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~~Drar PROGRAM REPORT 

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AKP-D-596 /11 ,,"J. J7 

: r· .. ·~ ~ .. '. 

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1 APRIL 1965 
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5 E C.RET 

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~~D Jill 

PROGRAM REPORT 

COpy NO. 17 

KP-D ORBITAL DOPPLER RADAR 

THOR! AGENA SATELLITE PROGRAM 

VOLUME II - TESTING 

, ADDII '0£& 

S 

OODVEAR AE"ROSPACE CORPORATION 

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LITCHFIELD PARK, ARIZONA 
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-ivSPECIAL HANDLING 

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TABLE OF CONTENTS 

LIST OF ILLUSTRATIONS · . . · . . . . . LIST OF TABLES . . . . · . . . . . . . . . . . . . · . Section 

I 

II 

III 

IV 

Title 

INTRODUC TION • • • · . . . . . . · . · . · . . . GROUND SUPPORT EQUIPMENT • • · . . . . . . 1. General........... · . . . . . 2. Detailed Description · · · · · • · · · · • • • · a. System Tester. · · · · · · · · • ~. Transmitter-Modulator Maintenance 

Tester . . · • · · · · · · · · · • · · · • 

c. R-F /I-F Maintenance Tester · · · • · · d. Reference Computer Maintenance Tester · e. Control Unit Maintenance Tester. · • · .. 1. Recorder Maintenance Tester · • · · • • · • 

GROUND-BASED RECORDING EQUIPMENT. 

1. 

2. 

3. 

4. 

General . . . . . . . . . . . . . . . . ,. . · . Ground-Based Recorder · . . . . . . Ground-Based Recorder Maintenance Tester 

a. Time-Averaging Synchronizer • • • 

ii". Video Sync-Pulse Generator. • • • • 

Prf Selection Simulator • · . ... . ~. ... 

· . . . . 

· . . • • • • • • 

· . a. Purpose. ~ • • • • • • • • • • • • 

b. Background Theory. • • • •.• •• • • • • • 

c. Video Simulator _. • • • • • • • • • · . . . · . GROUND TEST PROGRAM · . . . . . . . . . . . . . . 1. 

2. 

General. • . • · · · · · · · • · • • • · • • • 

Qualification Test Prog'ram • · • · · · • • 

a. Unit Design Approval Tests · 

· · · • • · 1?. System DeSign Approval Tests,. • • • • • 

c. Unit Environmental Tests • • • • · • • • d. System Temperature-Altitude Tests • • • • • 

e. Santa Cruz Tests. • • 1. Repeat · · · · · · · · · of System Temperature-Altitude Tests. 

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Section Title 

v 

VI 

VII 

3. Flight Equipment Test Program • • • • •• • • • • • • 

a. General..................... k. Unit Acceptance Tests • • • • • • • • • .• • • 

c. System Acceptance Tests • • • • • • • • • • • 

d. Payload Tests • • .. • • • • • • • • • • • • • • • 

PROBLEMS AND CORRECTIVE ACTIONS • • • • • • • 

1. General. • • • • • • • • • • · . . . · . . . · . . . Z. Transmitter-Modulator. • • • • • • • • • • • • • 

3. 

4. 

5. 

6. 

7. 

8. 

a. Vibration Problems ••••••••••••••• k. Pressure Altitude Problems. • • • • • • • • • • • 

c. Other Causes • • • • • • • • • • • • • • • • • • 

RF-IF. • • • • • • • · . . . Reference Computer 

Control Unit • .. • • 

• • 

Recorder · . . . 

• • • 

• • • 

• • • • 

• • • • 

· . . . · . System Problems. • • • • • • 

a. Timing • • • • • • • • • • • • • 

~. Line Filter • • • • • · . . . 

• • • • • • • • 

· . . . . . . . · . . . . . . . · . . . . . . . 

• • · . . · . . · . · . . • • • • • • • • 

c. Clutterlock • • • • • • • • • • • · . . d. Crt Bias Setting and Video Gain •• • • • • • • • • 

Problem Status at Launch • • • • • • • • • 11 • • • • • 

a. General . · · · • · · • • • • • · • • • · · • • • 

k· Transmitted Power Output • • · • • • • • • • • • 

c. Sensitivity Time Control Wave Form • · · · · d. Twt Output Interruptions • • • • • • • • • • • • • 

e. Accelerometer • • • • • • • • • • · • • • · • • • 

ORBITAL OPERATIONS • • • • • • • • • • • • • • • • • 

Mapped Areas · . . • • • • • • · . · . . . . • • · . 1. 

Z. 

3. 

Pre-Recovery Operating Modes • • 

Post-Recovery Operating Modes •••• 

• • • • • • • • 

· . . . • • • • 

ENVmONMENTAL FLIGHT DATA • • 

1. General.......... • • 

Z. Flight Temperature Data . • • 

a. 

~. 

c. 

General • • • • • • • • • 

Ascent Heating. • • 

Orbital Temperature Data • 

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Secti-on Title 

3. Flight Vibration and Shock Data 

4. Flight Pressure Data • • .. • • 

vm UNIT FLIGHT TEST ANALYSIS .. · 

1. General · · • · · · • • • • • 

Z. Transmitter-Modulator • • • • 

a. Genera.l • • • • · • · • • 

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d. Transmitter-Modulator Average Input Power 

e. F73 - Transmitter-Modulator Peak Output 

Power. • • • • • · · • • · • · · • · · · 

3. RF-IF. • • • · • • • • • • · · · · · • · · • 

a. General. · · • • · • · • • · • · • • • • b. Fll - R-F /I-F High-Voltage Supply Monitor 

c. F74 - Twt Pulse Monitor • • · • · • • • • 

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e. F72 - R-f Power to Transmitter Monitor. 

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4. Reference Com~uter • • · · • · · • · • • · • • 

5. Control Unit • • · • • · · • • • • • • 

a. General 

· · · · · · · • · • • · • • • • • • ~. Power Supply Output Voltages FI, F3, .and F4 • 

c. Primary Voltage Inputs F5, F7, 

6. Recorder • • · • • • · • · • · a. General. • • • • • • • • • ~. FlO - 10-Kv Power Supply · c. F55 - Ramp Output • • · · ~. F71 - Crt Llght Output • • • 

e. F49 - Film Motion • • • • • ]. Isolation Roller POSition · • 

SYSTEM FLIGHT TEST ANALYSIS. 

· 

1. 

Z. 

3. 

General • • • · · · · • · · • • 

Typical Radar Pictures 

· • • • · 

Automatic Gain Control (Agc) • 

a. General Operation · of Agc · · b • Age Operation During Pass 34 

.£. Agc Operation During Pass 30 

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Section 

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APPENDIX 

I 

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Title 

4. Attenuator Changes •. · • · · • • • • · • · • • · • • • 

a. Attenuator, G.eneral Operation • · • • • • · • • • • 

~. Attenuator Step 0 • • • • · • • • • · · • • • 

c. Attenuator Steps 2 and 3. · · • • • · • • • • · • • 

d. Attenuator Steps 3 and 4. · · · · • • · • · · · • • 

e. Attenuator Steps 5, 6, and 7. · · · · · • • • • 

5. Changes of Pulse Repetition Frequency • • · • • • • • · a. Effect on Mapping • • • • • • · · • • • • · · • • ~. Operator Recognition Problems • • • · • • • • • • 

c. Evaluation of Prf Selection Simulator • • • • • • • · 

6. Clutterlock · · · • · · • • • • • • • · · · · • • • • 

a. General • · • • • • • • · · · · · • • • I). Clutterlock Correction Angle · 

• • • · • · • · · · • 

c. Vehicle Motions • • • · • • · • • · · • • • · · • d. Clutterlock Commands · · • • · • • • · · · · e. Clutterlock Response - Pass 14 • • • • • • • • · · 1. Clutterlock Loop Qpen - Pass 56. • • • • • · • • · I.. Clutterlock Operation over Water - Pass 16 • • • · • ~. Clutterlock Performance -·Other Passes · · • · · • 

7. Azimuth Ambiguities · · · • • · • • · • 0 · 0 • · · • 

a. . General • • 0 0 · • • • · • 0 0 · • 0 • · · • • · ~. Azimuth Ambiguities in Radar Map. · • • • · · 

8. Range Ambiguity · · • • • • • • • • • • • • 0 0 0 • • 

a. General. • · • • 0 · 0 • · · .. 0 · · • • 0 • b. A mbiguous Interval • • • • • 0 • · 0 • • • • 0 • • 

9. Other Items of Interest • • • • • • • • • · • • · • 0 0 

a. General 0 • • • 0 0 0 0 • .' 0 · • · • • · • · 0 • ~. Target Swallowing • • · • • • • .. 0 • 0 • 0 · · • 

c. Sticking 0 • · · • · 0 • · • • · • • • • • d. Film Fogging · • • • · • • • • • · • • • · 

.. 0 

" · 

C ONC L USIONS • • • • • • 0 • • • · • • • 0 0 • • • • • 

1. General • • • • • • • • • 0 • • • • • 0 • • • • • • • 

2. Progl'am Results • • • • • • • • • 0 • 0 • • · • • · • 

3. Recommendations • • • • • • • • • • • • • • • • • • 

4. Future Applications • • • • • · • · • • • • • • • • • 

CHRONOLOGY OF PROGRAM HISTORY' · . . • • • • 0 o • 

ESTIKA TE OF RADAR BACK-SCATTERING 

C OE FFICIENTS • • • 0 0 • • • • • • • • • • • • • • • • 

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Figure 

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LIST OF ILLUSTRATIONS 

Title 

KP-II Radar System Tester • · . . · . . . · . · . . . Transponder. • • • • • • • · . . · . . · ... . · . . Azimuth Target Generator. • · . . · . . . . . . . . · . . Block Diagram of Azimuth Target Generator · . . . . . . . Data Film Analyzer . . . . . . . . . . . . . . . . . . . Transmitter-Modulator Maintenance Tester. • · . . . R-F /I-F Maintenance Tester · . . . . . . . . . . . . . . Reference Computer Maintenance Tester • · . . . . . . . . Block Diagram of I-f Simulator . . . . . . . . . . . . . . Control Unit Maintenance Tester. • · . . · . . . · . • • • 

Recorder Maintenance Tester • · . · . . . . Time-Averaging Synchronizer • . . · . . · . . • • • • 

Prf Selection Simulator . . . . . . • • • • • 

Sensitivity Time Control Compensation . . · . • • · . . . . Block Diagram of Prf Selection Simulator · . . . . . . . . Qualification Test Program • · . . . · . · . . • • • • 

Flight Test Program • • • • · . . · . . . • • • • • • • • 

Traces Depicting Radar Coverage of United States by 

the KP-U Radar. Deoember 1964 •••• • • · . . . . Typical Component Temperatur~ Cycle • • ~... • • • • · . . ZO Orbital Temperature History - Transmitter-Modulator 

Zl 

Unit. • . . . . . . . . . • . . . . • • . • · . . . Orbital Temperature History - R-F /I-F Unit ••••• 

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Title 

Orbital Temperature History - Reference Computer 

Unit. • • • . • . • • • . . . . • . .'. . . . • • • . • • 

Orbital Temperature History -. Control Unit • • • • • • • • 

Orbital Temperature History - Recorder Unit ••••••• 

Transmitter-Modulator Unit Monitor Signals · . . . . . . . R-F /I-F Unit Monitor Signals • • • • • • . . · . . . . . . Reference Computer Unit Monitor Signals. • · . . . . . . . Control Unit Regulated Voltages Monitor Signals. · . . . Control Unit Primary Power Monitor Signals · . . . . . . . Recorder Monitor Signals • • • • • • 

Recorder Film Motion Monitor Signals 

· . . . 

· . . . 

· . . · . . . • • • • • . . Topographical Map of East Chicago, ruinois 

Radar Return from East Chicago, nlinois. 

· . . . . . . . . . . . . . . 

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34 Typical Radar Maps • • • • • • ". • • • • • • • • • • •• 109 

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Topographical Map of Phoenix, Arizona . . . · . . • • • • 

Radar Return from Phoenix, Arizona • · . . . · . . · . . . Automatic Gain Control Operation . . . . . . . . . . . . . Gain Control Functions • • • • • . . • • • • • • • • • • • 

Automatic Gain Control and Attenuator Step Zero . ... . . . Receiver Attenuation versus Step Position . . .. . . . . Comparison between Attenuator Steps 2 and 3 • • • . . . . . Effects of Processing on Radar Imagery . . . . . . . . . . Range Shifts with Prf Changes . . . . . . . 

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Title 

Range Inversion • • • • • • • • • • • • • • • • • • • •• 133 

Osoillosoope Piotures of Tape Playbaok. · · · • · • • • · · 

Clutterlook Response • . . . • . . . . · • · • · • · • · • 

Deterioration of Mapping with Clutterlook Loop Detailed • • • 

Vehiole Motions and Clutterlook Response - Pass 16. · • • • 

Clutterlook Operation over Water • . . · • • • • • • • · " . Vehiole Motions and Clutterlook Response - Pass 30 · . . . Possible Range Ambiguity. • • • • • • • • • · . . . · . . Antenna Gain versus Step Error and Altitude · . . . . . . . Baok-Soattering Coeffioients versus Aspeot Angle • • • 

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Table 

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XVI 

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xvm 

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LIST OF TABLES 

Title Page 

System Tester Commercial Equipment • • • • • • • • • •• 14 

Transmitter-Modulator Maintenance Tester Commercial 

Equipment ••••••••••••••••••••••• 18 

Reference Computer Maintenance Tester Commercial 

Equipment. • . . . . • . . . • • . . • . • . • • . . . . Control Unit Maintenance Tester Commercial 

Equipment. • • • • • • • • • • • • • • • · . . . . . . . Recorder Maintenance Tester Commercial 

Equipment. • • • • • • • • • • • • • • • • • • • • · . . Transmitter-Modulator Problems • • • • • • • • • • • • • 

R-F /I-F Problems • • • • • • • • . . . • • · . · . • • • 

Reference Computer and Control Unit Problems •••• · . . Recorder Unit Problems • • • • • • • • • • • • • • • • • 

Orbital Pass Coordinates • • . . · . . ... . · . • • · . . Status of Radar Functions • . . . . . · . . • • · . . Mission Time Analysis • • . . . . . . . . . · . . . . . . Flight Instrumentation: Temperature Monitors · . . . . . . Flight Instrumentation: Vibration Monitor · . . . . . . . . Flight Instrumentation: Pressure Monitor · . . . . Unit Temperatures Durlng Ascent · . . . . . . . . . . . . Flight Vibration and Shock Data . . . . . . . . . . . . . . Flight Pressure Data • • • • • • · . • • • • 

Transmitter-Modulator Conditioned Signals 

R-F /I-F Conditioned Signals. • • • • • • 

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25 

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Table 

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Title 

Reference Computer Conditioned Signals • • • • • • • • • • • • 

Control Unit Conditioned Signals • . . . . . . . . . . . . . Recorder Conditioned Signals 

Back-Scattering Coefficients. 

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SECTION I - INTRODUCTION 

AKP-U-596 

The initial design and development phase of the KP-U doppler radar is documented 

in Volume I of this report. This volume - Volume II - documents the ground test 

and flight test phases. 

The sections on ground testing include a description of the specialized test equipment which was used, the tests that were performed on the equipment, the problems 

encountered, and the corrective actions that were taken. 

The sections on flight testing describe the operations which were carried out during 

the mission and present a detailed analYSis of the radar system performance. 

A summary of testing events on a monthly basis is included as Appendix I. 

Estimates based on flight data of the radar back-scattering coefficient, 0"'0' are presented in Appendix U. 

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SECTION U - GROUND SUPPORT EQUIPMENT 

1. GENERAL 

AKP-U-596 

The KP-U radar ground support equipment is composed of a system tester and 

five maintenance testers (one for each radar unit). The system tester is capable of detecting a malfunction in the system and locatiug the malfunction to 

one of the unit maintenance testers. The tester is capable of controlling the 

system and exercisiug all modes of operation, including radar data recording .. using simulated targets suppUed by the tester. An optical correlator (data 

film analyzer) is supplied along With the .system tester for demonstratlDg 

system resolution. 

The unit testers are capable of locating malfunctions to :within replaceable boards 

or subassemblies. Each tester supplies power for the unit and controls its 

mode of operation. Monitoriug capabilities for unit alignment are also available • 

All system testers and unit testers are one-, two-, or three-bay military 

cabinets. All cabinets are dark gray with light gray front panels. The front 

panel nomenclature is black silkscreeliing. 

The testers operate on 115 v :1:10 percent, 60-cycle power. The system and 

Transmitter-Modulator testers require three-phase voltage. All other testers 

require single-phase primary power. The unit testers supply prime power to 

each unit but the system tester does Dot have the capability of supplylng'the 

radar set with prime power. All testers control mode switching, simulate 

inputs as needed, and provide monitoring points for testing. Commercial test 

equipment is built into each tester to be used in conjunction with the test 

pOints. 

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2. DETAILED DESCRIPTION 

a. System Tester 

(I) General 

SECRET 

SPECIAL HANDLING 

SECTI.QN D 

The system tester is pictured in Figure I. The data film analyzer, 

also a part of the system test equipment, is described in following 

paragraph (8). 

(2) Power Control Chassis 

This chassis produces the 28-vdc power for tester operation and con ... 

troIs the 60-cycle power to the various pan'els and commercial test 

equipment contained within the tester. 

(3) System Control Panel 

The system control panel simulates the normal command control 

function to the radar set. Power control commands of warm-up, 

pre-operate. and operate can be initiated. In addition, system control 

functions such as prf r~te, receiver attenuation, built-in-test, clutterlock· time constant, clutterlock ground, film drive ON-OFF, and time 

delay override are provided. 

(4) Test Point Panel 

The chassis provides access to radar test signals on the front panel 

of the tester. 

(5) "etering Panel 

The metering panel provides six meters (M-I through M-6) for contin";' 

uous monitoring of selected test points. 

(6) Azimuth Target Generator and Transponder Panels 

These two panels operate together functionally to simulate the generation of doppler targets as a function of vehicle movement. The 

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Figure 1 - System Tester 

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AK P-II-5CJj 

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SECTION D 

transponder is pictured in Figure 2. The unit is shown removed from 

the system tester and mounted in its case for portable field use. 

(a) Azimuth Target Generator 

The azimuth target generator (Figure 3, lower portion) is used 

to modulate microwave signals in the microwave section of the 

transponder. The azimuth target generator produces an output 

simulating the doppler equation 

Zv sin e 

fd = ~ 

where 

e = angle measured from zero doppler 

v = velocity of vehicle. 

~ = wave length of transmitted signal. 

(1) 

As the velocity of the vehicle v and the wave length of the transmitted frequency are constant. the doppler frequency at any 

point is only a function of the angle or constant doppler line. 

The rate of change of doppler frequency as the vehicle moves 

past a given target can be approximated by 

(2) 

where R = slant range to. target. This doppler signal is produced 

by mixing the output of two oscillators, one a fixed 5-kc oscillator 

and the other a voltage-controlled oscillator (vco) which is driven 

by a ramp through a frequency excursion of 3 kc to 7 kc at a 

linear rate. The resulting mixer output is a swept frequency Z kc . t 

through 0 to 2 kc simulating a doppler target • 

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Figure Z - Transponder 

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S:E'C R ET 

AKP-U-596 
twj 

f 

... 

CD 

W 

, 

> III 

n ell 

.... - s .. 

• ~ Z I 

Q) 

:. I ~ Z ~ ~ .. -'= -

z "" Q 

~ c:s 

CD 

... 

ID 

0 

"" 

... 

COAX LOAD POWII 

fIANSMITTR-

.aMl . TUII! 

"'-DI FIIONr NoNIL 

AmNUotoTOI (0ImUI') 

INM 

(TEST POINT) 

Mt=~} I 

$TALC) 

SINGLE SIDE ...... ND MODUlATOIS 

/A27_si;;"' S KC -- AZIIIUTII TAIIOIT .-~= c"HAN; N~ - - ., I 

I 

I 

I 

I 

..,.. 

~--~lr~ ____ -i~ ______ l.~~~lc~og~N~E~cr~~- : NUM1E12 _L I 

I 

~ ......n . .!.. : 

I t 

OSCIUATO.D ~PO'fl.~5Y~ .__ . . _ , -.. I I 

I : 

I I 

I 

I I - I 

___ .~&1. ___ J 
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To produce a second doppler target offset from the first target, 

a second oscillator operating at 5 kc :i: Afd ,where Afd is the 

frequency difference between the two oscillators, is mixed with 

the same vco output as was used for the first doppler target. 

AId can be calculated from 

Af = Zv X desired azimuth displacement d T .. range to target • (3) 

During operation the DOPPLER ON-OFF switch supplies power 

to one of the 5-kc oscillators and the ramp and vco to produce 

one doppler target. Placing the resolution switch to the ON 

pOSition will supply power to the offset oscillator (5 kc + f

d ) 

which will produce the second doppler target. The frequency of 

. the 5-kc and offset oscillators can be measured from the frontpanel test points using the counter. Adjusting the offset control 

will vary Afd and thus provide various azimuth displacements. 

For each doppler target two outputs are used, the two differing 

only in phase. Thus there are four outputs which go to the 

microwave section. Each doppler target has two outputs, each 

shifted by 90 degrees so that a single side-band signal can be 

produced in the microwave section. 

(b) Transponder Panel 

The transponder panel (Figure 3, upper portion) receives a transmitter pulse from the system under test,modulates the pulse with 

the doppler wave form received from the azimuth target generator, 

delays the signal to simulate range. and returns the resulting 

. signal to the system. 

The input signal from the system encounters a transmit-receive 

(t-r). tUbe which blocks the pulse from entering the receiver 

channel. The signal, attenuated ZO db by the cross guide coupler, 

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passes through a fr,ont~panel mounted. 70-db variable attenuator. 

After paSSing through the attenuator, part of the signal is fed 

into the crystal detector on the front panel marked POWER 

DETECTOR. The rest of the signal continues down the wave 

guide and splits into two parts. Channel No's I and Z. where 

each part or channel encounters an orthomode (balanced) mixer. 

At the orthomode mixers each signal is mixed with a local oscillator reference which is generated inside the transponder and 

modulated in a single side-band modulato! by the doppler wave 

form. The result of mixing the two signals is a third signal 

whose frequency is 70 mc + fd (refer to Equation (1». This 

signal is fed into the i-f section where it is delayed. A switch 

in series with the output of the Channel No. Z orthomode mixer 

allows the doppler target in Channel No. Z to appear as displaced 

in azimuth or range from the doppler target in Channel No. I. 

