THE UNIVERSITY O F MI CHI GAN COLLEGE OF ENGINEERING Department of Electrical Engineering Space Physics Research Laboratory Engineering Report No. 2 ENGINEERING DESIGN OF A PITOT-STATIC PROBE PAYLOAD Go F. Rupert ORA Project 05776 under contract with: NATIONAL AERONAUTICS AND SPACE ADMINISTRATION GODDARD SPACE FLIGHT CENTER CONTRACT NO. NAS5-3335 GREENBELT, MARYLAND administered through: OFFICE OF RESEARCH ADMINISTRATION ANN ARBOR April 1967

TABLE OF CONTENTS Page LIST OF TABLES v LIST OF FIGURES vi ABSTRACT x 1o INTRODUCTION1 1o 1 History 1 1,2 Design Objectives 1 2. SYSTEM DESIGN 2 201 Description 2 2,11 Tracking 2 2o12 Aspect 5 202 Densatron System 5 2,21 Amplifier 10 2,22 Range Selector Circuit 12 2.23 Power Supply 12 2.3 Instrumentation Section 12 2531 Control Deck 14 2,32 Battery Supply Module 23 2 33 Shift Register 24 2.34 Magnetometer Deck 24 2.35 Special Circuits 30 2,351 Thermistor Supply 30 25352 Calibration Regulator 30 2.353 Calibrate Timer 34 2,354 Pedestal Inverter Supply 39 2,4 Telemeter System 39 2.41 RF Link 42 2. 42 Data Formats 45 2.43 System Component Discription 49 20431 Transmitter Deck 49 20432 Commutator Deck 49 2,433 SCO Deck 49 2.434 Telemetry Antennas 58 2o5 Ground Control System 58 2.51 Umbilical System 61 2o52 Control Console 65 iii

TABLE OF CONTENTS (Concluded) Page 3. MECHANICAL DESIGN 67 3.1 Probe Section Assembly 67 3.2 Telemeter Section Assembly 78 3.3 DOVAP Assembly 81 3.4 Despin Module 81 4. TESTING 88 4.1 System Tests 88 4.2 Environmental Tests 89 4.21 Vibration 89 4.22 Thermo-Vacuum Test 89 4.23 Dynamic Balance 90 5. SYSTEM PERFORMANCE 93 5.1 Aspect Data Accuracy 93 5.2 Telemetered Data Accuracy 93 5.21 Response 93 5.22 Noise 94 5.221 Densatron Output Noise 94 5.222 Telemeter System Noise 94 5.23 Voltage Errors 96 5.24 Error Summary 97 5.3 Reliability 97 5.4 Summary of System Performance 98 6. LAUNCH OPERATIONS 99 7. FUTURE CONSIDERATIONS 106 7.1 Wind Measurement 106 7.2 Amplifier System 106 7.3 Ionization Gages 107 8. APPENDIX 108 8.1 Nose Cone Heating 108 8.2 Engineering History of Pitot-Static Probe Launchings 117 8.3 Equipment Specifications 125 8.4 Thermal Vacuum Test 149 9. REFERENCES 159 iv

LIST OF TABLES Table Page Io Gage Ledex Program Format 21 IIo Telemeter Ledex Program Format 22 III. Pull Away System Cable Assignments 23 IVo PAM Channel Assignments 47 V, Magnetometer Channel Assignments 48 VI. Densatron Amplifier Noise Characteristics 95 VII. Launch Operations Time Schedule 100 v

LIST OF FIGURES Figure Page 1. Pitot-Static Probe rocket before launch. xi 2. Rocket configuration. 3 35. Typical flight profile. 4 4o The Densatron. 6 5~ Densatron schematic 7 6A. Tritium source ionization gage. 8 6B. Current/pressure characteristics of tritium source ionization gage. 9 7. Amplifier block diagram. 11 8. Range selector functional diagram. 13 9. Instrumentation block diagram. 15 10. Instrumentation with shift register-side view. 16 11. Instrumentation with magnetometers-side view. 17 12. Column deck arrangements. 18 15A. Gage Ledex switching format, 19 13B. Telemeter Ledex switching format, 19 14. Control deck —front view. 20 15o Battery module-front view, 25 16. Battery module wiring. 26 17A. Magnetometer deck-front view. 27 17B. Magnetometer mounted on shift register. 28 vi

LIST OF FIGURES (Continued) Figure Page 180 Magnetometer deck wiring. 29 19o Thermistor circuit characteristics, 31 20. Thermistor supply schematic, 32 21, Calibration regulator schematic. 33 22. Calibration regulator compensation network curve, 35 235 Calibration regulator temperature curve. 36 24, Calibrate timer schematicO 37 25. Pedestal supply schematic. 40 260 Pedestal dupply load characteristics. 41 27, Pedestal supply temperature response. 41 28. Telemetry system data flow diagram. 46 29. Flight paper record; NASA 14.251. 50 30, Transmitter deck-side view, 51 31o Transmitter deck and battery module-top view, 52 32, Transmitter deck wiring. 53 330 Commutator deck-front view, 54 34. Commutator wiring, 55 355 SCO deck wiring, 56 360 SCO deck, 57 57A. Model 2 003 telemetry antennas, 59 37Bo Impedance/frequency characteristic for model 2,003 telemetry antennas, 60 vii

