABLE 2.4-6. INSULATION MATERIALS 

HIGH-TEMPERATURE MATERIALS 

Shields Spacers 

Aluminum Foil 

Stainless Steel Foil 

Copper Foil 

Nickel Foil 

Molybdenum Foil 

Tantalum Foil 

Borosilicate Glass Fiber Paper 

Refrasil 

Carbon Cloth 

Zirconia Paper 

LOWER TEMPERATURE MATERIALS 

Aluminized Mylar 

Aluminized Kapton ("H" Film) 

The reason for choosing a composite blanket over a single material is to minimize weight. 

The major disadvantage of a composite blanket is the difficulty of predicting analytically 

the optimum number of layers for each of the two materials. 

2.4.1.7 Areas Requiring Additional Analysis 

The remaining tasks essential to completion of the external system thermal analysis are 

summarized below: 

a. Coating evaluation 

1. Selection of materials 

2. Evaluating of tolerances of property values resulting from application 

plus environmental degradation 

b. Effect of door cycling on tracking mirror and fixed fold assembly gradients 

1811-ET-- SPECIAL HANDLING 

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c. Insulation evaluation 

1. Material evaluation 

2. Configuration optimization 

3. Design of mounting method 

4. Tolerance evaluation 

d. Failure modes 

1. Open door 

2. Failed fixed fold assembly mount heater 

e. Ascent heating analysis 

1. Evaluate impact of new shroud design 

2. Evaluate time period following shroud ejection and prior to orbit injection 

3. Evaluate impact of orbit injection maneuver 

1. Orbit Adjust Thruster Firing 

1. Analysis of several potential shroud configurations 

2. Evaluation of system recovery time 

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2.4.2 TELESCOPE ASSEMBLY 

2.4.2.1 Design Philosophy 

The Optical performance of the Acquisition Tracking System is sensitive to optical distortion 

caused by thermal gradients in the optical elements. The design philosophy developed to 

minimize temperature-induced distortions is: 

a. Thermally decouple tube from external environment by means of a highly 

reflective finish on the tube exterior 

b. Minimize contact area between optical elements and their mounts 

c. Use of low thermal conductivity materials, where structurally feasible, to 

mount the various optical elements in place. 

In addition to the above considerations, the absence of a gravitational field precludes any 

significant convection heat path resulting from the presence of a gas. The analyses performed to date indicate that an active control system is not required and that a judicious 

selection of the exterior surface finish plus isolation mounting of the individual optical 

elements constitutes an acceptable design solution. 

2.4.2.2 Temperature Constraints 

Precise temperature constraints have not yet been generated although the optimum temperature is around 70°F, the temperature at which the elements are fabricated. Having the 

elements operate at temperatures slightly higher than this results in relative movement of 

the elements with an accompanying defocusing. 

2.4.2.3 Environment Description 

The telescope passes through three widely differing environments: 

a. One section of the component extends into the living crew area where the temperatures are maintained between 66° and 80°F by an active air flow system. The 

forced convection essentially maintains the portion of the telescope in this area 

between the temperature levels previously cited. 

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b. The major portion of the telescope assembly passes through two consoles (either 

consoles 2 and 3 or 7 and 8) containing electronic dissipating components and cold 

plate heat exchanges in a stagnant gaseous environment of approximately 5 psia. 

c. In the area between the birdcage structure and the pressure shell, an annulus 

exists where living crew area air is recooled and moisture is condensed in an 

air-to-liquid heat exchanger. This annulus acts as an airflow duct, and a forced 

convective environment exists. 

2. 4. 2. 3. 1 Console Temperature Envelope 

The temperature envelope developed for the telescope is shown in Figures 2.4-18 through 

2.4-22. The concept of a cylindrical envelope on which the boundary temperatures could be 

specified was developed in order to minimize the amount of reanalysis required, prompted 

by relocations of black boxes. The most extreme temperatures expected on these boundaries 

have been specified. 

2. 4. 2. 3. 2 Atmosphere Characteristics 

Table 2.4-7 shows the nominal design properties for either a 100 percent 02 atmosphere 

at 5 psia or a 70 percent 02/30 percent He atmosphere. The 70 percent 02/30 percent 

He atmosphere is the baseline design condition. 

2. 4. 2. 4 Temperature Profiles 

2.4.2.4.1 Outer Section (Window and Objective Elements) 

The temperature gradients across the window and the objective elements are critical in 

evaluating the optical degradation of the telescope system. As previously stated, the design 

objective of minimal temperature gradients can be met basically by virtue of the fact that 

gaseous conduction is low and that radiation coupling to the environment can be minimized 

by means of a low emittance surface finish on the exterior of the tube. These three elements have been analyzed along with a 25-inch-long section of tube. An adiabatic boundary 

was assumed at the end of the tube. Each element was axially trisected to form three disks. 

Each disk was in turn divided radially into 25 areas. The window housing was represented 

2-260 SECRET SPECIAL HANDLING 
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TEMPERATURE 

ENVELOPE 

-DDFigure 2.4-18. Effective Temperature Zones for 

Telescope Temperature Envelope 

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Figure 2.4-22. 
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TABLE 2.4-7. NOMINAL DESIGN PROPERTIES 

HELIUM/02 PURE 02 

Partial Pressure 02 3.5 psia 5.0 psia 

Partial Pressure CO2 5.0 mm Hg 5.0 mm Hg 

Partial Pressure H2O 7.5 mm Hg 7.5 mm Hg 

Partial Pressure He 65.0 mm Hg -- 

Molecular Weight 24.75 31.93 

Gas Const (R) lb-ft lb-ft 62.43 lb-°R . 

48.4 lb-OR 

Density 0.0216 lb/ft3 0.0278 lb/ft3 

C

p 

0.2675 Btu/lb-°R 0.2240 Btu/lb °R 

Cv 0.1868 Btu/lb-°R 0.1614 Btu/lb °R 

k 0.0253 Btu/hr ft oF 0.0151 Btu/hr ft °F 

4 0.0487 lb-in. /ft-hr 0.0485 lb/ft-ft 

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by an additional five modes, and the retaining rings were likewise broken up into five modes 

each. The boundary condition assumed was a 4° radial gradient imposed on the outermost 

window disk. A 70°F boundary temperature was assumed. Figure 2.4-23 shows the nodal 

breakdown used in the analysis. Table 2.4-8 gives a summary of the results of this 

analysis. 

TABLE 2.4-8. SUMMARY OF OBJECTIVE LENS ELEMENT 

TEMPERATURE GRADIENTS 

ELEMENT NO. 1 MAXIMUM 

TEMPERATURE (°F) 

MINIMUM 

TEMPERATURE (°F) 

MAXIMUM 

GRADIENT (OF) 

Disc 1 71.587 69.243 2.344 

2 71.436 69.371 2.065 

3 71.296 69.492 1.804 

ELEMENT NO. 2 

Disc 1 71.274 69.579 1.695 

2 71.193 69.602 1.591 

3 71.104 69.664 1.440 

It should be emphasize that DACO is required to maintain the peripheral temperature 

gradient around the window to ± 1°F. 

2.4.2.4.2 Elbow Region 

A steady state analysis has been made on the elbow region to determine gross steady state 

temperature levels of individual elements and to indicate temperature problem areas to be 

examined in a transient analysis. The assumed environment was an 02/He environment. 

Two cases were analyzed: 

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Figure2. 4-23. Nodal Netwo

rk of OuterElements 
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Case 1. Fold mirror back face plus inner surface of mirror cover with an 

emittance = 1.0 

Case 2. Fold mirror back face plus inner surface of mirror cover with an 

emittance = 0.1 

A summary of the temperature pattern obtained is shown in Table 2.4-9. The nodal identification of the elbow region is shown in Figure 2.4-24. 

An analysis is currently in progress to determine a more detailed temperature distribution 

through the low power group. 

2.4.2.4.3 Pechan Prism 

A detaiiei steady state analysis of the Pechan prism has been completed based on a 70°F to 

140°F boundary temperature. A schematic of the nodal arrangement is shown in Figure 

2.4-25. The average heat dissipation from the Pechan drive motor is 0.24 Btu/hr. An 

adiabatic boundary was assumed at either end of the Pechan structure. A summary of the 

results is presented in Table 2.4-10. 

As is evident from Table 2.4,101 the Pechan prism is essentially an isothermal mass, with 

a nominal temperature of 98.6°F. The maximum radial gradient is on the order of 0.01°F. 

This corresponds to a negligible optical path difference (OPD). The field flattener has a 

nominal temperature of 98.5°F with a maximum radial gradient of 0.15°F. The OPD for 

the field flattener is 1/20 A for a temperature gradient of 0.15°F. This by itself will not 

significantly affect the image quality. However, if all of the elements were to produce wavefront patterns which were to add up in phase, then some noticeable degradation would occur. 

2.4.2.4.3 Zoom Assembly 

The zoom assembly is also exposed to unsymmetric boundary temperatures. In addition, 

the eyepiece contains a heat source, namely the peripheral display. Table 2.4-9 shows 

a temperature gradient of 1.58°F in the zoom structure. The temperature gradients in 

the individual elements have not yet been determined. 

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Figure2. 4-24. NodeDesig

n

ations of Elbow 
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TABLE 2.4-9. TEMPERATURE DISTRIBUTION IN ELBOW REGION 

NODE DESIGNATION CASE 1 CASE 2 

1 Tube 100.0 100.5 

2 96.2 96.7 

3 99.1 99.6 

4 99.3 99.8 

5 Elbow 98.7 99.2 

6 97.4 98.6 

7 97.4 98.6 

8 95.8 97.0 

9 Zoom Structure 99.9 98.8 

10 Elbow 98.4 96.8 

11 96.0 96.8 

12 98.1 98.5 

13 95.0 97.3 

14 Low Power Elements 97.4 98.2 

15 Fold Mirror 96.6 97.6 

16 Pechan Structure 98.5 98.6 

17 97.6 97.7 

18 Pechan 98.3 98.1 

19 Field Flattener 98.6 98.2 

21 Zoom Element 98.4 97.8 

22 Pechan Drive 97.6 97.7 

23 Zoom Elements 98.6 98.0 

24 Cover 81.4 81.3 

25 Mirror Cover 97.2 97.7 

26 4 

98.0 98.4 

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TABLE 2.4-10. TEMPERATURE DISTRIBUTION IN PECHAN PRISM 

PECHAN PRISM FIELD FLATTENER 

NODE NO, TEMPERATURE (°F) NODE NO, TEMPERATURE (°F) 

1 98.65 21 98.55 

2 98.65 22 98.47 

3 23 98.40 

4 98.64 24 98.55 

5 98.65 25 98.47 

6 26 98.41 

7 

8 

9 98.64 

10 98.65 

11 

12 

13 

14 

15 

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2.4.2.5 Conclusions 

The analysis completed to date has indicated no serious optical distortion resulting from 

temperature gradients imposed on the various optical elements. 

2.4.2.6 Additional Analysis Required 

Following is a summary of work remaining to be completed: 

a. Couple models of individual sections together to evaluate total system. 

b. Evaluate transient effects 

c. Examine gradients in all of the critical elements. 

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2.4.3 THERMAL DEVELOPMENT TEST PLANS 

2.4.3.1 General 

Optical system characteristics can be superimposed on one another. For this reason, 

the acquisition optics testing will be done on two distinct component assemblies. The 

optical system can be divided into two distinct zones: the portion of the system external 

to the pressure shell and that portion of the system internal to the pressure shell, with 

the window being common to both regions. Testing, consequently, will be done in two 

distinct phases: The development model of the tracking mirror, fixed fold assembly, and 

window will be mounted in a simulated shroud and subjected to a simulated space environment; a development model of the telescope assembly will likewise be tested, subjected 

to a simulated environment representative of the interior of the pressure shell. 

2.4.3.2 Externally Mounted Equipment 

A. Environmental Simulation 

B. Earth and Albedo and Simulation - No precise techniques for the simulation of the 

combined effects of albedo and Earth emission. Several schemes have been investigated 

for potential means to provide a simulated environment. These are listed below along 

with a brief description of some of their inherent advantages and disadvantages. 

a. High-Pressure Xenon Arc Lamp 

Advantages: Best source of those considered for solar simulation. 

1. Poor spectral match. High proportion of energy emitted in narrow wavelength band (0.8-14). 

2. This type of lamp currently not adaptable to use in a vacuum chamber. 

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b. Iodine Quartzline Lamp 

Advantages: 

1. Provides closest approximation to the combined spectrum of earth emission 

and albedo flux. 

2. Rapid response time; can be cycled to represent variations in incident 

irradiation. 

3. This type of lamp has seen considerable vacuum service. 

4. Simple controls. 

5. Lower operating pressure - 1-2 atmospheres compared to 20 atmospheres 

for Xenon lamp. 

Disadvantages: 

1. Difficult to achieve uniform flux; array has to be folded in order to give 

proper field of view between the several optical surfaces and the simulated 

earth. 

2. Relatively high cost compared to I/R heater panel. 

c. I/R Heater Panel 

Advantages: 

1. Relatively inexpensive 

2. Requires no elaborate calibration 

Disadvantages: 

1. Poor overall spectrum match 

2. Relatively slow response time 

Figure 2.4-26 shows the mean spectral distribution of albedo radiation. Plotted on the 

same figure is the spectral distribution of a 3250°K black body characteristic of the 

effective color temperature of the iodine quartzline lamp at full power. The latter 

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.2o 

,19 

.18 

.1 

.16 

wi.14 

.13 

(11.12 

.11 

14.10 

Z.o8 

J.07 

a.06 

t4.05 

a. 

Lii 

.03 

.02 

.01 

0 

Figure 2.4-26. Mean Spectral Distribution of Albedo Radiation 

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curve has been normalized such that the integrated value gives an emissive power equal 

to the combined emissive power of the earth emission plus albedo. Figure 2.4-27 gives a 

plot of the spectral distribution of the approximate earth emission. The effective absorptance for an aluminized surface typical of the tracking mirror front surface and for a 

thermodynamically black surface (typical of the remainder of the enclosed system) for 

several energy sources is tabulated below: 

As is evident from Table 2.4-11, there is a very significant difference in the effective 

absorptance of an aluminized surface from an I/R source as compared with the actual 

combined earth and albedo flux. The present test scheme calls for the use of an array of 

iodine quartzline heaters to provide a simulated earth and albedo. Solar, molecular heating and ascent effects can be acceptably simulated by affixing a series of thin heater 

blankets to the remaining surfaces of the on-orbit shroud. 

C. Pressure and Temperature Simulation - The Itek vacuum chamber facility in which 

these tests will be conducted is capable of providing a vacuum of 1 x 10-5 mmHg, a 

pressure more than adequate to represent the vacuum of outer space. Temperature 

simulation of outer space is accomplished by means of a LN2cooled inner shroud. 

D. Transient Testing - A series of tests will be run under the condition of a simulated 

environment. There are three major tests planned: 

a. Static Thermal Test - This test would be for the express purpose of measuring 

temperature patterns within the external system for a non-operational mode 

under simulated orbital heat fluxes. 

b. Optical Test - This test would be a purely optical test in which interferometric 

measurements would be made on the external optical assembly. 

c. Operational Test - This test would be used to generate data on power consumption, tracking stability of the tracking mirror. 

Various orbital parameters will be varied along with investigating viewing periods of 

varying duration. 

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Figure Deleted 

Figure 2.4-27. Approximate Spectral Distribution of Earth Radiation 

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TABLE 2.4-11. COMPARISON OF CHARACTERISTICS OF DIFFERENT 

EARTH AND ALBEDO EMISSION SOURCES 

SOURCE 

EFFECTIVE ABSORPTANCE 

ALUMINIZED SURFACE BLACK PAINT 

Actual Albedo plus Earth 0.106 0.95 

Solar Flux 0.153 0.95 

Earth Flux 0.131 0.96 

Iodine Quartzline Lamp 0.093 0.94 

I/R Source 0.049 0.96 

Xenon Arc Lamp 0.186 0.95 

E. Steady State Soak Tests - The external assembly will be maintained at various uniform 

temperature levels in the range of 0-100°F. Interferometric measurements will be 

made to determine the distortions resulting from the system operating over this wide 

range of temperatures. 

2.4.3.3 Internal Mounted Equipment 

A. Simulation of Zero-Gravity Effects - The absence of a gravitational field eliminates 

one very major heat transfer mode; viz, natural convection. Any accurate thermal testing of the telescope must correct for the presence of natural convection effects. Two 

concepts have been explored to simulate a zero-gravity field: 

a. Wrap the exterior surface of the telescope with a lightweight foam. 

b. Test the telescope assembly in a reduced-pressure atmosphere ( ,--,3 2.5 cmHg.) 

These techniques will be discussed further in order: 

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B. Foam Enclosure - Low density freon-blown polyurethane foams are available having 

thermal conductivity comparable to either the 02or 02/He atmospheres previously 

described. These foams would simulate the conductance of the actual gaseous environment. The major disadvantages of foams are two: (a) eliminates radiation coupling between elements internal to telescope and externally between the tube and the simulated 

environment; (b) will not simulate the transient characteristics of either the 02 or He/02 

atmospheres. 

C. Reduced Pressure Testing - At Grashof numbers < 2000, convection is suppressed in 

an enclosed gas space. At 2.5 cmHg pressure, thermal conductivity of the enclosing 

gaseous atmosphere is reduced by 5-7% from that at atmospheric pressure. For a test 

condition at atmospheric pressure, the temperature differential measured on the telescope tube could be as much as 25% lower than that compared with a zero-gravity 02 /He 

atmosphere and upwards of 47% lower than that predicted for a zero-gravity 02 

atmosphere. This very obviously establishes the need for this kind of testing. 

D. Individual Element Testing - Individual elements will be tested while subject to a 

simulated environment. These elements will be selected on the basis of analysis having 

shown these elements to be potential problem areas. 

E. Telescope Assembly Vacuum Testing - A model of the telescope assembly will be 

tested in a simulated environment in a vacuum chamber maintained at reduced pressure 

to determine temperature gradients in the various optical elements. A series of heaters 

will be mounted on a framework enclosing the telescope to program the expected temperature behavior of the telescope environment. 

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2.5 ATS STRUCTURES ANALYSIS 

The following section of this report shows the structural analysis performed by GE on the 

ATS portion of the acquisition subsystem. The purpose of the analysis was to verify the 

structural integrity of the subcontractor's design. The analysis indicates that the subcontractor's design is adequate from the structural standpoint. It also indicates that the subcontractor's design is not optimized thus making a saving in weight probable. It should be stated 

again that all the structural analysis shown in the report was performed by GE using subcontractor drawings as a basis. The report also determines the possible system misalignments due to factors such as lg load relief and mechanical joint slippage. A discussion is 

also included on the area of dynamic analysis. The complete system dynamic analysis 

planned by GE with subcontractor assistance is discussed in detail. The status of the subcontractor's weight is also discussed in detail. 

The ATS structures analysis section is broken down into the following subsections: 

2.5.1 Structural Analysis of Telescope Under Powered Flight Loads 

2.5.2 Structural Analysis of Tracking Pedestal Under Powered Flight Loads 

2.5.3 Structural Analysis of the Protective Shroud Under Powered Flight 

Loads and On-Orbit Operating Loads 

2.5.4 Deflection Analysis of Telescope and Tracking Pedestal Due to 1G Load 

Relief 

2.5.5 Determination of Optical Path Misalignments Due to Mechanical Joint 

Slippage 

2.5.6 Determination of Misalignments Due to Shell Pressurization and 

"Hot-Dogging" 

2.5.7 Dynamic Analysis and Loads Determination 

2.5.8 ATS Weight Analysis 

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2.5.1 STRUCTURAL ANALYSIS OF THE TELESCOPE UNDER POWERED FLIGHT LOADS 

Introduction 

The structural analysis of the telescope which follows is intended to verify the basic structural integrity of the design and to determine the validity of the subcontractor's weight 

estimate by determining the degree of optimization of the telescope structure. Also, 

recommendations will be made concerning areas and methods of potential weight reductions. 

Conclusions 

The analysis indicates that the subcontractor's structure is able to withstand the loads due to 

powered flight accelerations with an adequate margin of safety. It also shows that the 

structure is over-designed, thus indicating the possibility of a reduction in structural weight. 

Description of Structure 

The telescope structure is a tubular structure made of aluminum. In the analysis, beryllium 

was considered for certain portions of the structure. The purpose of this was to demonstrate 

the feasibility of using beryllium in the structure. The walls of the tubular portion of the 

telescope (between shear fitting and penetration fitting) are currently 0. 040-inch thick. The 

elbow which has to carry the weight of the Pechan prism and the zoom/eye piece by a combined bending and torsion is made of aluminum and has a heavier gage wall. This is necessitated by higher loading. The structural housing for the Pechan prism and the zoom/eye 

piece are also of a heavier gage wall. This is necessitated by high local loads primarily 

due to drives. The structure is mounted to the Lab Module pressure shell at two places. 

The objective end of the telescope is mounted to the DACO shell penetration fitting. This 

joint was originally planned to be designed so that there would be no capability of transmitting a torque. This concept is reflected in the analysis. A torque transmitting joint is 

currently being studied. Such a joint would allow a better seal and may be easier to fabricate. The other joint is the shear-tie between the telescope shell and the vehicle shell. This 

joint is located at the field joint between the elbow and the constant diameter portion of the 

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telescope. The shear-tie is so designed that it will carry loads due to telescope twisting 

and also vehicle longitudinal (X direction) loads. The shear-tie is designed to transmit a 

minimum amount of load in the telescope logitudinal direction. This was also reflected in 

the analysis. 

2-286 

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2.5.2 STRUCTURAL ANALYSIS OF TRACKING A PEDESTAL UNDER POWERED FLIGHT 

LOADS 

Introduction 

The structural analysis of the tracking pedestal is intended to verify the structural integrity 

of the subcontractor's design and to determine the validity of the subcontractor's weight 

estimate. This is done by determining how optimized the structure is and making recommendations as to areas of potential weight savings. 

Conclusions 

The analysis shows that the structure is capable of carrying the loads imposed by powered 

flight accelerations. The high margins of safety calculated indicate that the structure has 

not been optimized. This means weight reductions may be feasible. 

Description of Structure 

The tracking pedestal structure is a beryllium structure which will be machined from hot 

pressed beryllium block. The structure is made of four basic pieces; the pedestal, the 

motor-encoder housing, the yoke, and the mirror bezel. This analysis does not include the 

mirror bezel; the other portions are included. The pedestal is a short cylinder with a threepoint gusseted mounting plate. The connection to the vehicle shell is accomplished by using 

three DACO provided mounting studs. The connection between the motor housing and the 

pedestal is made by using mechanical fasteners. The motor-encoder housing houses the 

roll motor and encoder. The pitch gyro electronics will be mounted externally on the 

housing but this is not reflected in the analysis presented since at the time this analysis 

was performed the electronics was mounted to the mirror bezel. The roll shaft runs the full 

length of the housing and is supported by double row bearings at the forward and aft end. 