This switch is the AZIMUTH-RANGE switch located on the front 

panel. (The method whereby the azimuth displacement is varied 

was discussed in the preceding paragraph (a) on the azimuth target generator.) The range displacement is controlled by the length 

of a coaxial cable connected to two coaxial connectors marked 

RANGE on the back of the transponder chassis. 

The two channels are added together to form one ~1gna1 and sent 

to the delaying portion of the i-f section. Here one of two different delays can be selected by means of the STALO-RESOLUTION 

switch located on the front panel. The switch is thrown to the 

STALO-RESOLUTION position when a frequency-stability test is 

being run and to the RESOLUTION pOSition when a resolution test 

is being run. The STALO position delays the pulse 1960 microseconds for long-time stability tests. The RESOLUTION position 

produces four pulses delayed 30.' 50. 70 and 90 ~icroseconds. 

The signal is then applied to a third orthomode mixer (or modulator) where itismixed with the original continuous-wave local 

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oscillator signal and filtered to produce a signal identical to the 

input signal except that It contains the doppler information. This 

signal then goes through a variable attenuator and the unfired 

transmit-receive tube to the output. 

(c) Derivation of Constants for Transponder 

Referring to Figure 4. the azimuth target generatorportloD of 

the transponder. it is desired to solve for the ramp period (t), the 

separation in cycles between the two oscillators (~fd) and other 

similar parameters of the generator. 

Starting with the basic equation relating doppler frequency. vehicle 

velocity. and wave length. 

f - Zv sin e 

d - ~ 

differentiating with respect to e 

where 

vdt 

de = It · 

Substituting in Equation (S) 

dfd = Zv c~s 9 (~) 

dfd ZvZ ZvZ 

dt = n cos e e! R~ (since e Is small). 

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(4) 

(S) 

(6) 

(7) 
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SINE J MIXER I ~ I 

OSCILLATOI PHASE 

5KC S .. FT8 

COSINE I M1XB l 

I J 

I I-KC l SINE 

I FILTH I 

J I-KC COSINE 

I FILTEI 

FILTEIS FOI 

l...aSlDE .... ND 

SECTION II 

10 SlNGU-SlDE-MND 

WCrtJWAYE MODUlAlOI I TAIGET NlMiWEI I 

t;J"t1 

IIAMI' 

GENEllA10I 

OSCILLATOI 

5KC + Aw 

I VOlTAGECONl'lOWD 

OSCILLAtol 

SINE I MIXER 

I 

PHASE 

SHIFTEI 

COSINE ~ MlXEI 

I 

3-7KC 

I' J 

I I 

. 1 J 

I I 

I-KC 

FILTEI 

I-KC 

FILTEI 

SINE 

' COSINE I TAIGET NUM1E12 

TO SlNGLf-SlDE-lAND 

WCrtJWAYE 

MODULAlOI 

Figure" - Block Diagram of Azimuth Target Generator 

-IZSPECIAL HANDLING 

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AKP-U-596 

Assuming a vehicle velocity of 25,500 ft/sec, antenna depression 

of 55 degrees, and an altitude of 130 nautical miles, 

.6.fd 2(25.5 X 103)2 X 0.82 

.6.t = 130 X 6076 X 0.1025 

= 13,200 cycles/second. (8) 

USing a total frequency sweep of 2 kc - 0 - Z kc, or 4 kc, then 

the ramp period is 

4000 

t = 13. 200 = O. 303 sec • (9) 

To offset one target 50 feet from the first target. the difference 

in frequency will be 

.6.fd = (2V c~s e) .6.a 

= (~) cos e (~) .(10) 

and, for small values of a, 

(cos a e! I) 

= 2 X 25.5 X 103 X 50 X 0.82 

0.10Z5 130 

= 26 cycles. (11) 

For a beam width of 0.34 degree to the half-power points, from 

Equation (4), 

= 2v sin 9 _ 2v sin O. 170 

fd ~ - i 

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AKP-I1-596 SECTION II 

2 X Z5. 5 X 103 sin O. 170 

- 0.10Z5 

= 1450 cps (at half-power points). (IZ) 

The ramp amplitude is chosen to provide a frequency range of 

5 ~Z kc at the output of the voo. The simulated doppler sweep of 

Z ko - 0 - Z kc is processed by a low-pass filter to give a onehalf power roll-off at 1500 cps simulating an aotual target doppler 

return. 

(7) Commercial Test Equipment 

Table I lists the commeroial test equipment items which. in addition 

to the specially designed equipment already described, are housed 

in the system tester. 

TABLE I - SYSTEM TESTER COMMERCIAL EQUIPMENT 

Item Manufacturer Model no. 

VTVM Hewlett-Packard 410BR 

Electronio oounter Hewlett-Packard 5Z3CR 

Delay generator Hewlett-Paokard Z18AR 

X-band test set Hewlett-Paokard 6Z4CRX 

Differential voltmeter John Fluke 803BR 

Oscilloscope Tektronix 585 

. Speotrum analyzer Polarad TSA-W 

Scope camera Tektronix C19 

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SEGRET 

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SECTION D AKP-D-596 

(8) Data Film Analyzer 

The data film analyzer (Figure 5) serves as a simple optical correlator 

for demonstrating the resolution of the synthetic phase histories of 

targetiJ provided by the transponder and azimuth target generator. The 

main optical elements of the data film analyzer are the light source, 

filter system, condensing system, azimuth lens, data film. range lens, 

transform plane slit, and microscope. A removable magazine section 

is provided for insertion of the data film strip to be analyzed. 

!!.. Transmitter-Modulator Maintenance Tester 

The Transmitter-Modulator maintenance tester (see Figure 6) consists of 

specially designed test panels, commercial test equipment, and auxiliary 

equipment housed in a single-bay test rack. The specially designed equipment includes a microwave equipment chassis, control and monitor chassis, 

and a high-voltage power supply. 

(1) Microwave Equipment Chassis 

The microwave chassis contains the wave guide equipment necessary 

to test the Transmitter-Modulator. The microwave equipment consists 

of two basic sections: a calibrated continuous-wave r-f source to provide a signal input to the klystron, and a transmitter output power 

and pulse characteristic monitor. 

(2) Control and Monitor Chassis 

The control and monitor chassis contains the power switching, interlocks, and indicators necessary for safe operation of the transmitter; 

a pulse generator to provide the prf trigger to the transmitter; meters 

and test points for monitoring signal and power voltages and wave 

forms; and the low-voltage power supplies. The controls and switches 

for the high-voltage power supply are also included in this chassis. 

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Figure 5 - Data Film Analyzer 

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Figure 6 - Transmitter-Modulator Maintenance Tester 

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(3) High-Voltage Power Supply 

SECTION D 

The power supply is a standard, three-phase, full-wave, unregulated 

choke input with a pi-section filter capable of delivering 0 to 5000 vdc 

at 500 ma with very low ripple. 

(4) Commercial Test Equipment 

. Table D.lists the commercial test equipment in the T~ansmUter,", 

Modulator maintenance tester. 

TABLE n - TRANSMITTER-MODULATOR MAINTENANCE TESTER 

COMKER~IAL EQUIPMENT 

Item 

Test oscilloscope 

Microwave power meter 

Inverter. ZOOO eps 

Klystron power supply 

.!t. R-F /I-F Maintenance Tester 

Manufacturer 

Tektronix 

Hewlett-Packard 

Behlman-Invar 

Hewlett-Packard 

Model no. 

RM45A 

430CR 

RI61A 

·716A 

The R-F /I-F maintenance tester (Figure 7) consists of specially designed 

test panels. commercial test instruments. and auxiliary test equipment 

housed in a double-bay test rack. The specially designed test panels include a stalo check chassis, controller. noise generator, d-c power supply, 

and vhf attenuator panel. 

(I) Stalo Check Chassis 

The stalo check chassis is used to determine the phase stability of the 

local oscillator in the R-F /I-F unit. The dispersed X-band pulse from 

the R-F /I-F unit is delayed, then returned to the R-F /I-F unit as a 

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Figure 7 - R-F !I-F Maintenance Tester 

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target return for detection and compression. When the compressed 

pulse is synchronously d~modulated (phase-detected), the stalo 

stability can be measured by observing amplitude variations in the 

detected pulse. 

(2) Controller Chassis 

The controller cbassis contains the power switching and indicators 

for operation of the R-F/I-F unit, meters and test points for monitoring signals and power voltages, and wave forms and signalgenerating circuits required to test the R'~F /i-F unit. 

(3) Noise Generator Chassis 

The noise generator chassis incorporates an X-band noise tube, 

mount, and power supply with a wave guide attenuator to provide a 

calibrated variable excess noise source. When making noise figure 

measurements, the noise generator is connected to the receiver by a 

coaxial cable and coaxial-to-wave guide adapter. The i-f signal output of the receiver is connected to an r-f voltmeter through an i-f 

attenuator. With the noise generator off and the receiver on. the i-f 

attenuator is set for a convenient reference on the r-f voltmeter. The 

i-f attenuation is increased by 3 db. the noise source turned on, and 

the noise level attenuator adjusted for the same reference level output on the r-f voltmeter. For this setting the signal noise equals 

the input noise and the nO.ise figure of the receiver may be calculated. 

(4) D-c Power Supply Chassis 

Four commercial power supply modules mounted on a common chassis 

supply d-c power to the tester and.·to the R-'F /I __ F unit under test. 

d. Reference Computer Maintenance Tester 

The Reference Computer maintenance tester (Figure 8) consists of 

specially designed test paJl'l8ls, auxiliary equipment, and commercial 

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Figure 8 - Reference Computer Maintenance Tester 

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SECREt 

speCIAL. HANDLINO 

AKP-II-596 SECTION U 

test equipment housed in a dual-bay test rack. The specially designed 

panels include a power supply unit. power control unit. and a simulator 

control unit. 

(l) Power Supply Unit 

The power supply unit supplies the power for the Reference Computer 

and for the tester. 

(2) Power Control Unit 

The power control unit controls the input power to the tester. 

(3) Simulator Controller Unit 

The simulator controller unit furnishes the following controls and 

signals for the Reference Computer. 

(a) Control of D-c Power 

The d-c power is controlled by a d-c power switch which controls 

power to the Reference Computer. 

(b) Mode-Setting 

PRF SELECT is accomplished by a four-pole. 16-position rotary 

switch. The 30-millisecond timer on the timer relay board. when 

actuated by the CIL RESET switch. generates a 30-millisecond 

28. 3-vdc pulse for the CLUTTERLOCK ON and CLUTTERLOCK 

TIME CONSTANT SELECT circuits of the Reference Computer. 

The BIT switch energizes the BIT circuit of the Reference Computer. 

(c) Conditioned Signals 

The conditioned signals from the' Reference Computer are monitored by the CONDITIONED SIGNA,L SELECT and GROUPING 

SELECT switches and the CONDITIONED SIGNAL METER~ 

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(d') Reference Computer and Tester Signals 

• Reference Computer signals and tester signals used to control 

the Reference Computer are monitored and/or terminated on 

the front panel. 

• 

• 

COUNTER SELECT allows selecting of test signals from the 

Reference Computer for frequency counting. 

(e) Clutterlock Integrator Testing 

The integrator linearity and clutterlock response to a step 

function are tested by the use of a timed variable amplitude 

test pulse. 

The VCO-CAL switch allows test1~g of the clutterlock operational 

amplifier as an Integrator or as a 1:1 d-c amplifier. 

(f) I-f Simulator 

The I-f simulator portion of the simulator controller unit 

provides a simulated i-f output (a chopped" 70-mc single sideband signal) for the closed-loop operation of the clutterlock 

circuit of the Reference Computer. Figure 9 is a block 

diagram of the i-f simulator. The 70-mc signal. offset by 

prf/4 from the Reference Computer. is modulated with a 

"doppler" frequency (fd ) in a single side band (ssb) modulator 

and then chopped at a 33-kc rate to provide simulated i-f pulses. 

Refering to Figure 9, the output of a vco, whose output frequency 

can be varied by means of a potentiometer from 40 to 60 kc. is 

compared in two synchronous demodulators with a SO-kc cosine 

and a SO-kc sine signal from a fixed SO-kc oscUlator. The two 

outputs of the synchronous demodulators are at the difference 

frequency and are In quadrature. These two signals serve a8 a 

reference in the ssb. The ssb output is 70 mc + prf/4 :l:fd• 

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lTV ~ I FILTEl I-~_""" SYNCHIONOUS _ LOW PASS I---

TI »-KC DEMODI.l.ATOI FIlTEl 

COSINE 

I ------..... 

SIMULATED IF 

rum 

»-KC PHASE 

I 

I 

I 

I 

I 

- SINGLE - HI III 

ro- SIDE lAND ~ DlODEt--__ _ OSCILLATOI SPlma MODtJ.ATOI GATE 

I 

I 

_ FUR 

~ SYNCtW)NOUS f-r- LOW MSSI--I __ .... 

»-KC SIN DfMODtJ.ATOI FILTEI 

.JlIlfl (50KC *'01> SINE 

VOLTAGE PHASE 

~OII--"'SPlITTEI (50KC*,,,COSINE) 

7O-MC OffSET I""" "" 1 

+23.5 -23.5 

(FIIOM WBENCE COMPUTER) 

TlIGGEI 

---j rloo "SEC ---j rlOO" SEC 

r ----rLfl- ----BJIIlII 

.... ,... 

-

---..t 33-KC I. ______ .... 

GATE GENElATO ..... ------t MULTIVIllATOI ... 

2226-22 

Figure 9 - Block Diagram of I-f Simulator 

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This output of the ssb modulator is next chopped at a 33-kc rate 

by gated 33-kc pulses. These pulses are gated on during a 100-

","sec pulse which is triggered at prf rate. These gated 33-kc 

pulses in turn gate the 70 mc + prf/4 zfd output of the ssb modulator in a diode gate. The output of this diode gate is a coherent 

i-f signal which simulates the i-f return from the radar. 

(4) Commercial Test Equipment 

,; 

Table Ullists the commercial test equipment in the Reference Computer maintenance tester. 

TABLE m - REFERENCE COMPUTER MAINTENANCE TESTER 

COMMERCIAL EQUIPMENT 

Item Manufacturer 

Oscilloscope Tektronix 

Counter Hewlett-Packard 

Signal generator Hewlett-Packard 

VTVM Hewlett-Packard 

R-f voltmeter Boonton' Electric 

Multimeter ANPSM-6 (Mll-M455ZC-USAF) 

.!.. Control Unit Maintenance Tester 

Model no. 

585 

5ZZB 

'608C 

410B 

91C 

The Control unit maintenance tester (Figure 10) consists of specially 

designed test panels. auxiliary equipment. and commercial test equipment housed in a single-bay test rack. The specially designed panels 

include a power supply load panel. power control panel. and a-c power 

oscillator panel. 

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SPECIAL HANDLING 

Figure 10 - Control Unit Maintenance Tester 

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(I) Power Supply Load Panel 

AKP-D-596 

This chassis contains three variable power resistors which act as loads 

for the three power supplies in the Control unit. Provision is made for 

shorting the outputs of the supplies to ground for overload recovery 

testing. 

(2) Power Control Panel 

The power control panel of the tester allows'for monitoring. switching 

and controlling all the input power to the Control unit. It also provides 

for measurement of the output voltages and currents of the Control unit 

power supplies. Provision is also made for reading all the monitor 

voltages from:the Control unit. 

(3) A -c Power Oscillator Panel ' 

This chassis contains two fixed-frequency power oscillators of 400 and 

2000 cycles. The outputs of the oscillators are connected so that either 

may be switched to drive the Type 351A Invertron power amplifier (refer 

, to Table IV) which supplies the a-c inputs to the Control unit. 

(4) Commercial Test Equipment 

Table IV lists the commercial test equipment in the Control unit maintemance tester. 

TABLE IV - CONTROL UNIT MAINTENANCE TESTER 

COMMERCIAL EQUIPMENT 

Item Manufacturer Model no. 

Oscilloscope Tektronix RM32 

Differential voltmeter John Fluke 803BR 

A-c power supply Behlman-Invar Type 351A 

D-c power supply Behlman-Invar Type 275W-8 

Fixed frequency oscillator Behlman-Invar Type 1-400 ~. 01 percent 

Fixed frequency oscillator Behlman-Invar Type 1-2000 =0.01 percent 

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5 E e R Et , SPECIAL HANDLING 

SECTION II 

f. Recorder Maintenance Tester 

The Recorder maintenance tester (Figure 11) consists of specially designed 

test panels. commercial test equipment. and auxiliary test equipment housed 

in a commercial single-bay rack. The specially designed equipment consists 

of a Recorder control and test chassis. d-c power supplies. an auxiliary 

film take-up cassette, and a recorder service stand. 

(1) Recorder Control and Test Chassis 

The Recorder control and test chassis provid~s for Recorder power 

switching. film drive control and monitor of Recorder d-c voltages. 

conditioned signals. and wave forms. In addition to these control 

and monito.ring functions the chassis contains a number of testsignal generating circuits and provides a sweep trigger to the 

Recorder. 

(a) Sweep Trigger Generator 

The sweep trigger generator provides a pulse for the ramp generator in the Recorder. This circuit consists of a variable frequency 

multivibrator. blocking oscillator. amplifier. and limiter. 

(b) Range Mark Generator 

The range mark generator provides a train of pulses that is 

synchronized with the swe.ep in the Recorder. These range 

marks are used to measure crt sweep linearity. 

(c) 50-Cps Generator 

The output of the 50-cps generator is used to provide pulses 

which may be recorded on the film to serve as a measure of 

uniformity of film speed. As the film speed is five inches per 

second. ten pulses will appear per inch of film. 

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Figure 11 - Recorder Maintenance Tester 

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(d) 1. 6-Kc and 3.2-Kc Generators 

SECTION D 

This circuit consists ,of a full .. wave bridge rectifier and a pair of 

synchronized oscillators tuned to 1.6 kc and 3.2 kc, respectively. 

The outputs. which are sampled by five-microsecond pulses, are 

used to determine the azimuth resolution of the Recorder. Considering the vehicle velocity to be 25,000 feet per second. these 

test frequencies correspond to spatial frequencies of 15.6 and 1.8 

feet per cycle, ~espectively. 

(e) Prf/Z Generator and Adjustable Video Gate Generator 

The prf/Z circuit consists of a bistable multivibrator, which is 

triggered by the sweep trigger. and a diode limiter. This output is also used to determine azimuth resolution. 

The sweep trigger is also used to trigger a delay circuit which 

produces a delayed video gate. 

The video gate is a five-microsecond video test pulse which may 

be varied to occur from 5 to 80 microseconds after the start of 

the sweep. A video amplifier is used to drive the cable and 

IOO-ohm input to the video amplifier in the Recorder. 

(f) 6-Kc, IO-Mc. and 14-Mc Generators 

The unblanldng gate from the Recorder is used to gate on a 

Hartley oscillator to produce. depending on which of three 

tuned circuits is chosen. one of these three test frequencies. 

The output is used to check the range resolution capability of 

the Recorder. Considering the slant range interval to be 13. 5 

microseconds. these test frequencies correspond to spatial 

frequencies of 11 ~. 43, and 30 feet respectively. 

(g) IO-Cps GeneratOr 

This circuit consists of a free-running multi vibrator and an 

emitter follower. The output may be switched to anyone of 

six data lamps in the Recorder. 

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(Z) D-c Power Supplies 

AKP-D-596 

This chassis provides the +Z8. 3~vdc. +Z3.5-vdc. -Z3.5-vdc. and 

+300-vdc voltages to the Recorder. 

(3) Auxiliary Film Take-Up Cassette 

This unit was described in Volume I. page 14Z. 

(4) Recorder Test Stand 

A Recorder test stand accompanies each tester and is used to support 

the Recorder and Auxiliary take-up cassette during test. 

(5) Commercial Test Equipment 

Table V lists the commercial test equipment in the Recorder· maintenance tester • 

TABLE V - RECORDER MAINTENANCE TESTER 

Item 

Oscilloscope 

Volt/ohm milliameter 

A-c power supply (3) 

COMMERCIAL EQUIPMENT . 

Manufacturer 

Tektronix 

Simpson 

B ehlman-Invar 

RM15 

Z70 

161A 

Model no. 

Fixed frequency oscillator Behlman-mvar Type 400 cps :1:0. 0 I perc ent 

Bridge oscillator 

Photomultiplier 

microphoto meter 

General Radio Type 1330-A 

American Instrument Co 10-31Z 

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. SECTION m - GROUND-BASED RECORDING EQUIPMENT 

1. GENERAL 

The ground-based recording equipment located at the tracking station is used 

to provide a permanent photographic record of the video information telemetered 

from the vehicle. As this information should match the information recorded in 

the vehicle, it is only necessary to provide a KP-D Recorder and a means of 

. triggering it to record the telemetered video. The ground-based recording equipment therefore consists of aRecor~er and a Recorder maintenance tester which 

is modified to provide for extracting the video sync-pulse from the composite 

video. A prf selection simulator is also provided for training traQking-station 

personnel (refer to paragraph 4 following). 

z. GROUND-BASED RECORDER 

3. 

The Ground-Based Recorder is identical to the flight Recorder described in 

Volume I, Section VII, paragraph '5. In use it is mounted on the Recorder test 

stand which is supplied with the Ground'-Based Recorder tester" 

GROUND-BASED REC ORDER MAINTENANCE TESTER 

This tester is identical to the Recorder maintenance tester described in Section 

n except that it contains an additional assembly: the synchronizer and test 

generator chassis. 

- - The synchronizer and test generator chassis contains two circuits: the timeaveraging synchronizer and the video sync-pulse generator. 

The .composite video containing the video sync-pulse and the video signal is 

fed to the time-averaging synchronizer. The synchronizer ~etects the video' 

sync-pulse and generates a sweep trigger at the same prf. A manually operat.ed switch is provided for changing the frequency of the synchronizer to the 

approximate frequency of the video sync-pulse • 

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SECTION III 

The video syno-pulse generator produces a test video syne-pulse that is used 

to check the synchronizer. A 16-position rotary switch is provided to change 

the prf of the pulse generator from 8215 to 8735 cps in approximately 35-cps 

steps.· 

!:.. Time-Averaging Synchronizer 

The time-averaging synchronizer has the composite video signal as an 

input. detects when the video sync-pulse occurs. and generates a sweep 

trigger output that can be more stable than the input. When the synchronizer has no input or is first turned on. automatic search circuits vary 

the lock-on Circuits about the prf number selected on the front panel. 