The roll shaft is attached to the yoke by means of mechanical fasteners. The yoke is a 

fabricated box structure which supports the mirror and bezel, the roll gyro and electronics, 

the pitch gyro, and the pitch motor and encoder. The pitch motor and encoder are mounted 

2-307 

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on opposite sides of the yoke and are attached to the ends of the yoke arms by means of 

mechanical fasteners. The roll gyro and electronics are mounted to the yoke close to the 

roll shaft. The pitch gyro is mounted to the mirror bezel. The bezel is a beryllium pan 

into which the mirror is potted. The bezel has fittings to which shafts are attached which 

form the pitch shaft of the mirror. In the analysis, the mirror and bezel are considered to 

be a lumped mass with a rigid member forming the pitch shaft. 

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2.5.3 STRUCTURAL ANALYSIS OF THE PROTECTIVE SHROUD UNDER POWERED 

FLIGHT LOADS AND ON-ORBIT OPERATING LOADS 

Introduction 

The structural analysis performed on the protective shroud was intended to demonstrate 

to the subcontractor that a much lighter weight shroud is attainable with the use of lighter 

materials and some design changes to make the design more efficient. The shroud design 

proposed by the subcontractor was not analyzed due to the fact that the design weight of 

23.5 pounds was unacceptable. An effort was made to derive a minimum weight preliminary design for transmission to the subcontractor. The results of this analysis show that 

it may be possible to design a shroud weighing approximately 14 pounds. The weight could 

possibly increase slightly due to thermal problems but hopefully the design could be further 

optimized toward the 14-pound weight. The shroud was checked for powered flight loads 

as well as a worst case on-orbit load due to attempted operation with a motor seizure. 

Description of Structure 

The shroud design analyzed is a basic beryllium framework with beaded fiberglass skins to 

give additional rigidity to the structure. The end plates are shown as fiberglass but could 

be made of beryllium if thermal problems dictate this. The envelope is the same as the 

subcontractor and the deflections incurred are no worse than the subcontractor's design, 

the reason for this being the sizeable increase in stiffness due to the beryllium framework 

and beaded skin versus the subcontractor's aluminum structure (e. g., E for beryllium = 

42 x 10

6 psi versus E for aluminum = 10 x 10

6 psi). 

Design Loads 

Shell acceleration from the Martin Second Load Cycle will be used. The conditions considered will be for the maximum axial load condition and for the maximum lateral load 

condition. 

2-319 

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Maximum axial load condition: 

Axial Ws: +5.6 

-2.7 

Lateral g's: Negligible 

Maximum lateral load condition: 

Stage H Ignition 

Axial g's: +2.8 

+0.59 

Yaw g's: +0.41 

-0.24 

Pitch g's: +0.80 

-0.44 

Launch 

Note: All accelerations are limit and ultimate = 1.4 x limit. 

The basic configuration of the proposed shroud is shown in Figure 2.5-1. For the purpose 

of analysis the weight of the shroud structure will be assumed to be 10 pounds. 

HINGE POINT 

BERYLLIUM FRAMEWORK 

HINGE POINT 

BEA DE D 

FIBERGLASS 

SKIN FIBERGLASS 

END PANEL 

LAUNCH LOCK 

BERYLLIUM RING SECTIONS 

Figure 2.5-1. Basic Configuration of Proposed Shroud 

2-320 

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TABLE 2. 5-1. MATERIAL COMPARISONS 

O 

s4 

C C 

a) 

a 

C) 

6 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

--tEeRET- SPECIAL HANDLING 

Analyze the beryllium framework: 

15.5" 

48" 1,• Id16" 

Consider the following loading case: 

The loading applies to an axial acceleration. Consider the load to be taken out at 

each of the three'corners of the rectangular framework. 

To determine the load distribution consider the curved beams to be continuous straight 

beams on three supports with a uniform load. 

2-322 

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- q 4. + 4 4 44 4 4 4 4 • 

R1 1.4 L R2 10 R1 

3 

R1 16 = qL 

10 R2= 16 = qL 

We have the following: .•. R2 = 5/8 (5 pounds) = 3.125 pounds 

R

1 = 0.9375 pound 

3.125 LB 

0.9375 LB 

0.9375 LB 

3.125 LB 

0.9375 LB 

Consider the following: 

0.9375 LB 

1.5625 LB MME= 1.5625 '15.52 + 8.02 

= 1.5625 304 I 

15.5

2

+8.02 

= 1.5625 (17.4) 

= 27.2 in. -lbs 

////////// 

2-323 

SPECIAL HANDLING 
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Determine Z

req 

for an angle made of AMS 7901 beryllium 

F

ty 

@ 400°F = 22950 psi 

22950 - 27.2 (5.6 g's) 152.5 

Z

REQ Z

REQ 

152.5 Z

REQ 229.5 - x 10-2 = 6.65 x 10-3 

Try using a 7/8 inch by 7/8 inch by 0.040-inch thick angle 

Section properties: A = 0.066 in. 2 

I = I = 0.0051 in. 4 

xx yy 

Z = I/C = 0.051/(0.875 - 0.249) 

Z = 0.0051/0.626 = 0.0081 in. 3 

cr = M/Z = 152.5/0.0081 = 18850 psi 

M. S. = 22950-1 = +0.22 (LIMIT) 18850 

F 

Limit will be critical condition F - 1.43 > 1.4 

tY 

Determine deflection: 

6 PL3 1.5625 (5.6) (17.4)3 

3E1 (3) (42 x 106

)(0.0051) 

(1.5625)(5.6)(1.74)2

(17.4) 

- 0.072 inch 126 (51) 

2-324 

--S'EP,it['fi' SPECIAL HANDLING 
M. S. - 26400 1 = +LARGE (ULT) 473 

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Check Column Action of Longitudinal Member 

Consider Case 7 (page 305) in Roark: 

77.2EI L = Where L = 48 inches 

PCRIT (1.122)

2 

I = 0.0051 

(9.85)(42 x 106

P )(0.0051) - (0.875) (1.76) ( 9.85) 

CRIT

(48) = 2900 

6.25 = 15.2 lb/in. where FACTUAL = 48 - 0.13 lb/in. x 5.6 = 0.73 lb/in. 

PCRIT 

Check lower longitudinal angles for the following condition: 

ArTh m 

R2 

11"1 

M = 15.5 (3.125) 

M = 48.5 in. -lb 

48" Z = 2(0.0081) = 0.0162 

M 48.5 

.•. a = = = 3000 psi b Z 0.0162 

Check shear: 

Q = E Ay = 0.626(0.04)(0.313) = 0.00785 

VQ (1.5625)(5.6)(1.54) 

T = T = 338 psi (LIMIT) 

bb 

(0.040) 

T

ULT 

= 1.4 x 338 = 473 psi 

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Check crippling of angle: 

2 

CHIT 

t 

Cr = K E 

2 b for Be 1- v

2 1.0 

1 -v 

2 

. . a CRIT = KE 13 (—) = 3.29(42 x 106

) 0.626 

2 

cr CHIT = (3.29)(42)(0.41) x 104 = 56.6 x 104 psi 

2 

T

E t = K CHIT 2 > cr cRITfor K= 4.40 

1 - v 

.•. Crippling will not be critical. 

Applying the maximum axial load factor of 5.6: 

a

b = 5.6 x 3000 = 16800 psi 

. . M.S. - 22950 1 - +0,36 (LIMIT) 16800 

There will also be a stress component due to the axial load which will be negligible and 

also a beam-column effect which will not be considered at this time. 

Consider the Case of Launch Acceleration Where We Have a Lateral Load 

Determine magnitude of lateral accelerations: 

2-326 

-SUM- SPECIAL HANDLING 
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YAW-0.41 g 

Acceleration acting at shroud 

G

R = 0.41 cos 26° + 0.8 cos 64° 

G

R = 0.41(0.898) + 0.8(0.438) 

G

R = 0.368 + 0.35 = 0.718 g 

G

T = 0.41 sin 26° - 0.8 sin 64° 

G

T = 0.41(0.438) - 0.8(0. 898) 

G

T = 0.1795 - 0.71 g 

G

T = 0.539 g 

We will get the following loading condition: 

(Launch Accelerations) (The accelerations are applied to the 

loads shown. ) 

2-327 

2. 625 LB 

2.625 LB 2.625 LB 

-S-NRET- SPECIAL HANDLING 
NRO APPROVED FOR 

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Determine Bending Stresses in the Longitudinal Beam 

Z = 0.0081 in. 3 

A = 0.066 in. 2 

qL2 _ (0.15)(48)2 (0.112)(48)2 

M = - 43.2 + 32.3 = 75.5 in. /lb 

ql 8 8 8 

M P 75.5 = 9330 + 133 = 9463 psi 

a Z -1-A = 0.0081 + 0

8

.

.

0

7

6

5

6 

22950 -1 = +1.42 (LIMIT) M.= 9463 

Shear will not be critical. 

Check Bending in Members Forming End Triangle 

Loading: 

:3.75 LB :;.75 LB 

Section properties of angles 

8.75 LB A = 0.066 in. 2 

J = I = 0.0051 in.4 

yy 

Z = 0.0081 in. 

16"--8.1 

M due to 8.75-pound load = 4.375 1115.52+ 8.02 = 4.375(17.4) 

M = 76.1 in. /lb 

M due to 3.75-pound load = 3.75(8) = 30 in. /lb 

M due to 2.69-pound load = 41.7 in. /lb 

2-328 

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.•. MTOTAL = 147.8 in. /lb 

M 1780 

ab = = = 9150 psi Z 0

4

.81(2) 

Determine Tensile Loads and Stresses 

Consider components to be additive (they may not be so): 

15.5 

P 17.4 17

8 

4 = (3.75) + (2.69) = 3.34 + 1.24 = 4.58 pounds . 

• . at 

= P/A = 4.58/0.066(2) = 35 psi 

= 9200 psi M. S. = 229501 - +1.40 (LIMIT) 9200 

Determine Weight of Framework 

L = 4(17.4) + 2(16) + 3(48) = 70 + 32 + 144 = 246 inches 

.•. V = LA = 246(0.066) = 16.2 in.3 

W = p V = 0.067(16.2) = 1.10 pounds 

Determine effect of a 100 in. /lb torque applied at one end of the shroud and the other end 

is locked (no launch locks). 

LOCKED END 

100" /LB 

16" 

48" 

-IfeRH--SPECIAL HANDLING 

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Compare Torsional Rigidity of Members to Bending Rigidity 

Consider the 7/8 inch by 7/8 inch by 0.040-inch thick angles: 

Developed length - 1.65 

K = $bt3 = 0.333(1.65)(0.04)3 = 8.8 x 10-4 

Consider the following: 

where: XA = - 6 ao (Ref: Peery page 457) 

6 aa 

6 ao = Deflection due to applied loads 

6 aa = Deflection due to unit redundant 

6 ao = KG TL 

8 

(5)(48) x 10-2 = 0.273 inch (.8) 

ML TL as = 2 (--Ei- + 2 KG ( 

(2)(16) -6 (2)(48) x 10-2 6 aa - x 10 + (42)(0.0051) (88.)(20) 

6 aa = 32 x 10-2 

(42)( 

8 

51) 8 x 10-2 = (0.015 + 0.545) x 10 4 -2 

8 

2-330 SECRET SPECIAL HANDLING 
0.2" 

5.0" 

0.016" THICK (t) 

ko— 2" —04 1.0" ke-- 2" ___•4 

1-0.2" 

4- R 0.725 

a - 43.0 

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6 as = 0.56 x 10-2 

O. 273 XA - - 0.56 x 102 = 48.75 49 in. /lb 

.•. The twisting in 1-2 is 51 in. /lb. 

Determine the Torsional Capability of the Shell (Shroud Skin) 

Assume the skin to be made of 181-Volana Glass Fabrk (Ref: page 79 to 81, MILHDBK-17): 

Mechanical Properties: (@ 400°F/one-half hour soak) 

Flexural strength - 48000 psi 

Tensile strength - 35000 psi 

Modulus of elasticity - 2.9 x 106 psi 

Assume shear strength to be 60 percent of tensile strength: 

.*. F = 21000 psi 

U 

Also assume u (Poisson's ratio) = 0.12 

- 

2.9 x 106 

G 2(1+0) 2.24 - 1.3 x106 

Consider the following skin cross-section: 

==6 3=1=:= 117 

Determine section properties: 

--S-EGRET- SPECIAL HANDLING 2-331 
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Determine Io and y for bead: 

2 sin a ) I = R3

t a+ sina cos a 

I = (0.725)3 (0.016) { 43.6 + sin 43.6° cos 43.6° 2 (sin 43.6) 

57.3 43.6/57.3 

I = 0.38(0.016) [ 0.764 + 0.5 - 1.25] 

I = 0.006(0.014) = 84 x 10-6 in. 

y s

1 = R 

a 

- a-) = 0.725 (1 0 764 - 0.725 (1 - 0.904) 

. 

y

1 = 0.725(0.096) = 0.069 in. 

A = 2 a Rt 87.2(0.725)(0.016) _ 

- 0.01765 in. 2 57.3 

Determine total section properties: 

A y Ay I

o 

1 0.032 0.008 2.561-4 0.68251-6 

2 0.032 0.008 2.561-4 0.68251-6 

3 0.018 0.131 23.4 1-4 84.0 1-6 

28.521-4 1-6 0.082 85.365 

d 

0.0264 

0.0264 

0.097 

Ad2 

22.41-6 

22.41-6 

169.01-6 

213.81-6 

—28.52 -2 

y - 8.2 x 10 = 3.44 x 10-2 

.•. INA = 299 x 10-6in. 4 

2-332 

-SECRET SPECIAL HANDLING 
T 258 

e = b = (12) = 54 inches 57.3 i3bt3L G 

(0.333)(54)(0.0893

)(20 x 106) 

144 

0 = (54)(20)(705) 

12" R 

e - 

(1.0)(48) 

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RELEASE 1 JULY 2015 -SIM SPECIAL HANDLING 

Determine equivalent flat section: 

I = 299 x 106= 5t3 

t

3 = 716 x 10-6 

t = 8.9 x 10-2 = 0.089 inch equiv 

ANGLE OF TWIST: 

12 

0 = 1.89 x 10-4 rad/lb 

Angle of twist for Be framework: 

TL (0.51)(2.4) 

e = x 10-2 = 0.139 x 10-2 rad/lb KG (8.8) 

TL TL 

bt3G KG 

1.89 x 10-4 Ts = 0.139 x 10-2 T

F 

T

s 

= 0.0735 x 102 T

F = 7.35 TF 

T

S + TF = 100 .•. 8.35 TF = 100 

T

F = 12 in. /lb 

ItreRff SPECIAL, HANDLING 2-333 
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RELEASE 1 JULY 2015 

Determine torsional shear stress in the framework: 

6.1(3) _ 18.3 x 104 

cx bt2 1.65(0.04)2 26.4 

T = 6950 psi (LIMIT) 

T = 1.4 x 6950 = 9725 psi (ULT) 

:. M.S. = 264001 - +1.71 (ULT) 9725 

Determine torsional shear stress in the shell: 

- 

264 - 6.2 x 102 

bt2 42.6 x 10-2 

T = 620 psi (LIMIT) 

T = 1.4 x 620 = 870 psi (ULT) 

100 0 . M. S. = 2870 1 - +LARGE (ULT) 

Check shear crippling: 

Shell: 

)2 

TCRIT = 0.1E -t-- + 5E t (-- 

(0.016 1+6 ( 256 ) = (0.1)(2.9) + (5)(2.9) TCRIT 12.0 256 

T + (2.9)(51.2) = 387 + 148 = 535 psi TCRIT 0.12 

= 

2-334 

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A crippling problem exists. Consider increasing the thickness to 0.020 inch. 

_ (2.9)(20) + (2.9)(80) = 484 + 232 = 716 psi CRIT 0.12 

Due to the beef-up the shell will take more load thus for brevity consider the shell to carry 

the full 100 lb/in. 

(1 0 0 

.•. T = 620 8

--- = 705 psi (does not reflect increase in tequiv) 

8 

. . Increase wall to 0.020 inch 

Determine weight of shell: 

BEADED AREA Unbeaded area = 39 x 43 = 1675 in. 2 

 V = 1675(0.02) = 33.5 in. 3 

WT = 0.068(33.5) 

WT = 2.280 pounds 

54" 

48" 

Weight of Beads: 

L = 54(9) + 48(11) = 486 + 528 = 1014 in. 

V = AL = 0.018(1014) = 18.25 in. 

WT = 0.068(18.25) = 1.25 pounds 

. . Weight of skin = 3.53 pounds 

1-H4ff- SPECIAL HANDLING 

2-335 
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RELEASE 1 JULY 2015 —MO- SPECIAL HANDLING 

Determine weight of rings. (Consider them to be the same as stiffeners.): 

A = 0.066 in. 2 

L - 54(4) = 216 in. (assume 4 rings) 

.. V = 14.25 in. 3 

WT = 0.067(14.25) = 0.96 pound ti 1 . 0 pound 

Weight of ends (assume double end): 

A = /TR2 = 77(144) = 452 in. 2 x 4 = 1808 in. 

V = 1808(0.016) = 29.0 in. 

WT = 0.068(29) 2.0 pounds 

.•. Total weight = 1.1 + 3.5 + 1.0 + 2.0 = 7.6 pounds 

Assume additional structural weight to bring total to 9.0 pounds. 

Consider drive to weigh 3.0 pounds and insulation and coatings to weigh 2.0 pounds. 

Shroud weight is 14.0 pounds 

2-336 

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2.5.4 DEFLECTION ANALYSIS OF TELESCOPE AND TRACKING PEDESTAL DUE TO 

1G LOAD RELIEF 

Introduction 

A deflection analysis of the telescope and tracking pedestal was performed to determine 

the effects of optical path alignment due to lg load relief. The results of this analysis are 

shown on the following pages. 

SECRET SPECIAL HANDLING 2-337 
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2-338 
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CK. 

DATE 

• 

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CK. 

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DATE 

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2.5.5 DETERMINATION OF OPTICAL PATH MISALIGNMENTS DUE TO MECHANICAL 

JOINT SLIPPAGE 

Introduction 

The amount of mechanical joint slippage has been determined and how this joint slippage 

will affect the optical path. On the tracking mirror a misalignment between the center of 

the mirror and the pressure bulkhead has been determined. For the telescope the misalignment between the eye-piece and objective lens has been determined along with the 

optical path misalignment between the eye piece and tracking mirror. It should be noted 

that the amount of mechanical joint slippage is a maximum worst case value. It assumes 

worst tolerances, single fastener joints, and total movement in an adverse direction. It 

is strictly a hypothetical number and should not be used for design purposes. 

4EM- SPECIAL HANDLING 

2-345 
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RELEAS 

AGE / OA: 

MODEL 

REPORT 

row.. 1•*13 III (•-SW1 

2-346 
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REPORT 

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DATE 

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DATE 

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REV. 

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2.5.6 DETERMINATION OF MISALIGNMENTS DUE TO SHELL PRESSURIZATION AND 

"HOT-DOGGING" 

Introduction 

The effects of shell pressurization on system misalignments has been investigated. The 

results of this analysis are presented in this section. 

The value of angular misalignment between the tracking mirror and the external fixed-fold 

mirror was determined using the computer program used to determine the shell stiffness 

matrix discussed in Section 2.5.7. Considerable additional analysis is required in this area 

to support error analysis. 

Preliminary analysis has also been performed to determine the effect of vehicle thermal 

bending or "hot-dogging" on system alignment. Results-to-date indicate that "hot-dogging" 

may not have a significant effect. 

A possible reason for this is the fact that the micrometeoroid shield acts as a buffer. 

Therefore, any thermal distortions are felt by the shield only and very negligible distortions are seen by the pressure shell itself. 

2-351 

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AR. 

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2-352 
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—SEDRET— SPECI HANDLINGb AGE 0 URINAL', ELECTRIC MODEL 

ECIE04 4i&APF sir_ 

FIT 441ELL liNE,110.. 

unkt. izzi,lrev _ 

.1_ 

POPthe 1 SIMI II (111-11111) 2-353/2-354 
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2.5.7 DYNAMIC ANALYSIS AND LOADS DETERMINATION 

Introduction 

Structural dynamic analysis of the ATS portion of the acquisition subsystem will be performed by GE. The analysis will utilize the models of system components provided by the 

subcontractor along with a GE generated model of the Lab Module pressure shell. This 

analysis will be used to determine overall systems response due to various disturbances 

along with the determination of loads due to these disturbances. 

Discussion 

The structural dynamic analysis performed by the subcontractor has been reviewed in 

detail by GE. It is felt that, considering the stage of the design effort, the subcontractor's 

analysis meets his requirements. However, considerably more dynamic analysis is required in order to verify total system performance. The subcontractor does not have the 

capability or the required inputs to perform such an extensive analysis. Therefore, GE 

will perform the analysis with certain information obtained from the subcontractor. This 

information will be in the form of detailed models of the telescope, fixed-fold mirror, and 

tracking pedestal. These models will have node points at each of the major pieces of glass 

or groups of glass so that the effects of disturbances on system performance can be monitored and evaluated. Stiffness matrices for each of these components will be generated by 

the subcontractor and then incorporated into an overall system matrix, which will include 

the coupling effects of the vehicle shell. The vehicle shell will be represented by a 36 x 36 

stiffness matrix which represents six degrees of freedom at the six system mounting 

points. This matrix was generated by GE because the associate contractor could not 

generate the needed information prior to June 1968. The information needed was the 36 x 

36 matrix for this analysis along with the local shell flexibilities which were needed by the 

subcontractor for design purposes. The information was initially generated in the form of 

a 36 x 36 influence coefficient matrix and then inverted to obtain the stiffness matrix. The 

influence coefficient matrix was obtained by modeling the DACO Lab Module pressure shell. 

2-355 

SPECIAL HANDLING 
10.72 10.72 

±1.8 10.72 

110.0 

+3.0 

-1.0 

+3.0 

-1.0 

Max. X load 

Max. Y load 

Max. Z load 0.72 ±1.8 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

The model represented an 80-inch long portion of the Lab Module shell. The pressure 

bulkhead at Station 517 was represented by a series of radial springs which represented 

the bulkhead flexibility. The forward end of the shell was considered to be simply supported. The matrix was generated using structural analysis programs on the GE 635 computer. The disturbance mentioned previously will come from many sources. Among the 

more important ones will be: tracking and slewing of both the acquisition and main optics 

tracking mirrors; ACTS firings; opening and closing of the thermal cover and protective 

shroud, and activation of the acquisition system controls in order to change power, etc. 

The system model will also be used to evaluate the response to pyrotechnic shock and 

random vibration so that realistic design loads can be used. At the onset of the design of 

this system the subcontractor had used a steady-state load of 35 g's in each direction to 

design the structure. This was obtained from DR1100 which stated that the system was to 

see a 35g shock pulse for 2.6 milliseconds. This was a very conservative, if not unrealistic approach, therefore it was felt that it was necessary to clarify the loading so that a 

realistic structure could be obtained. To accomplish this, powered flight accelerations 

were obtained using load factors from the Martin Second Cycle Loads Report modified to 

show trends obtained from the 3A load cycle. The "bird-cage" was used as a basis for 

determining these loads. The powered flight load factors are as follows: 

Limit loads (ULT = 1.4 x LIMIT) 

The X, Y, and Z directions are in the vehicle coordinate system. 

2-356 

--S-EGHT- SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

The shock loading was also clarified by giving the subcontractor the shock curve (Figure 

2.5-2). This curve was obtained from SAFSL 10003 and reflects a reduction of peak shock 

of 3.5. The curve was also modified by extending the curve past the 100 cps cut-off shown 

in SAFSL 10003. 

The Lab Module shell stiffness and flexibility matrices are also included after the shock 

curve. 

1000 

100 

10 

1 

10 100 1000 10,000 

FREQUENCY 

Figure 2.5-2. Shock Curve 

2-357 

-IfeRR SPECIAL HANDLING 

INPUT - G's 
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1 

5.07640002E-05 

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-9.19330000E-06 

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1.44710000E-47 

1.21670000E-07 

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9.43690001E-10 

1.17190000E-07 

8.14160002E-08 

3.94099999E-07 

1.34840000E-06 

9.21880002E-07 

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2.27060000E•08 

4.15620001E-08 

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1.171,90000E-07 

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4.788200.01E-08 

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4.40380001E-06 

1.05920000E-07 

7:30709999E708 

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7.85850003E-10 

5.343600018..09 

2.92100000E709 

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71.61360001811 

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73.84669998E48 

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713,81350002E711 

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73,650300028-08 

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1.46920000E-09 

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3.78979998E-08 

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73.61320001E700 

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2-371 

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8.71470000E-07 

4.22009998E+06 

5.57729997E-07 

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1.49530001E+07 :2,76350001E+08 

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5.03380000E+08 1,6090001E+07. 

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2-372 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

9 

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100010001F-08 

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RELEASE 1 JULY 2015 "----SMET- SPECIAL HANDLING 

17 

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5.5/729997E-07 -9.45800005E•09 

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2-375 

SECRET SPECIAL HANDLING 
21 

22 

23 

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1.44710000E-07 

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2-376 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

-1.91780000E-01 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

-MU- SPECIAL HANDLING 

P5 

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2-377 

SECRET SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

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2-378 

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3,4050000OF-09 

34 

-4.2PG.900'2TF-7 

i 7- 

-1.3823999F-41E 

-2.96010000E-09 

1.49200000E-06 

-1.7F47000IF-06 

2.73)72'0200E-07 

.V I` 

-3.44550002F-0P 

F.54670001E-09 

7.42969997F-OF 

-2.01010000E-07 

-9.0!7P999FF-09 

-'.71740001E-09 

2.4F35999F-1e 

-11.547440 E-9"9 

F.2949999/7E-08 

-5.46590000E-09 

-3.53109999E-0P 

6.36079997F-06 

3.1251000IF-05 

-4.9"3PF9F-0!= 

-5.73969999F-I1 

3.0193000VF-P9 

1.3546004TF-OP 

-6.626F009TF-P9 

-1.6F090070E-VP 

9.S972999FF-09 

4.64270000E-0F 

2.0503200IE-07 

4.ZE76999FF-OF 

3.99E80000E-08 

8.12299998E-10 

-2.5613000IF-07 

-3.12349999F-09 

-4.06900000E-10 

5.243801002E-0S 

5.34589999E-00 

F.08330003E-09 

2.77070000E-09 

-5.66010000E-09 

-1.11900000E-09 

1.16000000E-07 

2.11060000E-07 

1.84450000E-09 

3.53700000E-06 

4.30969999E.606 

3.40500000E49 

-1.F0630000E-09 

-1..44280000E-07 

4.17175(7001E-10 

6.75540002E-09 

-1.442F0000E-07 

'-F.87,109997E-10 

-6.9F219997E-08 

3.45590001E-06 

_P.05030000t-07 

3.99S80000E-4'F 

-6.03419997E-09 

-P.404E0000E-OF 

3.95400002F-09 

1.79549.999E-70F 

-1.SPRO0000E-07 

2.63350000E-07 

6.82691001F-06 

1.31s9301100E-P3 

3E 

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2.5.8 ATS WEIGHT ANALYSIS (See Section 4.4 for total subsystem weight status) 

2.5.8.1 The weight of the system has been under very close scrutiny because the system 

weight is currently well above the weight acceptable. Various steps are currently being 

taken to determine what a realistic system weight will be so that a final weight figure can 

be negotiated with the customer. The structural analysis in Sections 2.5.1 through 2.5.3 

was performed with the identification of overweight areas in mind. The results of the 

analysis are being used to show the subcontractor where weight may be reduced and if 

necessary, how to reduce weight. Much of this increase can be attributed to misinterpretation of environmental information such as loads and temperatures. A concerted effort 

is being made to clarify this information so that a representative weight can be obtained. 

Currently, the most troublesome weight areas are the protective shroud, the gimbal 

inertias and sizeable weight increases in the telescope. 

A. Protective Shroud 

The latest subcontractor weight estimate of the protective shroud is 23.5 pounds. The 

shroud weight has doubled over an earlier estimate of 11 pounds. The reasons given by 

the subcontractor for the weight increase is that the thermal environment is such (350°F) 

that aluminum is the only lightweight material that can be used. This has caused the 

greatest increase since beryllium or magnesium were being considered before. It is felt 

that this is a poor solution, since at the temperatures quoted, beryllium has more than 

ample strength and the ductility of beryllium at this temperature is quite favorable along 

with the modulus of elasticity. Another reason given is that there is a requirement for 

redundant drives which necessitates an increase in drive weight. This requirement will 

cause some increases in weight but with a more efficient drive design the weight increase 

can be kept at a minimum. The other factor causing weight increases is the fact that a 

better definition of the design has been obtained. This is an unavoidable increase but should 

be able to be kept to a minimum. GE estimates that it may be possible to design a shroud 

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that will weigh a minimum of 14 pounds. This estimated weight considers the redundant 

drives and a better design definition. One other potential problem area on the shroud is 

the effects of the plume from the ACTS. The plume will impinge on the shroud and preliminary thermal analysis indicates that temperatures of 1000°F are possible. One solution would be to install a shield over a portion of the front end of the shroud. This shield 

should weigh less than a pound and would reduce the temperatures to more acceptable 

levels. 

B. Gimbal Inertias 

Currently the inertias of the rolling and pitching masses are below specification values. 

Inertias below specification levels will cause a reduction in smoothness which would cause 

jitter when the mirror is moved. The logical means of increasing inertia is to add mass 

at some distance from the reference axis. The subcontractor has attempted to do this but 

has caused sizeable increases in system weight which are totally unacceptable. Therefore, 

the subcontractor has been directed to perform studies on repositioning gyros, etc, to 

reduce the amount of weight needed for bringing the inertias up to specification value. The 

studies performed have been reviewed with the subcontractor but they have not progressed 

to a definite conclusion. Work is also being performed to determine the lowest inertias 

that would be tolerable. Coupled with the inertia problem is the fact that the roll and 

pitch masses are not balanced. To alleviate this problem the subcontractor, under GE 

direction, will make provision for adding balancing weights during ground testing. The 

weights will then be removed and the system locked for powered flight. At the present 

time it appears that some additional weight will be required to raise the inertias to an 

acceptable value but the exact amount is unknown. A five-pound allowance for such a 

weight had been carried previously so the subcontractor is requesting an additional 28 

pounds for inertia make-up. This number is part of the current weight estimate. 

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C. Increased Telescope Weight 

The weight of the telescope has increased by seven pounds per side or 14 pounds for the 

total system. The reasons given by the subcontractor for this increase has been increased 

glass weight (a sufficient mounting edge had not been considered) and also by the addition 

of the Q-mirror and its related hardware along with better design definition. These increases cannot be discounted but it appears that it may be possible to reduce the weight 

of the telescope to a point where there will be a zero net increase or decrease. There are 

certain areas that may offer promising weight reductions. The telescope wall thickness 

is now 0.040 inch. It appears possible that this thickness can be reduced to 0.030 inch to 

0.032 inch as indicated by structural analysis. The subcontractor is investigating this 

aspect. The reduced wall thickness can also be applied to the Pechan prism housing and 

the zoom/eye piece housing. Another area being considered is the substitution of beryllium 

for aluminum in the telescope structure. The subcontractor has been very hesitant to do 

this since he claims that the coefficient of thermal expansion of the aluminum compensates 

for changes in the index in refraction of the glass due to temperature gradients. However, 

the subcontractor has never backed up this statement analytically and has not investigated 

other materials. He was directed to perform these studies and to formally report these 

results to GE so that a material change can be initiated, if possible. 

2.5.8.2 The current ATS weight estimate is 375 pounds. This weight includes the increased telescope weight, additional weight for inertia make-up, and the GE estimate for 

the shroud. It also includes weight for a sun-sensor which was not carried previously and 

a two-pound addition to the tracking pedestal harness due to DACO's refusal to relocate 

their harness penetration. In Table 2.5-2, a breakdown of the 375-pound weight is 

compared with that of the mutually agreed upon 323-pound weight in August 1967. 

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TABLE 2.5 -2. ATS WEIGHT BREAKDOWNS (Sheet 1 of 2) 

DETAILED 323 LB AND 375 LB BREAKDOWNS 

ITEM 323 LB 375 LB 

Telescope Objective End 52.50 58.00 

Telescope Low Power Group 54.50 43.58 

Telescope Elbow Assembly 14.80 

Telescope Pechan Assembly 13.00 14.00 

Telescope Shear Mount 2.50 3.58 

Telescope Zoom Assembly 11.54 

Telescope Eye Piece Assembly 27.00 6.00 

Telescope Harness and Connectors 11.50 

Telescope Total 149.5 163.00 

External Fixed-Fold Mirror 30.5 30.5 

Tracking Pedestal 108.0 110.0 

Additional Weight for Inertias 

and Balancing 

5.0 33.0 

Tracking Pedestal Total 113.0 143.0 

Inner Shroud 22.0 28.0 

Pressure Window 8.0 8.5 

Sun-Sensor 0.0 2.0 

System Total 323.0 375.0 

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TABLE 2.5-2. ATS WEIGHT BREAKDOWNS (Sheet 2 of 2) 

COMPARISON OF 323 LB, 329 LB, 344 LB, AND 375 LB BREAKDOWNS 

ITEM 323 LB 329 LB 344 LB 375 LB 

Telescope 149.5 149.5 149.5 163 

External Fixed-Fold Mirror 30.5 30.5 30.5 30.5 

Tracking Pedestal 113.0 113.0 120 143 

Inner Shroud 22 28 36 28 

Pressure Window 8 8 8 8.5 

Sun-Sensor 0 0 0 2 

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2.5.8.3 Potential Weight Reduction Areas Under Study 

There are a number of weight reduction areas currently under study. These studies are 

being performed by both the subcontractor and GE. These items are listed along with a 

brief description. 

a. Reduction of Shock Loading Due to Extension of DR 1100A Shock Curve Below 

100 Hz - The subcontractor is currently cutting off the shock curve at 100 Hz 

which means that for any structural frequency below 100 Hz, the shock value for 

100 Hz is being used. This is very conservative since the frequency of the 

structure is in the 30 to 40 Hz range. It has been determined by discussion with 

Aerospace Corp. engineers that the curve should be extended to 0 g's. This will 

cause a reduction in shock loading which should cause a reduction in structural 

weight. 

b. Reduction of Telescope Wall Thickness to Possibly 0.030 Inch - This is covered 

in a previous portion of this section. 

c. Reduction of Mirror Thickness From 1.0 inch to 0.75 inch - Studies are currently 

being made to reduce the fixed-fold and tracking mirror thicknesses to 0.75 inch. 

The biggest problem appears to be tolerances. The subcontractor claims that a 

reduction in mirror thickness will cause a tightening of tolerances to the point 

where it may be impossible to fabricate. Another mirror study is being made on 

the possibility of going to an egg-crate mirror which would cause sizeable decreases in weight. It should be noted that the weight savings may not be applicable to the tracking mirror since this may aggravate the inertia problem. 

d. Usage of Beryllium in the Telescope Structure - This is covered in a previous 

portion of this section. 

e. Design Optimization - This can be accomplished by performing detail structural 

analysis to determine overdesigned areas and then taking the necessary steps to 

reduce the weight and eliminate the overdesign. This process will evolve as 

more details of the design are defined. 

2.5.8.4 Potential Weight increase Areas 

There are currently some areas which will require close scrutiny due to their potential 

impact on the system weight. Some of these areas are near solution with a minimum 

weight increase while others are still pending. 

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a. Increased Shroud Weight Due to ACTS Plume Impingment - Covered previously 

in this section; may cause only a small increase in structure and insulation. 

b. Out-of-Specification Inertias - Covered previously in this section; currently 

being studied with some increase likely. 

c. Increased Structural Weight Due to Increased Structural Stiffness Necessitated by 

Dynamic Disturbances from ACTS Controls, etc. - This area is discussed in the 

dynamic analysis section but could cause serious problems. 

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2. 6 BORESIGHTING (See also 2. 3. 2) 

The equations for the boresighting procedure are derived and a degraded mode of operation investigated. The material is broken down in Paragraphs 2.6.1 and 2.6.2. 

2.6.1 NORMAL MODE EQUATIONS 

To point the main optics by using the ATS, the alignment between the ATS optical system 

and the primary optical system must be known. This ATS/Primary optics alignment is 

found by on-orbit boresighting. 

The form of the boresighting equations depends upon the types of errors that are considered in the ATS and primary optics. 

Considering the ATS and the primary optics as two rigid boxes, the alignment error between them can be written as ax' a

y

, and a

z

. 

ATS "BOX" PRIMARY OPTICS "BOX" 

OPTICAL AXIS (a , a , a ) 

x y z 

LOS 

LOS 

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If

x 

 , a 

y 

 , and a are the only major errors from the ATS system to the primary optics 

tracking mirror, the boresighting equations can be derived as follows: 

If a

x' y' and a are misalignments between the ATS and the primary optical systems, 

the primary optics coordinates x , can be related to the ATS coordinates, x

a

, by: 

x

a 

Y

z a 

a 

1 a 

z 

-a 

y 

= -a 1 a

z x 

a -a 1 

y x 

Ps] 

x 

p 

Yp 

z 

p 

( 9) 

where 

Cos e 0 -Sin 

S S 

0 1 0 (10) 

Sin a 0 Cos 6 

s 

If {Vin} and {V. are the line of sight vectors in terms of the ATS and primary in a, 

p

} 

optics coordinates, the "images" in both systems can be written as: 

Vimage A = 

T] 

[Mai [ 4)g] Vin1 (11) 

and 

V. image p 

[131 [Md {Vin p (12 

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where 

Cos ;lig Sin tag 0 

-Sin ii) g Cos tag 0 

0 0 1 

(13) 

( s is the ATS gimbal skew angle.) 

1 bz -by 

-bz 1 0 

by 0 1 

(by and bz are the misalignments of the primary optical axis.) 

M

a and M are matrices consisting of sines and cosines of the gimbal angles. 

-s2a c2a sw -c20 cw a a a a a 

M

a 

c2a sw

a 

c

2w

a 

+ s2w 

a 

sal' 

 a 

sto

a 

ccl.)

a 

(1-s2a

a

a ) 

-c20 cw sto

a 

cw

a 

(1-s2a 

a

) s2w

a 

+ c2w

a 

s2 CT a a 

a 

s2a 

-c2cr sw 

P p 

c2a cw 

P P 

(sg and c g are used for Sin g and Cos g) 

[b 

M 

p 

(14) 

(15) 

(16) 

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(e s) (M) (3) 0 = 

P 11J (20) 

a 

B

A 

a 

C

a 

= _a 

z 

a 

Y 

a

z 

1 

-a 

x 

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The "image" of the line of sight is a 

-1 

V. = 0 image 

a 0 

1 

V. 0 image P 0 

ray along the x axis. 

(17) 

(18) 

Next, solve equations 11 and 12 to obtain the line of sight vector, and make the following 

substitution; 

A

a 

B

a 

C

a 

= 1 V in 

a 

= 

[osT] 

V 

1 

0 

0 

(19) 

and 

V. in 

P 

Equations 19 and 20 are related to each other by the transformation in equation one. 

-a \ A Y 

a x 

: 

p 

C 

11 P 

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p 

BA P 

C 

P 

= (e) 

i 

(21) 
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or: 

0 -C B 

P p 

C 0 -A 

P p 

-B A 0 

P p 

a 

a

z 

(21) 

These equations can now be used with measured ATS and primary TM gimbal angles to 

determine the alignment between the ATS and main optics. (However, it should be mentioned that equation 21 represents only 2 independent equations. Thus, more than one set 

of measurements will be necessary to solve for a x, a 

y 

 and a

z

. 

Let the subscript one refer to one set of measured gimbal angles and two refer to another 

set. Then: 

0 

C131 

0 

C

p2 

-Cpl 

0 

-C p2 

0 

131 

-Apl 

B

p2 

-Ap2 

aclx

a

z 

A

al 

B

al 

Aa2 

B

a2 

A

pl 

B p1 

A p2 

B p2 

(22) 

To solve this set of equations, two methods an available. One, drop one equation, and 

solve for ax, a , and a

z 

by standard matrix inversion; and two, perform a maximum 

Y 

likelihood solution for a , a , and a. The second method will yield the best accuracy 

x y z 

since the measured gimbal angles will contain some error, and hence this is the desired 

method of solving for a x, a y, and a

z

. 

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This maximum likelihood solution for the ATS alignment is: 

K 

r KT B

A1 

- B 

- Ap2 A

a2 

B

A2 

B

p2 

where: 

x 

y 

z 

A

al 

Apl 

p1 

A 2 

(23) 

(K) = 

0 -Cpl pl 

C

pl 

0 -A 

pl 

0 "Cp2 p2 

_Cp2 0 -A 

p2 

(24) 

The disadvantage of this second method is that the equations are more complex and therefore will require more computer time for solution. 

An error analysis performed on this boresighting procedure has shown an improvement of 

about 7 percent in the pointing accuracy of the main optics in the slave mode when the least 

squares solution is used for foresighting. Hence the least squares solution is the currently 

planned method of obtaining the ATS/Prime Optics boresighting. 

The error analysis conducted for ATS boresighting has shown that with apportioned values 

of errors that were used, the prime optics can point within 6 minutes (slave mode) and the 

ATS can point within 8 minutes (automatic mode neglecting ephemeric and target location 

errors). However if the ATS on orbit misalignment exceeds the apportioned values, the 

pointing errors increase. This increase would then require an increase in the number of 

ATS alignment errors found by the boresighting procedure. 

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2.6.2 FAILED PELLICLE MODE 

This paragraph presents the procedure required for boresighting the ATS to the main optics 

in case of failure of the visual optics in the primary optical system. (Failed pellicle mode). 

The standard procedure for boresighting each ATS to the main optics is to center a target 

simultaneously with both the main optics and ATS, and record the gimbal angles in both 

systems. When this has been done for two targets, the gimbal angle information can be 

used to determine the alignment between the ATS system and the primary optical system. 

The procedure becomes more complicated for the failed pellicle mode. For this mode, the 

crew must: 

a. Locate and center the target with the ATS that is being boresighted. 

b. Command the main optics to point to the same target--assuming perfect alignment 

between the main optics and the ATS. 

c. Photograph the target, and record the ATS and main optics gimbal angles at the 

time of photography. 

d. Record the primary optical axis alignment. 

e. Repeat the procedure for a second target. 

f. Develop both photographs. 

g. Inspect both photographs to determine the primary TM gimbal angles that would 

have caused the target image to appear on the center of the film format. Enter 

these angles into the computer via the keyboard. 

If the target does not appear near the center of format on the pictures taken in step c, the 

accuracy obtained by this boresighting procedure may be degraded. This in turn, implies 

that the boresighting procedure may have to be repeated for each ATS to obtain sufficient 

accuracy. 

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h. Insert the gimbal angles obtained in step g, the ATS gimbal angles from step c, 

and the optical axis alignment from step d into the computer program used for the 

standard boresighting procedure. 

2.6.2.1 Nomenclature 

by, bz - angular offset of the primary optical axis in the pitch and yaw direction (by and 

bz are obtained from on board alignment monitors) 

a - tracking mirror pitch gimbal angle 

w - tracking mirror roll gimbal angle 

e

s - pitch angle of the primary optical axis and tracking mirror (Os = 2 degrees 

nominally) 

k - optical magnification 

S( ) - used for sine ( ) 

CO - used for cosine ( ) 

2.6.2.2 Analysis 

Assume that a photograph has been obtained and a well defined target appears on it. (Note 

that this assumes that the alignment between the ATS and primary optical system must be 

no worse than 30 arc min. Since the MO semi field of view is approximately 32 arc min. , 

a 30 arc min. alignment error between the ATS and primary optics would cause the target 

image to appear at the edge of the film format.) Also assume that the primary TM gimbal 

angles and optical axis alignment were recorded at the time the picture was taken. It is 

then desired to find the TM roll and pitch corrections that would cause the target image to 

appear on the center of the film format. 

The photograph can be represented as shown in the following sketch. The axes y, z define 

the center of the film format. 

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'VI 

z 

TARGET 

Let A y1 and A z1 represent the displacement of the target image from the center of the film 

format. 

Let V in represent a ray entering the primary optical system. The image of this ray is: 

V image = (b) (Mp) (0s) Vin (1) 

where 

(1 b

z 

(b) = -b

z 

1 0 (2) 

0 1 

Y 

S2 or 

M= -C2cr SW (3) 

C2CT CCO 

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cos e

s 

0 -Sin S - 

0 1 0 

Sin e s 0 cos e 

If the optical system is pointed perfectly, and V in represents the line of sight, the image 

is a vector along the x axis. When errors are present, the image will have components 

along the y and z axes. 

Then, one can write: 

V image (desired) = (b) (M 1 P

) desired (6S

) V LOS -- {() 

0 

and 

i 

V image (actual) = (b) (M ) v = Y P actual S LOS k Z 

■■■ 

Equating (6s) V LOS from both equations given 

T 1 1 T (M p) desired (b) 0 = (MP

) actual (b) k Y 

0 k Z 

Since (b) is obtained from the alignment monitor, (M p ) actual is derived from measured 

gimbal angles; and A y, z are obtained from the photographs, equation (7) can be used 

to solve for the desired gimbal angles. 

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e

s 

(4) 

(5) 

(6) 

(7) 
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If b and b are assumed to be small, equation (7) can be approximated to: Y z 

S2 a 

-C2cr Su) 

C2o.Cco desired 

(MP) actual (bT

) 

1 0 1 

kLY } + {0 

k bZ 0 

(8) 

From equation 8 the desired pitch and roll angle can be easily obtained. 