After lock-on the search circuits are turned off and the synchronizer 

can automatically follow small changes in prf. 

(1 ) General Description 

The composite video signal (see Figure 12) is fed through a galnchange circuit G-l which is normally such that the gain. prior to 

lock-on. is high. The video signal then goes to a diode· bridgesampling oirouit S-I whioh is gated on by a one-microsecond 

blocking osclllator BO-I. The output of the sampling oircuit is 

near zero except when the blocking oscUlator causes the sampling 

circuit to sample during the sync pulse. The sync pulse is a onemiorosecond positive pulse followed immediately by a·one-microseoond negative pulse. 

The output sampling oircuit drives a vco and helps to set its frequency. The output of the vco in turn is used to trigger BO-I. 

The trailing edge of BO-I also triggers a 0.25-mioroseoond 

blocking oscillator BO-2 whioh causes sampling oircUit S-2 to 

sample the sync pulse. After lock-on. S-2 samples only during 

the negative portion of the sync paIse and causes switching amplifier 

A-2 to cut off the operation of sawtooth generator ST-I. Before 

lock-on. ST-I is sweeping the vao approximately :1:50 cycles at an 

8 cps rate. 

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COMPOsm 

VIDEO IN GAIN 

0--............ CHANGE - G-I 

SAMPLE 

CIRCUIT 

5-2 

SWITCH '-- AMf'lIFIB 

A-2 

6 EC RET' 

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,---;1 r19 "SEC 

SAMPlE 

CIRCUIT 

5-1 

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DELAY "1...r :lFIa 

MlI.n- fMITTII 

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SWEEP L 

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Figure lZ -. Time-Averaging Synchronizer 

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(Z) Lock-On 

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SECTION m 

When the synchronizer has an input but is not locked on, and the vco 

manual frequency control is set approximately correct, lock-on is 

accomplished in the following manner. The vco is free-running but 

not at the correct frequency. Blocking oscillator BO-I is firing and 

causing the gated bridge sampling circuit S-I to sample but its fUtered 

output is zero. Therefore, the vco does not receive any bias source 

from amplifier A-I. The trailing end of BO-I triggers BO-Z but the 

filtered output of the gated bridge sampling circuit S-Z is also zero. 

Switching amplifier A-Z is set in such a manner as to set the gainchange circuit G-I at high gain, and also to allow the sawtooth generator ST-I to cause the frequency of the vco to sweep in a sawtooth 

manner about the frequency of the composite video trigger. 

As soon as blocking oscillator BO-I causes S-I to sample a portion 

of the video-sync pulse, the circuit quickly locks on to the sync pulse, 

the gain-change circuit decreases the gain, the search sawtooth 

generator is turned off, and the vco operates at the same prf as the 

input. 

(3) Sweep Trigger 

The output of the vco is used to trigger a delay multivibrator. Tile 

trailing edge of the delay multivibrator is used to develop the sweep 

trigger pulse in the succeeding amplifier and emitter follower stage. 

b. Video SynC-Pulse Generator 

A test pulse was needed to check the operation of the synchronizer and, 

because of the unique shape of the pulse. the video sync-pulse generator 

was designed to generate this wave form. The frequency is controlled by 

a 16-position switch circuit that generates prf's from 8Z 15 to 8735 cps in 

approximately 35-cps steps. 

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4. PRF SE LECTION SIMULATOR 

a. Purpose 

AKP-U-596 

The prf selection simulator (Figure 13) is used for training trackingstation personnel in the selection of prf for best radar performance. For 

a given altitude, considering the elevation beam pattern and depression 

angle, there will be an optimum prf setting •. FOr a different altitude, the 

prf for optimum performance will change. The prf is selected by visual 

examination of the return followed by a manual adj~stment of the prf selection switch. Prior to an actual mission the ope:rators can become familiar 

with the techniques involved by use of the prf selection siuiulator. 

b. Background Theory 

Refer to Figure 14. Assume that the radar is at point S transmitting a 

beam toward the ground. A pulse leaves the radar and arrives at point 

SI at range R1• This point will be somewhat down on the beam pattern 

and the radar return for an average target will be of amplitude E1• For 

point Sz the return amplitude will be maximum since the target is at the 

center or maximum power point of the antenna beam.' For point S3 the 

return will again be down. The resultant return for a uniform target 

distribution will be peaked as shown in curve 1. The peaking effect 

will be compensated for by the sensitivity time control (ste), curve Z, 

provided the prf is properly selected to center the radar return about 

the mid-pOint of the stc curve. 

In Volume It Figure 16, the stc wave form was shown to occur at a fixed 

time following the transmitted r-f pulse. The prf is adjustable from 8Z 15 

(step 15) to 8735 (step 0) pulses per second in 16 steps. This represents 

a change of 114.5 to lZI.7 microseconds or about I/Z microsecond per 

step. Since about 16 prf pulses are sent out before the first pulse is returned from the ground, this results in the ste being delayed or advanced 

approximately eight microseconds per prf step relative to the radar return. 

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Figure 13 - Prf Selection Simulator 

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OUTPUT VOLTS 

(2) STC GAIN CURVE 

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(3) IlADAIIETUIN COMPENSATED rI'( STC 

AKP-D-526 

!;(I)IIAOAIIIfTUIIN (VOLTS) 

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IANGE- 13 

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--~.. ,"--- .'-...:' c PIF LOW 

/ . -. ...---.- ........ 

--~~---- '... #-_ d PItF HIGH 

-.:a 

- .. .----- .......... -

• 2226-28 

Figure 14 - Sensitivity Time Control 

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SECTION m 

Figures 14(b), (c), and (d) show the radar return (upper envelope only) at a 

constant altitude (based on an assumed uniform target distribution) for low. 

high and optimum prf. Prf refers to actual rate of transmission of radar . 

pulses. Similar results are obtained (except for slight changes in the radar 

return envelope) for other than nominal altitudes as the prf is stepped to either 

side of the optimum value. Figure IS shows the prf selection simulator in 

block diagram form and the envelope of the output wave form of the Simulator,. 

The noise generator G-l (Figure IS(a» generates a random noise voltage 

which simulates the target return. The frequency range of the noise is . . 

apprOximately ZSO kc to 10 mc. The noise voltage is fed to two modulators 

M 1 and MZ• The first modulator modifies the noise voltage with the stc 

wave form. The output of the first modulator serves as an input to the 

second modulator. The modulating signal for the second modulator is the 

resultant of an addition of a sine wave and a low-variable-frequency wave 

form. The sine wave provides the simulated antenna beam envelope, and 

the low frequency, a variable "terrain II input. 

The output of the second modulator is amplified in a four-stage amplifier 

A-I to obtain the desired signal level. The amplified modulator output is 

inhibited for five microseconds every 114 to IZI microseconds. This is 

accomplished by clamping to a d-c level the output of one of the amplifiers 

with the On-Gate generator G-Z. The amplified output of the modulator is 

fed to summer S-Z where it is added to the sync pulse. 

The combined signal (modulated noise and sync pulse) is then passed through 

a limiter L-I. The output of the limiter is adjustable from the front panel 

of the noise generator. The output of the limiter is Output 1. Output Z on 

the front panel is the output of the limiter after it has passed through an 

additional stage of amplification (see Figure IS(b». 

A noise adjustm~nt located on the front panel provides a means of combining 

the output of the second modulator with additional noise •. 

The prfselection switch located on the front panel is a ten-deck switch. The 

"A" deck is used to supply one of 16 d-c levels to a vco. The oscillator 

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Figure 15 - Block Oiagram of Prf Selection. Simulator . 

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SECTION m 

output is used to generate the prf trigger at anyone of 16 selected rates. 

The prf trigger is the basic trigger and all other signals are generated 

with respect to this trigger. The ALTITUDE switch selects one of eight 

voltages for routing to the delay generator. Thus the d-c level selected 

as an input to the delay generator is determined by the setting of altitude 

and prf. The output of the delay generator provides a trigger which in 

turn generates a square wave. The square wave generates the sine wave 

mentioned above. 

The negative-going edge of the sync pulse is delayed 10 microseconds from 

theprf trigge~. The start of the stc wave form is delayed approximately 

20 microseconds from the negative-going edge of the sync pulse. The 

negative-going edge of the sync pulse is differentiated and is called "scope 

sync". 

£. Video Simulator 

The video simulator was an airborne version of the prfselection simulator. 

The video simulator provided an output wave form representative of only 

the correct prf setting at nominal vehicle altitude. The video Simulator 

was used as a signal source to provide modulation of a Lockheed Type vm 

transmitter. Both the simulator and the transmitter were placed on board 

a jet aircraft, which was used to simulate the video link output of the orbital 

vehicle during collimation of the 60-foot receiving antenna at the tracking 

stations. While collimation checks were being conducted, the tracking 

station personnel became familiar with the signal and the tuning techniques 

that were required during orbital operations. 

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SECTION IV - GROUND TEST PROGRAM 

1. GENERAL 

AKP-U-596 

This section covers the tests which were performed for both qualification and 

acceptance of the radar equipment. Each test is briefly described with particular emphases placed on the problems encountered. The material in this 

section provides a background for the more detailed discussion in Section V 

of problems and corrective actions. 

z. QUALIFICATION TEST PROGRAM 

The qualification tests were begun at Goodyear Aerospace and completed at 

LMSC. The tests were performed on a radar system (Serial No.1) which 

was set aside for this purpose. The tests were performed at both unit and 

system levels. The qualification test program will be discussed with the aid 

of the flow chart in Figure 16 • 

a. Unit DeSign 'Approval Tests 

Following fabrication. each unit was subjected to an Adjustment and Calibration Procedure (ACP) as a preliminary to the Design Approval Test Procedure 

(DATP). The DATP consisted of a low level vibration test and a thorough 

functional test. The units came through these tests with very few problems. 

The individual ,unit maintenance testers (refer to Section U: Ground Support 

Equipment) were used to perform the tests. 

b. System DeSign Approval Tests 

Following the unit DATP's, the units were tied together in a system configuration with the system tester (refer to Section U) and subjected to the 

system ACP and then the system DATP. 

The system design approval tests were selected to verify that the system 

was operating as designed and that it would deliver the required radar 

data with sufficient quaUty to meet the design goals. The results of system 

design approval testing illuminated some incompatibilities between the 

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UNIT I AND H DESIGN APPROVAL 

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SYSTEM I AND H DESIGN APPROVAL 

TESTS CALIBRATION TEST PROCEDURE . PROCEDURE 

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DESIGN APPROVAL H ENVIRONMENTALH DESIGN A

TEST PROCEDURE QUALIFICATION TEST PROC

PAYLOAD TESTL-..ITEMPERATURE-H ~~iA HTEMPER

PROCEDURE 02 r-IALTITUDE TESTS ALTI
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radar system and tester. Additionally indicated were other problems 

relating to system timing, conditioned signal outputs, recorder crt 

bias, and repeatabUity of fl1m exposure and development. 

All of these problems were corrected during the testing phase. In one 

area. the original design approval tests failed to uncover a serious 

system problem. This problem. in connection with the clutterlock operation. 1& discussed in detail in Section V. paragraph 7. £.. 

c. Unit Environmental Tests 

At the conclusion of the design approval tests. the Serial No. I radar 

system was shipped to LMSC for environmental qualification testing. 

There. the electrical portion of each unit design approval test was repeated to ensure proper unit operation after shipping. The individual 

units were then subjected to vibration. shock. and temperature-altitude 

environmental tests in accordance with the specification requirements. 

The problems which appeared during environmental testing are mcluded 

In the discussion of problems and corrective acUoBsin Section V following • 

(1) Vibration 

The units experienced few serious problems in the vibration tests 

but a number of attaching hardware problems appeared. A switch 

came loose on the Transmitter-Modulator. Two loose choke colls. 

a loosened connector, and a broken wire on a printed circuit board 

were found In the testing of the R-F/I-F unit. In the Recorder a 

choke coll vibrated loose. The Reference Computer and Control unit 

suffered no damage. The most serious problem occurred in the 

Recorder when the crt changed brightness. This necessitated a 

later redesign of the gun structure. 

(Z) Shock 

The problems encountered during shock testing were not serious and 

were easily corrected. The Transmitter and Contr~l unit. had no': 

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failures. In the R-F /I-F unit a cold solder joint was found after testing. 

The Reference Computer developed a short circuit in a coaxial insert. 

with an additional short circuit caused by human error during testing. 

The power supply studs came loose in the Recorder and the inserts 

pulled out of some lead weights on one of the isolation roller assemblies. 

(3) Temperature-Altitude 

Many problems which were encountered later in temperature-altitude 

testing at system level in the temperature-altitude-simulation (TAS) 

chamber did not show up in the TAS unit testing phase. A single highvoltage problem appeared in the Transmitter when arcing in the plotted 

connector of the pulse transformer caused erratic operation. In the 

R-F !I-F unit the pre-ampUfier displayed symptoms of temperature 

instability. An instability of r-f power output. possibly caused by a 

loose connection. disappeared when the pre-amplifier was replaced. 

The Reference Computer and Recorder underwent the tests without difficulty. In the Control unit a power supply overload recovery problem 

appeared which was traced to a diode which did "not meet all the design 

requirements as a function of input voltage and load conditions. The" 

problem was corrected by selecting an improved component. 

At the conclusion of the unit environmental tests. each unit was retested to the uni~ DA TP and the results compared with the results of 

the same procedure run prior to environmental testing. 

d. System Temperature-Altitude Test 

The units were subsequently assembled in the vehicle nose section and 

checked to Payload Test Procedure PLTPOZ. The complete nose was then 

placed in the TAS chamber for environmental qualification at the system 

level. Several difficulties were experienced with the Transmitter-Modulator during the temperature-altitude tests. T~e thyratron triggering 

circuit was shown to be voltage sensitive and design changes were made. 

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A klystron and inverse diode were inadvertantly destroyed because of a 

test error· and a pulse-forming network failed. After these malfunctions 

were corrected. twelve simulated· orbits were completed without a failure 

and the system was sent to a test range for field tests with the antenna. 

!.. Santa Cruz Tests 

The field series of tests was conducted at an LMSC antenna test range 

in the coastal mountains near Santa Cruz. California. The radar and 

antenna were mounted on a tower on one peak and a pair of corner reflectors placed on an adjacent peak. The tests. in addition to showing 

the resolution to be within specifications. provided a valuable check on 

round trip gains and losses through the radar system and antenna. The 

only malfunction was a t-r tube which failed during the tests and had to 

be replaced. 

f. Repeat of System Temperature-Altitude Tests 

Following the Santa Cruz tests the temperature-altitude tests 'were 

• repeated in the TAS chamber. Momentary interruption of the highvoltage power supply was experienced during several operate cycles. 

The difficUlty was traced to a faulty inverse current overload relay 

in the Transmitter-Modulator. 

• 

3. . FLIGHT EQUIPMENT TEST PROGRAM 

a. General 

The flight radar system was tested as shown in the flow chart of Figure 17. 

The test program included both unit and system tests. 

b. Unit Acceptance Tests 

The formal acceptance tests were started at Goodyear Aerospace Corporation. Each unit upon completion was coupled to its particul~r unit maintenance tester and subjected to a component Adjustment and Calibration 

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TESTS CALIBRATION TEST 

PROCEDURE PROCEDURE 

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Procedure (AC P) and then to an Acceptance Test Procedure (ATP) designed 

to prove that the unit was functioning correctly. 

c. System Acceptance Tests 

Upon completion of the unit acceptance tests, the units were cabled together 

with the system tester and subjected to the system ACP and then to the system ATP. Upon completion of the system ATP, the units comprising the 

system were shipped to LMSC. 

d. Payload Tests 

(I) Agena Nose Section Tests 

After arrival at LMSC, each unit was retested to the unit ATP to 

ensure operation after shipping. The units were then installed in 

the Agena nose section and tested to the first of the payload test 

procedures fPLTPOZ.). This procedure, which was similar 1il content 

to the system ATP, was designed to show proper operation of the 

radar set with the associated equipment in the barrel. A battery 

simulator was used as the primary power source. 

(Z) Agena Compatibility Tests and Anechoic Chamber Tests 

The nose section was next moved to the Agena assembly area for preliminary compatibility tests, anechoic chamber tests, and final compatibility tests. The anechoic chamber tests were conducted with the 

entire Agena assembled in the anechoic chamber and tested as an 

entity. self-powered by batteries and remotely controlled. No 

troubles were encountered with the payload and the tests were completed in June, 1964. 

(3) System Temperature-Altitude Tests 

In this test, the nose was installed in the TAS chamber and a series 

of tests run which were designed to simulate actual orbit times • 

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temperatures. and pressures. The problems which occurred during 

this test included one voltage breakdown each in the klystron, inverse 

and charging diodes, and pulse transformer. Also included were 

capacitance changes in some of the radio frequency interference filters and a short circuit in one radio frequency interference filter. 

These problems. which occurred from September through November, 

1964, are discussed in detail in Section V. At the time of completion 

of the test on 1 December 1964. a total of 152 hours had been spent 

at simulated altitude in the TAS chamber. 

(4) Prelaunch Tests 

Following the TAS chamber tests. the radar system was shipped to 

Vandenberg Air Force Base, California. for final tests prior to 

launch. A bearing failure in the data film supply cassette resulted 

from a human error. This malfunction was corrected and the satellite was launched on December 21, 1964. 

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SECTION V - PROBLEMS AND CORRECTIVE ACTIONS 

1. GENERAL 

AKP-D-596 

The problems which appeared during the ground testing of the radar set mostly 

originated as a direct result of the high-altitude environment and, to a lesser 

extent, to the high vibration levels. This section discusses the unit and system 

. failures and problems which appeared during testing. together with the correctivE 

actions which were taken to prevent recurrence. 

Z. TRANSMITTER-MODULATOR 

Failures of this unit usually resulted either from vibration or froom the combination of high-voltage and high altitude environment. 

a. Vibration Problems 

Early in the program it was decided that all of the units would be hard 

mounted to the vehicle structure. It soon became apparent that this 

approach was not adequate for the Transmitter-Modulator. Vibration 

amplifications up to 50 times at various frequencies between 100 and 1000 

cps were experienced. Accordingly, a center-of-gravity mounting using 

stiff vibration isolators was designed. This type of mounting minimized 

the vibration problem •. Table VI lists the various problems encountered 

in the Transmi~er-Modulator. Of these problems, items 1, Z, and 3 

can be classified as resulting from vibration. 

~. Pressure Altitude Problems 

Those problems which can be classed as voltage breakdown resulting from 

the high altitude environment include items 4 through lZ of Table VI. These 

failures were invariably initiated during exposure to reduced pressure and 

resulted from various causes such as improper bonding of potting materials 

at high-voltage points, voids in potting, air bubbles or reduced pressure 

in oil filled units, and contamination of potting. They were corrected by 

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extensive testing in altitude chambers. thorough failure analysis programs. 

and close process control • 

.Q.. Other Causes 

Items 13 through 16 were miscellaneous problems. Items lZ and 13 were 

design deficiencies requiring design changes. Items 14. 15 and 16 required 

no corrective action other tbau parts replacement since they were either 

secondary failures or resulted from use of substandard parts. 

3. RF-IF 

The majority of the problems in this unit. as in the Transmitter-Modulator. 

were the result of the high voltage in combination with the high altitude environment. In Table VU, items I. Z. and 3 are vibration problems and items 5. 6. 

7. 8 and 9 are high voltage/high altitude problems. Items 10 and 11 are failures 

involving only replacement of parts in test and monitor circuits. Items lZ. 13 

and 14 were design deficiencies requiring circuit redesign. 

4. REFERENCE COMPUTER 

This unit was very reliable throughout the program. Item I. Table vm. was a 

design deficiency associated with the clutterlock and built-In-test (BIT) circuitry 

and was corrected by circuit modifications and adjustments. (This problem is 

discussed in paragraph 7. £. •• ) Item Z was a workmanship problem. 

5. CONTROL UNIT 

This unit was very reliable and had only two problems (refer to Table VUI). 

The first was an end seal problem in the rfi filters which was solved by elimination of the filters from the circuit. These particular filters were found to 

be unnecessary. The second problem. failure of the +300 vdc supply to 

recover following a short circ'utt overload. occurred only for low input voltages or low temperatures and was corrected by a component type change. 

-5ZSPECIAL HANDLING 

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4 I 

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7 3 

8 5 

9 z 

10 4 

11 z 

.. " 

TABLE VI - TRANSMITTER-MODULATOR PROBLEMS 

Problem description 

Overload switch vibrated loose 

Klystron filament shorted 

Charging choke open, broken wire 

High-voltage breakdown, potted area 

around pulse transformer bushing 

Klystron breakdown, potted area 

around cathode 

Pulse-forming network, shorted 

capacitors 

Pulse transformer, internal breakdown, shaping coil to ground 

Thyratron high-voltage breakdown 

C rac-kedklystron window 

Inverse diode, high-voltage breakdown 

. CJlarging diode, high-voltage breakdown 

Corrective action 

Redesign switch mounting 

Reduce heater feed-through diameter 

to provide better support 

Change to stranded wire, flll with 

epoxy sand 

Change potting procedure 

Ellminate vendor use of silicon 

grease prior to potting. Contamination problem 

Redesign case to maintain oil·pressure at high altitudes. Use bellows, 

epoxy sand, stiffer case 

Redesign shaping coil mounting. Use 

fiberglass insulation. change tuning 

slug 

Modify metal container, change potting procedUre and material 

None. Test error. wave guide pressure 

Change to epoxy potting. Change case. 

X-ray diodes 

Change to ~poxypott1ng. Change .case. 

X-ray diodes 

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. TABLE VI - TRANSMITTER-MODULATOR PROBLEMS (Continued) 

Problem description 

Breakdown of high-voltage interconnecting leads 

Thyratron multiple triggering 

Apparent inverse overload and thermal switch operation 

C orrecti ve action 

Change method of shield braid termination 

Redesign triggering circuit 

Replace bad relay. By-pass inverse 

and thermal overload circuits for 

night 

Delay timer 'inoperative I Replace timer. By-pass delay timer 

for flight 

Inverse resistors, high-voltage break- I None. Secondary failures 

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I 

I 

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I 

I 

I 

I 

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- . .. 

TABLE vn - R-F/I-F PROBLEMS 

Problem description Corrective action 

Varactor frequency multiplier detunlng I Bonding of coll turns to coil form. 

as a result of vibration Bonding of capacitors. Use of foam 

potting 

R-f choke colls broken loose in I-f I Design change. better fastening 

switch because of vibration 

Broken wire on printed circuit board 

because of vibration 

Design change, board attachment 

Cold solder joint to relay. Discovered Resolder 

in vibration testing 

2-kv (twt) supply voltage breakdown Use of potting 

Varactor frequency multiplier voltage Use of foam potting 

breakdown 

TWT high voltage breakdown I Vacuum potting of high voltage leads 

Voltage divider high voltage breakdown Redesign for better spacing. Change 

to epoxy potting material 

Spiking indicated on 2-kv supply moni- DeSign change 

tor 

Defective coaxial test connector Replace 

Defective temperature monitor Replace 

Defective twt output monitor circuit 

Defective r-f power monitor 

Poor wave form, r-f to transmitter 

monitor 

Design change 

Design change 

Redesign signal conditioner circuit 

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Number 

of 

fallures Problem description Corrective action 

Reference Computer 

- Clutterlock circuit not operating proper- 1. Filter addition, On-Gate Une 

ly and BIT loop gain incorrect. Design Z. Addition of ground plates and 

deficiency straps on coaxial connectors 

3. Circuit changes 

4. Change BIT loop gain 

- Short in coaxial connector Repair. (Workmanship problem) 

Control Unit 

Z Rfi filters, oil leaks Elimination of filters from circuit 

- :l:Z3 vdc supply failure to recover from Circuit design change (component 

short circuit overload. Design defi- change) 

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6. RECORDER 

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AKP-D-.596 

The Recorder unit had very few problems (refer to Table IX). Item I was a 

light level shift in the crt induced by vibration or shock. This problem was corrected by a vendor redesign of the crt gun structure. Items Z and 3 were 

also vibration problems. Items 4 and 5 were intermittent and shorted capacitors in the rfi filters. One filter type was eliminated from. the circuit and the 

other was simply replaced. Item 6 was an open magnetic brake coiL Item 1 

was a temperature sensitive light monitor circuit which was redesigned. Item 

8 was not a fallure since the drive speed was well within specifications. The 

meter drum drive was changed to further minimize short term fUm speed 

variations. 

1. SYSTEM PROBLEMS 

During the system design approval and acceptance tests at Goodyear Aerospace 

and during the following tests at LMSC, certain system problems arose which 

required system modifications. 

a. Timing 

During the design approval tests the chirped pulse (see Volume I, Figure 11) could not be made to fall within the klystron high-voltage pulse. 

The source of error was traced to a O. Z5-microsecond delay occurring in 

a pulse transformer in the transmitter. The transformer was replaced 

with an avalanche pulse circuit which eliminated most of the delay. A 

O. I-microsecond delay line was later added in the transmitter trigger 

line to further improve the timing between the chirped and klystron pulses. 

b. Line Fllter 

When the radar system was assembled with the 400-cps inverter power 

source in the Agena, the filters contributed substantially to a poor power 

factor which in turn created a severe distortion of the input wave shape to the 

Recorder as a function of load changes. The distortion was particularly 

objectionable as the crt regulated high-voltage supply output and therefore 

the sweep amplitude is affected by distortion. The sweep amplitude t.ended 

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1 

1 

1 

1 

1 

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TABLE IX - RECORDER UNIT PROBLEMS 

Problem description C orrecti ve action 

Light level change in crt because of Redesign of crt gun by vendor 

vibration 

Inserts in lead weights on twiddle roller Material change, lead to stainless 

assembly pulled out during vibration s,teel 

Choke on printed circuit board damaged Component bonded to board 

dur'ing vibra.tion 

Rft filter oil leakage problem Eliminate filters from, circuit 

Rft filter shorted Replace filter 

Magnetic brake •. open col1 ' Replace co11 

Photocell in crt light monitor tempera- Change photocell type, change 

ture sensitive circuit 

Variations of film drive speed Change meter drum drive to minimize variations 

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to change as the drive motor. which also uses 115 vac. 400 cps. was 

turned on and off. The rfi filters were replaced with insulated feed 

throughs and the power factor correction capacitor values were changed 

for the drive-on condition. A near-unity power factor was finally obtained which greatly reduced the wave shape disto·rtion. 

c. Clutterlock 

The original system tests failed to point out a problem involving excessive 

error in the clutterlock loop. The problem was finally brought into focus 

late in the system test program at LMSC. Test work at Goodyear Aerospace pinpointed the principle source of error as leakage of 70-mc signals 

(both c-w and prf/4 phase stepped) between the Reference Computer and 

the RF-IF. These leakage signals were entering th~ RF-IF onthe OnGate line and on the other interconnecting cables. The ~hields of these 

cables were "hot" because of inadequate grounding in the Reference 

Computer. 

The corrective action consisted of the addition of braided ground straps and 

grounding plates on each printed circuit board connector containing shielded 

coaxial pin connectors. In addition. a small potted r:"f rejection filter was 

installed iil the On-Gate line just prior to its output to the R-F /I-F unit. 

These changes proved adequate to give the clutterlock loop of flight system 

No. 1 an operational capability at least 10 db better than that required. 

d. Crt Bias Setting and Video Gain 

As the optimum brightness setting of the crt could not be verified by a 

series of flight tests as in a piloted aircraft radar. a great deal of study 

and ground experimentation was required in advance of the Agena launch. 

The crt bias level and video gain settings are major factors in obtaining 

high over-all system performance. In the KP-II system three recorders. 

one airborne and two ground based. were used. each requiring proper 

setting prior to flight. 

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SECTION V 

To establish the crt bias point. it was first necessary to measure the crt 

transfer function of bias volts versus light output. A microphotometer 

was used to make fUms having incremental steps of light output for film 

developing tests. Various developer types and processing times were 

evaluated to provide the desired range of film densities (or square root 

of film transmissivities. '\}T>. The values of"iF were then plotted as a 

function of crt bias to produce an S-shaped curve. The high transmissivity 

end of the curve is determined primarily by the cut-off characteristics of 

the crt and the low transmissivity end by the maximum density of the fUm. 

From the theoretical standpoint of linear recording. the bias point should 

coincide with the VT being equal to 50 percent. However. measurement 

of the relative amount of fUm noise as. a function of iT showed that film 

noise could be reduced by choosing the bias point slightly lower than the 

50 percent value. 

The video gain was then selected to linearly record. small signals and to 

allow large signals to be limited by the crt-film characteristics. This 

choice was made to ensure adequate ground painting •. 

To resolve these problem areas, close coordination was required between 

Goodyear Aerospace, LMSC, and, especially, 

the Air Force Satellite Photographic Processing Laboratory (SPP L). 

8. PROBLEM STATUS AT LAUNCH 

a. General 

There were four radar problems remaining just prior to launch. These were 

evaluated and judged to be not serious and are discussed here for completeness. 

b. Trans mUter Power Output 

The transmitter power was 0.5 db below the minimum specified in the test 

specification (speCification value was 74. 7 dbm; actual was 74. Z dbm). 

This could have been corrected by a re-adjustment to increase the drive 

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power from the twt, but the problem did not warrant removal of the R-F!I-F 

unit to make last minute adjustments. 

c. Sensitivity Time Control Wave Form 

The sensitivity time control wave form was outside of the specification 

limits in two respects: the ratio of maximum to minimum amplitude 

should have been 2.0 :0. 1 but was actually 2.2; and the delay of the 

sensitivity time control start should have been 31 microseconds maximum 

but actually measured 35 microseconds. It was decided that this condition 

would not degrade the final maps. 

d. TWT Output Interruptions 

The T AS chamber tests exhibited a number of momentary spikes on the 

-2 kv monitor FII, and the r-f power to transmitter monitor F-72 during 

high altitude operation. Extensive retests of the R-F /I-F unit in an altitude chamber at Goodyear Aerospace indicated the problem to be nondestructive in nature. 

e. Accelerometer 

Tests showed the accelerometer in the Transmitter-Modulator unit to be 

inoperati vee As this accelerometer was used only for flight test instrumentation, no corrective action was taken • 

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1. MAPPED AREAS 

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SEC TION VI - ORBITAL OPERATIONS 

AJ(P-U-596 

The areas mapped during the night test are shown in Figure 18. Each trace on 

the map represents a mapped interval which is approximately 10 nautical miles 

wide (5. 95 nautical miles in slant range) and several hundred miles long. Table 

X gives the approximate latitude and longitude (determined by examination of 

the map film in all but one pass) of the start and finish of each pass. 

Z. PRE-RECOVERY OPERATING MODES 

For the first seven payload operate periods (passes 8. 9, 14, 16, Z4. Z5,and 30), 

no attempts were made to determine the radar system limits. as the primary purpos~ of the mission was to demonstrate the feasibility of operating a satelliteborne radar system and early experiments could have placed the primary goal in 

jeopardy. Immediately after pass 8 the data obtained from the wide band telemetr: 

link were sent by courier to be processed and correlated at the 

_ This step was taken to verify that the radar was operating properly. During 

these first seven active passes only OPERATE. PRF a;ndOFF commands were 

sent to the radar system. The necessary prf commands were determined from 

the A-scope presentation in the screen room at the traoking ·stations. 

3. POST-RECOVERY OPERATING MODES 

On pass 33 the data· film from the satellite was recovered. On the next active 

pass (40), the automatic gain control (agc) circuitry in the radar was commanded 

off and fixed amounts of receiver (r-f/i-f) attenuation (step Z and step 3) were 

selected. 

Again, on pass 41. the radar was operated with several steps (0. 3, 4, 5) of 

fixed attenuation. 

No experiments were conducted on pass 46 but on pass 47 the prf was stepped 

through all of its 16 positions. On pass 56 a command to short the integrator 

output (clutterlock circuitry) was sent. On pass 57 clutterlock time constant 

number Z was selected and on pass 7Z additional fixed attenuator settings were 

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SECTION VI 

selected. Table XI shows the status of the various radar functions. which 

were subject to commanding from the tracking stations, during the active 

orbital life of the vehicle. Table XII is a time analysis of the mission. 

Pass Pass 

no. direction 

8 S-N 

9 S-N 

14 N-S 

16 N-S 

Z4 S-N 

Z5 S-N 

30 N-S 

40 S-N 

41 S-N 

46 N-S 

47 N-S 

56 S-N 

57 S-N 

7Z S-N 

* S-Look data 

TABLE X - ORBITAL PASS COORDINATES 

Start 

North West 

latitude longitude 

38° 

3Zo 

4Zo 

41° 

39° 

33° 

47° 

40° 

3Zo 

45° 

44° 

41° 

3Zo 

46° 

56.0' 86° 58. S' 

Z8.9' 11Zo 33.5' 

I 

03.0' 79° 51.0' 

Z7.5' lZ4° 48.5' 

Z3.5' 88° 49.0' 

11.5' 114° ZZ.5' 

Z3.5' 85° Z9.8' 

40.1' 90° 08.0' 

4Z. Z' 116° 46.0' 

37.5' 86° ZI.5' 

47.5' 108° Z3. l' 

41.8' 91° 37.8' 

51.6' 118° Z8.7' 

17.5' 90° 50.0' 

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North 

latitude 

45° 47.5' 

39° 46.5' 

34° 46.Z; 

35° 55.1 

46° 51.3' 

4Zo 00.9' 

.. 

34° 09.0' 

46° 54. I' 

41° Z4.Z' 

34° 06.5' 

33° 57.5' 

47° 38.Z' 

43° 34.Z' 

51° 01.Z'* 

End 

West 

longitude 

8Zo 5Z.5' 

109° 06. Z' 

76° IZ. I' 

lZ1° 49.0' 

84° 19. 0' 

109° Z4.6' 

78° 01.5' 

86° ZO.5' 

llZo Z7.0' 

80° 10.0' 

10Zo 4Z.5' 

87° 58.0' 

113° ZZ.5' 

87° 15.6'* 
1 

.... , .. :; .~':.~. -::r~" . 

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SECTION VI ____ .....:AK=P-U-596 

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Flpre 18 - Traces Depicting Radar Coverace of United States by 

The KP-U Radar, Deoember 1964 

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TABLE XI - STATUS OF RADAR FUNCTIONS 

Automatic Fixed 

Pass rain attenuator 

no. control steps 

8 Yes None 

9 Yes None 

14 Yes None 

16 Yes None 

Z4 Yes None 

Z5 Yes None 

30 Yes None 

33 

40 App. Z, 3 

Ssec 

41 No 4, 5, 4, 

3, 0 

46 Yes None 

47 Yes None 

56 Yes None 

57 Yes None 

7Z App. O. 7 

36sec 

Prf initial Prf 

(at OPERA TE on) changes 

step step 

14 13-lZ11-10 

10 5-6-7-8-9 

10 9-8 

10 None 

11 10 

11 10-9 

8 9-8 

Reoovery of Data Film 

8 None 

9 None 

8 7-6 

6 All 

7-15. 0-6 

6 7-8 

9 8 

8 7-6 

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Clutterlock Clutterlock 

time integra:tor 

constant output 

TC No. 1 Normal 

TC No. I Normal 

TC No. 1 Normal 

TC No. 1 Normal 

TC No. 1 Normal 

TC No. 1 Normal 

TC No. I Normal 

TC No. I Normal 

TC No. 1 Normal 

TC No. 1 Normal 

TC No. 1 Normal 

TC No.1 Shorted 

(approx 7 sec; . 

TC No. Z Normal 

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TABLE XU - MISSION TIME ANALYSIS 

Payload pre- Total payload 

Payload warm- operate time Qperate (Tw +T~ +To) 

Active pass up time (Tw> (T ) (time (Tp> cyc e 

no. (min) (mPn) (min) 

8 5.5 2.5 2.0 

9 4.5 3.8 2. 1 

14 4.5 3.6 2. 1 

16 4.5 2.4 1.6 

24 4.5 1.4 2.2 

25' 4.5 3.4 2.8 

30 4.5 2.2 3.7 

40 4.5 2.4 1.8 

41 4.5 4.0 2.4 

46 4.5 2.7 3.2 

47 5.1 0.2 3.0 

56 4.5 2.3 1. 7 

57 •• 5 4.1 2.9 

72 4.5 3.3 1.4 

Totals 64.6 38.3 32.9 

Total payload cycle time (T + T + T ) = w p 0 2.3 hr 

Time from launch to pass 8 = 11.4 hr 

Pass time = 93.5 hr 

Total orbit time = 107. 2 hr 

(from launch to 

end of pass 72) 

·-68-

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(min) 

10.0 

10.4 

10.2 

8.5 

8. 1 

10.7 

10.4 

8.7 

10.9 

10.4 

8.3 

8.5 

11.5 

9.2 

135.8 

SECTIQN VI 

T1Dl~ to next 

active pass 

(hr) 

1.3 

7.7 

2.8 

11.4 

1.3 

7.7 

14.4 

1.3 

7.7 

1.3 

12.9 

1.3 

22.4 

93.5 
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SECTION VU - ENVIRONMENTAL FLIGHT DATA 

1. GENERAL 

AJCp-n-596 

. The various units of the KP-U radar system were instrumented so that component and unit characteristics could be monitored under the thermal and 

dynamic environments encountered during ascent and orbital conditions. 

Sensors were installed to measure unit and component temperature, vibration, 

shock, and pressure responses under the flight conditions. 

Thermistors (Fenwal 918. 919. and General Electric DO 53) were used as the 

temperature sensors in the equipment. These resistive elements are used 

in circuits energized by the vehicle's Z8. 3-volt regulated power supply. Resi stance of these thermistors is inversely proportional to the component 

temperature. The conditioned signal of the monitor circuit thus decreases as 

the body temperature increases, and vice versa. These sensors were mounted 

on particularly high heat dissipating or temperature sensitive components all 

well as on unit case~. Skin temperatures of the conical support and the cylindrical vehicle section were monitored. Temperatures on the cylindrical section 

we~_e measured in the area of the Transmitter-Modulator and the Control unit. 

Piezoelectric accelerometers were used to measure equipment dynamic 

response. These devices were used in conjunction with ampUfiers for supplying conditioned signals to the telemetry subcarrier oscillators. The monitor 

system has a frequency response from 0 to ZOOO cps and is capable of measuring acceleration levels up to 15 g (zero to peak). Accelerometers were 

mounted on the primary structure of the Transmitter-Modulator, R-F/I-F, 

and the Recorder units. These sensors were oriented in the vehicle.longitudinal (X-X) direction. 

Pressure sensors were installed in the vehicle payload compartment and in the 

equipment to measure in-flight pressures. Five Pirani gages were used, two 

sensors were placed in the vehicle payload compartment. and one sensor each 

in the Transmitter-Modulator. Recorder. and in the interconnecting wave guide 

between the Transmitter-Modulator and BF-lF. A complete list of the various 

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monitor sensors used to determine equipment environmental characteristios 

are given in Tables XIn, XIV, and XV. 

Flight temperature, vibration, and pressure information was relayed to ground 

tracking stations by narrow band telemetry. Flight results and analyses of 

these results are presented in the following paragraphs. 

Z. FLIGHT TEMPERATURE DATA 

a. General 

Equipment temperatures were monitored during vehicle ascent and during 

each active (equipment operating) orbit. Temperature measurements 

were made primarily during the equipment OPERA TE-ON mode although 

temperature data were also obtained before and after the active phase of 

several orbits. Temperature data taken before the OPERATE-ON mode 

included data from the payload WARM UP and PRE-OPERA-TE modes. 

Most components contained within the various units, with the exception 

of the Transmitter-Modulator, utilized full power during both the PREOPERATE and OPERATE-ON modes. 

b. Ascent Heating 

Aerodynamic heating occurring during ascent resulted in slight heating of 

the radar equipment. From an initial temperature of 55-60 degrees F 

(precooled on launch pad) the unit temperatures specified in Table XVI 

were experienced. 

The Recorder unit located in the vehicle conical section (estimated maximum 

average skin temperature - 460 degrees F) reached the highest temperature. 

This maximum temperature occurred approximately Z800 seconds after 

lift-off. The Transmitter-Modulator and the R-F !I-F units located in 

the vehicle cylindrical section reached their maximum temperatures 

approximately 4000 seconds after lift-off. The maximum recorded 

ascent temperature of the Reference Computer' and Control units occurred at 5400 seconds after lift-off and was probably stUl increasing 

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TABLE xm - FLIGHT INSTRUMENTATION: TEMPERATURE MONITORS 

Unit 

Control 

Reference Computer 

Recorder 

RF-IF 

Transmitter-Modulator 

Telemetry 

number 

FZO 

FZl 

FZZ 

FZ3 

FZ4 

FZ5 

Fz6 

FZ7 

FZS 

FZ9 

F30 

F31 

F3Z 

F33 

F34 

F35 

F36 

F37 

Locations 

Power supply 

Unit case 

Video amplifier 

Gate sequencer 

Sine x/x pulser 

Timer A 

Audio synchronizer demodulator 

Unit case 

Deflection amplifier 

High-voltage power supply 

Unit case 

Traveling wave tube (twt) 

Frequency multiplier 

Transmit-receive tube (t-r tube) 

Klystron collector 

Klystron body 

Unit case 

Pulse transfo.rmer 

TABLE XIV - FLIGHT INSTRUMENTATION: VIBRATION MONITOR 

Unit 

Recorder 

RF-IF 

Transmitter-Modulato r 

Telemetry 

number 

FS3 

FS4 

FS6 

Locations 

Primary structure - Longitudinal 

axis 

Primary structure - Longitudinal 

axis 

Primary structure - Longitudinal 

axis 

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TABLE XV - FLIGHT INSTRUMENTATION: PRESSURE MONITOR 

Telemetry 

number 

FIOO 

FIOI 

FIOZ 

FI03 

FI04 

Pressure monitored 

Vehicle payload compartment 

Vehicle payload compartment 

Recorder unit - Internal 

Transmitter-Modulator unit - Internal 

Wave guide between the Transmitter-Modulator 

and the R-F/I-F units 

TABLE XVI - UNIT TEMPERATURES DURING ASCENT 

Unit 

Transmitter-Modulator 

RF-IF 

Reference Computer 

Control 

Recorder 

Temperature 

5400 seconds 

after 11ft-off 

(F) 

80 

85 

75 

80 

9Z 

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Maximum 

ascent 

temperature 

(F) 

86 

93 

75 

80 

98 

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slightly. This condition occurs because of the unit geometry and location 

and local vehicle skin condition. 

c. Orbital Temperature Data 

A summary of unit and component orbital temperature data is presented 

in Figures 19 through 24. Figure 19 shows typical temperature cycles of 

two components. One component is energized during both the PREOPERATE and the OPERATE-ON modes, and the other component dissipates 

power primarily during the OPERATE-ON mode. Component cooling is 

shown for 1260 seconds after the operate-off command. 

The orbital heating characteristics for the Transmitter-Modulator, R:'F /I-F, 

Reference Computer, Control, and Recorder units are given in Figures 

20 through 24. These plots show the component temperature rise during 

the OPERATE-ON mode for each active orbit. It is noted that three 

temperature monitors - F24, F25, F26 -are not plotted. These thermistors, located within the Reference Computer unit, indicated a maximum 

orbital temperature of 91 degrees F with cyclic variations similar to other 

monitors within the Reference Computer unit (F22 and F23). 

No temperature response curve is presented for the klystron collector 

(Transmitter-Modulator, F34). The temperature sensor for this component was calibrated over the temperature range 350-575 degrees F. 

Conditioned sig~al data indicated that the klystron collector temperature 

was below the 350 degree F calibration minimum during all operating 

phases. No information on temperature variation with time was obtained. 

In analyzing temperature data, a slight temperature drop was found to 

occur when the equipment was energized. This abnormal temperature 

variation is believed to be caused by a decrease in voltage output of the 

regulated 28. 3-vdc power supply during the equipment operating modes. 

This is the voltage source for the temperature monitoring circuits and 

thus temperature compensation was required. Temperature data were 

normalized accordingly. 

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~ 

V V I 

~ 

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-- _ . 

.... -

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-_ ... --- -.. 

(:I 5331030) 31t'1.1 W3dW3.1 (:I S331030) 3Wl.1 W~3.1 

Figure 19 - Typical Component Temperature Cycle 

-74-

SPECIAL HANDLING 

SECRET 

SECTION VII 

I i 

I, 

§ 

si 

§ 

.. R 

~ .. R 

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fit' 

0 

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F35-KL YSTRON BODY 

110 

-I&. MAXIMUM ~LLOWAILE TEMPERATURE '- 2SO DEGREES F 

0 

1&1 100 e. 