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2.7 IMAGE DEROTATION AND MANUAL CONTROL STICK RECOUPLING 

The equations relating to derotation of the target image via the Pechan prism, and the decoupling of MCS signals, is presented in this section. The material is subdivided as 

follows; 

2.7.1 Introduction 

2.7.2 Notation and Definitions 

2.7.3 Coordinate Systems 

2.7.4 Tracking Mirror Matrix 

2. 7.5 Image Derotation 

2.7. 6 Image Motion Decoupling 

2.7.7 Approximations to the Derotation Equations 

2.7. 8 Standardized Equations 

2. 7. 9 Image Derotation 

2.7.10 Image Motion Decoupling 

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2.7 IMAGE DEROTATION AND MANUAL CONTROL STICK DECOUPLING 

2.7.1 INTRODUCTION 

Image derotation is the alignment of the image such that it appears on the visual display 

the same as it would appear if it were observed in-track. The reason for image derotation 

is for visual adaptation. 

Decoupling refers to coordinating the fore and aft motion of the control stick to vertical motion 

of the image on the visual display and right and left motion of the control stick to horizontal 

motion of the image on the visual display. The decoupling equations relate image motion at 

the center of the visual display to gimbal angle rates necessary to produce such image 

motion. 

The image velocity sensor (IVS) measures how well the target is being tracked by sensing 

any motion of the image relative to the line-of-sight (LOS). The gimbal angle rate errors 

necessary to track the target are then given by the decoupling equations. 

The equations for these functions are derived in this section. 

2. 7.2 NOTATION AND DEFINITIONS 

LOS Line-of-sight 

Subscript a Associated with ATS 

Subscript p Associated with primary optics 

C Cosine 

S Sine 

XBYBZB Body reference system 

XpYpZp Primary optics reference system (nominal) 

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X

a 

Y

a 

Z

a 

ATS reference system 

X' Y' Z' Primary optics reference system (misaligned) 

P P P 

Os Total cant angle (primary optics = 2 degrees) 

x

g Gimbal cant angle (ATS = 9 degrees) 

a , co Gimbal angles - Smith notation 

0 , A Gimbal angles - Balch notation 

[M] 

Mirror matrix (tracking mirror) 

{P} 

Object point 

1:"} Image point 

{1;'} Image velocity 

y Derotation angle 

y IVS Angle to IVS sensor orientation 

b

y

, b

z 

Angular misalignments of primary optical axis 

i y' i z Components of measured (IVS) or desired (control stick) image 

velocity 

Ad. , A&) Gimbal angle rates corresponding to 4 and L 

Y z 

K

o 

Constant due to magnification, loop gains, etc. 

0} Unit normal to tracking mirror 

{T } In-track unit vector 

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— 2.7.3 COORDINATE SYSTEMS 

The coordinate systems in which the gimbal angles are measured for the ATS and primary 

optics are shown in Figures 2.7-1 and 2.7-2 respectively. Their relation to the body 

reference system is shown in Figure 2.7-3 in which it can be seen that: 

XP XB 

Yp =

I 

{

es YB 

4p ZB 

where (1) 

[

Ce

s 

0 Ses 

0 1 0 

SOS0 ce

s 

:a 

Z

a 

XB 

YB 

ZB 

where (2) 

Fg1 

[

Cxg Sx

g 

 0 

-Sx

g 

 Cx

g 

 0 

0 0 1 

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Led 

and 
X 

p 

• 

Figure 2.7-1 

45° a p 

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{TNia} 

(Caa+ Scra) 

,sirr -(Caa _ sad sw a 

(Caa- Scra) Cw a 

LINE OF NORMAL 

SIGHT 

p 

OPTICAL 

AXIS 

{ 

-(Ca - Sad ' 

1■1 2 -(Cap P 4- Scr )So.) P P 

(Cap+ SCY )C(4 P 

Figure 2.7-2 

2-406 

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Figure 2.7-3 

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2.7.4 TRACKING MIRROR MATRIX 

The mirror matrix which transforms object space into image space is a function of the 

direction cosines of the normal to the mirror. The normal to the mirrors are given in 

Figures 2.7-1 and 2.7-2. Therefore the mirror matrices, 

for ATS are: 

-S2°.a 

Sw aCara 

-Cw C2cr a a 

Sw a

Cacr a 

C2w

a 

+ S2w

a 

S2 cr

a 

Sw a Cwa (1 - S2cr

a

) 

-Cw C2 

a aa 

Sw aCw a(1 - S2cra) 

S2 w 

a 

+ C2w 

a

S2cr

a 

(3) 

for primary optics are: 

[Dap] 

S2cr 

P 

-Car Ste 

P P 

C2cr Cw 

P P 

-C2cr SW 

p p 

C2w -S2w S p p ar p 

Sw

PCwP P 

(1+S2) 

C2cr Cw 

P P 

SW 

P P 

Cw (1 + Sar

P 

 ) 

S2c4) -C2(.1.) S2cr 

P p p 

■./. 

(4) 

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YB 

ZB 

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2.7.5 IMAGE DEROTATION 

The purpose of image derotation is to orient the image such that a vertical line in the image 

plane corresopnds to an in-track line in object space 

For ATS: 

Since the image plane is the YBZBplane, the image points are given by 

Pcgi T [Ma] xg] {Pa} 

(5) 

The image of the in-track unit vector in the image plane is found from (5) to be (Figure 2.7-4). 

S w

a

C2 cr

a

C2x

g 

- Sx

g

Cx

g 

(C2ct.:

a 

+S2 cr

a 

(1 + S2wa) ) 

-Cw

a

C2cr

a

Cx

g 

-Sx

g

Sw 

a

Cw 

a 

(1 - S2cra) (6) 

Y

a 

YB 

Z

a 

+ ROTATION 

IN TRACK 

LINE 

ZB ZB 

BE FORE AFTER 

ROTATION ROTATION 

Figure 2.7-4 

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Defining positive rotation of the image to be given by the right hand rule about the X/3axis, 

the image derotation angle for the ATS is given by: 

Y 

 = TAN-1

a 

Sco a C2cra C2xg - Sx gCxg (C2w 

a 

+ S2cra(1 + S2(4) a)) 

(7) 

-Cw

a

Car 

a

Cx

g 

- Sx

g

Sw 

a

Cw

a 

(1 - Sara) 

where 

-90o s y

a 

s 900 

Since the actual image rotation is twice the derotation prism rotation, the command to the 

derotation prism would be y a/2 to orient the image. 

For primary optics: 

There is an on-board alignment monitor associated with the primary optics that measures 

the angular offset of the optical axis about the Y and Z axes. These measurements, 

deonted by b and b y 

z

, are used to construct a transformation matrix between the XpYpZp 

and XL)('PP Z' coordinate systems (Figure 2.7-5). Since the image plane is the Y'PP Z' plane, P 

the image points are given by 

= [ ID ] [MPH esIPP 

where (8) 

[b] 

1 b

z y 

-b 

-bz 1 

by 0 1 

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The image of the in-track unit vector in the image plane is found from (8) to be (Figure 

2.7-6). 

(9) 

C2crpCco pCes- 52

co pses+ C2w pS2apSes+ by(S2crpC0 - C2crpCwpSes) 

Defining positive rotation of the image to be given by the right hand rule about the Xi; axis, 

the image derotation angle for the primary optics is given by: 

yp = TAN -1 p p pS -C2cr Sco COs-Ste Ccep Ses (1 + S2ap )-bZ p (S2cr Ces -C2a SO tau) 

C2cr Co.) p 

 ce s -s2 

p

Se

s 

 +C2W

p

S2ap

ses

+bY (S2apci9S

-C2aP Cu)P ses ) p 

(10) 

{:} -C2a p Su) COs- Sco Cce SeS p 

(1 + S2a )-bZ (S2a COs- C2ap Cu)p 

 seS ) p pp p 

X' 

p 

1 bz -by-1 1 Xr, 

. -b 1 0 Y 

z P 

b 0 1 Z 

_ N' P 

Z

p ' Z p X' - ACTUAL OPTICAL AXIS 

X - OPTICAL AXIS IF NO MISALIGNMENTS 

p 

Figure 2.7-5 

2-413 

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Y' P 

ilbv+ ROTATION 

Y' 

P 

Z' P 

X' P 

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BEFORE 

ROTATION 

Z' 

P 

AFTER 

ROTATION 

Figure 2.7-6 

Again the command to the derotation prism should be y p/2 to orientate the image correctly. 

Equations (7) and (10) are derived assuming the image of the tracking mirror is not inverted 

before reaching the visual display. Sould the image be inverted an odd number of times i.e. , 

1, 3, 5 , then the derotation angle is given by (180 - y) where y is given by (7) or (10). 

2-414 

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VISUAL DISPLAY VISUAL DISPLAY SENSOR DISPLAY 

Z

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2.7.6 IMAGE MOTION DECOUPLING 

The purpose of image motion decoupling is to relate vertical and horizontal image motion to 

changes in the gimbal angles. These equations give the gimbal angle rates necessary to 

correct for image motion detected by the IVS, and also provide the coordination between control stick motion and image motion on the ATS and primary optics visual display. Denoting 

the induced image motions by the control stick or the image motions detected by the IVS as 

yand 

z

, the decoupling equations are derived as follows. (See Figure 2.7-7.) 

For ATS: 

The position of the image is given by (5) and the velocity of the image is given by the chain 

rule as follows: 

[ 6 {PL} 6 IP' 

aco

a 

A• 

a 

6 cr a 6 w

a 

Figure 2.7-7 

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Now it is assumed that and z 

are the velocity components of the image of the optical 

axis; that is the image point at the center of the visual display. If any other point were 

chosen for decoupling, the off-axis image velocities would have to be considered in the decoupling equations. Since the mirror matrix is involutory ( [M] = [M] -1), the object 

whose image is along the optical axis is given by: 

{Pa} = [xg] 

T -1 

[Mal [xg] I: 

(12) 

Taking the partials indicated in (11) and substituting (12) for { Pa} yields: 

a {Pa} a [Mai 

- [Xg] T a cr a [M 

a 

 [x g] -1 a a 

a 0 

a 13a.} = [x 

w g 6 w 

a via,1 [ma] [x a] o 

a 0 

a 

(13) 

The products of the partial derivative of the mirror matrix and the mirror matrix are given 

by: 

3 [Ma] 

0 -Sw 

a 

Sw 

a 

0 

-Cw 0 a 

0 Cwa

C2a 

a 

-C w a a C2a 0 

-Sw aC2cr a (1 -S2a) 

Cw 

a 

0 

0 

Sw 

-(1 

a

C2a

a 

-Sara) 

0 

(14) 

(15) 

a 0. [Ma

] = 2 

a 

a [ma

] 

w

a 

[Ma] - 2 

2 -416 

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0 

= -2Sw

a 

2Cw Cx 

a g, 

0 

C2aCw 

a a 

SwaC2vaCxg+Sxg(1-S20a) 

i x 

z 

P' 

a 

Ov (16) 

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Now evaluating the partials in (13) and substituting them into (11) yields: 

Since and are the only components under consideration (16) reduces to 

-2Sw 

a 

C2a 

a

Cw 

a Ac; a 

2Cw 

a

Cx

g 

SwaC2craCx0+ Sxg(1 - S2a a) w 

a 

The inverse of Equation (17) is actually the desired expression since the gimbal angle rate 

• • 

commands are to be derived from /

Y 

 and / zwhich correspond to control stick position. Also image derotation has to be considered in the final equation which is: 

ova K

o 

K

4 -K3

] [ Cy 

a 

where 

K

1 

• 

-2Sw 

a 

K2 = 2Cw

a

Cx

g 

K

3 • C2a

a

Cw 

a 

K

4 

• 

Sw 

a 

C2cr 

 a

Cx

g 

+ Sxg (1 - S2cr 

D = K

1 K

4 -K2 K

3 

K

o 

= Constant depending upon optical magnification, loop gains.... 

y

a 

= Image derotation angle 

2-417 

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z 

A (.I) 

a D -K2 K1 -Sy 

a 

(17) 

Ya Cya frY 

(18) 
(21) 

a J 11/113} 

w 

[m] [ld 0 

p 0 

a {P; } 

a co p 

3 [MP ] 

[Dad 

0 

C ce 

p 

C2cr P 

C2cr P S p 

—C2cr Cw 

P P 

0 

(1 + 52 a ) 

P 

—C2o Sw 

P P 

—(1 + S2cr

P 

 ) 

0 

3 co 

2-418 

(23) 

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For primary optics: 

The analysis for decoupling in the primary optics follows the same logic as for decoupling 

in the ATS. Equation (11) becomes: 

{1'4 = 

3 

pp 

 ca 13, bar 

3 v p ace Ath (19) 

The object whose image is along the optical axis is given by: 

= [es] T [

MI

] [id T 

0 

1

0 

(20) 

Taking the partials indicated in (19) and substituting (20) for {

P

P} yields: 

a ip 

( PI 

a a 

p 

[b] 

[m 

a cr [Mr] [b]

0 

1 

0 

Where the products of the partials of the mirror matrix and the mirror matrix are given by: 

a im pi M = 

a cr L Pi 

2 

0 —Sw Cw 

Sw 0 0 (22) 

-Cu) 0 0 

p 

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Now evaluating the partials in (21) and substituting in (19) yields the desired velocity 

components: 

2Sw Cw C2a + b (1 + S2 v ) 

P p Y p 

-2Cw Sw C2cr + b (1 + S2a ) 

p p z 

(24) 

As before, the inverse of (24) is actually desired since the gimbal angle rate commands are 

to be computed from i) yand :t z, that is: 

Ac; p 

O co 

K

o 

K

4 

-K2 

-K3

1 K

3

[Y.] 

Y y 

D 

where (25) 

Kl 2Sw p 

K

2 -2Cco p 

K

3

P C2a + by (1 + S2 ) 

P 

K = Sw C2a + b (1 + S2a ) 

p z 

H4 

p 

 = 

is a rotation matrix to be discussed 

For the primary optics visual display which has been derotated by the angle given in Equation 

(10), the rotation matrix is: 

[Yi 

Cy p Sp 

Visual [-Sy Cy 

Display 

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For the IVS, if the coordinate system in which the image velocities are measured is rotated 

about the X; axes (right hand rule defining positive) by an angle from the Yi'DZp axes 

then the rotation matrix in (25) is given by: 

[v] IVS SY 

[ CYIVS 

IVS 

- IVS SYi 

CY IVS 

(27) 

. . . 

. Notice that the IVS senses I and 1 which are due to Pa and Au) . The gimbal rate 

y z P P • command to correct the image motion is then the negative of these or -Acr P and -Act)

P 

 . 

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2.7.7 APPROXIMATIONS TO THE DEROTATION EQUATIONS 

Since the reason for derotating the image is for ease in visual adaptation by the crew, the 

question arises about the accuracy with which the image need be orientated. Neglecting 

gimbal cant in the case of the ATS and total cant and misalignments in the case of the primary 

optics equations (7) and (10) reduce to: 

y - Le 

a a 

yp - 

Investigating the range of gimbal angles, these equations introduce maximum errors of 15 

degrees in the ATS derotation and 2 degrees in the primary optics derotation. Therefore 

equation (28) is not recommended for approximating the ATS derotation angle. 

The obliquity angle is a good approximation to the ATS derotation angle. However, this 

quantity would have to be computed first since it is not available in the computer. 

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(28) 

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2. 7. 8 STANDARDIZED EQUATIONS 

Early work done by several groups resulted in several definitions of gimbal angles. The 

extensive work done by E. Balch has resulted in his notation being chosen to be the standard 

notation. The transformation between the notation used in this report and the standard 

notation is given by: 

a = 9 - 450 a = 450- 6 

a a p p 

w = 0 w = -0 

a a p p 

(30) 

-_, 

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2. 7. 9 IMAGE DEROTATION 

ATS: 

SC/ S20 C2x

g 

 - Sx

g 

 Cx

g 

 (C2O

a 

- C2e

a 

(1 + S2S2 a)) 1 1

a 

y

a 

= TAN 

-cn

a

529

a

Cx

g 

- Sx

g 

SC/

a

CO 19 

a 

(1 + c2 

—90o s y s 90° 

Primary Optics: 

yp = TAN 

S2e 

p 

cnPCBs- S21 se

s+

2

n p p 

C26 S8 

s

+bY (C28P Ce

s 

 -S26P al p ses ) 

p 

Or 

y TT + (within 2°) 90o s y s 270o 

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--1 a a 

+S26 SO Ce + SS2 CO Se (1 + C28 )-b (C28 Ce - S28 cn Se ) (32) -1 13I) s p Ps p z ps p p s 
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K

2. 7.10 IMAGE MOTION DE COUPLING 

ATS: Equation (18) 

. 

1 -2SO 

a 

K

2 = 2CO 

a

Cx

g 

K3 = S2e 

a 

 crl 

a 

K4 SO

a

S20

a

Cx

g 

+ Sx

g 

(1 + C2 8a) 

Primary Optics: Equation (25) 

K = —2SO 1 p 

K2 = -2 co, 

P 

K

3 = co s28P + b Y 

P (1 + C2 0 ) P 

K4 = -SO S20 +b (1 + C2 0 ) 

p p z p 

Also: 

a

s 

= 0

a 

ap = -0 p 

w = 

a 

O 

a 44 = 13 

(33) 

(34) 

(35) 

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SECTION 3.0 

CUE SYSTEM ANALYSIS 

The material in this section is subdivided as follows; 

3. 1 Introduction 

3. 2 VDP Optical Analysis 

3.3 Subcontractor Compliance 

3.4 Controls and Structure 

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3. 1 INTRODUCTION 

A full PDR report by Lear Sigler was submitted to the Air Force. This report provided 

analysis of the optics and controls and described the projector design in detail. This report presents the results of an independent GE optical analysis of the VDP and reviews 

Lear's comments concerning compliance with the GE component spec. 

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3.2 VDP OPTICAL ANALYSIS 

An analysis of the projection optics has been performed. The active elements which were 

included in the analysis are the film base, the right-angle prism, and the lens which is a 

Cooke triplet operating at F/3.55. The purpose of the analysis is to determine if the system 

is capable of meeting the required performance level and to evaluate the assigned manufacturing tolerance values. 

The on-axis modulation transfer function, MTF, is shown in Figure 3.2-1. For comparison, the MTF of an aberration-free, diffraction-limited system is also shown which 

represents an upper limit to the performance of a real system of the same numerical 

aperture. The abscissa of Figure 3.2-1 is the spatial frequency in line pairs per mm of 

the aerial image (which appears immediately behind the rear projection screen) and is related to the film spatial frequency through the magnification, m, of the system; i. e. , 

film frequency = m x (screen frequency). 

The ordinate is the modulation transfer function which represents the ratio of the image 

contrast to the object contrast; i.e. , 

Ci (U) 

MTF (2

) - C (mU) o 

where 

C. (U) = image contrast at image frequency 2, 

C

o 

(mU) = object contrast at object frequency m2, 

and 

C - B max - B min 

B max + B min 

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E. 0 

Z 

2 2 

2 2 --.. '*---- a gl. 

14 )4 

E-4 

0

00 

4., ‘r t-- co 0 

• o c'l II ca' 2. I I p., `-`-i 2 

4.1 z o ° 44 4.O 0 E. i.4 cr) ,-,-; 44 0 ,rD 

cn 4) 

• 

ra4 e4 .. 0 E. - co 

U 

P-4 '1' 

CO) 0 II I 1 U cr) co t- Q • 

Z up 

= Z 0 41 GO c:II 

O 0 c E p .D ° LW c ; P 4 -s „ E. 

IDEAL DIFRACTION LI

MITED (MTF) 

O al 4d <4 a

o 

 II 2 .-1 

wp ,4 c..) c ri) 

0 

■-a 

•-

44., 

c)

E. <4t:4 

z 0 44 c9 c) , 

II w 

0 

DIFFRACTION LIMIT 

LINEPAIRS PER MILLIMETER IMMEDIATELY BEHIND SCREEN 

o Ln et. 

0 c; d 

• 00 

a) 

0 

4.3 

0 

r.4 

44-1 

a) 

O 

• 

0- 

0 

Cd 

4.4 

fA 

8 Figure3. 2-1. 

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It is seen from Figure 3.2-1 that at 9.8 1p/mm the MTF value for an aerial image is 

approximately 0.45 and at 17.6 1p/mm the MTF value is 0.14. The two frequencies correspond to the film frequencies of 128 1p/mm and 228 1p/mm, respectively for a system 

magnification of 13.0. 

To include the effects of the screen on the overall system, the MTF of the projection 

optics must be multiplied by the MTF of the rear projection screen. Figure 3.2-2 shows 

both MTF curves and their product curve which represents the overall MTF of the system. 

Evaluation of the system must be made in terms of the minimum perceptible contrast which 

is first discernible by the eye. Measurements of this quantity by DePalma and Lowry, 

JOSA 52, 328 (1962), are presented in Figure 3.2-3. From the figure, it is clear that 

frequencies near 10 1p/mm cannot be perceived by a normal population observer with this 

system, since contrast values approaching unity would be required. On the other hand, a 

10x magnifier having an MTF as low as 1/2 would require contrast values exceeding a 

contrast of only 0.005. This value corresponds to an average brightness of 20 ft-lamberts 

viewed at a distance of 10 inches. 

Working the problem in reverse, we will calculate the minimum film contrast necessary at 

128 1p/mm on the film to yield a perceptible contrast on the screen using a 10x magnifier 

with an MTF equal to 0.50. The equation to use is: 

C min. percept. C - film (MTF)

mag. 

x (MTF) 

screen 

x (MTF)proj. 

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• 

PROJECTION SYSTEM MTF 

X OVERALL MTF 

LINE PAIRS PER MILLIMETER AT SCREEN 

O CD O r-f O ri) O Fi 

r

M 0 

••-■ 0 N 44-1 rn as to 0 

•

■-■ z 8 Figure3. 2-2. 

O a) cg, ts LO •er, GNI 

C; C; O O O C; 

•—• 3,Ll1I 

RFT- SPECIAL HANDLING 

O 

3-9 
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- 

- 

• 35 INCHES 20 FT 

X 35 INCHES 300 

0 10 INCHES 20 FT 

REF DEPALMA 

J. OPT. SOC 

-LAMB 

FT -LAMB 

-LAMB 

AND LOWRY 

AM 52,328(1962) 

0 

0 

- 

_ 

- 

- 

0 

. 

• 

0 

_ 0 X 

• 

X 

X 

00 

0 

0 

0 

_ 

,-- 

,- 

■-• 

1 I 

0 

) .5 

* 

1 iiiii 

• 

0 

0 X 

0 9 

i . 9 

I I 111111 

x 

glIPP 

x 

X 

1 1 111 II 

0. 10 1.0 10. 0 

OBJECT LP/MM 

Figure 3.2-3. Contrast Sensitivity for Sine Wave Test Objects as a Function of Object 

Spatial Frequency for Several Viewing Distances and Brightnesses 

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1.0 

CA 

0.10 

MINIMUM PERCEPTIBLECONTRAST 

0.01 

0.001 

0.01 
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where 

Thus 

(MTF) mag. = 0.50 

(MTF) screen = 0.80 

(MTF) prof. = 0.49 

0.005 C - film 0.195 

Product = 0.195 

= 0.025 

which corresponds to a relative contrast of 1.05/1.0 at 128 1p/mm on the film. 

For 228 1p/mm on the film, the MTF values become: 

(MTF) mag. = 0.50 

(MTF) screen = 0.66 

(MTF) prof. = 0.18 

C min. percept. = 0.012 

Product = 0.0595 

0.012 C - = 0.202. film 0.0595 

This corresponds to a relative contrast of 1.5/1 at 228 1p/mm on the film. 

An analysis of the tolerance values assigned to the component separations and lens thicknesses has been performed by use of a statistical approach. Referring to Figure 3.2-4, 

the lens thicknesses considered are of lenses I, II, and III and of the prism. The separation considered is the distances between the surfaces of lenses I-11, II-II, and between III 

3-11 

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a, 

3.61 ± 0. 08 

0 *C *a *o 

Visu

al Display Proje

ctor Proje

ction Lensand DeviatingPrism Figure3. 2-4. 

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3-12 SECRET SPECIAL HANDLING 
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and the prism. The tolerances shown in the figure were used in this analysis. If it can be 

assumed that the assigned tolerance value represents one standard deviation, then the propagation of errors for a system of several parameters, each with an assigned standard 

deviation, results in a system standard deviation which can be computed by the relationship 

= E 

2 6F 2 

cr 

2 

(X

n

) an n=1 

where Xnis the nth parameter, anis the standard deviation of the nth parameter, a is the 

standard deviation of F, and n is the total number of parameters. 

This calculation has been carried out, and the results are indicated in Figure 3.2-5. Also 

shown for comparison is the nominal MTF curve as calculated by LSI. It is seen to fall 

within the calculated standard deviation. 

The interpretation of the curve of Figure 3.2-5 is that if the tolerance values assigned to 

different parameters represents one standard deviation, then the system standard deviation 

at any frequency represents the range in which there is a 63 percent chance that the MTF 

value will fall. Also the chance that the MTF value will be more than one standard deviation 

below the nominal value is 19 percent. 

Shown on Figure 3.2-3 is the minimum perceptible contrast as determined from the Schade 

curve and the screen contrast resulting from film contrasts of both 6 to 1 and 1.6 to 1. 

The intersection of the output modulation curves and the Schade curve represents the visual 

cutoff of the system for the specified input contrast. Using a display brightness of 400 ftlamberts, the visual cutoff for a 1.6-to-1 contrast input occurs at 6.5 line pairs per mm on 

the screen or 84.5 line pairs/mm on the film. 

3-13 

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Figure3. 2-5.ToleranceEv

aluation 

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ti 

CV 

CO 

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It is unlikely that the average unaided eye will be able to resolve the screen image produced by 228 1p/mm on the film, regardless of the film input contrast. 

From inspection of Figure 3.2-2, it is possible to determine the minimum magnification 

necessary to resolve 17.6 1p/mm on the screen (228 1p/mm on the film at 6 to 1 contrast). 

At 17.6 1p/mm, the screen contrast is seen to be 0.05, allowing for a magnifier MTF of 

0.9, the resulting contrast becomes 0.045. The Schade curve (0.40 ft-lamberts) has a 

value of 0.045 at 2.6 1p/mm. The necessary magnification is therefore 17.6/2.6 = 6.7. 

A similar calculation indicates that for a display brightness of 400 ft-lamberts a minimum 

magnification of 17.6/4.0 = 4.4 is necessary. 

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3.3 SUBCONTRACTOR COMPLIANCE COMMENTS 

The following is a listing of the Visual Display Projector requirements set forth in the 

Component Specification. A brief summary of the estimated system performance and the 

method of analysis used to determine this estimate are also included. This material was 

submitted by Lear Siegler, Inc. in their VDP Design Review Report in November 1967. 

3.3.1 ACCESS TIME (SINGLE FRAME OPERATION) 

Requirement - The time to step one frame in either direction shall not exceed 0.35 

second. 

Analysis Method - Digital and Analog computer studies. 

Estimated Performance - 0.30 second or less. 

Variances - None 

3.3.2 OPERATING TIMES 

Requirement 

b. Projection lamp turn-off and Binary Code comparison shall be accomplished 

in 0.05 second after receipt of external command. 

c. Frame search rate shall be 100 frames per second. 

d. Acceleration, deceleration and fine positioning within a screen image accuracy 

of ±0.125 inch shall be accomplished within 0.35 second. 

e. Frames shall be searched, positioned, and displayed within 

number 

100 

of frames 0.35 + second. 

f. Switching to standby mode shall be accomplished within 0.05 second after 

external command is removed. 

g. Warm-up time shall be less than 10.0 seconds. 

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Analysis Method 

Paragraphs b. , f. , and g. , - Theoretical circuit analysis. 

Paragraphs d. and e. - Digital and Analog computer studies. 

Paragraph c. - Breadboard tests. 

Estimated Performance 

b. The propagation time for the VDP circuits is less than 1 millisec. However, 

a delay circuit must be incorporated at the input dependent upon the maximum 

"zero" time of the updating external command. If this max. zero time is less 

than 45 millisec, the requirement can be met. 

c. 100 frames per second minimum. 

d, e. 0.35 second maximum except as noted below. 

f. Switching time will be 1 millisec plus max. "zero" time of updating external 

command. It is assumed that the projector is displaying a frame at the time 

external command is removed. 

g. Less than 1.0 second. 

Variances 

d, e. When there are less than three consecutively numbered frames prior to the 

commanded frame, up to 0.10 second additional time may be required. 

3. 3. 3 DISPLAY SCREEN BRIGHTNESS 

Requirement - Brightness shall be adjustable from 0 to 15 foot-lamberts. 

Analysis Method - Theoretical based on published data. 

Estimated Performance - 0 to 15 foot-lamberts on optical axis. 

Variance - None 

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3.3.4 DISPLAY RESOLUTION (PROTECTOR) 

Requirement - A viewer with a 10x (max.) magnifier will be able to resolve up to and 

including 228 1p/mm at 6 to 1 contrast on the film (corresponds to 17.6 1p/mm on the 

screen). 

Analysis Method - LSI and subcontractor theoretical and computer analysis, also 

breadboard test data. 

Estimated Performance - Resolution of 228 1p/mm (on film) at 6 to 1 contrast ratio 

is possible with a 7x magnifier. 

3.3.5 ALPHA-NUMERIC DISPLAY TIME 

Requirement - The alpha numeric display shall be accomplished within 0.3 second 

after receipt of external command. 

Analysis Method - Theoretical analysis of circuit. 

Estimated Performance - Time to display will be less than 0.1 second. Since the 

standby control monitors only the projector external command lines, the alphanumeric display can only be used when the VDP is out of standby mode. Therefore, 

the time required to go from standby to full power need not be considered. 

3.3.6 SEARCH AND FINE POSITIONING POWER 

Requirement - Maximum power to be 31 watts, limited to 1.0 amp. 

Analysis Method - Theoretical power analysis from circuit schematics. 

Estimated Performance - Maximum power during high speed search will be 26 watts. 

Input current will be 1.2 amp. at 22V input and 0.84 amp. at 31V. Pulse power (less 

than 1 sec) will be 41.4 watts during acceleration for a period of 0.2 second. This is 

1.88 amp at an input voltage of 22V or 1.34 amp at 31V. 

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Variance - Maximum power is not exceeded during search but maximum current is exceeded by 0.2 amp. at 22V input. If maximum current is to be limited to 1.0 amp, even 

at 22V, then the power limitation is 22 watts, not 31 watts. General Electric Action 

Item D-12 is to clarify this requirement as well as to establish a pulse power requirement. 

3.3.7 SCREEN ILLUMINATION POWER 

Requirement - Maximum total power during screen illumination shall not exceed 30 

watts, current limited at 1 amp. 

Analysis Method - Theoretical power analysis from circuit schematics. 

Estimated Performance - Maximum power during screen illumination will be 22 watts. 

Input current will be 1.0 amp. at 22V or 0.71 amp. at 31V input. 

Variance - None 

3.3.8 ALPHA NUMERIC DISPLAY POWER 

Requirement - Maximum power shall not exceed 5 watts. 

Analysis Method - Theoretical power analysis from schematics. 

Estimated Performance - Maximum power drawn from the input line when only the 

alpha-numeric display is used will be 6.0 watts at maximum brightness. 

Variance - Maximum power is exceeded by 1.0 watts. The actual load for the display 

is less than 5 watts, but power supply efficiency is low when only the display is used. 

Therefore, during projection mode the additional power required for the alpha numerics 

will be less than 5 watts. 

3.3.9 STANDBY POWER 

Requirement - Maximum standby power shall not exceed 0.03 watt. 

Analysis Method - Theoretical analysis of circuit. 

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Estimated Performance - Maximum standby power will be less than 0.03 watt. 

Variance - None 

3.3.10 POWER INTERRUPT 

Requirements - The VDP shall withstand voltages from 0 to 20 volts for periods up to 

one second and return to specified performance within 10 seconds. 

Analysis Method - Theoretical analysis of system schematics. 

Estimated Performance - This requirement is met. 

Variance - None 

3.4 CONTROLS AND STRUCTURE 

Discussion and supporting analysis of VDP controls is presented in detail in the LSI Visual 

Display Projector Design Review Report (LSI-TD-755-1167-ID) dated November 1967. A 

description of VDP controls (panel) is also contained in this report in Section 5. 

3-21/3-22 

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SECTION 4.0 

SUBSYSTEM DESIGN INFORMATION 

The material in this section is subdivided as follows: 

4.1 Drive K Electronics 

4.2 Control Sticks 

4.3 Gyro Selection 

4.4 Weight and Power Status 

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4.1 DRIVE K ELECTRONICS 

4. 1. 1 INTRODUCTION 

The Drive K Electronics has been subdivided into seven subfunctions designated K-0, K-1, 

K-2, ---- K-7. This portion of the engineering analysis report outlines the functional 

requirements of each circuit function, briefly describes the circuit approach for each 

including preliminary schematics, and includes the engineering laboratory data which has 

been taken on these circuit functions. 

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4-3/4-4 
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4. 1. 2 K-0 ELECTRONICS 

In its primary function, the K-0 electronics provides the digital circuitry for the precise 

positioning of both the inner and outer axes of Drive K by means of a position servo system. 

It will accept information in digital form for both axes from the MDAU and integrate it to 

develop digital commanded position information. It will process digital signals received 

from position encoding devices in each axis to develop actual position information. A 

digital position error word for each axis will be derived in the K-0 electronics from a comparison of commanded and actual positions and sent to digital-to-analog converters in the 

K-1 electronics. That component will provide the analog drive signal for the position 

servo system. Actual position data will be supplied to the MDAU on request. 

4. 1. 2. 1 Functions (K-0) 

The K-0 electronics portion of the DKE (Figure 4.1-1) shall be capable of performing the 

following functions: 

a. Provide the input/output buffering required for the flow of data between the various 

functional areas of the DKE and components associated with but not part of the 

DKE package. The function of the buffer portion of the K-0 electronics is to 

provide the following capabilities: 

1. To accept the command rate data used for the control of a two-axis position 

servo referred to as the Main Drive. 

2. To accept and decode the position/rate signal (received as part of the command rate data input) used for control of the commanded scaling in the K-0 

electronics and to provide inhibit signals to K-3, K-5, K-6B and Gyros as 

shown in Figure 4.1-1. 

3. To accept and distribute the position command data required by the K-3 

electronics. 

4. To transmit position feedback data received from an associated component. 

5. To provide switching required to accept data from multiple sources. 

6. To transmit digital data to the telemetry subsystem. 

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NRO APPROVED FOR 

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OSITION/RATE SIGNAL 
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b. Digitally integrate the rate command data received by the input/output buffer to 

generate a position command signal for servo control in two axis. 

c. Process the feedback data from two associated, incremental encoders and to 

generate a binary digital word which is representative of this data. 

d. To digitally combine position command data inputs and encoder feedback inputs 

and to generate from these binary-coded position error signals which are transmitted to the K-1 electronics. 

e. To generate an interlock signal for K-7 from signals received from the associated 

incremental encoders. 

f. To accept primary and back-up stow commands and generate required stow signals. 

g. To accept and respond to a back-up sync command. 

4.1.2.2 Performance Characteristics (K-0) 

4.1.2.2.1 Input/Output Buffering (K-0) 

A. Digital Rate Command - The e input buffer receives 32 bits of rate command information from either of two identical components (MDAU 1 and MDAU 2) at the rate of 100 times 

per second. The data is strobed serially into the register by means of 32 shift pulses on a 

separate line originating in the same component as the data. The first data bit contains no 

information and is always transmitted as a logic "0". Bits 2 through 16 are 14 data bits 

plus sign (15 bits total) representing the inner axis rate command. Bit 17 indicates the 

position mode or rate mode scale factor to be used in the arithmetic operations. Bits 18 

through 32 are 14 data bits plus sign representing the outer axis rate command. The data 

is transmitted least significant bits (LSB) first. Within (TBD) μ sec following reception of 

the last bit, the K-0 electronics will be sent an execute pulse originating in the same component as the data. The K-0 electronics will use this pulse to set the scale factor command into the control logic and to alert the logic to the fact that rate command data will be 

shifted from the input buffer to the respective inner axis and outer axis 0 shift registers 

during the next timing cycle. The scaling for this command is as indicated in Table 4.1-1. 

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TABLE 4.1-1. RATE COMMAND SCALING 

PARAMETER POSITION MODE 

(HIGH RATE RANGE) 

16 RATE MODE 

(LOW RATE RANGE) 

OA RATE MODE 

(LOW RATE RANGE) 

e

c MSB 

6c 

LSB 

N 

26.7796 deg/second 

3.269 x 10-3 deg/sec 

14 bits + sign bit 

6.6959 deg/second 

8.1725 x 10-4 deg/sec 

14 bits + sign bit 

1.6737 deg/second 

2.043 x 10-4 deg/sec 

14 bits + sign bit 

B. Stow Command - Means shall be provided to accept a manually initiated stow command 

to back up the automatic stow command received on the rate command line. The stow command shall cause a 5°/sec rate to be registered until stowing has been accomplished. The 

backup stow command will originate in the control console. 

C. K-3 Command - The K-3 command is also a binary digital word received by the K-0 

electronics. This command is received once per position mode selection, and has scale 

factors as indicated in Table 4.1-2. 

TABLE 4.1-2. K-3 FUNCTION COMMAND SCALING 

eK3 MSB 22.5 deg 

(most significant bit) 

19 K3 MSB 0. 045 deg 

(least significant bit) 

N

K3 10 bits + sign bit 

D. Position Readout - The Drive K Electronics position readout from the encoders consists 

of two binary digital words transmitted from the K-0 electronics and having scale factors 

as indicated in Table 4.1-3. The encoder feedback words shall have a code bit in addition 

to the data bits to identify the specific load position being transmitted. 

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SECRETSPECIAL HANDLING 
e MSB 45 deg 

9 LSB 2.75 x 10-3 deg 

N 6 15 bits + sign bit 

Encoder 

Scaling: IA: 

OA 

360 

29 

(2) 

360 

29 

(2) 

deg/count 

deg/count 

Encoder Pulse 

Rate, Max: IA: 120 KHz 

OA: 150 KHz 

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TABLE 4.1-3. OA AND IA DKE POSITION READOUT SCALING 

a. Encoder Feedback Register Characteristics: The feedback registers in which 

the OA and IA position words are accumulated shall be scaled as indicated in 

Table 4.1-4. 

TABLE 4.1-4. OA AND IA FEEDBACK REGISTER SCALING 

IA AND OA 

e

FDBK MSB 45° 

eFDBK LSB 1.25 arc sec 

NeFDBK 

18 bits ± sign bit 

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b. Synchronization - The feedback processor shall include means to synchronize the 

OA or IA feedback binary digital register word to the OA or IA load position whenever the OA or IA passes through a specific reference point referred to as the 

"sync" position. The sync positions are: 

OA Sync: 0 degree 

IA Sync: +11.25 degrees 

Synchronization of the e

FDBK registers shall take place automatically, without 

interrupting any other functions, whenever the load passes through the sync 

position in the position mode. Automatic sync will be inhibited during the rate 

mode. 

c. Initialization - The feedback data processor shall generate signals which will 

cause the OA and IA to drive through the sync positions and to reset the rate command ( en) to zero and the position command register to the sync position within 

10 milliseconds after power turn-on and the receipt of a K-7 load mechanism CW 

limit signal. In addition to the automatic initialization which shall take place 

whenever power is turned on, a redundant backup initialization path shall be 

provided which shall operate on a manually initiated signal originating on a control console. 

4.1.2.3 Additional Buffering Requirements for K-0 

a. The input/output buffer shall transmit a constant 625 kHz clock output signal. 

b. The input/output buffer shall transmit digital telemetry signals to the telemetry 

subsystem. 

c. Generate inhibit signals to K-3, K-5, K-6b, and gyro control heaters as a function of the position/rate mode. 

d. The input/output buffer shall transmit an interlock signal to the K-7 electronics 

indicating that either the OA or IA stow limit is not present. 

e. The input/output buffer shall generate a separate indication of OA and IA sync by 

supplying a one second logic level to the control console upon receipt of the 

following: 

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1. A change in state (0 to 1 or 1 to 0) of the respective encoder sector line and, 

2. An initialize gate indicating the main drive is in the process of being 

synchronized. 

4.1.2.4 Digital Integrator (K-0) 

The digital integrator includes a 15-bit rate command register ( e shift register), a 32-

bit position accumulator shift register, and a full adder. The 0 shift register is periodically updated with rate command information from the 0 input buffer. The contents of the 

6 shift register are then repetitively added to the accumulator to derive commanded position information. 

When the rate mode of operation is commanded, the contents of the e shift register are 

serially added to the 15 LSBs of the accumulator register once every timing cycle. Information is restored in the e shift register by an end around shift and the results of the 

addition are fed into the MSB end of the accumulator register. Note that there are only 

15 bits of information in the e shift register compared with the 32 bits in the accumulator 

register. Thus, if the e information is positive, 0 's continue to be added to the accumulator data following the MSB of e data. If negative, l's are added. This is in keeping with 

the signed 2's complement binary notation. The result in the accumulator is commanded 

position information at the end of the timing cycle. 

Whenever the e shift register is to be updated with new information, the end around shift 

is interrupted and new data is entered from the e input buffer as the previous word in the 

0 shift register is being added to the contents of the accumulator register. Both inner 

axis and outer axis e shift registers are updated simultaneously. 

In the position mode, the operation is the same as before except that 0's are added to the 

2 LSBs of the accumulator data and then the e data is added to bits 3 through 17 of the 

accumulator data. 

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The digital integration shall be performed at a rate of 19.5 kHz + 0.01 percent. The 

accumulator for the integrated rate command ( e command register) shall be scaled to 

accept the effect of the LSB in the rate command word. For each axis, the scaling is as 

shown in Table 4.1-5. 

TABLE 4.1-5. POSITION COMMAND SCALING 

e

c 

MSB 45 deg 

e

c 

LSB 4.1910 x 10-8 deg 

N 

ec 

31 bits + sign bit 

4.1.2.5 Feedback Processor (K-0) 

The feedback processor includes a five-bit up-down counter, a shift register buffer for 

the counter, a full adder, a 20-bit position feedback shift register le feedback), and interface logic. 

Each encoder (IA and OA) provides data on two lines. One line carries pulses when the 

encoder travels in one direction, and the other line carries pulses when the encoder travels 

in the opposite direction. Pulses appearing on these lines are mutually exclusive. If the 

up-down counter is counting in the up direction corresponding to the positive direction of 

rotation of the encoder and the encoder changes direction, the interface logic will sense the 

first pulse appearing on the opposite line and switch the counter to count pulses on this line 

in the down direction. The contents of the counter in the down mode represent position 

changes in signed 2's complement binary notation. 

The contents of the up-down counter are added to the e feedback shift register once every 

timing cycle. This is accomplished by inhibiting the pulses on the encoder input lines for 

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one 625 kHz clock period. During the inhibit time, the contents of the counter are transferred to the shift register and then the counter is reset. The inhibit signal is released 

and the contents of the shift register are serially added to the contents of the 9 feedback 

shift register. The results of the addition are returned to the 9 feedback shift register 

via an end around shift. The encoder input lines are inhibited for less than 1/4 of the 

shortest encoder pulse cycle so that no count is lost. 

When the encoder sync pulse is detected, the up-down counter is immediately reset and 

alerts the control logic that the 9 feedback shift register is to be reset. This is to insure 

that the contents of the 9 feedback shift register accurately represent the encoder position. 

The counter continues to count, from zero, the encoder pulses occurring after the sync 

until the next inhibit period. During the next inhibit period, the 8 feedback register is 

preset and the contents of the counter are transferred as before to the associated shift 

register. Subsequent to the inhibit period, the shift register contents are added to the 

contents of the 9 feedback shift register. The sync logic will be disabled during the rate 

mode of operation. 

The K-0 electronics will detect and accept requests from an MDAU for position feedback 

data, and within 100 μ sec, will respond with a "ready" signal if data is ready. During 

the 100 g sec period, the contents of the 9 feedback shift register will be shifted nondestructively to an output buffer register. Between 1.6 and 3.2 sec after the appearance 

of the ready signal, the K-0 electronics will accept from the MDAU a series of 625 kHz 

pulses to send position feedback data from the buffer register to the MDAU. The 17 bits 

of data sent back will include a sign bit and a tag bit to indicate whether it is inner axis or 

outer axis data. The type of data will alternate with successive request signals. 

4.1.2.6 Position Error Adder (K-0) 

This unit digitally combines the output of the 9 integrator-adder representing commanded 

position information and the contents of the 9 feedback shift register once each timing 

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cycle. The control logic will select the appropriate thirteen bits plus sign of each word to 

be added during either the rate or position mode of operation. The scaling for position 

and rate modes is indicated in Table 4.1-6. 

TABLE 4.1-6. OA AND IA ERROR WORD SCALING 

e

E MSB 0.41842 deg 

e

E LSB 1.02 x 10-4 deg 

N 

e E 

13 bits + sign bit 

This data (0E) shall then be transmitted to the two D/A converters of the K-1 electronics, 

so that eE is positioned in the LSB location of the K-1 D/A converter. 

LSB 

4.1.2.7 Timing and Control Circuitry 

All the necessary clock signals required for operating the K-0 electronics in a synchronous manner are derived from this unit. The prime source is a stable 1.25 mHz 

oscillator. This unit will process the output of the 1.25 mHz oscillator to derive two 

625 kHz (800 ns wide) clocks phase displaced by 90° from each other. From these clocks, 

four additional 625 kHz clocks are derived. Each is 400 ns wide and is displaced 400 ns 

from the preceding one. With this clock arrangement, time races in the control logic are 

avoided and reliable operation insured. These clocks are used in conjunction with the 

outputs of a 5-stage synchronous counter pulsed at a 625 kHz rate to provide control signals 

during each timing cycle. This logic circuitry will combine internal and external signals 

with signals from the timing generator to insure the proper sequence of operations of the 

K-0 electronics within the basic timing cycle. The timing cycle consists of 32 pulses of a 

continuous 625 kHz clock, and all operations are repetitive from cycle to cycle. The control logic will also furnish appropriate sequences of shift pulses to the registers in the 

various units to control the transfer of data. It will also supply logic and timing functions 

to the D/A converters in the K-1 electronics. 