1&1 

lie 

~ 90 i 

1&1 

L 

~ 80 III 

~ 

70 

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0 

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tL 100 

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F36-UNIT CASE 

PREDICTED MAXIMUM TEMPERATURE - ISS DEGREES F 

- - -

~ L- _ ~ -~ I.. - I-

-

F37-1ULSE TRANSFORMER 

MAXIMUM ALLOWABLE TEMPERATURE - 185 DEGREES F 

-

.... - -~ -". -~ -~ 

- -". 

OPERATING TIME 

AKP-D-596 

~ " ~ L1 ~ 

I -r"""'" 4fII 

"--

I""'" - -I- ~ 

-I- -

4fII " -

". 

- -

(MINUTES) 2.0 2.1 2.1 

14 

1.6 2.2 

16 24 

2.8 3.7 

25 30 

1.8 2.4 3.2 3.0 1.7 2.9 1.4 

PASS NO. 8 9 «) 41 46 47 56 51 72 

<1000-18 

Figure 20 - Orbital Temperature History - Transmitter:Uodulator Unit 

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F33-TRANSMIT-IECEIVE TUIE 

120 

MAXIMUM AlLOWABLE TfMPERATlJtE". f 

-

Y- 110 

O 

III 

e. 100 III 

~ ~ 

~ 90 

I 

III 

~ 80 

70 

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III 

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~ 70 

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~ 

~ 70 

185 DEGREES F 

1/ 

I ~ ~ II ~ j I J 

I J J I 1 I I I 

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F3HREQUENCY MULTIPLIER 

MAXIMUM ALLOWABLE TEMPERATURE· 149 DEGREES F 

.... ....... ... I"" ... I"" - I- - -~ 

-

F31-TRAVELING WAVE TUIE 

MAXiMUM ALLOWABLE jEMPEIATlJtE:' 185 DEGREES F 

" --I- .--- ~ 

, -""'" ... - , / , 

SECTION VB 

I , 

V , ~. 

J I J ~ , J I I I 

I- - I- - 10- ... 10- ... ... ... ~ 

~ ... " ... I.-

, ~ .." ... r" 

OPERATING TIME 

(MINUTES) 2.0 2.1 2.1 1.6 2.2 2.8 3.7 1.8 2.4 3.2 3.0 1.7 2.9 1.4 

PASS NO. 8 9 14 16 24 25 30 «) 41 46 47 56 51 72 

.fOOO-19 

Figure Zl - Orbital Temperature History - R-F !I-F Unit 

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I. 

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F22-VIDEO AMPLIFIER 

MAXIMUM ALLOWABLE TEMPERATURE - 212 DEGREES F 

130 

- I 

'( 

"-

0 120 ~ it I I i ~ ~ 

-

~ 110 

i 

I 100 

IU 

to90 

80 

110 

-

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F23-GATE SEQUENCER 

MAXIMUM ALLOWABLE TEMPEl UURE - 212 DEGREES F I L I I 

1/ I ~ '( 'I I I ) 1/' J 

/ , V I I 

I I 

F27-UNIT CASE 

PREDICTED MAXIMUM TEMPERATURE - 110 DEGREES F 

~., ",. -- ~- -- -~ -~ 

""" 

AKP-II-5 96 

!J 

~I '{ 

' I 

". I 

J 

'I ~ 

I j 

1I ' I 

, ., 

,., ... -~ , -

~ 50 

OPERATING TIME 

(MINUTES) 2.0 2.1 2.1 1.6 2.2 2.8 3.7 1.8 2.4 3.2 3.0 1.7 2.9 1.4 

PASS NO. 8 9 14 16 24 25 30 «» 41 46 47 56 51 72 

Figure ZZ - Orbital Temperature History - Reference Computer Unit 

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100 

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F2O-POWER SUPPlY 

...... 

u. PlEDICTED MAXIMUM TEMPERATURE - 165 DEGREES F 

S 90 Q 

SECTION VB 

-

~ ~ , ~ ~, ..... , ~-- -- I ... "..---""" , " -- ., "., ~ .. ~ ~ ... :l 

i 

80 , , 

i 70 

'" ~ 

60 

F21-UNIT CASE 

100 ...... PREDICTED MAXIMUM TEMPERATURE - 130 DEGREES F ... 

C> 90 !f 

-

~ 80 ~ 

~ 

~ 70 ~ 

'" ~ 

60 

OPERATING TIME 

-- ~ , 

(MINUTES) 2.0 2.1 

PASS NO. 8 9 

, ~ , ~, ~ ~ , ~~ ~ ~ , , ~ " ~ 

2.1 

14 

1.6 2.2 2.8 3.7 

16 24 25 30 

1.8 2.. 3.2 3.0 

<CO .1 46 47 

~ ~ 

, 

1.7 2.9 1.4 

56 51 72 

Figure Z3 - Orbital Temperature History- Control Unit 

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F28-DEFLECTION AMPlIFIEI 

110 

- PREDICTED MAXIMUM TEMPERATURE - 175 DEGREES F u.. 

~ 100 

0 

-

w 

! 90 too 

~ 

~ 80 ! 

,1 ~ , ~ 

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I ", 

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w 

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~ 

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w 90 

!5 

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i too i'O 

MAXIMUM ALLOWABLE TEMPERATURE - 185 DEGREES F 

.,. ~ 

... ~ .,. .-~-.--.- -~ 

F3O-UNIT CASE 

PREDICTED MAXIMUM TEMPERATUlE - 134 0 GREE F 

~.-

-""" -~-~ -~ -~-~ -""" -

OPERATING TIME 

,.-

... 

AxP-II-596 

~ • ." ~ 

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r , I ". 

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- (MINUTES) 2.0 2.1 

PASS NO.8 9 

2.1 

14 

1.6 2.2 2.8 3.7 1.8 2.4 3.2 3.0 

16 24 25 30 «) 41 46 Q 

1.7 2.9 1.4 

56 51 72 

..000-21 

Figure Z4 - Orbital Temperature History - Recorder Unit 

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SECTION VD 

Examination of flight temperature data indicates that the equipment remained well within predicted and allowable maximum temperature limits. 

Temperature history data show normal accumulation of heat and a gradual 

heating trend. 

Variations in the orbital operating temperature data and apparent inconsistencies in the temperatures experienced at the start of the various cycles 

are explained somewhat by considering the effect of the pre-operate mode 

on the equipment thermal response. Figure 19 showed that components 

experience the greatest rate of heating shortly after being energized. 

Furthermore. components in Figures Z 1 through Z4. with the exception 

of the t-r tube. a~e fully energized during the pre-operate mode. Therefore. the temperature at the start of a particular orbital operating cycle 

(operate-on mode) and the rate of heating (temperature slope) is directly 

dependent upon the length of the pre-operate condition. A comparatively 

short pre-operate period results in a lower initial temperature as well 

as a steeper temperature slope. 

3. FLIGHT VIBRATION AND SHOCK DATA 

Vibration and shock response of the equipment was monitored during vehicle 

launch and ascent for approximately 330 seconds. Satisfactory data were 

received via the telemetry link from sensors in the R-F /I-F and Recorder 

units. No meaningful data were obtained from the monitor in the TransmitterModulator. It was determined that this accelerometer was not functioning 

properly during prelaunch checkout. No repair or replacement was effected. 

A summary of the vibration and shock data for the R-F /I-F and Recorder 

units is presented in Table XVII. Vibration or shock levels. duration and 

frequency characteristics. where applicable. are given for various stages 

of vehicle ascent. 

Both shock and vibration responses were measured during ascent. The vibration spectrum encountered was. in general. complex. It was not possible to 

determine the frequency components of a particular response spectrum in most 

. -80-

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l • 

TABLE XVU - FLIGHT VIBRATION AND SHOCK DATA 

Recorder Response R-F /I-F Unit Response 

System Ley.e~ Duration Leve~ Duration 

time Function {gfs) (sec) Remarks (g's) (sec) Remarks 

68935.62 Lift-off 6 3.5 :1:3 g. 15 cps sine 2 3.5 :I:Z g. 15 cps sine 

with high fre- with high frequency vibra- quency vibration. 

tion, :1:3 g super- :1::1. 5 g superimimposed posed 

68951.22 0.54 Mach 1 (max) 6 High frequency 2(max) 6 High frequency number vibration. vibration 

Random amen 

.. plitude 

m 

n • ... 

%~ 

68959 to 2.5 (max) - High frequency 0.5 to - High frequency 68963.3 vibration. 0.15 vibration 

Random ..... 

z· amplitude 

0 

roo 

Z 

Q 

68963.3 to 1 - High frequency 

69001. 1 vibration. Negligible response 

Random 

amplitude 

69001. 1 Booster Negligible response Negligible response separation 

6901Z to 2 - 20 cps'. sine 2 - 20 cps, sine 

69086.5 

69086.5 Main engine 6 0.2 Shock pulse. 6 0.2 Shock pulse, cut-off triangular triangular 

69095.4 Vernier 4 O. 1 Decayingvi- 1 O. 1 Decaying vibraengine bratory tory burst. 

cut-off burst. Decay Decay time: 

time: O. 1 sec O. 1 sec 

* Zero-to-peak acceleration 

~cz LIIIiii m :;a OOno 

~§;> ~,. (I) ." 

~ (I) ." 

""""'" - :;a 17.:!! :!! 0 

..,.,,~ ~ ~ me 

-<" ···0 CDn:;a 

c....;=:;a C::;am 

.- .... .- ;:; ~ o (I) 

~ m 

~ "a 

• = 

UI • 

-.0 

0' 

!C 

m 

n 

• roo 

~ S! 

Z 

Q 
fit 

~ 

"' n • r-, 

%00 .N 

%' 

o 

ri 

TABLE xvn - FLIGIJT VIBRATION_AND SUOCK DATA (Continued) 

'.. . - Recorder Response R-F !I:"F Unit Response 

System Leve, Duration Leve~ Duration 

time Function (g's) (sec) Remarks (g's) (S8C) Remarks 

69101.1 Agena 6 o. 1 Decaying vi- 3 o. 1 Decaying vibraseparation bratory tory burst. 

burst. Decay Decay time: 

time: O. 1 sec 0.1 sec 

69105.3 Fairing 6 0.08 Decaying vi- 1.5 0.08 Decaying vibraseparation bratory burst tory burst 

69115.3 Ulage 0.5 0.5 Vibration: 0.5 0.5 Vibration: 

rocket 60 cps 60 cps 

firing 

691Z8.5 Agena Z 0.5 Decaying 1.5 0.5 Decaying shock 

burn shock 

69368.7 Agena 9 - Decaying 7 - Decaying shock. 

engine shock. First pulse dura· 

cut-off' First pulse tion approx 

duration 0.014 sec 

approx 

0.016 sec 

. Decaf time: Decay time: 

o. 15 sec o. 15 sec 

* Zero-to-peak acceleration 

~ i.ID C Z 

"" m ::u 1ee ~ ~ en "V 

1- ~;I U!!::!! 0 ~m< Cfi;I c m omC 

z-<-n ····0 CDO::u 

eS;::u 

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C'l 

~ . .... 

o 

~ 

<: 

= 

sc: 

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~ r-

~ % 

SZ 

~ 
• 

• 

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cases. Vibration amplitude (acceleration level) also varied considerably 

(random amplitude) for a particular response. Values given in the summary 

are the maximum values that occurred. Vibration response betwe.en the 

listed times was small and of little consequence. Most equipment vibration 

occurred within 60 seconds of lift-off. Maximum vibratory acceleration 

levels were reached at lift-off and during the transonic range of vehicle 

ascent. Several shock pulses and short bursts of decaying (highly damped) 

vibration (probably shock induced) were encountered at various other stages 

of vehicle ascent. 

Comparison of vibration levels encountered In the Recorder and R-F/I-F units 

(assuming comparable vehicle Input to the two units) shows the effectivenf;tss . . 

of the isolation system used on the R-F /I-F unit. As a result of this system, 

the high frequency vibration response of the R-F /I-F unit was 50 to 80 percent 

less than the acceleration levels encountered in the hard-mounted Recorder unit. 

The equipment was designed and qualified to more stringent environmental 

requirements than those encountered during vehicle ascent. The ascent 

vibration levels. although somewhat comparable to design requirements, were 

of such short duration that the occurrence of equipment damage is extremely 

unlikely. Relatively high vibratory acceleration levels occur for only 4 to 6 

seconds during ascent as compared to the qualification .test duration of 3 

minutes per vibration frequency octave. Shock levels were also well below 

design criteria. 

4. FLIGHT PRESSURE DATA 

Pressure Information obtained under night conditions was extremely limited. 

No meaningful orbital pressure data were obtained. The Pirani vacuum gages 

used to monitor pressure have an operating range from Z to 100 microns. 

These gages proved to be inadequate for this application. Conditioned signals. 

from all sensors indicated that the pressure went below the calibrated range. 

Off-scale readings occurred at the times listed In Table XVDI • 

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SECTION yu 

TABLE xym - FLIGHT PRESSURE DATA 

Telemetry number 

FIOO 

FlO! 

FIOZ 

FI03 

FI04 

Time of off-scale reading 

300 seconds after lift-off 

300 seconds after lift-off 

Between pass no. Z and 

pass no. 8 

300 seconds after lift-off 

3600 seconds after lift-off 

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SECTION VUI - UNIT FLIGHT TEST ANALYSIS 

I. GENERAL 

AKP-U-596 

This section analyzes the flight performance of each of the five units of the 

radar set. The analyses were performed by studying the telemetered signals. 

Each of the telemetered signals from the radar was examined and the signals 

which were most indicative of the status of the units are discussed in detail 

in the subsequent paragraphs. 

Z. TRANSMITTER-MODULATOR 

.!.. General 

The conditioned signals furnished for telemetering by the TransmitterModulator unit are listed in Table XIX. 

TABLE XIX - TRANSMITTER-MODULATOR CONDITIONED SIGNALS 

F-number Signal 

Fl3 Klystron filament voltage 

Fl4 High-voltage supply 

FI6 Inverse current 

Fl7 Klystron wave form 

F56 Operate Tran smitter-Modulator 

Two other signals, the F7Z input to the Transmitter-Modulator and the 

F73 output from the Transmitter-Modulator, indicate the input and output 

power levels and are furnished by the R-F /I-F unit • 

From these signals, the klystron filament voltage, high-voltage supply, 

input to Transmitter-Modulator, and output from Transmitter-Modulator, 

have been chosen for discussion and are plotted in Figure Z5. The klystron 

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30 

LOWD LIMIT 29.5 KW 

III 

1= f 20 

41' 1--.. 41' --, 

9 

2 10 

o 

-... --- ~-- ,.... 

UPPEI LIMIT ".6 KV 

-10--I- - I- -

LOWER LIMIT ".0 KV 

SECRET 

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F73-PfAK OUTPUI' POWEI 

I- 41' 

I- __ 

--- ~ 

-~-

F12-AVEIAGE INM POWER 

-I- -~- '-

'"'" 

Fl+oHlGH-VOLTAGE SUPPLY 

to--I- - I- - 1-. '- -

~~ 

'- -

FI3-KLYSTION FILAM&lT VOLTAGE 

7.6 UPPER LIMIT 7.5 V 

III -I-

~ o 7.2 

> -I-

-I6.8 LOWER LIMIT 7.0V 

OPERATING TIME 

SECTION vm 

.... J--~-,. 

\ 

~ 

\. ~ 

I'" 

• .. 1 --

I- - '- -to-- to- • 

, 

" 

-I-

(MINUTES) 2.0 2.1 2.1 

PASS NO. 8 9 ." 

«JOO-25 

1.6 2.2 2.8 3.7 

16 2" .25 30 

1.8 2." 3.2 3.0 1.7 

«» "1 '" "7 56 

2.9 I." 

51 72 

Figure 25 - Transmitter-Modulator Unit Monitor Signals 

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wave forll1 and inverse current signals. although useful for trouble shooting. 

are not discussed as the unit behaved normally. The operate TransmitterModulator signal indicates the status of certain control signals in the unit. 

b. Fl3 - Klystron Filament Voltage 

The klystron filament voltage data were furnished for only four passes. 

The plot shows that the klystron filament voltage was within the prescribed 

limits of 7 to 7.5 volts. 

£.. F14 - Transmitter-Modulator High-Voltage Supply 

The Transmitter-Modulator high-voltage supply plot shows that the supply 

voltage was well within the limits of 4 to 4.6 kv except for the last pass. 

The supply voltage decreased rapidly on the last pass because of the 

battery voltage falling below the regulating range of the high-voltage 

power supply. 

The Transmitter-Modulator high-voltage aupplyreached the operating 

level approximately ten seconds after the operate command was given. 

d. Transmitter-Modulator Average Input Power 

The r-f drive power furnished to the Transmitter-Modulator unit by the 

R-F !I-F unit, is adjustable at the R-F !I-F unU. The r-f drive power is 

normally set to a value just below saturation for each klystron. The value 

of the r-f drive power required by the klystron varies with eaoh individual 

tube. As delineated in Section V. paragraph 8. k.. and in paragraph .!. 

following. the klystron drive power was set slightly below the level required 

for the particular klystron used in the flight Transmitter-Modulator. 

On pass 47 the average input power decreased linearly. then rose abruptly 

and decreased again. This behavior was caused by the ohanges in prf during 

the pass. as average power varies directly with prf. The pass started 

with prf step 6 (8519 ops) and was decreased.in consecutive steps to prf 

step 15 (8215 cps). which accounts for the linear decrease in average 

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power. The prf was then changed from step 15 to step 0 (8ZI5 to 8735 cps), 

which accounts for the sudden increase in average power. The prf was 

again decreased in consecutive steps to step 6, which accounts for the 

decrease in average power after the Jump. Changes in average power 

on other passes where the prf was varied were not as readily discernible 

as on pass 47 where the prf was varied through the entire range. 

On pass 7Z there was an apparent increase in the average input power 

to the klystron; however, since the vehicle batteries were low and the 

telemetry link was not functioning properly, this is believed to be an 

incorrect indication. 

e. F73 - Transmitter-Modulator Peak Output Power 

The Transmitter-Modulator peak output power data were corrected to 

incorporate the final test measurement before launch. The final test 

measurement was 74. Z dbm which corresponded to a conditioned signal 

of 1. 3Z volts ilardUne and 1. 29 volts telemeterlng. This measurement 

was taken as a calibration point, 'and the received data were corrected 

by adding o. 3 db. The calibration point represents a peak power of Z6. 5 

kilowatts. This is below the system specification minimum value" of Z9. 5 

kilowatts because of insufficient drive power (refer to preceding paragraph!!. ). 

The data plotted also take into account the r-f pulse width and the prf. The 

pulse width measured at the 70 percent point was 0.9 microsecond and was 

assumed constant. The prf for the plotted power points was taken from the 

operating log. 

The telemetered power data are baSically in terms of average power. The , 

average power reading automatically reflects any changes occurring in 

klystron beam voltage (amplitude and pulse width), filament voltage, 

r-f drive, and prf. The peak output power would be expected to remain 

constant provided that the klystron beam voltage, filament voltage, and 

r-f drive remained constant. As can be seen from the data plotted in 

Figure Z5 the klystron filament voltage, r-f drive power, anc:f high-voltage 

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supply were not the same for each pass. Consequently. some variation 

in peak power is expected from pass to pass. It could be a·ssumed that 

if the factors mentioned above were constant during the pass that the 

peak power should remain constant. However. the received data showed 

a slight tendency for the power to increase from the beginning to the end 

of the pass because of a slight tendency towards an increase in power as 

the klystron warms up to operating temperature. 

It was considered desirable to present the output power data in terms of 

peak power. This was obtained from the relation. 

P

av = PW X PRF 

where 

P p = Peak output power 

P = Average output power av 

PW = R-f pulse width at 70 percent amplitude 

PRF = Pulse repetition frequency. 

(13) 

The high-voltage supply reached the normal operating level approximately 

10 seconds afte.r the operate command was given. In the data plot. the 

first point for each pass was taken at approximately ZO seconds after the 

operate command was given. 

The plot shows that the peak power decreased from pass 40 to pass 7Z. 

The decrease corresponds somewhat with the lower high-voltage supply 

for these passes. In passes 56 and 57 the input power was slightly below 

average. Pass 47 shows the effect of the prf change discussed in paragraph £. 

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On pass 72 the output power started low and dropped off rapidly during the 

pass. This corresponds quite well with the drop off in high-voltage supply 

during the pass. 

3. RF-IF 

a. General 

The conditioned signals which are monitored in the R-F!I-F unit are 

listed in Table XX. 

TABLE XX - R-F !I-F CONDITIONED SIGNALS 

F-number 

Fll 

F53 

F72 

F73 

F74 

F75 

Signal 

High-voltage supply monito" 

Gain control monitor 

R-f power to Transmitter 

R-f power from Trans mitter 

Twt pulse monitor 

100-Mc output power monitor. 

Of the above signals, FII high-voltage supply, F72 r-f power to 

Transmitter, F74 twt pulse, and F75 100-mc output power are discussed 

in detail in the following paragraphs and plotted in Figure 26. F53, gain 

control, is discussed in Section IX: System Flight Test Analysis. F73 

was fully discussed under the Transmitter-Modulator, paragraph 2. e. 

!!.. FII - R";F !I-F High-Voltage Supply Monitor 

The 2000-vdc R-F !I-F high-voltage telemetered plot shows that the high 

voltage was very near the required twt cathode voltage ·of 1975 vdc during 

each of the fourteen passes. The voltage remained well within the specified operating limits for stable twt amplification. Telemetered data 

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1«) F72-AVEIAGE R~ POWER TO TRANSMlTTEI 

-~ -~.- ~ 

III 120 

5 

~ 100 

3 

-~ -.. -1--I- - ~ 

~ 80 

F75-1CJO-MC OUTPUT POWER 7S 

70 -,.. • I- -1- • I- -

6S 

LOWER liMIT .oC6 MW 

60 

F74-TlAVRING WAVE TUBE PWE 

_ 2.7 

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i2 1.9 

UPPER LIMIT 3.0 VDC 

-~ . I- - I- • 

LOWER LIMIT 2.0 VDC 

UPPE LIMIT 2.1 KVDC 

-~ ---I- -

LOWI tR LIM T 1.9 KVDC 

OPERATING TIME 

1--~ -1--I- - 1--1--

Fll- HIGH-VOLTAGE SUPPLY 

-- ~ -- -,..... --

.AKP-D-596 

\ " 

r 

-!- ~ j -

~ -I- - 1-. -, 

~ 

I- - l-

. -I- - I- .. 