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4.1.3 K-I ELECTRONICS 

4.1.3.1 Description 

The K-1 electronics (see Figure 4.1-2) will operate in a primary mode or a backup mode. 

In the primary mode, the K-1 electronics will provide drive signals to the Inner Axis gyro 

(IA) and to the Outer Axis gyro (OA). The IA signals will be developed by a D/A conversion of position error ( 6E) data received from the K-0 electronics. The OA signals will 

be developed by a D/A conversion of rate (6) command information received from the K-0 

electronics. The output of the D/A ladder networks (IA and OA) provide the inputs for the 

buffer and compensation amplifiers. These amplifiers provide the power gain required to 

drive the gyro torquer coils (IA and OA). The amplifiers are identical in each axis. 

The requirements on the amplifiers are very stringent, requiring low drift, low output 

noise, and high gain linearity. To meet these requirements, it was necessary to use 

discrete part amplifiers (see Figure 4.1-3) rather than microcircuit amplifiers. A 

description of the amplifier characteristics follows: 

The amplifier has two stages of differential gain, a buffer stage to match impedances, and 

a class B output stage to provide the five watts of power required. The input differential 

stage uses a 2N2453A transistor consisting of two matched transistors on the same chip to 

obtain good parameter matching and excellent temperature tracking at the input stage. The 

second differential stage also uses matched transistors with similar characteristics. The 

amplifier uses current feedback to provide a high output impedance to the load and results 

in the transfer function of output current for input voltage required by the specification. 

The amplifier gain is 169 ma/volt in each axis. The drift at the output is less than 2.0 ma, 

2 cr and the linear range of operation is up to 200 ma into a 100-ohm torquer coil load. 

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K-1 ELECTRONICS 

PRI ME 

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BUFFER GYRO D/A AND 

COMPENSATION 

TA 

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a b. 

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co 

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SECRET SPECIAL HANDLING 

Figure 4. 1-3. DriveK Buffer Amplifier, K-laIA (K-laOA) 

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In the IA backup mode, the D/A conversion of IA, eEdata from the K-0 electronics will 

result in signals to drive the K-la IA, amplifier. In the OA backup mode, the D/A conversion of OA e Edata from the K-0 electronics will result in signals to drive the K-lb 

OA amplifier. Mode selection will be determined by signals from the console. 

The K-lb amplifier is shown in Figure 4.1-4. The amplifier shown has a closed loop voltage gain of 23.2. Breadboard tests of the output drift show a 19. 2 my change in the output 

over the range from -18°C to +70°C. 

The plot of the open loop gain (GH) and phase margin is shown in Figure 4.1-5. The closed 

loop gain is 1 + GH. If the minimum GH product is 800, a variation of the forward gain 

(G) from infinity to 18,800 (giving a GH of 800) will produce a total gain variation of 0.125 

percent. The gain is therefore controlled primarily by the ±1 percent feedback resistor 

which maintains a 3.5 percent tolerance on the gain. 

The stability of the amplifier was determined by using the ECAP computer program and 

worst case transistor parameters. The phase margin at the crossover, with all transistors 

at their worst case maximum value, is 40 degrees. 

4.1.3.2 Requirements Summary (K-1) 

The K-1 electronics shall perform the following functions: 

a. Accept digital position error data or rate command data and timing signals from 

the K-0 electronics and convert these signals to analog voltages. Logical and 

timing functions of the D/A converter will be supplied by the K-0 electronics. 

b. Provide the required transfer functions to generate gyro command and K-2 input 

signals. 

c. Provide analog status signals of critical circuit parameters to K-9 electronics 

(telemetry conditioning and AGE). 

4-18 

SECRET SPECIAL HANDLING 
Oo Pi 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 --SititEt SPECIAL HANDLING 

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---SENET-SPECIAL HANDLING 

4-19 
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4-20 

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RELEASE 1 JULY 2015 SPECIAL HANDLING 

4.1.3.3 Performance Characteristics (K-1) 

a. D/A Conversion - The D/A converters of the K-1 electronics shall generate voltages proportional to the OA and IA position errors and OA rate command received 

from the corresponding channels of the K-0 electronics. Each converter shall 

have a constant scale factor gain (volts/quantum level) and shall be consistent 

with Table 4.1-6 for K-la IA, K-lb IA, and K-lb OA; and Table 4.1-1 for K-la 

OA. Timing signals for the D/A converters shall be derived from KO. Each 

converter shall be capable of processing word rates of 19.5 kHz. D/A conversion 

delay shall not exceed the inverse of word rate. 

b. Electronics Gain - The following steady state gains shall be provided: 

A K-l

a

IA 72.5 RAD 15 percent MPS 

OLTS K-lb VRAD IA 13,600 15 percent 

AMPS-S R EC K-1

a

0A 0,220 ±5 percent AD 

K-1b OA 1730 VOLTS

RAD ±6 percent 

These gains shall remain the same for position and rate modes. Compensating 

buffer amplifier gains shall be made to keep the gain constant for position and 

rate mode digital scaling. A maximum of 100 milliseconds between digital scale 

changes and analogue scale changes is allowed. 

c. Dynamic Response - The following transfer functions shall be provided in K-1 

electronics: 

K-la IA 

K-lb IA 

4-21 

SPECIAL HANDLING 

S/10 + 1 

S + 1 

S/20 + 1 

(S/200 + 1) (S/100 + 1) 
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K-1A OA 1 

S/30 + 1 

K-lb OA S/10 + 1 

(S/100 + 1) (S/1000 + 1) 

The break frequencies shall be generated within a 110 percent tolerance. Additional phase lag due to breaks above 30 rad/sec shall not contribute more than 

6 degrees phase lag at 30 rad/sec. 

d. Linear Range and Saturation - The gain specification given above shall apply 

within the following output ranges: 

K-la IA, K-la OA: 10.1 to 200 ma into a load of 100 ohms ±10 percent with 

a stability of better than 1 percent. In the region of 0.04 ma, the gain shall 

not be less than zero. The saturated output current shall not exceed 240 ma 

in magnitude. Recovery, to 90 percent of its saturated value, shall occur in 

less than 2 milliseconds when the input is switched to zero from a saturating 

level. 

K-lb IA, K-lb OA: ±17.5 Volts. Saturation shall not exceed a value which 

will damage K-lb or the following stage, K-2. Recovery to 90 percent of 

saturation shall occur in less than 2 milliseconds when the input is switched 

to zero from a saturation level. 

e. D. C. Offset and Long Term Drift: 

K-la IA 11 ma 

K-la OA 110 m amps 

K-lb IA ±380 my 

K-lb OA ±57.3 my 

The long term drift applies to frequencies below 0.2 rad/sec. 

f. High Frequency Noise (W Z 0.2 rad/sec) - TBD 

4-22 

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4.1.3.4 Stability Analysis (K-1) 

A stability analysis was performed on the K-la buffer amplifier shown in Figure 4.1-3. 

The open loop gain of the amplifier varies from 10,000 V/V to 80,000 V/V when driving 

a 115-ohm load. The closed gain required for this amplifier is 16 V/V (145 ma/V). The 

open loop gain variation will cause only a 0.12 percent amplifier gain change. Most of the 

amplifier gain variation can be assigned to the input and output impedance variations. 

A stability analysis was also performed on the K-lb backup amplifier shown in Figure 4.1-4. 

This amplifier is very similar to the K-la amplifier discussed above and has an open loop 

gain of 18,000 V/V to 100,000 V/V. The closed loop gain required for this amplifier is 

58.5 V/V for the inner axis and 5.05 V/V for the outer axis. The open loop gain variation 

will cause less than a 0.1 percent gain change for the closed loop operation of the inner 

axis amplifier. The gain change will be less for the outer axis amplifier. As in the case 

of the K-1 buffer amplifier, most of the gain variation can be assigned to the input and 

output impedance variations. 

4.1.3.5 Test Results 

The K-la amplifier has been breadboarded and subjected to the following tests: 

a. DC gain and linearity test 

b. Drift test due to temperature 

c. Noise rejection 

d. Open loop frequency response 

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The DC gain and linearity test was made with the amplifier load equal to 115 ohms. The 

test was run at three different temperatures: -20°C, +25°C and +70°C. The test data is 

summarized below: 

-20°C +25°C +70°C 

Gain 16.496 16.486 16.471 

Zero Offset - 0.0555 volts - 0.065 volts - 0.070 volts 

The DC gain obtained from this data is 1.655 V/V ±0.04 percent (144 ma ±0.04 percent) 

and meets the specification requirements. 

The temperature drift obtained from this data is less than ±10 my (+90 μ A) for temperature variations from -20°C to +70°C. Separate testing has not been performed on the K-lb 

amplifier. However, due to the close similarity of design, this amplifier can be expected 

to perform as well as the K-la type. 

A test was conducted to determine the effect of power supply line noise on the output noise 

of the amplifier. The following outputs were obtained when a 20 my signal was introduced 

on the negative power supply lines: 

f (cps) E

o 

(my) 

20 0.2 

60 0.4 

100 0.6 

200 1.4 

300 1.1 

400 1.0 

500 1.0 

600 1.0 

A test was conducted to determine the open loop frequency response of the amplifier and 

the results are shown in Figure 4.1-6. 

4-24 

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RELEASE 1 JULY 2015 SECREF SPECIAL HANDLING 

0 0 40.1 

••••■ 

am. 

0 

FREQUENCY (CYCLES/SEC) 

-; 

► 

o In 

N tr, 1•••• 

(Ha) Km) 

-S-Eaft SPECIAL HANDLING 4-25/4-26 
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RELEASE 1 JULY 2015 

SPECIAL HANDLING 

4.1.4 K-2 ELECTRONICS 

4.1.4.1 Description 

The K-2 electronics shown in Figure 4.1-7 control the analog current needed to drive the 

torque motor stator windings of the main drive. This circuitry consists of a compensation 

amplifier and a power amplifier for each of the two drive axes. The inputs to the compensation amplifiers come from the K-1 electronics, and the power amplifiers are used to drive 

the torque motors. 

The compensation amplifier is a 524AL microcircuit and the power amplifier uses discrete 

parts with a bridge circuit in the output stage. The load is 6.4 ohms for the outer axis 

and 19.4 ohms for the inner axis. 

4.1.4.2 Performance Characteristics (K-2) 

a. Gain - The following gains shall be provided in K-2 Electronics. 

K2IA Comp. Amp. 28.6 V/V 15 percent 

K2IA Power Amp. 0.0281 amps/volt ±5 percent 

K2OA Comp. Amp. 14.3 V/V 15 percent 

K2OA Power Amp. 0.132 amps/volt 15 percent 

b. Dynamics: 

K2IA Comp. Amp. 

K2IA Power Amp. 

K2OA Comp. Amp. 

K2OA Power Amp. 

(S/100 + 1) (S/40 + 1) 

(S/1000 + 1) (S/4 + 1) 

1 

(S/1000 + 1) 

(S/100 + 1) (S/40 + 1) 

(S/1000 + 1) (S/4 + 1) 

1 

(S/1000 + 1) 

4-27 

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IA MODE SELECT FROM K8 

SHAPED POSITION TK2 IA . 

IA GYRO OUTPUT 

ERROR FROM Kla 

(IA) 

GAIN 

AND 

COMPENSATION 

TORQUE 

MOTOR 

DRIVER 

OA MODE SELECT FROM K8 

SHAPED POSITION POSITION 

ERRO FROM Klb 

(OA) 

OA GYRO OUTPUT 

—Ow 

GAIN 

AND 

COMPENSATION 

TO OA 

TORQUE 

MOTOR 

TORQUE 

MOTOR 

DRIVER 

7 IC; OA 

_1 

Figure 4.1-7. K-2 Electronics Block Diagram 

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4-28 

TO IA 

TORQUE 

MOTOR 

K2 ELECTRONICS 
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The critical frequencies shall be met within ±10 percent. 

Additional critical frequencies shall not contribute more than 6° phase lag under 

100 rad/sec for the compensation amplifier and 6° below 2000 rad/sec for the 

power amplifier. 

c. Linear Range and Saturation - The gain and frequency values and tolerances stated 

in the above paragraphs shall be met within the range of operation indicated in the 

table below. In the power amplifier crossover region, the gain shall remain 

positive. 

K2IA Comp Amp ±17.5 Volts 

K2IA Power Amp ±(0.001 ± 0.228) Amp 

K2OA Comp Amp ±17.5 Volts 

K2OA Power Amp ±(0. 01 ± 1.19) Amp 

The load on the power amplifiers are: 

K2IA Power Amp 36 ohms ±10 percent 

K2OA Power Amp 7.6 ohms ±10 percent 

For the compensation amplifiers, the saturation level shall be determined at a 

level that will not damage the following stages. 

For the power amplifiers, the saturation levels shall be: 

K2IA 0.565 ±15 percent Amp 

K2OA 2.60 ±15 percent Amp 

d. DC Offset and Long Term Drift 

K2IA Comp Amp 500 my 

K2IA Power Amp 2 ma 

K2OA Comp Amp 50 my 

K2OA Power Amp 4 ma 

The long term drift applies to noise frequencies below 0.2 rad/sec. 

4-29 

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4.1.4.3 Stability Analysis (K-2) 

A stability analysis was performed on the K-2 power amplifier (Figure 4.1-8). The open 

loop gain of the amplifiers varies from 200 V/V to 600 V/V when driving a 7.6-ohm load. 

The closed loop gain required for this amplifier is 1 V/V (0.32 AN). The open loop gain 

variation will cause only a 0.35 percent amplifier gain change. Most of the amplifier gain 

variation can be assigned to the input and output impedance variations. 

The K-2 compensation amplifier (Figure 4.1-9) consists of a 524AL microcircuit amplifier 

which has a specified open loop gain of 1200 V/V to 2000 V/V. The closed loop gain 

required is 28. 6 V/V. The open loop gain variation will cause only 1.0 percent amplifier 

gain change. Most of the amplifier gain variation can be assigned to the input and output 

impedance variations. 

4.1.4.4 Test Results on the K-2 Electronics 

Breadboards were built for the compensation amplifier of Figure 4.1-9. The following 

tests were then conducted on the amplifier: 

a. Frequency response test 

b. Maximum output for 1 percent distortion 

c. DC offset versus temperature 

-30°C +25°C +80°C 

Output 

DC Offset 

(Input Grounded) -3.7 my -2.2 my -0.6 my 

d. Linearity test (see Figure 4.1-10) 

e. Rejection of power supply noise (see Figure 4.1-11) 

I SPECIAL HANDLING 

4-30 
a 

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co 

co 

co 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

to+28V UNREG 

cc 

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g 

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4-31 

—SEGREf SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

4-32 

ci I 

cq 

X 0 0,1 

III 

a, 

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--SERET-SPECIAL HANDLING 

Figure4. 1-9.K-2 Comp

ensationAmplifier 

-SICRET- SPECIAL HANDLING 
0 1-1 

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RCENT OF FULLSCALE INPUT 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 Ainfr SPECIAL HANDLING 

(Las,laq sacunom) llomia „Ladino INHOHHcl 

-SEC-Itff-SPECIAL HANDLING 

4-33 
(i-c1) MAI - amt.' incusio 

4-34 

SPECIAL HANDLING 

SPECIAL HANDLING 

0 0 

0 0 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

Figure 4. 1-11. 

FREQUE

NCY (

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ft 

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0 

C 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 'SUM- SPECIAL HANDLING 

A breadboard of the power amplifier (Figure 4.1-8) was constructed, and similar tests 

were conducted plus an additional test of the output noise. An output noise value of 0.5 ma 

was measured when the amplifier was connected to a 7-ohm resistive load. 

Finally, the two amplifiers were mated and the following tests were conducted on the total 

K-2 electronics. 

a. Linearity vs temperature 

-15°C +25°C +70°C 

Output 

DC Offset 

(Input Grounded) +0.46 my +0.57 my +0.70 my 

Gain (ma/volt) 11.7 11.6 11.3 

b. Rejection of power supply noise 

c. Output Noise measurement 

For this test, the combined amplifier was loaded by a 7-ohm resistor in series with a 75-

millihenry inductor. The input was terminated in a 1000-ohm source. The output noise 

was measured on an oscilloscope as being less than 1 my peak-to-peak. This is equivalent 

to approximately 150 microamperes at the output. 

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4.1.5 K-3 AND K-4 ELECTRONICS 

Description 

The K-3 and K-4 electronics (see Figure 4.1-12) provide control signal outputs to a dc 

torquer within the required gain, dynamic, and other performance characteristics outlined 

in the specification. 

Each,of the K-3 and K-4 electronics is comprised of a microcircuit compensation amplifier 

(see Figure 4.1-13) and a discrete part power amplifier (Figure 4.1-14). The compensation amplifier is a 524AL microcircuit and the power amplifier consists of two differential 

stages of gain followed by an output bridge which supplies current to a 51.3-ohm load. 

Feedback in the power amplifier is taken from the current sensing resistor to provide the 

high-source impedance required by a current amplifier. 

The K-3 compensating amplifier has unity gain and drives the power amplifier which has 

a gain of 0.418 amp/volt. 

The K-4 compensating amplifier also has a gain of unity and drives the power amplifier 

which has a gain of 0.0964 amp/volt. 

4 -3 7 

SPECIAL HANDLING 
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K3 INHIBIT FROM KO 

K-3 I 

D/A COMM'D 

FILTER 

SUMMER 

AND 

COMPEN 

•• POWER 

AMPL 

K-3 POSITION 

COMM'D 

(DIGITAL) TO K3 MOTOR 

ase 

POSITION FDBK (ANALOG) POSITION FDBK 

TO MDAU NO. 1 AND NO. 2 FROM K3 POSITION TRANSDUCER (ANALOG) 

K-4 

4-38 

K-4 POSITION COMM'D 

(ANALOG) 

K4 HI/LO SIGNAL 

TO MANUAL CONTROL 

STICK ASSEMBLY 

COMM'D 

FILTER 

SUMMER 

AND 

COMPEN 

POWER 

AMPL 

TO K4 MOTOR 

.---111■ 

POSITION FDBK 

FROM K4 POSITION TRANSDUCER 

Figure 4.1-12. K-3 and K-4 Electronics Block Diagram 

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0 

oo 

c 

0 

rYi 

GV 

CV 

--vvv—I I 

SPECIAL HANDLING 

F-4 

a, 

z 

Figure 4. 1-13. Comp

ensation Amplifier 

4-39 
L-1 

0-0 

0 0 C 

to 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

E 

0 

ti 

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0 0 

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SPECIAL HANDLING 

0 

4-40 
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4.1.6 K-5, K-6 AND K-7 ELECTRONICS 

Description 

The K-5, K-6 and K-7 Electronics shown in block diagram fashion in Figures 4.1-15 and 

4.1-16, are comprised of a number of logic controlled power switches. Some of these 

switches are bipolar and are used to drive motors bidirectionally. A circuit was designed 

(Figure 4.1-17) for bipolar power switch, and preliminary testing has been done on a 

breadboarded circuit. 

4-41 

SECRET SPECIAL HANDLING 
INHIBIT FROM KO 

COMMANDS 

K5 CW & CCW 

TRIP 

COMM'DS { FROM TEST SWITCH 

ON CONSOLE 

K6A 

TRIP TRIP 

LOGIC POWER 

SWITCH sr 

FROM SUN SENSOR COMMAND TO K6A TRIP 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

-SEM- SPECIAL HANDLING 

K5 

POWER LOGIC SWITCH 

TO K5 

MOTOR 

ex• AND 

BRAKE 

RELEASE 

FROM K5 CW & CCW LIMIT SWITCH 

olf 

BIVALUED DISCRETE POSITION 

INDICATION TO MCSA 

TO K6A ALARM 

(CONSOLE) 

Figure 4.1-15. K-5 and K-6A Electronics 

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K6B POSI TI ON 

COMMAND 

( FROM KO) 

K6B 

POWER C SWITCH 

TO K6B MOTOR 

CW AND CCW LIMIT SWITCHES FROM K6B 

PRI LAUNCH LOCK 

RELEASE (FROM LMMC)1 

BU LAUNCH LOCK 

RELEASE (FROM CONSOLE) 

K7 RELEASE 

COMMAND 

K7 FWD DISCONNECT 

COMMAND 

K7 AFT DISCONNECT 

COMMAND 

PRI K7 COMMAND 

(LMMC) 

BU K7 COMMAND (CONSOLE) 

K2 STOWED (FROM KO1 

110. 

K7 ELECTRONICS 

DRIVERS 

DRIVERS 

DRIVER 

DRIVER 

DRIVER 

FWD 

POWER 

LOGIC 

SWITCH 

AFT 

POWER 

SWITCH 

K2 LAUNCH LOCKS ► 

K7 LAUNCH LOCKS 

K7 RELEASE 

► 

K7 FWD DISCONNECT 

K7 AFT DISCONNECT 

IP■ 

K7 FWD MOTOR 

K7 AFT MOTOR 

K7 CLOSED 

TO LMMC K2 

POWER RELEASE 

FWD CW AND CCW LIMIT SWITCHES 

AFT CW AND CCW LIMIT SWITCHES 

K7 ALARM TO CONSOLE 

Figure 4.1-16. K-6B and K-7 Electronics 

SECRET SPECIAL HANDLING 

4-43 
Figure4. 1-17.K-5 Switch Circuit 

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4.2 MAGNIFICATION CONTROLLER (See also Section 1.3.3.8) 

Functional Characteristics 

The Magnification Controller, commonly called the left hand control stick (LHCS), provides 

analog and discrete signals to control ATS high and low range magnification, ATS zoom, 

selection of main optics magnification levels, switch the MCS to control main optics, ATS 

finger-operated voting switches, and cue hold command. 

Operation 

Analog signals are generated by the LHCS to provide ATS zoom control. Movement of the 

stick through its range causes zoom operation. Upon reaching the midpoint of the ATS control range, a discrete signal commands ATS high range, and the zoom control resets to the 

lowest zoom position. Further movement of the stick causes ATS zoom to 127x. 

Upon reaching ATS 127x, the stick can be moved to the left, generating a discrete which 

commands MCS main optics control. The stick can then be moved to one of four positions 

in detents, which command the four main optics magnification levels. 

Finger-operated switches on the stick provide target reject, inactive, and active votes. 

Twisting the stick (20 degrees or less) generates a cue hold discrete command. 

Status 

The Magnification Controller system requirements have been defined. A preliminary design 

concept has been generated and a brassboard is being constructed for use in the R-38 simulation facility. Flight crew comments will be utilized in a redesign of the initial concept. 