I' 

~ ~ 

(MINUTES) 2.0 2.1 2.1 1.6 2.2 .2.8 3.7 1.8 2." 3.2 3.0 1.7 2.9 1." 

PASS NO. 8 9 1" 16 2" 2S 30 «) "1 .oC6 47 56 51 72 

GJO-26 

Figure 26 - R-F /I-F Unit Monitor Signals 

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indicate that the high voltage increased approximately 40 volts during the 

final operating pass. However, this apparent increase in high voltage 

was caused by a decrease in the +Z3.5 vdc connected to the high-voltage 

monitoring circuit and does not represent a true increase in high voltage. 

Therefore. as the high-voltage power supply apparently did not drop out 

of regulation because of the decreased 115 vac, it can be assumed that 

the output voltage of the high-voltage power supply remained essentially 

constant throughout the final operating pass. 

c. F74 - Twt Pulse Monitor 

The telemetered twt pulse data show that the twt pulse remained relatively 

constant during the entire life of the system and was well within the speci-

. fled operating limits. The different operating levels are caused by changes 

in the prf and variations in the telemetering Unk. The decrease in the 

operating level during the final pass was caused by a decrease in the 

+Z3. 5 vdc supply and indicates the general decaying condition of the 

vehicle batteries. 

d. F75 - 100-Mc Oscillator Output Power Monitor 

The curve for the average output power of the IOO-mc oscillator (see 

Figure Z6) demonstrates that the 100-mc output level remained well 

above the 46-mw input power level required for satisfactory frequency 

multiplier operation. The decrease in 100-mc output power during the 

final pass was caused by low Z3. 5-vdc inverter supply output voltages 

and again indicates the low charge level of the vehicle batteries. 

e. F72 - R-f Power to Transmitter Monitor 

The average output power of the R-F /I-F unit (Transmitter-Modulator 

input power) was slightly below the calculated system specification 

value. As previously delineated in paragraph Z: Transmitter-Modulator, 

the klystron input drive level is adjusted by a variable attenuator in the 

R-F /I-F unit. This adjustment is made at the system level and varies 

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for each individual klystron. Although the R-F /I-F unit has a much greater 

capability; the output power of the flight unit was set slightly below the 

required drive level of the klystron (see Figures ZS and Z6). This condition 

was known prior to lift-off and resulted from the deciSion not to remove 

the R-F /I-F unit covers and re-adjust the klystron input drive power when 

a new Transmitter-Modulator unit was installed in the flight vehicle. 

4. REFERENCE COMPUTER 

The conditioned signals monitored in the Reference Computer are listed in 

Table XXI. 

TABLE XXI - REFERENCE COMPUTER CONDITIONED SIGNALS 

F-number 

F46 

F47 

F48 

F60 

F61 

F6Z 

F63 

F64 

F6S 

F66 

F67 

F68 

Signal 

Clutterlock relay 

BIT relay 

Clutterlock time constant no. 1 

Integrator output 

Agc 

70-mc reference 

70-mc offset reference 

70-mc single side band 

Sweep trigger 

O.06-microsecond 70-mc pulse 

HO-mc pulse 

On-Gate 

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Of the above signals. three have been selected as being indicative of the status 

of the unit: F6Z 10-mc reference. F64 70-mc single side band, and F66 

O. 06-microsecond 10-mc pulse. These signals are plotted in Figure Z1. 

Each signal remained essentially constant throughout the night. On pass 7Z, 

sUght variations occurred because of changes in the +Z3.5 vdc reference and 

the variations in the telemetry signals. The outputs, however, are stUI 

within the established limits of operation. 

The remaining signals were analyzed but are not plotted. The first three -

F46, F41 and F48 - are relay position monitors. The. relays behaved normally during tlie flight. The next two signals, F60 and F61, are discussed 

in detail in Section IX. The remaining Signals, F63, F65, F61, and F68, held 

constant within their prescribed Umits. 

5. CONTROL UNIT 

a. General 

The signals available for analysis from the control unit are listed in 

Table XXII. 

TABLE XXII - CONTROL UNIT CONDITIONED SIGNALS 

F-number 

FI 

F3 

F4 

F5 

F1 

F8 

F9 

Fll 

FIZ 

Signal 

+Z3.5 vdc 

-Z3.5 vdc 

+300 vdc 

+Z8.3 vdc 

liS v 400 cps Phase AB 

115 v 400 cps Phase CB 

115 v Z kc 

Pre-operate monitor 

Transmitter-Modulator timer 

override monitor 

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2.0 UPPEI LlMIT'2.0 VDC 

-

~ -~-I-

~ 1.0 

0 

Z 

S 

.... 

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-' IU 

it 

-' 

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F62-~ IEFEIENCE MONITOI 

-1--1----I-

-' 

~ F64- 10-11£ SINGLE SIDE lAND MONITOI 

Z 2.0 

0 UPPER LIMIT 1.5 VDC 

iii 

1.0 -10--I- -~ -~ -1--1-' 

-

~ 

~ 0 

LOWER LIMIT 0.5 VDC 

0 

Z 

ii 

e 

z 

4.0 

F66- 0.06 SINE XIX ~ PULSE MONITOI 

i UPPER LIMIT 3.5 VDC 

-' IU 

it 

-' 

~ 3.0 -'"'" -

~ 

0 

iii 

LOWER LIMIT 2.5 VDC 

2.0 

OPERATING TIME 

---I- - I- - -

AKP-U-596 

-~ -~- 1--~ 

-~ -~ -I- - .... 

-----~ -'-

(MINUTES) 2.0 2.1 2.1 

14 

1.6 2.2 2.8 3.7 1.8 2.4 3.2 3.0 1.7 2.9 1.4· 

PASS NO. 8 9 16 24 25 30 .to 41 46 ..., 56 51 72 

Figure 27 - Reference Computer Unit Monitor Signals 

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SECTION vm 

From these signals. Fl +Z3.5 vdc. F3 -Z3. 5 vdc. and F4 +300 vdc. have 

been chosen as output signals which show the status of the unit. In 

addition. F5 +ZS.3 vdc. F7 115 v 400 cps Phase AB. andF8 115 v 400 cps 

Phase CB. are included in the analysis since they. show the condition of 

the primary power input. 

k. Power Supply Output Voltages Fl, F3. and F4 

A study of the variations in the telemetered outputs from the Control unit 

regulated power supplies reveals that each voltage output (+Z3. 5 vdc. 

-Z3.5 vdc. and +300 vdc) turned on for each orbit at pre-operate within 

the :1::1 percent regulation specified (see FigureZ8). When the radar was 

switched into operate. however. a jump in each of these voltages was 

indicated and for some orbits the indicated values were outside of the 

:1::1 percent regulation limits. Since the normal performance of these 

supplies within their regulating range is well within the specification. 

and since the ~oad presented to these supplies does not change appreciably 

when switched to operate. these yariations are believed to be caused by 

variations in the telemetry link. 

c. Primary Voltage Inputs F5, F7. and F8 

One exception to the above analysiS occurred during pass 71.. Figure Z9 

shows the variation of the input voltages. During pass 7Z each of the 

input voltages fell off sharply indicating the end of battery life. During 

this pass the primary voltages. and also the Control unit po:wer supplies. 

went out of regulation as the battery voltage went down. 

The conclusion reached from the indicated telemetry data during the preoperate mode. and from the quality of the data during each operating mode. 

is that the Control unit operated within_ specifications for all orbits. 

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~ +24.5 

~ +23.5 .... 

o 

> +22.5 

_ -2".5 

~ 

~ -23.5 

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+310 

R 

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o 

>+290 

OPERATING TIME 

" ,. .- , 

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, 

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SPECIAL HANDLING 

F1 + 23.5 VDC 

.. ~ ~ ". r -~ , , 

F3 - 23.SVDC 

.. " ..... ~ .... ~ r- 1-

F" + 300 vee 

~, , .. , 

, -r- - , 

A,KP-U-596 

:, " , " 

~, , , ..... ~ 

, , ..... "- - - " - - " 

--- , ~ 

- .... , , ~ , 

(MINUTES) 2.0 2. J 2.1 1.6 2.2 2.8 3.7 1.8 -2.4 3.2 3.0 1.7 2.9 1.4 

PASS NO. 8 9 14 16 24 2S 30 «) 41 <46 47 56 57 72 

Figure Z8 - Control Unit Regulated Voltages Monitor Signals 

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F5 +28.3 VDC 

SEcTloNym 

IIIIIIIII1111 * I 

F7 115 V .., CYCLES PHASE AI 

III1 fllllllll{1 

g 125 

~ 115 &It 

!:i 105 

g 9S 

~ 125 

l1S 

~ 105 

o 

> 9S 

OPERATING TIME 

.. .... 

-~ 

(MINUTES) 2.0 2.l 

PASS NO. 8 9 

F8 115 V ., CYCLES PHASE CI 

-"-

F9 115V 2 KC 

-.. 

2.1 1.6 2.2 2.8 3.7 

14 16 24 2S 30 

1.8 2.4 3.2 

.4C) 41 46 

.. 

~ 

.. 

, 

~ 

3.0 1.7 2.9 1.4 

47 56 51 72 

Figure 29 - Control Unit Primary Power Monitor Signals 

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6. RECORDER 

a. General 

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AKP-I1-S9' 

The Recorder unit furnished the conditioned signals listed in Table XXID 

for telemetering. 

TABLE xxm - RECORDER CONDITIONED SIGNALS 

F-number Signal 

FlO IO-kv supply 

FI8 Deflection current 

F49 Film motion 

FSI Operate Recorder 

F5Z Isolation roller position 

F55 Ramp output 

F70 Video output 

F71 Crt light output 

From these signals five were selected for detailed discussion. These 

are FlO IO-kv supply. F55 ramp output. F71 crt light output. F49 film 

motion, and F5Z isolation roller position. The first three signals show 

the condition of the sweep circuits and crt. The last two signals 'show 

the condition of the film drive system. 

Of the remaining signals. the deflection current monitor samples the 

positive-going portion of the deflection current wave form. The output 

was correct during the test. The operate recorder signal indicates the 

position of the film drive relay. The video output signal will be discuseed 

in Section IX. 

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b. FlO -IO-Kv Power Supply 

SECTION vm 

This monitor circuit is a resistor divider network across the 10-kv 

power supply. As shown in Figure 30. the power supply output voltage 

remained constant during the active passes. thus indicating proper 

operation. The function is plotted only through pass 30 .. The film was 

recovered on pass 33. 

£.. F55· - Ramp Output 

The ramp generator produces a linear voltage ramp which is proportional 

to the slant range to the target. The ramp monitor signal was well within 

the limits which were set up prior to the flight test (0.9 to 1.3 vdc). 

d. F71 - .. Crt Light Output 

The crt light monitor uses a photosensitive resistor in a voltage divider 

network. The output. which varies with light level on the face of the 

crt. is fed to an emitter foUower. This level was set prior to launch 

with no video present at a relative brightness level of 64 percent as 

measured on a microphotometer. The crt brightness varies with the 

amount of video present which is also a function of the 'terrain and of 

the i-f gain (attenuator step setting). For the pre-recovery passes 

shown in Figure 30. the radar set was in the agc mode (passes 8 through 

30 inc lusive). 

e. F49 - FUm Motion 

The film motion monitor indicates the pOSition of the cam-operated 

switch on the output isolation roller assembly. (Refer to Volume I for 

a detailed description of the function of the switch in controlling fl1m 

speed.) The microswitch opens and closes as the cam cycles either 

side of center position producing a 0.375 percent change in f11m speed. 

Figure 31 shows the cycling of the switch. Fast speed is indicated on 

the graph by the higher voltage (4.3 vdc). Film speed control was 

normal throughout the mission. 

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-.: 

g 

Z 10.5 KV i 

UPPER LIMIT 10.5 KV 

.... 

IU~ - 10 KV 

~~ .... 'I~ 9.5 KV 

C) LOWER LIMIT 9.5 KV 

V; 

-

~ 

~ 

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UPPER LIMIT 1.3 VDC 

z 

i 1.0 -~ 

.... 

IU 

~ 0.5 

.... LOWER LIMIT 0.9 VDC 

~ C) 

iii 

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Z UPPEl LIMIT 70 PERCENT IU 

~ 70 

~ -.: 

0 

.... 

z 

i 60 

.... 

x 

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I I 

C) 

~ 

lie 

u SO LOWER LIMIT SO PERCENT 

OPERATING TIME 

(MINUTES) 2.0 

PASS NO. 8 

2.1 

9 

5 E e R ET 

SPECIAL HANDLING 

FIO-IO-KV MoNITOR 

F55-ItAMP MONITOR 

-~ -~ 

F71-CRT LIGHT MONITOR 

-- ~ 

2.1 

14 

-- ~ 

1.6 

16 

-~ 

--II 

J " 

2.2 

U 

Figure 30 - Recorder Monitor Signals 

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2.8 

25 

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_ S.o ~ 

~ 

..: ".0 e 

z 3.0 

~ .... 

III 2.0 ~ .... 

.... 

~ 1.0 Z 

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29,500 29,520 

SECRET 

SPECIAL HANDLING 

F"9-fILMMOTlON MONITOR 

ACTIVE PASS NO. 9 

FS2-ISOLATION IOLLER POSITION MONITOR 

ACTIVE PASS NO.9 

29,540 29,560 29,580 29,600 

SYSTEM TIME (SECONDS) 

29,620 

Figure 31 - ·Recorder Film Motion Monitor Signals 

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f. F5Z - Isolation Roller Position 

AKP-II-596 

The isolation roller monitor indicates the position of a potentiometer 

driven by the input isolation roller assembly (refer to Volume I, 

pages 139-140, for detailed description of operation). Figure 31 

shows that the potentiometer moved immediately after initiation of film 

dri ve to stabilize the film tension at Z. 5 pounds. Film tension control 

was normal throughout the mission. 

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SECTION IX - SYSTEM FLIGHT TEST'ANALYSIS 

1. GENERAL 

AKP-U-596 

This section presents an analysis of the radar system operation during the 

orbital flight test. The purpose of this analysis is to establish that the 

radar functioned properly throughout the test. No attempt is made to analyze 

the radar performance in terms of resolution or to discuss over-all picture 

quality. These results are presented in a separate report by 

In the following pages. a short subsection is presented showing some typical 

radar mapping results of the flight. Following this. various functions which 

are impor~nt from an over-all system standpoint are discussed by reference to the telemetry and radar map data. The radar pictures. unless 

otherwise noted. are enlarged Z.6 to 1 from the original data film size. 

The map scale is 3. 6 miles per inch in range and Z. Z miles per inch in 

azimuth. 

Z. TYPICAL RADAR PICTURES 

Figure 3Z shows a section of a 1953 topographic map of East Chicago. 

nlinois. Figure 33 is a picture of the same area made with the KP-U 

satellite-borne radar enlarged 4.8 to 1 from the data film size. Several 

landmarks are identified by numbers on the topographic map and by corresponding numbers on the overlay to the radar map. 

Figure 34 shows other radar maps taken with the KP-U· radar. Figure 

34(a) is the radar return from Richmond. Virginia. and Figure 34(b) the 

return from Wurtsmith Air Force Base on Saginaw Bay in Michigan • 

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Figure 32 - Topographical Map of East Chicago. Illinois 

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.,.1.,"-"-'-" "" ,,,', ""'-""'l"'lI; ..... I111'.,"'.i ___ - .. i ...... i'-'4A~ E<3-R.E"T 

ir----

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1. IWNOIS CENTRAL 

RAlLiOAD 

2. OAKWOOD CEMETARY 

3. HEAVY INDUSTIY 

.t. 8EAKWATEI 

5. CHICAGO SKYWAY 

6. INDIANA HAItIOI 

7. LAND FlU 

8. OIL STORAGE AREA 

9. IlVEIIAIGE 

10. CALUMET FIEEWAY 

11. TRAIN 

12. UTTLE CALUMET 

IIVEI 

}o'igure 33 - Radar Return ·from East Chicago. Illinois 

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-- ORIIT DIII£CTlON 

-- ORIIT DIRECTION 

1. AllPORT 

2. GlAND CANAL 

3. SHOPPING CENTEl 

4. CENTRAL AVENUE 

4000-37 

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f 'IC"""" ••• ,IG,N .. 

AKP-II-596· 

NEAl SLANT lANGE, 

155 NAUT MI 

. PASS I" 

N TO S 

22 DECEMBER I~ 

FAR SLANT lANGE 

161 NAUT MI 

. .·.,~_~f~: ;':': .. id{-.';'!~/,: " NEAR SLANT IlANGE . 155 NAUT MI 

'. '" •.. , .... ,-,,.. 

(I) WURTSMITH AFI, MICHIGAN 

5. STATE FAIIGIOUNDS 

6. tlGH liSE IULDINGS 

7. ILACK CANYON FlEEWAY 

8. GlAND AVENUE 

Figure 34 - Typical Radar Maps 

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161 NAUT MI 

9. THOMAS lOAD 

10. ALUMINUM PLANT 

11. US tlGHWAY 80 

12. OIL STORAGE AREA 
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Figure 35 shows a 1953 topographic map of part of Phoenix, Arizona. 

(Note: Many changes have occurred in Chicago and Phoenix since the 

topographic maps were made.) Figure 36 is the radar return from the 

same area, enlarged 7 to 1 from the data film. The upper part of the 

radar map is somewhat dark as a result of a one-step misadjustment 

of the prf. This will be discussed further in paragraph 5 following. 

3. AUTOMATIC GAIN CONTROL (AGC) 

a. General Operation of Agc 

The automatic gain control permitted the receiver gain to vary as a 

function of the amplitude of the signal returned to the receiver from 

the terrain being mapped. For large returned signals, such as those 

from residential or manufacturing areas, the receiver gain was automatically reduced. For low signal returns, such as those from d~sert 

or sea areas, the gain of the receiver was increased. The dynamic 

'range of the agc was measured prior to the flight and found to be 35 db. 

b. Agc Operation During Pass 34 

The effect of the agc action on the map film can be seen in Figure 

37(a). The increases of receiver gain over open water areas produce lighter areas on the fUm. These light areas are seen to correspond to the decreases in agc voltage on the graph, Figure 37(b). 

c. Agc Operation During Pass 30 

Measurements on a non-flight system showed that the agc time constant was 3Z milliseconds for a 10-db step increase in received 

power and Z6 milliseconds for a 10-db step decrease in received 

power. A land-water boundary occurred at system time 57460 

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Figure 35 - Topographical Map of Phoenix. Arizona 

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Figure 36 - Radar Return from Phoenix. Arizona 

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1 

CI' 

." 

m 

n 

,. 

,.. I 

...... %-

~U1 

Z I 

C 

'" !: ... 

-' Z 0 

Q > 

v 

(!) 

<i( 

... 

> 

;:: 

<i( 

-' 

~ 

~ -----~ ... 

.. 

--'------

23,926 23,928 

04000-40 

(A) PORTION Of LAKE MICHIGAN SHORELINE 

(B) PLOT OF F53 MONITOR (SAME TIME SCALE AS MAP ABOVE) 

....... ~ 

'" .. / - "-- ~ ~ 

23,930 23,932 23,934 

SYSTEM TIME (SECONDS) 

Figure 37 - Automatic Gain Control Operation 
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seconds during pass 30. From Figure 38(a) it is seen that the monitor voltage decreased by 0.78 volts from Z.OZ to 1. Z4 volts, corresponding to a decrease in received power of approximately 10.7 

db (see Figure 38(b». Figure 39(a) is a photograph of the. landwater boundary and approximates a step function in terms of power 

received by the satellite-borne radar system. As the land-water 

boundary did not represent a perfect step function and as the 

telemetry samples are too far apart for accurate time measurements, 

it is not possible to measure the agc time constant with ac~uracy. 

It is interesting to note that the, sharp increase and subsequent 

decrease in receiver gain at system times 574Z6and 57431 (see 

Figure 38(a» occurred as the vehicle crossed Saginaw Bay in Michigan. 

The decrease in gain at system time 57471 was caused by the high 

returns from Cleveland, Ohio. The small'decrease at system time 

57478 was caused by returns from the city of Akron, Ohio • 

. 4. ATTENUATOR CHANGES 

a. Attenuator, General Operation 

The seven fixed steps of receiver attenuation covered a range from zero 

to 33.5 decibels. Provision was made to have these steps available for 

use in the event that the 'agc circuitry failed to operate properly. The 

L.,. data film recorded over the wide band data link at the New Boston 

Tracking Station was processed immediately after pass 8 to check the 

i-"agc operation and other important system parameters. Analysis of 

early tele~etry data and the map film from pass 8 showed that there 

was no need to use any of the fixed attenuator settings. However, 

after recovery of the data fUm from the satellite on pass 33, it was 

decided to obtain some data with the attenuator in various settings. 

Figure 40 shows the amount of fixed receiver attenuation for each of 

the possible attenuator settings (excluding step 1 which is agc). 

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-

~ 

~ 

III 

C) 2 ~ 

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(A) FS3 GAIN CONTROL MONITOR VERSUS TIME 

ATTENUATOR IN STEP I - AGC PASS 30 

l 

.... I~ 

0 

> 

., 

e ~ ~ - f , hr'- -

z r ~ fiJ 

~ 

0 

--' 

. .A 

r 

I 

-., 

~ -66 

2 

2 -76 

> iii 

!d -86 

-106 

57,380 

o 

51,. 

~ 

51,420 51,440 51,460 

SYSTEM TIME FOR PASS 30 (SECOM>S) 

(I) RECEIVED POWER 

vasus 

FS3 MONITOR VOLTAGE 

51,480 

------ ./ ~ 

~ V 

l 

0.5 1.0 1.5 2.0 

F53-MONITOR VOLTAGE (VDC) 

Figure 38 - Gaili Control Functions 

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""" 

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_. 

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__ OUiT DIIECTION 

- OIIIT DIIECTION 

4000-41 

(A) LAKE HIE LAND-WATB IOUNDAIV 

AurOWoTIC GAIN CONlIOl 

(I' CONFUSION. MOUNTAIN lANGE, urAH 

AnENUATOIIN STEP 0 

AKP-I1-596 

NEAl SLANT IlANGE 

155NAUT MI 

." PASS 30 

NTOS 

23 DECEMIEI 1964 < 

FAI SLANT lANGE 

161 NAUT MI 

NEAl SLANT IANCH 

165 NAUT MI 

PASS 41 

STON 

24 DECEMIEI 1964 

FAR SLANT lANGE 

171 NAUT MI 

Figure 39 - Automatic Gain Control and Attenuator Step Zero 

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35 

25 

--l1li 

e. 

z 

Q 

I-

~ 20 

Z 

au 

lI-

< 

laC 

III 

> 

W 15 

M 

laC 

a 

au 

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) 

V 

/ 

I 

J 

AKP-U-596 

STEP AnENUAnON (DB) 

0 0 

1 AGe 

2 1.7 

3 12.6 

4 19.8 

5 26.4 

6 ~.2 

7 33.5 

V 

V 

10 

5 / ~ V 

o 

o 2 3 5 6 

ATTENUATOR STEP POSITION 

Figure 40 - Receiver Attenuation versus Step Position 

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Attenuator step 0 allowed the receiver to make use of its full gain. Steps 

2 through 7 reduced the receiver gain 'by fixed amounts. 

b. Attenuator Step 0 

Figure 39(b) showed a section of the map which resulted after step zero 

was selected. In step zero. the gain was excessively high. Examination 

of the results for other attenuator steps shows that the gain for step 3 

was near optimum. Based on this selection. the receiver gain in step 0 

is approximately 12.6 db above the optimum level. For this map. the 

correlator was set up to provide the best picture possible for the bighgain conditions. Although some degradation has occurred. the radar map 

is still useful. 

£.. Attenuator Steps 2 and 3 

Figure 41 (a) shows a section of the radar map made over Green Bay, 

Wisconsin. when step 2 was selected. Some loss of resolution occurred 

as, based on the above selection of step 3, the gain was approximately 

10. 5 db too high. Note the ice formations in Green Bay. Figure 41 (b) 

shows Grand Island for attenuator step 3 which is near the optimum gain 

setting. 

!t. Attenuator Steps 3 and 4 

Figure 42 shows the radar map at the time of switching from step 4. to 

step 3. Figure 42(a) has been processed to favor step 4 and Figure 4.2(b) 

to favor step 3. This comparison demonstrates the versatility of the 

processor. 

~. Attenuator Steps 5, 6. and 7 

The map film made with the attenuator in step 5 mainly shows the return 

from larger targets such as mountains and major cultural areas. ' Plowed 

fields and generally level terrain show little return. 

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--- ORBIT DIRECTION 

__ ORBIT DIRECTION 

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(A) GREEN lAY, WISCONSIN 

ATTENUATO. IN STEP 2 

(8) GRAND ISLAND, LAKE SUPERIOR, MICHIGAN 

A TTENUA TOR IN STEP J 

AKP-II-5?6 

NEAR SLANT RANGE 

164 NAUT MI 

-"", -~-PASS 40 - -

STO N 

24 DECEMBER 19(,.4 

FAR SLANT RANGE 

170 NAUTMI 

NEAR SLA. H RANGE 

164 NAUTMI 

PASS 40 

S TO N 

2>4 DECEMBER 1964 

FAR SLANT RANGE 

170 NAUT MI 

Figure 41 - Comparison between Attenua tor Steps 2. and 3 

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.. ~ 

-- ORIIT DIRECTION STEP 3 

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STEP. 

(A) PROOSSED TO FAVOR STEP. 

STEP. 

(I) PROCESSED TO FAVOR STEP 3 

Figure 42 - Effects of Processing on Radar Imagery 

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NEAR SLANT RANGE 

16SNAUT MI 

. , 

PASS 41 

STO N 

24 DECEMIEI ''''' 

FAR SLANT lANGE 

171 NAUTMI 

NEAR SLANT lANGE 

165 NAUTMI 

PASS 4' 

STO N 

24 DEOMIEI ''''' 

FAR SLANT lANGE 

171 NAUT MI 
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No map film was recorded with the attenuator in step 6, but on pass 72., attenuator step 7 was selected. The map fl1m shows no targets at all for step 7. 

5. CHANGES OF PULSE REPETITION FREQUENCY 

a. Effect on Mapping 

(1) General Operation 

When the prf setting is correct, a uniformly illuminated map is 

obtained. Actually. because of the shape of the antenna beam, 

targets at mid-range are illuminated with more radar energy than 

targets at near and far ranges, but the radar returns are modified 

by the sensitivity time control (Volume I, page 32.) to attenuate the 

mid-range returns and produce a uniform map. 

By changing the prf, the tracking station operators were able to 

adjust the time position of the sensitivity time control function for 

proper compensation. This was usually accomplished within a few 

seconds after control was turned over to the recorder operators. 

During pass 47 the prf was stepped through all 16 of its settings 

to show the effect of large changes. 

(2) Range Shift with Prf Changes 

Figure 43(a) shows a portion of the radar map when the prf was set 

three steps high (left side of picture). After the prf change (step 14 

to step 13, right side of picture) the terrain is more nearly uniformly 

illuminated. It can be seen that even with the prf setting three steps 

high, the map is still well defined. The primary effect of the prf 

change on the map film is the range shift, made evident by the discontinuity in the highway (see arrow). 

This shift in range is also very pronounced in Figure 43(b), which 

shows a discontinuity in the Colorado River (see arrow) produced 

by a prf change from step 8 to 9. In this figure the prf went from 

two steps low to one step low. 

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The river is shifted toward the near-range edge of the map after the 

prf change. The distance between the satellite and the near-range 

terrain is given by the equation: 

where 

R -.£.2L +K 

sn - Z PRF 

= slant range distance from near-range targets to 

satellite 

c = velocity of light 

N = the integer number of interpulse periods between 

transmission and reception of the· radar pulse 

PRF - pulse repetition frequency 

K =. a constant-small compared to P:F • 

(14) 

A step change from 8 to 9 actually represents a decrease in prf from 

8449 to 8415 pps. Hence, Rsn is increased by 0.66 nautical mUe, . 

assuming a vehicle altitude of approximately 138 nautical miles and 

a depression angle of 55 degrees (N = 17). That is, the near-range 

edge of the map is o. 66 nautical mUe farther away from the satellite 

after the prf change. Therefore, the river actually appears closer 

to the near-range edge of the map film after the prf change. 

(3) Effect of Improper Dlumination 

No photographs showing the prf four steps off optimum are included ··but 

in. this case the illumination was so low at some ranges that no targets were discernibl~. 

The unequal illumination resulting from a prf five or more steps from 

optimum produced objectionable results. 

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.:,: ;,~", .... /~ . 

:~t.,}f.. 

-., ..• ''SaI~!,,"~ 

-- ORIIT DIRECTION 

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(A) LAND MASS NEAR LEBANON, ILLINOIS 

-- ORIIT DIRECTION 

(I) COLORADO RIVER, UTAH 

Figure 43 - Range Shifts with Prf Changes 

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.' 

NEAR SLANT RANGE 

168 NAUTMI 

•.. "'-' -PASS 8 . 

S TO N 

22 DECEMBER 1964 

FAR SLANT RANGE 

174 NAUT MI 

NEAR SLANT RANGE 

164 NAUT MI 

PASS 9 

STO N 

22 DECEMBER 1964 

FAR SlANT RANGE I~NAUTMI 
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(4) Range Inversion for Large Prf Errors 

For very large prf errors, as were encountered on pa's8 i4,-;twhen 

the prf was intentionally switched throughout its entire- r'ange,'; targets at the near-range edge of the map film are not necessarily the 

targets that were at the shortest slant range from the vehicle. The 

. prf switching adjusts the time position of the recorder _sweep to 

coincide with the return from a given transmitted pulse.; W.hen the 

prf is set too far off, the trailing edge of one pulse and.~bele~ding 

edge of the next pulse are recorded sequentially. and range inversion is produced. 

Figure 44 of Pueblo. Colorado. is an excellent example of this 

range inversion phenomenon. The center area of the topographic 

map does not appear; the upper and lower areas defined by,.t~e 

dotted lines are interchanged in the radar display. TherailrQad 

tracks 0 are pointed out in both Figures 40(a) and 40(b) to o~ient 

the reader. The barely visible crossroads CD and the curve in 

the river CD are to the east of Pueblo in Figure 44(a), and to the 

west of Pueblo in Figure 44(b) • 

. (5) Transients in Wide-Band Data 

The effect of prf change on the data fil~ recorded on the ground 

usually showed a data loss each time the prf was changed. This 

data drop-out occurs during the time required for the grQund:-based 

recorder crt sweep to lock in with the synchronizing pulse from the 

vehicle. 

b. Operator Recognition Problems 

The wave shapes viewed on the "A" scope. located in the screen rooms of 

the tracking stations. could be easily evaluated by the operator as to the 

need for prf changes. Errors in setting of the prf could be determined 

after viewing the "A" scope presentation for only a few seconds. The 

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integration effects of the human eye playa large part in ~e recognition 

,,' o,f. th,e~A" scope wave shape. If the eye is assumed to have a persistency 

of only 50 milliseconds. the picture presented to the brain represents an 

average of the returns from over 400 radar pulses. 

Some photographs of the "A" scope presentations are shown in Figure 45. 

The exposure time on the camera was long enough to simulate the integra-

.. titin 'effects of the eye. Note that some of the definition of the original rrA" 

, scope presentations is lost in the photographs. 

Equipment was available at the tracking station for photographing single 

sweep traces. It was found that these photographs could not be interpreted 

by the operators. and that they provided no information as to the correct 

prf as they did not include the effects of integrating a large number of 

returns. 

Sporadically. small groups of targets caused the wave form to have a noticeably higher or lower amplitude than the average ,signal level. However. these 

higher or lower pOints were only a few microseconds in duration and the overall shape of the wave form was clearly defined. Occasionally. lakes or other 

bodies of water produced Ifholes If (near zero returns) which lasted for a few 

seconds and which sometimes moved laterally across the range interval displayed on the 0 scilloscope. During these occasional short periods. it was 

not always possible to ascertain whether the prf setting was correct. 

£. Evaluation of Prf Selection Simulator 

This simulator was used as a training aid in advance of the orbitaloperatiCll 

for familiarizing the tracking station operators with the problems which 

would be enoountered later in setting the prf. 

As might be expected, there were a few minor differences between the actual 

wave shape of the returns, as viewed on the "A" scope by an operator. and 

the wave shapes as generated by the simulator (refer to Section III. paragraph 

4). The major difference was that the simulator produoed a wave form representative of a uniform terrain. Actually, small groups of targets caused the 

radar display to have a noticeably higher return during some portions. 

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(A) PUEaO COLOIADO 

LAIGf .., EiioI (SIX STEPS OfF) 

(. TOPOGWHICAL ~P OF PUEILO, COLORADO 

Figure 44 - Range Inversion 

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FAI SLANT ItANGE 

I52NAUT MI 

PASS 47 

NTOS 

24 DEClMlEI 1-' 

NEAl SLANT RANGe 

155 NAUT M. 
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0.5 VOLT, 20 MICROSECOND 

PER MAJOR DIVISION 

0.5 VOLT, 20 MICROSECOND 

PER MAJOR DIVISION 

0.5 VOLT, 20 MICROSECOND 

PER MAJOR DIVISION 

0~5 VOLT, 10 MICROSECOND 

PER MAJOR DIVISION 

SECRET 

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. (A) PASS NO.8 

PRF APPROXIMATELY ONE STEP HIGH 

(8) PASS NO.8· 

PRF CORRECT-

(C) PASS NO.8 

PRF APPROXIMATELY THREE STEPS tJlGH 

(D) PASS NO. 24 

MOMENTARY NEAR-ZERO RETURN 

Figure 45 - Oscilloscope Pictures of Tape Playback 

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T1ME~ 

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The simulator wave shape for the conditioh of the prf being one step off 

was readily recognized. 

It is concluded that the prf selection simulator is a very useful tool for 

familiarizing tracking station personnel with the wave shapes that are 

encountered during orbital operations. It is also useful as a signal source 

during checkout of the wide-band equipment at the tracking stations. 

6. CLUTTERLOCK 

a. General 

The clutterlock loop compensates for angular displacement of the antenna 

beam with respect to the zero doppler direction. The clutterlock correction signal is telemetered via the F60 integrator monitor. This Signal 

voltage must be converted to an angle in degrees. The angle must then 

be compared with the error introduced into the antenna beam position by 

vehicle motion. The vehicle yaw and pitch motions both contribute to 

clutterlock error. The contribution of roll is minor and will not be considered. 

b. Clutterlock Correction Angle 

The F60 integrator monitor signal is proportional to the clutterlock 

offset frequency. The proportionality factor is f (cycles per second) = 

1000 X volts dc. The clutterlock angle in degrees is obtained by substituting the offset frequency fd in the equation: ~ 

. f,l).. 

Sin 4» = -z;- · 

c. Vehicle Motions 

(1) Pitch Reference 

.(15 ) 

The three axes of the vehicle were referred to integrating gyros. The 

pitch attitude was programmed by an input to the pitch gyro at a fixed 

rate of four degrees per minute (360 degrees per 90-minute orbit) to 

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keep the vehicle longitudinal axis generally perpendicular to the local 

vertical. The pitch attitude was also corrected by sensors which provided a continuous reference to the horizon. At the time of initiation 

of the operate command. the horizon sensors were switched off and 

the vehicle was allowed to continue along according to the four-degreeper-minute program. 

(Z) Yaw Reference 

A fixed yaw bias of :l:Z. 44 degrees was provided to align the antenna 

beam with the relative ground velocity vector. The polarity was 

switched plus or minus depending on whether the vehicle was in a 

northbound or southbound pass. 

(3) Bias Uncertainty and Deadband 

There was a predicted uncertainty of :1:0.4 degree for eaoh axis. This 

_ bias -uncertainty resulted from gyro drift. horizon sensor uncertainty. 

and other tolerances. A deadband of :1:0. 18 degree was designed into 

the attitude correction system. The vehicle tended to hang at one limit 

of the deadband in pitch. probably as a result 9f a- residual magnetic 

torque. 

(4) Angular Displacement of Antenna 

To a close appro~mation. the resultant angular displacement of the 

antenna caused by yaw and pitch motions can be shown to be: 

where 

8

a 

= resultant angular displacement of antenna in azimuth 

direction 

8

y 

= -vehiole yaw angle 

8

p 

= vehicle pitch angle 

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SECTION IX AKP-U-596 

R 0 = ground range to target 

Rs = slant range to target 

h = vehicle altitude. 

d. Clutterlock Commands 

e. 

The clutterlock circuitry was designed to close the loop simultaneously with 

the operate command. However, the circuits were not designed to open the 

loop with the off command. The opening of the loop after each pass should 

have been accomplished by sending a clutterlock short command but it was 

not done. Consequently, the clutter lock loop was open and free to drift 

during pre-operate for all passes except 8 and 57. No harm was done 

however. as the integrator does not appreciably drift. 

Clutterlock Response - Pass 14 

For pass 14, the clutterlock loop acted immediately, upon applic,ation of 

transmitter high voltage, to cancel out an error of about 0.37 degree of 

antenna angular displacement resulting from an error in vehicle attitude. 

Figure 46(a) shows the section of the radar map and ~ graph of the corresponding clutterlock correction angle plotted to the same time scale 

as that of the map. The high voltage was applied at system time 57897 

seconds. At this time no targets are visible on the map because of improper sampling of the doppler spectrum. Within a few seconds, the 

clutter lock loop had compensated for the antenna displacement error and 

the ground painting was good. A clutterlock time constant of about Z. 5 

seconds is indicated (see Figure 46(b» which is in excellent agreement 

with the design value of Z. 5 seconds for the fast time' constant which was 

used for this pass. The radar display is somewhat degraded in this instance because of the reproduction technique employed. 

f. Clutterlock Loop Open - Pass 56 

During pass 56, the clutterlock loop was opened by means of 8; shorting 

relay at the integrator output. The short was applied seven seconds 

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SECTION IX 

after the pass began. The short in the loop caused the system to map 

along the line of zero doppler rather than the line of constant doppler 

which was centered in the illumination of the physical antenna. Figure 

47 is a section of coastUne mapped in this condition. The undesirable 

effects are quite evident. The results obtained are easily explained if 

it is recalled that the boresight of the antenna and zero doppler line are 

not usually coincident. When the angular difference is as small as 0.25 

degree. the doppler spectrum of the desired targets is not properly 

illuminated and the level of azimuth ambiguities is greatly increased. 

The fact that the system mapped properly. just previous to opening the 

clutterlock loop. verifies that the smeared image is caused by the absence of synthetic beam steering. 

I.. Clutterlock Operation over Water - Pass 16 

The pitch angle. yaw angle. resultant antenna displacement angle (Equation (15». and clutterlock angle (Equation (14» are plotted for pass 16 in 

Figure 48. During this pass the operate command was initiated over 

water. The clutterlock loop operated immediately. but with a slower 

response than that over land to correct the vehicle attitude error. Over 

water. the only signal for clutterlock to operate on is the return from 

wave patterns. Figure 49 is a radar map of the coastline in the vicinity 

of the Eel River in Northern California crossed about 12 seconds after 

the operate command. 

h. Clutterlock Performance - Other Passes 

Pass 30 is another good example of clutterlock over-water performance. 

In this case the operate command was initiated over Lake Superior. The 

pitch angle. yaw angle. resultant antenna displacement angle. and clutterlock angle are plotted in Figure 50 for this pass. At the start of passes 

30. 40. and 56 (as in pass 14). the clutterlock had to compensate for 

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SECTION IX 

+0.3 

Ui" 

w 

w +0.2 GO! 

C> 

w 

0 

- +0. J w 

..... 

C> 

Z 0 < 

~ 

U 

0 -0. J ..... 

GO! 

w 

I5 -0.2 ..... 

u 

: 

I 

I 

~r-- - J 

I 

1 

: /' 

V 

I 

-0.3 I 

I 

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(A) LAND MASS NEAR TITUSVILLE, PENNSYLVANIA 

(8) CLUTTERLOCK ANGLE 

~ ............ 

--7 r------ -------

L 

57,897 57,899 57,901 57,903 57,905 

4000-49 

SYSTEM TIME (SECONDS) 

Figure 46 - Clutterlock Response 

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SECTION IX 

ODIT DIRECTION -.. 

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(A) SHORELINE MAPPED WITH 

CLUnERLOCK CIRCUIT DISABLED 

(B) CONFIGURATION OF 

ACTUAL SHORELINE 

AKP-II-596 

FAI SLANT lANGE I~ NAUT MI 

PASS 56 

STON 

24 DECEMIEl 1964 

. NEAR SLANT lANGE 

164 NAUT MI 

Figure 47 - Deterioration of Mapping with Clutterlock Loop Disabled 

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SECTION IX 

.. " .! .,'J 

-to. 5 

0 

..- -D.S C) 

AU 

e 

AU 

... 

C) 

Z 

< 

-to. 5 

0 

-D.S 

-to. 5 

o 

-D.S 

68,600 

-4000-51 

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01SO - PITCH GYRO OUTPUT 

0151 - YA.W GYRO OUTPUT 

• tRESULTANT A.NTENNA. DISPLACEMENT 

• t--..... 

CLUTTERLOCK ANGLE 

~ ~ 

• 

68,6SO 68,700 68,7SO 68,800 

SYSTEM TIME (SECONDS) 

Figure 48 - Vehicle Motions and Clutterlock Response - Pass 16 

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relatively large errors in vehicle attitude. Initially, for these passes, no 

targets were present on the map film and after a few seconds the mapping 

was normal. For all of pa$s 57, the clutterlock time constant was set to~ 

time constant number two (5 seconds). The time constant could not be . 

measured because the vehicle attitude was apprOximately correct at the 

start of the pass, but the loop performance appeared to be normal. 

7. AZIMUTH AMBIGUITIES 

!.. General 

The angular locations of ambiguities are functions of prf and vehicle 

velocity (refer to Volume I, Section II, paragraph 3, and Section m, 

paragraph 4, for a general discussion of azimuth ambiguities). 

As· was shown in Volume I (page 10). the angle to the first ambiguities is 

given by the relationship 

where 

~ = carrier wave length 

F = prf 

V = vehicle velocity. 

'Y _ll 

I - 4V (16) 

These ambiguities are unfocused and do not produce discrete ambiguous 

images in the final map, but do act to increase the map noise level. 

The second ambiguities which appear at Z 'Yl are focused and produce 

ambiguous images in the final map. 

The system was designed to keep the level of azimuth ambiguities below 

the detectable limit in the final output map. However, under certain 

circumstances, azimuth ambiguities can become evident. This phenomenon occurred during pass 30 over Cleveland, Ohio. 

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--OII1T OIIECTION 

--OIIIT DIRECTION 

(A) Ea IlVEI EMPTYING INTO PACifIC OC!AN 

(HOITHEIN CALIFORNiA) 

r 2'Y1---, 

(I) AZIMUTH AMBIGUITIES WHICH OCCURRED DUlING 

HIGH GAIN SETTING OVEI WATEI 

(CLMLAND, OHIO, SHOIELlNE) 

Figure 49 - Clutterlock Operation over Water 

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NEAl SLANT lANGE 

156 NALIT MI 

PASS 16 

NTOS 

22 DECEMIEI 1964 

FAI SLANT lANGE 

162 NALIT MI 

NEAl SLANT lANGE 

155 NALIT MI . 

PASS 30 

NTOS 

23 DECEMIEI 1964 

FAI SLANT lANGE 

161 NALIT MI 
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..0.5 

o .,j ~ 

-0.5 

..0.5 

0 .~ 

-0.5 

..... 

fa 

0 

...... 

UoI 

-' C) 

Z 

~ ..0.5 

0 

-0.5 

..0.5 

o • ) 

-0.5 

57,3~ 

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DI~ - PITCH GYRO OUTPUT 

0151 - YAW GYRO OUTPUT 

RESUlTANT ANTENNA DISPlACEMENT 

CLUTTERLOCIC ANGLE 

5/,«10 5/ , .. ~ 5/,!JXJ 51,5!iJ 5/,6IXJ 

SYSTEM TIME (SECONDS) 

Figure 50 - Vehicle Motions and Clutterlock Response - Pass 30 

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" M-i>.. 'Azi muth Ambiguities in Radar Map , - - -

SECTION IX 

For pass 30, the pd was set at 8415 pps when Cleveland was mapped. At 

,'this time some azimuth ambiguities were present (see Figure 49(b». The 

relative velocity of the. vehicle with respect to stationary targets in the 

Cleveland area was approximately 25,080 feet per second. These values 

set the angle to the first focused ambiguity, 2 'YI , to be 0.01712 radian. 