This component is barely out of the paper design stage. 

4-45/4-46 

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4.3 VENDOR SELECTION TRADE-OFFS AND CRITERIA FOR THE ATS RATE GYRO 

The following factors influenced the selection of Honeywell as the gyro vendor: 

a. The Drive A gyro had been under contract for about four months, and we wished to 

have as much commonality with it as possible to minimize costs. 

b. The gyro transfer function and PSD envelope were specified by the N.G. & C. 

Requirements and Analysis Group, and were different from those of the A gyro. 

c. The thermal environment was specified by the Systems-Thermal Analysis Group 

as 0 to 80°F (as opposed to 60 to 80°F for the A gyro). 

d. The allowable space envelope was specified by the N.G. & C. Subsystems 

Engineering group. 

Thus, the problem became one of trying to resolve items b, c , and d with item a. Item a 

dictated that the contract be placed with Honeywell and utilize the same basic instrument 

as specified for the A gyro. However, our previous fact-finding on the A gyro, completed 

in January, had shown that Honeywell was the only vendor who had a suitable "developed" 

gas bearing gyro, and was willing to accept the MOL requirements. Thus, there was 

really no other vendor to consider. (A gas bearing gyro is required because of noise and 

reliability considerations.) 

Given the fact that space to mount a single package was not available, it was determined 

that two smaller packages could be made to fit the space available. A natural way to divide 

the component was to have the gyro in one package and the support electronics in another, 

connected with an interconnect cable. 

With a much larger dynamic temperature environment (0 to 80°F) than that of the A gyro, 

a different thermal design was required. This took the form of raising the gyro temperature to the 160-170° range, the maximum desirable from reliability considerations. This 

higher operating temperature allows power savings over the Drive K environmental range 

compared with a design operating at a lower temperature. 

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4-47 
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The transfer function requirements being different, and the operating temperatures of the 

A and K gyros having to be different because of their respective environments, if the same 

fluid were used in both gyros, the gyros would require different damping gaps to achieve 

the same damping coefficient. Alternately, different fluids could be used with the same 

damping gap in each gyro. This was preferred from cost and commonality considerations, 

and simplified the design of the gyro with respect to its transfer function. The noise requirements were also different. Fortunately, in both cases, the gyros appear to be able 

to meet the noise requirements, and work with Honeywell is underway to use a common 

method of testing. 

4 -48 SECRET SPECIAL HANDLING 
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4.4 ALPHA SUBSYSTEM POWER REQUIREMENTS (See Also Section 2.5) 

Figure 4.4-1 depicts the power flow within the alpha subsystem. The diagram shows the 

power flow for one manned station. Two identical stations are provided per subsystem. 

The diagram is drawn to highlight the power across contracted interfaces, namely Drive 

K Electronics, Itek (Sub A), Visual Display Projector and Gyroscopes. In each box is 

given the standby and peak power dissipated. The boxes in the Drive K Electronics signify 

major functional modules; the boxes in the Itek (Sub A) signify actuators, sensors, displays, 

or feedback transducers as defined by the legend. In the gyro, its mechanical elements, 

power conditioner (p.c.), electronics, heaters and heater controls are indicated; the Visual 

Display projector lists the power dissipated in several significant modes. 

Table 4.4-1 tabulates the information from Figure 4.4-1 into four basic Alpha Subsystem 

Modes and mission average. The modes are Shroud Open, Slew, Track and Shroud Close 

modes. This represents a preliminary evaluation of the development status of the subsystem versus its allocated requirement not to exceed 480 watts peak or 70 watts average. 

The estimated figures from the table are 458.82 watts and 51.99 watts, respectively. The 

peak power allocation can be attained by selective inhibiting of the peripheral display, and 

gyro heaters during slew. Since the performance is expected to be within the allocation, no 

such recommendation is made at this time. 

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TABLE 4.4-1. ALPHA SUBSYSTEM PEAK AND AVERAGE POWER SUMMARY SHROUD OPEN 

a 

w 

a

a

l 

0 

g 

54 

W 

0 

0. 

MISSION AVERAGE 

0 WATTS 

13.0 1.3 0.42 

40.0 4.0 1.29 

5.0 0.01 

5.0 0.01 

22.1 0.03 

11.0 0.70 

16.6 16.6 0.02 

9.0 9.0 0.62 

0.5 0.5 

0.5 0.5 

0.01 0.01 

0.01 0.01 

5.0 5.0 0.29 

2.0 2.0 0.42 

61.5 3.12 

8.5 8.5 0.83 

5, 5 5.5 5.5 1.08 

2.5 2.5 2.5 

4.0 4.0 

69.5 20.5 20.5 5.03 

36.9 36.9 

1.7 1.7 

1.7 1.7 

1.7 1.7 

17.3 5.4 3.69 

2.19 

2.19 

2.1 

1.0 

1.6 1.6 

7.2 7.2 

38.0 38.0 8.42 

152.21 229.41 199.32 18.21 

25.99 

458.82 

468.0 

51.99 

70.0 

Torque Motor IA 

Torque Motor OA 

Prism Actuator 

Zoom Actuator 

Power Changer Actuator 

Sun Shield Motor 

Shroud Actuators 

Encoders (2) 

Prism F.B. 

Zoom F.B. 

Power Changer F.B. 

Sun Shield F.B. 

Peripheral Display 

Sun Sensor 

Gyro Warm Up Heater IA 

Gyro Control Heater IA 

Gyro Electronics IA 

Gyro Spin Motor IA 

Gyro Torquer IA 

Gyro, Electronics, Heater OA 

Drive K Electronics KO, K3, K4, K5, K6, K7, K9, K10 

DKE K2 Modules IA 

DKE IA Torquer Power Amplifier 

DKE K2 Module OA 

DKE OA Torquer Power Amplifier 

DKE Power Amplifier, Prism 

DISE Power Amplifier, Zoom 

DKE Power Amplifier, Power Changer 

DICE Power Amplifier, Sun Shield Motor 

DKE Power Amplifier, Shroud 

DKE Power Amplifier, (2) gyro torquers 

Visual Display Projector 

Totals, Peak per Crew Station 

Total, Average per Crew Station 

Worst Case Peak per System 

Worst Case Peak Allocation 

System Average 

System Average Allocation 

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4.4.1 ALPHA WEIGHT STATUS 

The weight bogies for the subsystem were negotiated in October with exception of the Acquisition Optics components. 

These negotiated figures are shown in the table below along with the present status numbers. 

The Drive K electronics weight has grown significantly and may exceed 3.5 lbs. MDAU 

weight deltas are also not reflected in the following tabulation. 

ALPHA WEIGHT DATA - FOR GE ITEMS ONLY - (2 SUBSYSTEMS) 

CURRENT NEGO* 

Gyros and Electronics 20.2 22 

VDP (Excluding GFE Modules) 45.4 50 

Control Sticks (Left Hand) 5.0 5 

Harness 42.0 42 

Headrest 5.0 5 

Thermo Insulation 3.0 3 

Support Structure 13.0 13 

Console Signal Controller (Delta) 6.0 6 

Power Controller (Delta) 5.0 5 

Console Signal Controller - Per. Display 6.0 6 

Peripheral Display Harness 4.0 4 

Alpha Numeric Harness 4.0 4 

Total (lbs. ) 158.6 165 

*Negotiated Weight as of October, 1967, 

except Drive K electronics 

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SECTION 5.0 

CONSOLES AND DISPLAYS 

This section describes the present layout of the consoles and displays panels. The material 

is subdivided into two categories; Consoles and Displays Requirements and Panel Layout. 

5. 3 "B" Panel Controls and Displays 

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5.1 CONSOLES AND DISPLAYS REQUIREMENTS 

The CEI requirements relevant to the consoles and displays subsystem and the aquisition 

subsystem are: 

a. Provide the controls, status indicators, and alarm initiating signals as required 

for proper operation, monitoring, and control of the GE AVE. 

b. Provide dual visual displays, cue displays, and controls on each reconnaissance 

console as required for proper operation of the acquisition group and the tracking mirror. (This requirement for dual visual displays has been changed to 

reflect the current acquisition subsystem design of a single telescope and pointing assembly at each console.) 

5.2 PANEL LAYOUT - C & D Panels 

Figure 5-1 depicts the current operational panel layout for the acquisition subsystem while 

Table 5-1 lists all the controls and displays on Panels 8C and 8D. 

..■•••••■ 

Those proposed controls and displays which are related to the operation of just the Alpha 

Subsystem and a description of their functions are as follows: 

* 1. ATS Track Light - Alerts the crew that the ATS is in a rate track mode and to 

begin ATS tasks. 

* 2. ATS Telescope 2Provides for inspection of targets for weather coverage, activity 

status and optical feedback for manual steering control of the ATS telescope 

tracking mirror. 

*Indicates that this component is a direct part of the Acquisition Subsystem. 

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TABLE 5-1. LIST OF CONTROLS AND DISPLAYS 

* 1. ATS Track Light 

* 2. ATS Eyepiece Location 

3. MO Track Light 

4. MO Eyepiece Location 

5. Zero Reset 

6. Computer Update 

7. Position Rate Control Stick 

* 8. Cue Display Screen 

9. Rate Manual Switch 

10. Exposure Adjust Switch 

*11. Cue Brightness 

12. X Format Inhibit Light Switch 

*13. Cue Execute 

*14. Single Frame Step 

15. Manual Shutter 

*16. Manual Cue Selector 

17. Preferred Photo 

18. Wrong Module 

19. End of Reel - (Deleted) 

*20. Cue System Electrical Power Control (Mode Selector) 

21. Cue Test - (Deleted) 

*22. ATS Hold Switch 

*23. Cue Focus 

*24. Cue Film Module Insert 

*25. Magnification Control Stick (Voting Switches and cue hold) 

26. Position Mark 

27. Shutter Mode 

*28. Peripheral Display Brightness Control 

*29. a - N Display (Cue) 

*30. Track Path Switch 

31. Camera Selection Switch 

32. MO Brightness Control 

33. IVS Saturate Light 

34. Time to Next Group 

*35. Target Number 

36. Timing Setting Control 

37. Operate Switch 

38. Numeric Display (Event Timer) 

39. Mode Switch 

40. Master Caution 

41. Master Warning 

42. a -N Brightness Control 

*Aquisition S/S related items. 

5-4 

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Figure 5-1. Panels 8C and 8D Component Configuration 

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ATS Peripheral Display - Tentative Functions include: 

a) An non-override light (Group 2 or 3) indicates the target being taken by the MO 

is a non-override target. 

b) A primary light (Group 2 or 3) indicates the target in the ATS FOV is the 

primary target of the target group. 

Groups 2 and 3 display coded patterns utilized for target identification and 

information. 

d) A visual optics assigned light (Group 2 or 3) indicates the MO is assigned to 

this console and final target selection has been made and tracking has commenced. 

e) A decision time wipeout indicates the time remaining to the end of decision 

time for the target group. 

1) Lights in the decision time wipeout indicate the recommended dwell time for 

target viewing prior to slew to the next target. 

7. Deleted 

* 8. Cue Display Screen - Provides a projection surface for the cue display. 

*11. Cue Brightness - Provides manual control of the image light intensity projected 

on the display screen. 

*13. Cue Execute - Initiates the manual cue selection process. 

*14. Single Frame Step - Allows individual frame stepping, in either direction, without 

resetting the manual cue selector. 

*16. Manual Cue Selector - Provides a manual capability to select any desired cue. 

*20. Cue System Electrical Power Control - Provides control of the cue display 

input power and selection of the VDP control mode. 

*22. ATS Hold Switch - Provides a manual ATS target hold capability beyond the 

normal decision time and delays subsequent ATS slewing the next target group. 

* Indicates that this component is a direct part of the Acquisition Subsystem. 

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*23. Cue Focus - Adjusts cue focus for maximum clarity. 

*24. Cue Film Module Insert - Positions the cue film module in the VDP. 

*25. Magnification Control Stick 

a) Provides Main Optics and ATS telescope magnification control. 

b) Provides a control stick gain signal proportional with telescope magnification 

of the ATS and the Main Optics in order to achieve a constant eyepiece scene 

rate throughout various powers of magnification. 

c) Provides three logic voting switches for computer analysis and final target 

decision. 

d) Provides a cue hold function. 

*28. Peripheral Display Brightness Control - Provides manual light intensity control of 

the ATS peripheral displays. 

*29. Alphanumeric Display - Displays information not provided by other convenient 

displays, which are pertinent to the target. 

*30. Track Path Select - Provides a manual selection of the ATS track path for 

the respective console. 

*35. Target Number - Provides a ready identification to the target presently displayed 

in the ATS FOV for crew reference and comment identification. 

*42. Alpha-Numeric Brightness Control - Provides manual control of the a- N 

brightness. 

As part of the operational controls associated with the ATS telescope eyepiece itself, a 

filter control is provided for image contrast enhancement. This filter wheel has four 

sections for neutral density or spectral filters and is controlled by a pointer knob on the 

side of the eyepiece. 

In addition, the derotation mechanism, zoom lens, low power groups, and blanking shutter 

have backup override controls which minimize the possibility of a catastrophic failure mode. 

* Indicates that this component is a direct part of the Acquisition Subsystem. 

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These controls are behind the panel and have a requirement that they be operated with a 

pressurized gloved hand (Located behind 8C and 2C). 

The derotation mechanism has a manual override and index which enables the astronaut to 

manually position the Pechan prism within plus or minus five degrees of the neutral or 

central position in the event of a control signal or drive motor failure. This is a 1-5/16 

inch round knurled knob located behind the panel and must be turned several times to position 

the prism mounts (Located behind 8C and 2C). 

The zoom lens has a manual override which enables the astronaut to manually change from 

low zoom to high zoom or vice versa in the event of a control signal or mechanism failure. 

The control itself is a two position pointer knob located behind the panel to rotate the gears 

which position the lens (Located behind 8C and 2C). 

The magnification (power) change mechanism has a manual override capability which enables 

the astronaut to manually change from low power to high power and vice versa, in the event 

of a control signal or drive motor failure. The control is a key type of device which is 

turned to the desired position by turning a gear train which positions a cam controlling the 

lens position. Access to this control is from Douglas Console 3C and 7C because of its 

location in the system. 

The blanking shutter has a manual override which enables the astronaut to manually remove 

the blanking devices from the optical path between the target scenes and eyepiece. This 

will be an operation in the event of a control signal or mechanism failure which prohibits 

display of the target image. The control is a round knob which is turned 90 degrees to reset 

the shutter and is located behind the panel (Located behind 8C and 2C). 

The remaining control which has a direct astronaut interface is the launch locks which are 

accessible from Douglas' Console 3C and 7C. These launch locks hold the lens in a folded 

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position and must be released before the mission can be performed. There are also launch 

locks located outside the Lab Module and are activated by pyros. This may be initiated 

either by commands or manually. 

5.3 "B" PANEL CONTROLS AND DISPLAYS 

The following controls and displays have been identified for the Alpha subsystem on panels 

8-B, and 2-B. These items are not required for normal ATS operations, but are made 

available for equipment checkout and contingency operations. 

1. Controls to: 

a. provide backup stow capability for acanners (each STS) 

b. provide manual backup to drive scanners to sync position (each axis of each 

ATS) 

c. select between redundant encoders (each ATS) 

d. select between redundant control loops (each axis of each ATS) 

e. enable/disable scanner amplifier power (each ATS) 

f. enable/disable derotation drive power (each ATS) 

g. enable/disable zoom drive power (each ATS) 

h. enable/disable magnification drive power (each ATS) 

i. test blanking shutter (each ATS) 

j. enable/disable blanking shutter power (each ATS) 

k. enable pyro circuits for door failure mode correction (each ATS) 

1. fire pyro circuits for door failure mode correction 

m. provide back-up control to drive door fully open or closed (each ATS) 

n. mechanically disengage either/both door drive motors, an irreversible 

process (each ATS) 

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o. electrically select/disable door drive motors (each motor) 

p. actuate door spring-release mechanism to swing door free from scanner 

interference, irreversibly (each ATS) 

2. Display to indicate: 

a. scanner is stowed, pitch and roll (each ATS) 

b. analog position of door (one side): each ATS 

c. door is fully open or closed, (each ATS) 

d. scanner has achieved sync position (each axis of each ATS) 

e. door is not operable (each ATS) 

5-10 

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SECTION 6.0 

RELIABILITY, CONTAMINATION CONTROL AND MAINTAINABILITY 

The material in this section is subdivided as follows: 

6.1 Reliability Requirements and Evaluations 

6.2 Contamination Control 

6.3 Maintainability 

6.4 Eye Hazard Analysis 

6.1 REQUIREMENTS AND EVALUATIONS 

6.1.1 SUMMARY 

Based upon the Mission Reliability design criteria specified in Section 6.1.3, the following 

conclusions were reached: 

a. The system design concept appears adequate to meet the GE-AVE mission reliability 

requirements. 

b. The crew has been utilized to perform backup operations to the automatic equipment which is in agreement with the GE-AVE Reliability Policy. This should be 

pursued further to assure that this backup capability is optimized and that maintainability through repair or interchangeability of ATS-1 and ATS-2 components is 

possible. 

c. Additional studies appear necessary to assure that the design provides adequate 

crew protection for malfunctions in the sun blanking control functional assembly. 

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6.1.2 SYSTEM RE LIABILITY REQUIREMENTS 

The ATS system reliability requirements have been developed utilizing the SPDR (SS-MOL1A) and the GE-AVE CEI (CP1000A1) as the guideline documents. Following is a summary 

of the pertinent system reliability statements: 

a. SPDR, paragraph 3.3.4.3.3.11 states essentially that; 

1. The design shall be such that after any single failure, the navigation and control system shall be capable of supporting a productive mission. 

2. Redundancy shall be utilized and it shall be a design objective that this redundancy be achieved with minimum additional equipment configured so that 

failure may reduce capability but not abort the mission. 

b. CEI - CP1000A1, paragraph 3.1.2.1 states: 

1. "The probability that the GE-AVE will meet the requirements of the mission 

shall be 0.97 (MA) If 

2. "Where practicable, the GE-AVE shall be designed so that after any single 

failure, the GE-AVE shall be capable of supporting a productive mission. " 

3. Failure criteria for the evaluation of the system GE-AVE reliability are 

defined in terms of the primary mission objectives only; i. e. a failure to meet 

primary mission objectives is defined as: 

(a) Photographic mission terminated prior to 30 days 

(b) Less than 90 percent of the planned photographs obtained due to a GEAVE malfunction 

(c) Degradation in the GE-AVE that causes a degradation in resolution greater 

than 10 percent on more than 10 percent of the ZI horizontal measurement 

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6. 1. 3 MISSION RELIABILITY DESIGN CRITERIA 

Based upon these GE-AVE system level reliability requirements the following mission 

reliability design criteria have been established for the ATS subsystem: 

a. The primary mission objectives are achievable with one of the two ATS subsystems 

performing in accordance with specification requirements. (Note: To minimize the 

MOL effectiveness degradation due to loss of a single ATS, it is desirable to be 

capable of putting all primary targets on high priority targets as one ATS system. 

This is the current plan. ) 

b. The probability that one of the two ATS systems is performing, within specification 

limits, shall be at least 0. 999. 

c. A reliability requirement for each ATS system has been established as 0. 980 for 

the mission. This appeared attainable in the early design stages without utilizing 

block redundancy. (Note: Each ATS could be degraded to a 0. 968 reliability 

prediction and still meet the system requirements expressed in b. above. ) 

d. Block redundancy within each ATS subsystem should be eliminated to minimize 

weight if its only purpose is to achieve the 0. 980 reliability requirement. Block 

and/or functional redundancy within each ATS shall be utilized as required to meet 

the system level requirement of 0. 999. 

e. In the M/A mode, a functionally redundant operating mode exists that can negate 

the effects of a complete loss in the ATS function, i. e. the Automatic Mode with 

or without the crew utilizing the main optics eyepiece. This functionally redundant 

mode will permit achievement of a large percentage of the MOL objectives; however switching to the Automatic Mode due to equipment failure causes is considered 

as a failure of the GE-AVE M/A system to meet its primary mission objectives. 

f. The design shall be such that no single failure will cause the loss of both ATS 

subsystems. 

g. Provisions shall be made to utilize the crew to perform backup control, perform 

maintenance (interchange similar hardware), and perform diagnostic analysis 

where system effectiveness is improved without undue utilization of the crew's 

time. 

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6.1.4 DESIGN EXCEPTIONS 

The design meets all of the above requirements except the individual reliability apportionment of 0.980 to each of the ATS subsystems (predicted to be 0.975 without degradation and 

0.988 utilizing manual backup capabilities). The reliability data presented in the next section, when summarized at the ATS system level, establishes that: 

a. The probability of completing the mission with both ATS meeting specification 

requirements is 0.951. 

b. The probability of at least one of the two ATS systems meeting specification requirements is 0.9993. 

c. The probability of at least one of the two ATS systems operable with degraded 

performance, is approximately 0.9999. 

6.1.5 ACQUISITION SYSTEM RELIABILITY ESTIMATE 

The reliability apportionment to the Acquisition Subsystem is based upon satisfactory performance with one of the two sets operative and is 0.999. The analysis, as reflected in 

Table 6.1-1, presents the prediction for each set and that for operational redundancy. 

The estimate for the Optical Assembly, as shown, reflects prime configuration operation 

without manual backup. With manual backup, the prediction is increased from 0.9960 to 

0.9981. As presented in Table 6.1-1, the analysis does not reflect the planned capability 

of switching out and by-passing the gyro's in the event of failure. The gyro's and gyro 

electronics represent a significant portion of the failure risk, principally due to the longer 

operating time with a 12-minute warm-up period prior to each operating period. 

The failure rate applied to the Drive K Electronics reflects an upward scaling related to the 

recent increase in failure rate for the Drive A Electronics. In the case of the gyro and gyro 

electronics, the failure rate applied is that provided in the subcontractor's preliminary 

RFMA. It is considered to be conservative. 

-S-EGRR SPECIAL HANDLING 

6-7 
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TABLE 6. 1-1. RELIABILITY ESTIMATE ANAL

YSIS 

SPECIAL HANDLING 

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SPECIAL HANDLING 

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The failure rate applied to the Visual Display Projector presents a more conservative 

figure, 16.3 failure per million hours, as compared to approximately 8 fpmh provided by 

Lear Siegler in their proposal in response to RFP No. 134A. 

The predicted inherent reliability of the subsystem, even with the degree of conservation 

included, indicates that the subsystem apportionment will be met. 

6.1.6 SYSTEM FAILURE MODE AND EFFECTS 

6.1.6.1 Introduction 

This section summarizes data that is pertinent to the system reliability of the Acquisition 

Tracking Function. The information is presented in subsections 6.1.6.2 through 6.1.6.4. 