'Rela.tlng this to the terrain being mapped, the location of the first focused 

ambiguity occurs Rs X 2'Yl nautical miles from the actual target, where 

R = slant range to the target. In the case of Cleveland, this distance s. . . 

is ~. 7 nautical miles Which corresponds to. apprOXimately 1.3 inch on the 

map film shown. 

When the satellite approached Cleveland over Lake Erie, the receiver 

gain was 12 to IS db higher than when the city of Cleveland was actually 

mapped (see system time 57471 in 'Figure 38(a». This was because the 

agc action increased gain over areas of low return such as water, and 

reduced gain over high return areas. The ambiguous targets (one string 

of which is identified above the radar map) have an angular separation 

corresponding to 2 'YI • 

. When the receiver gain is reduced, as is evidenced by'lack of return 

from the water just off shore, no further focused ambiguities can be 

observed in the water even though strong land targets exist at the proper 

separation angle. There does not appear to be any evidence of the presence of unfocused ambiguities. 

It can be concluded that the criteria chosed for the control of azimuth 

ambiguities appear to be adequate except for conditions similar to those 

over the Cleveland land-water area. 

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SECTION IX 

8. RANGE AMBIGUITY 

a. General 

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As discussed in Section m of Volume I. it is possible under some conditions 

for range ambiguous targets to appear on the map film. A llkely example 

of this phenomenon is shown in Figure 51. The mountain along the railroad tracks between Enid and Heaton, Arizona (see arrow on Figure 51) 

does not appear on the most detailed topographical maps of the region. 

The return could be tlui.t from a low outcropping which does not appear on 

the map, but it is more likely that it is an ambiguous return from a nearby 

mountain outside of the mapped area. 

~. Ambiguous Interval 

At the time the targets in Figure 51 were being mapped. the prf was in 

error by one step. The prf used was 8381 pps which corresponds to an 

ambiguous ground range interval of approximately 16 nautical miles. 

Topographical maps of the area show mountainous terrain located at this 

distance which could account for the image. The level of range ambiguity 

suppression in this case was reduced somewhat because of the error in 

prf setting. 

9. OTHER ITEMS OF INTEREST 

a. General 

Phenomena known as target swallowing and range sticking were detected 

in some portions of the radar maps. Evidence of fUm fogging as a result 

of small light leaks was found. These effects are discussed in the following 

paragraphs. 

I!.. Target Swallowing 

The map of Phoenix (see Figure 36) shows this effect in one area. In the 

lower right hand corner of the map is a group of large targets from an 

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S'EOTION IX, 

oil storage area. The dark area extending to the right of the group is a 

swallowed area. 

Swallowing occurs when a very large target utilizes the entire dynamic 

range of the system to the exclusion of lesser targets at the same range 

but slightly different azimuth • 

.2,. Sticking 

Side-looking coherent, radar systems commonly produce long range streaks 

when large point targets are encountered. The general effect has been 

given the name "sticking" as the targets appear as sticks in range instead 

of points. In the KP-U radar the following malfunctions or misal1gnments 

could theoretically cause sticking: 

1. Direct leakage through the chirp networks during the 

recompression (dechirping) process. This could cause 

a stick on the near range side of the target. 

Z. Leakage of 70 megacycles through the sine x/x modulator 

into the i-f channel and thence to the transmitter. This 

could cause a balanced stick with the target iii the center. 

3. Demodulator unbalance before or after the sine x/x pulse 

caused by a base line shift in the sine x/x pulse. This shift 

can make it impossible to balance the modulator both 

before and after the pulse. A stick can appear either 

before or after the target, depending on which balance 

is chosen. 

The map of Phoenix (Figure 36) shows two large targets @With long 

sticks In the far range direction. These are attributed to modulator 

unbalance caus ad by base line shift. The sticks appear only on the 

very largest targets. 

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TERIAIN NEAl ENID, ARIZONA 

Figure 51 - Possible Range Ambiguity 

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AKP-II-596 

NEAR SLANT RANGE 

165NAUT MI 

PASS 9 

5TO N 

22 DECEMBER '964 

FAR SLANT RANGE 

171 NAUT MI 

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d. FUm Fogging 

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A fogging effect was evident approximately IS inches before the end of each 

of the recovered map films. Fifteen inches on the map film correspond 

to approximately 40 inches of data film. This figure places the spot of 

fogging some 40 inches from the Recorder crt. or somewhere between the 

Recorder and the take-up spool. Since the film is not running between 

active passes. any slight light leak in the vehicle could have produced the 

fogging. The fogging resulted in only an insignificant amount of lost data. 

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1. GENERAL 

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AKP-U-596 

SECTION X - CONCLUSIONS 

This report has documented the KP-II doppler radar program from inception 

through completion of an orbital test in an Agena satellite. All of the program 

objectives were realized during the first flight. The success Qf the mission in 

large measure resulted from the extensive environmental testing and the associated corrective action program. 

z. PROGRAM RESULTS 

The results of the KP-D radar program are summarized below: 

1. The use of doppler high-resolution radar for mapping of 

the earth's terrain from an orbital satellite is feasible 

within the present state-of-the-art. 

Z. The performance of doppler radar systems at satellite 

altitudes compares favorably with the performance of 

lower altitude aircraft-type radar systems. .In p.arUcular, satellite-borne systems are not subject to the 

limitation of resolution normally imposed by platform 

instability in aircraft. 

3. Satellite-borne radar systems, as expected, have the 

ability to map through all weather conditions and either 

by night or day, thereby overcoming the main limitations 

of photography. 

4. Wide-band video links and ground-based photographic or 

tape recording provide practical methods of data recovery. 

5. The design criteria used for the KP-U radar have provided 

a firm basis for design of future satellite radar systems. 

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SECTION :x 

6. The use of clutter lock for steering the synthetic beam was 

proved to be highly successful. 

7. High altitude radar back-scattering coefficients were shown 

to be "in close agreement with previous low altitude data (refer 

to Appendix fi). 

3. RECOMMENDATIONS 

A number of improvements could be developed for incorporation in the present 

equipment or in future systems. A partial list of these is given below: 

1. The present radar signal-to-noise ratio could easily be 

increased by the use of parametric amplifiers. Presently available unttscould improve the signal-to-noise 

ratio by 5 to 6 db. 

z. The range interval could be ext.ended by three methods: 

a. By eliminating the azimuth image ambiguity. 

This would extend the range interval by a 

factor of two. It could be accomplished by 

arranging the data processing method so as 

to eliminate the need for an azimuth offset 

reference. 

b. By the use of a longer antenna. However, 

the azimuth resolution is theoretically 

limited to one-half. the physical antenna 

length. 

c. By the use of multiple channels. This 

can be done by using separate antennas, 

each operating at different frequencies, 

and illuminating different range intervals. 

Alternately ,a single transmitter frequency 

can be used with multiple antenna sections 

being used for reception. 

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3. Range resolution could be improved to be more consistent 

with the present azimuth resolution. This can be readily 

accomplished by increasing the system band width and 

increasing the number of resolution elements of the recorder. 

4. On-board magnetic or dielectric tape recorders could be 

used for delayed read out of doppler data. 

AKP-U-596 

5. Automatic circuits could be incorporated in the payload for 

control of prf selection Independently of ground commands. 

The system could then be programmed for operation during 

any portion of the orbital period. 

6. The system could be repackaged for decreased weight and 

volume. For a system having the KP-U performance a 

reduction of weight from 340 pounds to 250 pounds is 

feasible. 

7. Real time electronic correlation and display of selected 

small areas are possible. 

4~ FUTURE APPLICATIONS 

Satellite-borne radars have wide application to both military and civilian 

requirements. These include location of military targets. bomb damage 

assessment. and general surveillance over land and water. Radar data can 

be used to improve present maps of the Earth, particularly for polar or other 

remote areas. Detailed mapping of the moon and the planets by orbital satellites using doppler radar is a next logical step. 

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APPENDIX I 

CHRONOLOGY OF DEVELOPMENT HISTORY 

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January 

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Al:P-D-596 

Design approval tests completed on first radar system. Resolution specifications 

met using system tester. Qualification model Recorder, RF-If, Reference Computer and Control units shipped to LYSC. 

Peak of fabrication manpower reached. 

February. and March 

Qualification model Transmitter-Modulator shipped to LMSC. 

Potting problem encountered in Transmitter-Modulator pulse transformer. 

Hydrogen thyratron containers ruptured by expansion of potting compound. 

Frequency multiplier in R-F /I-F unit failed during vibration tests. 

Radar System No. Z acceptance tests completed. 

April 

Transmitter-Modulator failed during qualification tests. 

Transmitter-Modulator trigger circult found to be sensitive to power supply input 

voltage. 

Qualification testing of radar System No. 1 started in temperature-altitude chamber. 

Manuals completed for all units and testers. 

Frequency multiplier redesigned using bonding and foam potting and successfully 

tested under vibration and low pressure. 

Voltage breakdown occurred in traveling wave tube at low pressure. Vendor directed 

to incorporate vacuum potting of high voltage leads. 

Light level change occurred in cathode ray tube during vibration tests. 

May 

Remaining testers shipped to LMSC • 

Ground-based equipment shipped to New Boston and Vandenberg AFB. 

Three pulse transformers failed during low-pressure testing. 

Problem of cathode ray tube light intensity change isolated to electron gun structure. 

Vendor directed to change desigJi. 

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APPIBDIX I 

Pulse-forming network experienced voltage breakdown during qualifioation testing 

in temperature-altitude chamber. 

Compatibility tests with 'lena vehicle successfully concluded in anechoic chamber. 

New bellows-type. oil-filled pulse transformer received from vendor. 

Two new cathode ray tubes with modified gun structure tested with satisfaotory reSUlts. 

Qualification radar system sent to LMSC Santa Cruz. California. test facility for 

combined radar and antenna tests. 

August 

Santa Cruz tests completed with resolution within specifications. 

Trans mUter-Modulator failed because of charging choke. 

September 

Qualification tests repeated on radar system in temperature-altitude chamber. 

Power factor change encountered as a function of pressure. 

Charging choke problem isolated to vibration. New design started using epoxy sand 

filler. 

Clutterlock loop problem recognized. 

October 

Data film drive problem isolated to nose oapsule film take-up. spool. 

Take-up film tension increased. 

Pulse transformers failed because of internal voltage breakdown. 

Clutterlock problems isolated. CoaXial connector shields and filter added. 

November 

Voltage breakdown problem in pulse transformer isolated to tuning inductor and 

design change incorporated. 

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APPENDIX I. 

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AKP-n-S96 ... 

Power factor change problem isolated to radiO frequency interference filters and 

corrective action taken. 

December 

Radar system tests successfully completed in temperature-altitude chamber. 

Vehicle launched on December Zl. Film capsule recovered on December Z:4. 

January 

Analysis of radar map data indicated excellent results. 

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APPENDIXU 

AICP-D-596 

ESTIMATE OF RADAR BACK-SCATTERING COEFFICIENTS 

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APPENDIX U - ESTIMATE OF RADAR BACK-SCATTERING COEFFICIENTS 

1. GENERAL 

Telemetry data of returned radar power were used to determine';the back-scattering coefficient, 0'0' for various types of terrain. The results presented 

here are based upon a particular set of assu~ption8; it is recognized that 

another study, using the same flight data, might yield somewhat different results. As only a small portion of the telemetry data were reducedinth1s study, 

it can be expected that a more complete study would yield a 'wider spread of 

values for the various terrain types. The results presented here are believed 

to be representative. 

z. CALIBRATION OF RADAR RECEIVER 

Prior to the orbital test, a calibration curve was obtained for the flight radar 

receiver relating the gain control mOnitor voltage, F53, to received 'power • 

This calibration curve was previously given in Section IX (see Figure 38(b), 

page 118). 

The calibration was made by reference to the thermal noil!i'e at the receiver 

input. The noise level was calculated from tile, relation 

S = KTB(NF) (A-I) 

where 

S = receiver noise power 

K = Boltzmann IS constant 

T = temperature degrees K 

B = receiver band width 

NF = receiver noise figure. 

An X-band traveling wave tube in series with a calibrated attenuator was used 

as a random noise power signal source. With the agc loop open and the stc 

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APPENDIX D 

. functioning ,the signal source was injected at the receiver input. The series 

attenuator was varied to double the receiver output (signal equal to receiver 

noise). Using this level as a point of reference, the curve of Figure 38(b) in 

Section IX was then obtained by varying the series attenuator and recording 

the voitage output at F53. (For the following calculations of 00' the curve was 

moved to the right by o. I volt to compensate for an in-fiight shift of the F53 

voltage from o. 75 volt to 0.85 volt for no signal input to the receiver. ) 

3. METHOD OF CALIBRATION AND ASSUMPTIONS 

The equation used in the calculation of 00 was obtained by re-arranging Equation 

(3Z) (Volume I, page Z9). This yields . 

P (411')3 R3 = __ ~~r~ __________ _ 

0"0 PtGZ~Z •• 5C~tsece (A-Z) 

The quantity P r was determined from the telemetered value of F53 voltage for 

each particular terrain using the calibration curve. The rangeR was a midrange value obtained from flight data. Pt was determined from telemetered 

information. The remaining quantities other than antenna gain G were assumed 

8 I -6 constant (~ = o. 10Z ft, + = 0.0064 rad, C= 9.85 X 10ft sec, ~t = 10 sec. 

and 9 = 530). 

Although the peak gain of the physical antenna was known, normalizing factors 

were required to account for the variation of gain across the ground area which 

contributed radar returns to the receiver. These normalizing factors were 

determined in the following manner. 

In the azimuth plane a rectangular beam pattern was assumed which had a 

width equal to the 3 db width of the actual pattern and a constant gain equal to 

the maximum gain of the actual pattern. This approximation is valid from a 

total received energy s'tandpointfor antennas (such as the one being used) with 

uniform illumination. 

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In the range direction. the received power P is sampled at the agc circuit 

r 

input by a gate 68 microseconds long. occurring ZI. 6 microseconds after 

reset (clock zero). An average antenna gain during this interval was calculated. The variations in gain from prf step errors and altitude changes were 

also taken into account. 

The resulting normalized gain values versus prf step error and altitude are 

shown in Figure A-I. Only two altitudes. one for northbound and one for 

southbound passes. were used since the altitude variations were small for the 

passes used • 

... 

o 

-

z 

~ -41 

40 

o 

/138 NAUT MI 

128NAUTMIJ 

2 

PRf STEP ERROR 

Figure A-I - Antenna Gain versus Step Error and Altitude 

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APPENDIX n 

4. EXAMPLE OFU' CALCULATION 

o 

For the calculation for pass number 9, system time Z9543, broken desert north 

of Phoenix, Arizona, 

9 = 530 , R = 171.3 mi, P = 85.5 dbm, Pt = Z5.3 kw r . 

G = 41.5 db, ~ = O. 10Z ft, C = 9.85 X 108 ftlsec, + = 0.0064 radian 

U' = Z.8Z X 10-Z (411";3 X (l7l.3 X 6080)3 

o Z5.3 X 103 (1. 4Z X 104)Z (0. 10Z)Z(0. 0064)(9. 85 X 108)(10-~) 0.5 (1.66) 

U' = 0.OZZ8 

o 

U' = -16.4 db. 

o 

5. COMPARISON BETWEEN KP-II RESULTS AND PREVIOUS EXPERIMENTAL 

DATA 

Figure A-Z shows experimental curves * relating U'o to types of terrain and 

aspect angle. (This figure is identical to Figure 10 in Volume I and is repeated 

here for reference. ) 

The results of calculations for U'o as obtained in the satellite doppler radar test 

are shown in Table A-I. The table is divided into two sections: the first seotion 

shows the comparison of KP-U measurements with the data of Figure A-Z; the 

second section shows the values of the various quantities used in the calculation 

of KP-II results. 

6. CONCLUSIONS DRAWN FROM KP-D RESULTS 

The results show good correlation with the previous results of Figure A-Z. All 

comparisons must, of course, be made at an aspect angle of 53 degrees. 

* Data on sea return are from Grant and Yaplee (refer to Proceedings of the I. R. E •• 

July, 1957). Balance of data was experimentally obtained as a result of Goodyear 

Aerospace Corporation flight tests in a PZV aircraft. . 

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a. Sea Returns 

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For sea returns, the results agree quite 'well with the Figure A-Z data. The 

Pacific Ocean calculation falls close to the value for the illustrated 15 to ZO 

knot point of the Grant and Yaplee data. A weather map for the day of the 

test showed a ZO to Z5 knot wind. 

The return from Lake Huron and Lake Superior is considerably higher than 

for the sea returns above. From the appearance of these lakes on the radar 

maps, they were assumed to be frozen. 

k. Broken Desert 

These figures agree very well with the previously obtained data • 

.2,. Heavy Vegetation 

The results given for a forested area of Northern Michigan and a coastal 

swamp, area in North Carolina agree quite well with the data of Figure 10. 

d. Cultural Areas 

The results for cities showed a return averaging around that for the residential area data of Figure A-Z. This is to be expected, however, because 

of the large area which is covered by the KP-D beam. The KP-D result is 

obtained by integration over an area approximately ten miles by one mile. 

The 0"0 result is therefore a composite of many terrain types as a large city 

can contain many different terrain samples: open fields, water, heavy vegetation, and residential, commercial and industrial areas. 

The sampling rate of the telemetry system (one sample per four miles) is 

much too low for accurate determination of 00 over such areas. For future 

flight tests a continuous channel should be used for recording the agc voltage. 

7. ACCURACY 

These results are subject to the measurement errors inherent in measurements 

taken from a complex electronic system during operation. However, the relative 

accuracy should be better than the absolute accuracy. 

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l 

i 

b" 

CIt 

,. ... 

Z 1ft ... n Q • 5 ,. I 

.... u :t -..I C) > • ~ 

.. ~ I 1M 

... ,. I::: 

-

Z ~ Q J: 

~ 

2226-29 

~ 

20, i· 

15.1----

10 ... ---

5 •• ---

0.1----

-5.1---.-....; 

-IO .. I---t----+--~ 

-15 ..... ---""1 

-2.5 ... ---..., 

-301 = 

-35 •• ---

~~!--~ 60· 70 80 

ASPECT ANGLE • (DEGREES) 

Figure A-Z - Back-Scattering Coefficients versus Aspect Angle 

> 

CZ 

.;n ~ ;a 

~~~ I.A ~ ~ -;a ::!!o 

m< Cm 

OmC 

~~cg 

COC);a 

c...~;a C:;am !<-trN m o ,. -. en 

N m 

> 

~ Ito 

II!t.1 

~ 

CIt 

.. 

1ft 

~ . 

.... 

.~ 

'Z' 

'0 . .... 

Z· 

Q 
CIt 

.. 

n "' • ,.. 

% • Z 

o 

,.. 

Z 

g 

.... 

-..J 

U1 

II .. 

Terrain 

typa 

8ea 

Lake. 

Lak •• :froaon) 

Broken 

de.en 

'arllllaad 

Se."y 

yacetaUon 

lIoantata. 

CeUllral 

ar.a. 

.. .. 

TABLE A-I - BACK-SCATTERING COEFFICIENTS 

Comparl.on datil aafer_.a 6l&a 

Baok-.oaHerln,ooofl1olant. 0' .. (db) 8y.tam R(mldl PI' Itt 

Looation KP-U rlpre 10 Paaa Uma (naut III ) (dblll) (kw peak) 

Atlantic Ocean -l4.l -Z3 to -31 14 58007 157.6 -9z.0 Z5.7 

Paolflc Ocean -H.I. t6 &l7Z5 157.8 -9Z.0 Zs. 7 

Lake Erie -l3.0 Not slven 30 57466 157.5 -91.0 z6. 3 

Lake Mlchl,an -19.7 Z4 2:3899 169.0 -88.0 Z6.4 

Lake Huron -IS. 9 Mot slven 8 1436Z 170.0 -84.0 Z6. 7 

Lake 8uperlor -15.8 30 57385 157.5 -84.0 Z6.3 

Nortll of PlleIenlx -16.4 -15 to -17 9 Z9543 171.3 -85.5 Z5.3 

Soutllern utah -14.7 Z5 Z9Z57 169. Z .-83.0 n.1 

Soatll of C IaIcaro -11.1 Not sly .. 14 Z3847 169.0 -79.5 Z6.4 

1II0laican -11.3 30 57540 157.5 -79.5 z6. 3 

Coa.tal a_DIp. N. Carolina -11.Z -8 to -13 14 57997 157.6 -79.0 Z5.7 

'or •• t. IIlolaI,aa -II. Z Z4 Z391Z 169.0 -79.5 Z6.4 

Weat Vlr,lnla -11.5 30 57535 157.5 -79.5 a6.3 

Arlaoaa -11.5 Not slY- 9 Z9547 171.3 -80.5 as. 3 

uealt -11.8 as z9ua 169. Z -80.5 27.1 

Clenland -6.7 Not apeoUloall,r 30 57470 157.5 -75.0 a6.3 

Clalearo -7.6 slye .. Sa. dll- Z4 .!3870 169.0 -76.0 a6.4 

PIleI_ix -10.5 OII •• lon. 9 .!9537 171.3 -79.5 as. 3 

Prf arror Aa .... . ..,. ....... , 

0 fl. 4 

Z to. , 

I 41.3 

0 41.6 

0 .... 1 4l.3 

I 41.5 

I 41.5 

0 41.6 

I 41.3 

0 41.4 

0 41.6 

1· 41.3 

1 41.5 

I 41.5 

I 41.3 

0 41.6 

1 41.5 

>ooz 

~mm:::u ~OOo 

tIj.!j: !j: ~ ~.en en :g -!!! !!! :::u 

"":::!! :::!! ° ,.H2I2I ~ _om O _z~"TI 

.... ° CDO:::u 

c...S;:::u C::::u m r--tro 

~ en ~ ~ m 

~ ~ 

I 

= I 

UI 

~ 

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APPENDIX ·n· 

Consideration of the calibration curve shoW's that .a two percent change in telemetry could cause -an error of 4 to 5 db in (To for low signal levels. For high 

level signals this error is reduced to approximately 1 db. The FS3 monitor 

circuit should be redesigned to reduce low level signal errors. 

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