6.1.6.2 Description of Function 

This subsystem consists of a pair of high performance telescopes; one for each astronaut, 

with a variable magnification range. The magnification is a minimum of 16X and a maximum of 127X. Each telescope has the capability for providing to the astronaut a scan of 

the ground before, behind and on both sides of the orbiting vehicle to enable him to enhance 

the performance of the main optics. Included in this subsystem are the control loops 

necessary to perform the following functions: slew to a programmed target control tracking 

of ATS optics, orient target image to align with vehicle velocity vector, control the magnification of the optical system, provide automatic sun blanking control for astronaut eye protection and control opening and closing of external door which protects scanner and folding 

mirror when not in use. Also included in the Subsystem is a Visual Display Projector which 

has the primary function of displaying, either automatically or upon astronaut command, 

stored cue information to assist the astronaut in locating and recognizing preselected 

targets. 

6-9 

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6 -10 

SPECIAL SUBSYSTEM - RELIABILITY BLOCK DIAGRAM 

Figure6. 1-1. ReliabilityBlo

ck Diagram 

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Figure 6.1-2 relates the interrelationships between the ATS, the astronauts and the other 

MPSS subsystems. It provides a pictorial presentation of the signals and commands between the subsystems. 

6.1.6.3 System Reliability Block Diagram 

Figure 6.1-3 presents the reliability logic of the ATS at the system level. It identifies the 

baseline operational paths to perform the ATS functions. Figure 6.1-4 presents the logic 

of one ATS system. In subsystem 6.1.6.4 more detailed logic diagrams are presented for 

each of the ATS functional assemblies. 

6.1.6.4 System Failure Mode and Effects Analysis 

Figures 6.1-5 through 6.1-16 in conjunction with Tables 6.1-2 through 6.1-10 provide a 

preliminary failure mode and effects analysis on a component input/output basis. The 

approach has been to analyze functional groupings of equipment regardless of where they 

appear in the subsystem component listings. The analysis is subdivided into the following 

functional assemblies: 

Scanner Controls and Drives 

Image Orientation 

Zoom and Magnification Control 

Sun Blanking 

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Scanner Passive Elements 

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TABLE 6. 1-8. 
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6-42 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING I Detection I 3 

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--S-E-RE- SPECIAL HANDLING 

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TABLE 6. 1-8. CUE SYSTEM (Cont) 
TABLE 6. 1-9.SCANNER PASSIVE ELEMENTS 

6 -44 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 —SECRET— SPECIAL HANDLING 

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—SEM- SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

---S-EGREE SPECIAL HANDLING 

Detection 

MAS/Crew 1 TLM 

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6- 5 

TABLE 6. 1-9. SCANNERPASSIVE ELEMENTS (Cont) 
TABLE 6. 1-10.OPERATIONAL INTERFACES 

G —4 6 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING Detection I MAS/Crew I 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING Detection 

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SPECIAL HANDLING 6-47/3-48 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

6.2 CONTAMINATION CONTROL 

6.2.1 PHILOSOPHY 

Performance characteristics of optical systems such as the Visual Display Projector and 

the Acquisition Tracking System are specified in terms of resolution, obscuration, etc. 

To meet the specified performance requirements it is necessary to fabricate and assemble 

the equipment under conditions that control the contaminants which affect the final optical 

properties, to package the assembled item in such a way that contamination of critical 

surfaces can be controlled when the assembled unit is exposed to an uncontrolled environment, and to provide for local cleaning and area environmental control when it is necessary 

to open an assembly for maintenance or installation. 

General Electric's philosophy is to define the performance requirements to the vendor, 

work with him in establishing facilities and procedures necessary to meet the requirements, and to provide in-house equipment and procedures to assure that contamination 

control requirements defined during manufactur will be maintained. 

6.2.2 VISUAL DISPLAY PROJECTOR (VDP) 

Contamination Control of the VDP has been specified for critical assembly areas to be 

supplied with air that meets FED STD 209A, Class 10,000, that the unit be capable of 

being externally cleaned so that it may be opened for maintenance work. Normal use of 

the VDP will be in an uncontrolled environment. Therefore, breathing ports are provided 

with filters the control entrance of contaminants while dust covers are required to protect 

the eye piece when it is not in use. 

Packaging of the VDP for shipment will include use of plastic bags for contamination control. The same procedure will be employed during storage of the unit. 

6-49/6-50 

-iRkff- SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

A review of all contamination control procedures for the VDP from factory-to-launch will 

be conducted by the Contamination Control Board to assure that consistent, integrated 

level of control is maintained. 

6.2.3 ACQUISITION TRACKING SCOPE (ATS) 

The Acquisition Tracking Scope (ATS) presents its own problem for contamination control 

in as much as the unit must be subjected to partial disassembly for installation into the 

Laboratory Module (LM) and the external mirror will be exposed to the room environment 

during testing and alignment at DACO. 

To obtain the performance required for the ATS System GE is working with the manufacturer to utilize his experience in obtaining the necessary clean room environments, now 

estimated to be FED STD 209, Class 100,000, and operating procedures. Packaging requirements for shipment will be developed which will maintain the clean condition of the 

unit from the manufacturer to the point of equipment installation into the LM. Installation 

of the ATS into the LM will be accomplished at DACO in an Environmentally Controlled 

Sheltered Area supplied with air that meets FED STD 209A, Class 100,000 as defined by 

SAFSL 10003 with oils and volatiles uncontrolled. It is planned that a portable clean tent 

that has a FED STD 209A, Class 10,000 air supply with oils and volatiles reduced to: 

NO2 0.03 ppm 

SO2 0.01 ppm 

Oxidents 0.01 ppm 

Hydrocarbon 0.03 ppm 

will be available at DACO and VAFB. Acetylene, Entane, Ethylene and Methane are not 

controlled. 

6-51 

Ifelifi SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SPECIAL HANDLING 

An Environmental Conditioning Blower unit, normally used to supply class 10, 000 air with 

oil and volatiles controlled to the Mission Module (MM), will be available to supply clean 

air to local areas if needed. 

For extended test periods at DACO that require the external optics to be exposed to the 

class 100, 000 room local plastic covers may be used to protect the mirror from excessive 

contamination during these periods. At other times the external cover will be kept in 

place to provide maximum protection of the mirror. 

Contamination Control at VAFB is maintained by: (1) having FED STD 209A, Class 100, 000 

air with oils and volatiles uncontrolled supplied to the MOL Environmental Shelter which 

encloses the LV; (2) having a portable class 10,000 clean tent with oils and volatiles controlled available for use, and; (3) by keeping the system covered when not being tested. 

6-52 

-IfOET-- SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 -IfeM- SPECIAL HANDLING 

6.3 MAINTAINABILITY 

6.3.1 CONCEPT 

6.3.1.1 Ground 

All of the subsystem components will be maintained on pad in the event of a failure with the 

exception of the optical assembly, the folding mirror and the support structure. Maintenance of the first two components require the use of the alignment facilities not presently 

planned for at VAFB. If either of these components fail, the present maintainability concept calls for shipment of the LM back to the factory for repair. Provisions for maintaining the support structure at VAFB will not be made since the probability of failure and the 

probability of failure detection are extremely low. Since the predicted failure rate for 

this component is 0.0001, non-detection of potential failure at the launch pad is not considered a serious problem. All other components will be repaired by removing the faulty 

one and replacing it with a good component. The faulty component will then be shipped, in 

most cases, to the factory for repair and then be cycled back as a spare. 

6.3.1.2 On-Orbit 

On-orbit maintenance is currently limited to switching to redundant equipment. The 

possibility of including provisions for some minor replacement of alpha subassemblies 

is under study. 

6.3.2 ANALYSIS 

In this analysis, detection of a failure is, in some cases, limited to those periods of time 

where a ground crew member has visual access to the particular component or when TLM 

is being received by the CITE. In cases where there is not TLM or visual access, such as 

some hidden structure, no failure detection is normally possible. 

6-53 

SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 -Sfei-ET- SPECIAL HANDLING 

Repair times given are the total times required to correct a failure. Time includes obtaining spares, AGE and test equipment, removing the faulty component, installing and 

validating the good one, and restoring the system to its original state. 

a. The two VDPs, located in console Nos. 2 and 8, can be replaced on pad. The 

exact operations to be performed depend upon details not yet firmed up. Consideration is being given to operations associated with disconnection of cold plate 

plumbing installed in the VDPs. This problem area will have to be resolved before a complete analysis of VDP removal can be made. Assuming a minimal 

amount of this type of operation, the VDPs can be removed, replaced and revalidated in 6 to 11 hours depending upon where in the launch cycle the failure 

occurs. Detection of a failure is limited to prior to T-120 minutes. 

b. The harness penetration set, a high reliability item (0.9999) may be repaired in 

8 to 16 hours. Detection of a failure is possible till T-10 minutes. 

c. The Drive K electronics contained in the consoles may be repaired in 6.5 to 18 

hours. Detection of a failure is possible until T-10 minutes. 

d. The shroud door and drive are replaced as a unit. Detection of a failure is 

possible prior to T-120 minutes. Maintenance times range from 10.6 to 14.7 

hours. 

e. The headrest on the console is highly reliable (0.9999). Detection of a failure is 

limited to prior to T-120 minutes. Repair can be accomplished in 5 to 10 hours. 

f. The magnification control stick may be removed and replaced if repair is required. 

Detection of a failure however is limited to prior to T-270 minutes. Maintenance 

time required is seven hours. 

6.3.3 EFFECTS OF MAINTAINABILITY ON OTHER SUBSYSTEMS 

At present, it appears that the design and procedure associated with the Alpha subsystem 

does not degrade the maintainability of other subsystems. A study is under way to insure 

that the lack of ability to remove the optical assembly does not prohibit maintenance of 

other subsystems. 

6-54 

-Sfe-Rf-T-SPECIAL HANDLING 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 RET SPECIAL HANDLING 

6.3.4 SUMMARY 

With the exception of the optical assembly and the folding mirror, ground maintenance 

activities can be performed with a minimum of problems and delay in time on the components of the Alpha Subsystem. 

The optical assembly has a high enough failure rate (0.9957) such that methods of repairing 

it on pad will be studied in order to increase the probability of meeting the launch window. 

The folding mirror is a highly reliable item (0.9999) and should present little problem. 

The same holds true for the structural members. 

TABLE 6.3-1. CHARACTERISTICS AFFECTING MAINTAINABILITY FOR ALPHA 

SUBSYSTEM COMPONENTS MAINTAINED ON PAD 

COMPONENTS APPROX. SIZE 

(INCHES) 

WEIGHT 

(POUNDS) QUANTITY ACCESS 

Support Structure N/A N/A N/A 

VDP 18 x 8.5 x 8 25 2 Console Nos. 2 and 8 

Harness N/A N/A (later) (later) 

Drive K Electronics 9 x 17 x 5 17.5 2 Console Nos. 2 and 8 

Shroud Door and Drive 36 x 48 x 12 (later) 2 LM Exterior 

Headrest 6 x 7 x 2 3 2 Console Nos. 2 and 8 

Magnification Control 7 x 7 x 10 6 2 Console Nos. 2 and 8 

Stick 

SECRET SPECIAL HANDLING 

6-55/6-56 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 SECRET SPECIAL HANDLING 

6.4 EYE HAZARD ANALYSIS 

A "ball-park" analysis has been performed to approximate the level of irradiance produced on the cornea and retina of the eye when viewing the sun spectral reflection of the 

sun, and certain other objects. It was the purpose of this analysis to determine if a sunblanking device was required for the ATS for protection against direct solar viewing, and 

also to attempt to determine if specular reflections of the sun from surface water presented 

a hazard. 

The direct viewing of the sun with the ATS was found to produce extremely high irradiance 

of the cornea of the eye which would certainly inflict damage. However, the nature of the 

MOL orbit and the hardware restraints on the scan field of the ATS make it virtually impossible to view the sun if the OV is earth stabilized. An automatic blanking shutter is 

thus required generally only as a backup for those cases where the OV is not earthstabilized and when the ATS is commanded to point away from the earth. 

The effect of viewing specular reflections of the sun from surface water was not so conclusively determined. The specular reflectance of water is very much dependent on the 

angle of incidence and the surface roughness of the water. Using a very conservative 

(high) value for surface reflectance (five percent) the irradiance level on the cornea was 

computed to be 0.2 watt/cm2

. The effect of such an irradiance level on the cornea has 

not been determined. 

To meaningfully assess the hazard presented by specular reflection from surface water, 

better data is needed to (a) determine the damage threshold of the eye in terms of irradiance levels and exposure times, and (b) to determine reasonable reflectance values for 

large bodies of surface water. 

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6. 4. 1 SUMMARY OF CONCLUSIONS 

a. The most damaging concentration of energy when viewing the sun, or a specular 

reflection of the sun, will occur on the cornea of the eye, not the retina. Retinal 

irradiance will be lower through the AO than when the sun is viewed with the naked 

eye from the earth. The irradiance of the cornea, however, will be much greater 

with the AO. 

b. The concentration of energy on the cornea is such that even specular reflection of 

ground water may present a hazard. Direct viewing of the sun will inflict a serious 

burn on the cornea! Corneal irradiance is most severe at high power. 

c. Since retinal irradiance is not high when viewing a specular reflection, the crew 

may be able to stare at a specular reflection without realizing cornea damage is 

occurring. 

d. While a sun sensor actuated shutter can prevent the sun from entering the field of 

view (FOV), it will probably not be possible to prevent specular reflections from 

briefly entering the field before the shutter closes. 

e. Diffuse reflecting clouds etc. will not damage the eye when viewed through the AO. 

6. 4. 2 DISCUSSION 

The severity of the eye damage problem can be evaluated by analyzing the irradiance of 

the cornea and retina produced when viewing various sources such as the sun, specular 

solar reflections, clouds, etc. "Ball park" calculations of this nature follow and are 

summarized in Table 6.4-1. Since the eye damage thresholds have not yet been determined, 

the corneal and retinal irradiances produced when viewing the sun with the naked eye are 

included to provide a familiar reference. For the same reason figures are included for 

the viewing of a specular reflection off of a lake with the naked eye. 

It may be surprising to note that irradiance level on the retina is less when viewing the 

sun through the high power AO, than when viewing the sun from the earth with the naked 

eye. It happens that the large aperture provided by the AO is more than offset by the 

magnification of the ATS which spreads the increased amount of light energy over a larger 

portion of the retina, and by the low transmission of the AO relative to that of the atmosphere. 

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TABLE 6.4-1. SUMMARY OF CORNEAL AND RETINAL IRRADIANCE LEVELS 

FOR VARIOUS LIGHT SOURCES 

SOURCE IRRADIANCE 

OF RETINA 

IRRADIANCE 

OF CORNEA 

TOTAL FLUX AT 

CORNEA (WATTS) 

1) Sun viewed with 

naked eye through 

one atmosphere 

(ground observer) 

0.3 watts watts 0.0009 0.018 

mm (assumes 4mm pupil) 2 

mm

2 

2) Specular Re flection 

of sun from water as 

seen with naked eye 

through one atmosphere (ground/aircraft observer) 

watts wat ts ,0.02 0.000045 0.0009 2 

mm 

2 

mm 

3) Sun viewed through 

AO at high power 

(127X) (zero atmosphere) 

s 0.04 watts wat ts 25 

mm 

8 2 

mm 

4) Sun viewed through 

AO at low power 

(16X) (zero atmosphere) 

watts watts 0.04 1.5 2 

mm 

0.5 

mm 

5) Specular Reflection 

from surface water 

as seen through AO 

(at high power) (2 

atmospheres) 

watts 0.8 watts 0.001 

mm 

0.21 2 

mm 

6) Diffusely reflecting 

cloud through AO 

(at high power) (2 

atmospheres) 

0.033 x 10-4 

watts 

watts 1.4 x 10-4 4.35 x 10-4 

mm 

2 

mm 

7) Diffusely reflecting 

cloud through AO 

(at low power)(2 

atmospheres) 

2.7 x 10-4 6.5 10-4 watts 

x 20.4 x 10-4 

mm 

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It is less comforting to note that the irradiance levels at the cornea are extremely high 

when viewing the sun in either the high or low power case, and may be significant even 

when viewing a specular reflection from a smooth body of water. This fact becomes palatable if it is recognized that the energy collected in the large aperture is concentrated to 

pass a 2 nun exit pupil near the eye, and spreads to a much larger area on the retina. 

The most disturbing possible source of eye damage is the specular reflections from the 

earth. It will be possible to prevent the sun from entering the FOV with either a sun sensor or computer actuated shutter. Specular reflections, however, will appear instantaneously in the FOV as the spacecraft moves into beams of reflected light from lakes, etc. 

These reflections will appear bigger, but no brighter, in the eyepiece than do similar 

reflections when viewed from airplane with the naked eye, but they will concentrate more 

than two orders of magnitude more energy on the cornea. 

It should be noted that the assumptions on which these calculations are based on are very 

conservative. The assumptions which may introduce the most error in the specular reflection irradiances are: 

a. The reflectivity of the water is taken as 0.05 which is probably not too far off for 

a perfectly smooth body of water. Any rippling, however, will reduce the specular reflectance as the diffusely scattered energy increases. A better estimate of 

the reflectivity could be obtained by measuring the reflectance of local lakes on a 

calm day as a function of incidence angles. 

b. The transmission of the optics was taken as 35 percent*, the current Itek estimate. 

This figure should really be applied solely to the visible region of the spectrum. 

Since glass blocks nearly of the ultra violet which comprises about 10 percent of 

the solar spectrum, and some of the infrared (IR), a somewhat lower transmittance may be more realistic. 

c. The transmittance of the atmosphere was taken as 0.65. The transmittance is a 

function of the angle of observation, weather conditions, etc. The chosen figure 

is believed to be realistic. 

*More recent Itek shows 29 percent to be a more exact figure for transmission. 

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d. The transmittance of the eye was taken as 0.5. It is believed that the actual transmittance ranges between 0.05 and 0.7 as a function of wavelength. Since the same 

figure was used for the case of the naked eye viewing the sun the comparisons 

between the retinal irradiances of Table 6.4-1 are valid. Corneal irradiance 

values are not affected by this assumption. 

e. No attempt was made to consider the spectral content of the energy incident on 

the eye in the analysis. 

6.4.3 CALCULATIONS 

The calculations which yielded the data in Table 6. 4-1 are presented on the following pages. 

a. Sun viewed with naked eye through 1 atmosphere 

- Assume an atmospheric transmission of 65 percent = pA 

- Assume the eye focal length to be 17 mm 

- Let S = the solar irradiance level = 140 MW/CM

2 

- Let O

s 

= the 1/2 cone collimation of the sun = 1/2 degree 

- Let the pupil diameter of the eye = 4 mm 

- 

Let T, the transmissibility of the eye = 0.5 

The irradiance at the cornea is then, 

I

c 

 = pA S = (0.65)(140 MW/CM2

) 

0.0009 W/MM2 

The diameter of the image on the retina is 

dr = 2f 8 = 2 x 17 x 0.0099 (O

s 

= 49° = 0.0044 rad) 

= 0.150 MM 

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The flux collected through a 4mm pupil is 

2 

F

c 

- 7 (a,4) x 140 mw/cm2 (0.65) 

= 11.6 mw 

The irradiance on the retina is then, 

T F

c 4 x 11.60 mw x 0.5 

TT dr

2

/4' 

7 (0.150)2 

= 0.27 w/mm2 

b. Sun as viewed through AO in orbit - high power case 

The flux collected is, 

2 

s F = p E IT 4 

[ 7 = (0.35)(0.00140) 102 25.42 

4 

= 25.00 w 

The irradiance level on the cornea is then, assuming a 2 mm exit pupil on the 

cornea 

F

c 25 I

c 

- - 8 w/mm2 

7 (2)2/4 IT (4.9)/4 

The image size on the retina is 

d = 2m e

s

f = 2 x 17 x 0.0044 x 127 

= 19 mm 

(f = 17 mm) 

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The irradiance of the level is then approximately 

= 

T F

c (0.5)(2.5)(4) ER 

c. Sun through AO - Low Power Case - In the low power case the irradiance on the 

retina remains the same as in the high power case since the decreased effective 

aperture (2.5 inches) is compensated for by a reduction in magnification. 

The irradiance of the cornea is reduced, however, the flux delivered to the cornea 

( ) 

by the AO is reduced by the square of the aperture reduction 2.52 

102 

6.25 I =

c 

100 x 8 = 0.5 w/mm2 

d. Specular Reflections of the sun from surface water as seen through the AO (High 

Power Case) - While a good figure for the reflectivity of smooth water was 

not available at this writing, it is certainly below 0.05. Assuming 0.05 and an 

atmospheric transmission of 0.65 the irradiance of the cornea and retina can be 

calculated from the similar values for direct viewing the sun with the AO 

I

c 

= (0.65)2 (0.05) (8) 

= 0.21 w/mm2 

I

r 

= (0.65)2 (0.05) (0.04) 

, 

= 0.001 w/mm

2 

e. Diffusely reflecting cloud through AO (High Power Case) - The radiance of a 

diffusely reflecting cloud with a reflectivity of 0.9 is 

(0.9) S (0. 9) (0. 140) mw 

B = = 4 x 10-4 

2 

CII1 mm

2 

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d

2

/4 TT (19)2 

= 0.04 w/mm2 

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The flux collected in the 1/2 degree AO field of view is, neglecting atmospheric 

losses, 

F = p irl3 a sin2 1/40 

= (0.35) 7T (4.4 x 10-3) 

(

10 4 2 ir) (25.4)2 

= 4.35 x 10-4 watts 

The irradiance of the cornea is then, 

4.35 x 10-4 

 w I

c 

- 

IT 2

2

/4 mm2 

= 1.4 x 10-4 w/mm2 

The diameter of the retina illuminated is the same as in b. above, d = 19 mm. 

The retinal irradiance is then, 

= 

TF _ (0.5) (4.35 x 10-4) 

d

2

/4 it (19)2 

, = 0. 033 x 10-4 w/nun2 

f. Diffusely Reflecting Cloud - Low Power - At low power the energy collected is 

increased by the larger FOV (4 degrees) and decreased by the reduction in 

aperture size 

Fc 

= pvBasin2 2° 

= (0.35) (g) 4 x 10 

= 204 x 10-5 

= 20. 4 x 10-4  w 

-4 

mm w2) 

(3.9 x 10 -2) 2 ( 2.52x 25.42 

Tr 4 

x4 

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lcis thus larger for the low power case by the ratio of the collected fluxes at 

low power and high power 

2.4 

Ic 40.35 x 1.4 x 10-4 = 6.5 x 10-4 w/mm2 

The retinal irradiance is increased by the ratio of the collected fluxes and the 

decreased magnification squared. 

)2 

4.35 30 (0.033 x 10-4 w/mm2) 

= 27 x 10-4 w/mm2 

Specular reflection of sun aft of water as viewed through naked eye. The irradiance produced in this case are reduced from the values derived for viewing the 

sun directly by the reflectivity of the water. 

I

C = (0.05) (0.4) = 0.02 w/mm2 

LR= (0.0009) (0.005) = 0.000045 w/mm2 

F = (0. 05) (0. 018) 

= 0.0009 w 

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