LIST OF FIGURES (Continued) Figure Page 38. Umbilical fly away release system. 62 59. Ground control console block diagram. 64 40. Ground control console-front view. 66 41A. Nose cone-rear view. 68 4lB. Disassembled nose cone. 69 41C. Nose cone assembly drawing. 70 42A. Nose tip-front view. 71 42B. Nose tube-side view. 72 435 Nose cone center section, 73 44. Nose cone transition section. 74 45. Unistrut assembly for instrumentation and DOVAP. 75 46. DOVAP section with hardware. 76 47. Nose probe assembly drawing. 77 48. Instrumentation section with electronics-front -view., 79 49. Top of instrumentation structure showing column support bracket, 80 50. DOVAP section mounted on Apache. 82 51. DOVAP transponder installed-side view. 83 52. DOVAP transponder installed-top view. 84 55. DOVAP section showing brackets for antenna coupling networks.85 54. Despin module assembled. 86 55. Despin module-showing weights, steel wire, and skins. 87 viii

LIST OF FIGURES (Concluded) Figure Page 56. Dynamic balance of nose cone at Wallops Island. 91 57 Dynamic balance of nose cone and Apache at Wallops Island. 92 58. Pressure calibration system-front view of control panelO 102 59. Pressure calibration system-front view, 103 60. Portable vacuum system, 104 61o Vacuum control unit for portable vacuum system. 105 62~ Curve of nose temperature AA6.340. 109 635 Curve of nose cone temperature-NASA 14o21o 110 64o Curve of nose cone temperature-NASA 14,21. Ill 65~ Curve of nose cone temperature-NASA 14.21. 112 66, Curve of nose cone temperature-NASA 14.21o 113 67, Curve of nose cone temperature-NASA 140251o 114 680 Curve of nose cone temperature-NASA 14o285. 115 69. Curve of nose cone temperature-NASA 14.289. 116 ix

ABSTRACT The engineering aspects of the Pitot-Static Probe experiment are presented with emphasis on the design and fabrication of the special instrumentation and circuits that were developed to make implementation of the technique possible. The fabrication of the mechanical system, testing procedures, and a prelude of launch operations are included. Future considerations towards system improvements and greater measurement accuracy conclude the report. x

LIST OF FIGURES (Concluded) Figure Page 56. Dynamic balance of nose cone at Wallops Island. 91 570 Dynamic balance of nose cone and Apache at Wallops Island. 92 58, Pressure calibration system-front view of control panelo 102 59. Pressure calibration system-front view. 103 60. Portable vacuum system, 104 61o Vacuum control unit for portable vacuum system, 105 62~ Curve of nose temperature AA6o340. 109 635 Curve of nose cone temperature-NASA 14o21o 110 64, Curve of nose cone temperature-NASA 14o21. ill 65o Curve of nose cone temperature-NASA 14o21, 112 660 Curve of nose cone temperature-NASA 1421o 1115 67, Curve of nose cone temperature-NASA 14,251o 114 680 Curve of nose cone temperature-NASA 14 285o 115 69~ Curve of nose cone temperature-NASA 14o289. 116 ix

I "~~~~~I'~K"'i~' if i:''''i':ii~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~i:8i'iiaiii~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~iiiii~~~~~~~~~~~~~~~~~~~~~~iiiill~~~~~~~~~~~~~~~~~~~~~~~~~~;i~~ NAA -66-334 Fig. 1. Pitot-Static Probe rocket before launch. xi

1, INTRODUCTION o 1 HISTORY The Pitot-Static Probe is a fully operational system capable of determining atmospheric density, pressure, and temperature in the altitude range between 30 km and 120 km. Basically, the technique involves making two direct pressure measurements on a rocket surface, one the pitot or impact pressure, the other being the static or ambient pressure (hence the term Pitot-Static Probe). The measured pressures are then used to mathematically derive the other parameters, Since the program's inception in early 1960, several launchings have been successfully carried out at various locations (see Appendix 8o2)o Initial development of the probe (following work by Ainsworth, Fox, and LaGowl) began with the support of the AFCRL Geophysics Research Directorate whereby a prototype nose cone was developed to investigate the feasibility of the experiment. Earlier work by this laboratory resulted in the development of the Densatron,2,3 a radioactive ionization pressure gage, two of which were subsequently used as the pressure transducers in the prototype nose coneo Other instrumentation needed to support the experiment were also developed and provisions were made for installing DOVAP4 instrumentation in the nose cone, A test launching of the prototype nose cone aboard a Nike-Cajun Rocket (AA6.340) was conducted in October41960,at Ft. Churchill, Manitoba, Canada. The flight objectives were realized and further work, with the support of NASA, Goddard Space Flight Center, towards the development of a practical system was initiated. 1o 2 DESIGN OBJECTIVES Besides making the basic measurements, primary objectives in the PitotStatic Probe design were to provide an experiment capable of being launched anywhere with minimum ground support equipment required and without loss of data accuracy. Every effort has been directed towards building a reliable system that requires little field maintenance. The payload design also included provisions for gathering engineering data that would increase knowledge of the measurement and facilitate system improvement Wherever possible, flexibility was included in the design to allow periodic updating of the instrument that would reflect the latest developments in the technique or instrumentati ono 1

2. SYSTEM DESIGN 2.1 DESCRIPTION The Pitot-Static Probe experiment is launched aboard a two-stage NikeApache rocket which is capable of reaching an altitude of approximately 140 km using the Pitot-Static configuration (Fig. 2). The nose cone serves as the actual measurement probe as well as housing the instrumentation. The pitot or impact gage orifice is located in the 305-in. diameter hemispherical nose tip. Ten equally spaced holes 5/15-in. in diameter located approximately 32 in. behind the tip are the inputs to the ambient pressure chamber. Useful pressure data areo gathered Only on the upleg portion of the flight. A typical flight profile is shown in Fig. 35 From the standpoint of useful data, the ambient pressure measurement begins almost immediately after Apache burn-out at an altitude of 20 km and continues to 85 km, while the impact measurement starts near 50 km and terminates near 120 km. During the measurement period, a number of support functions, which are described below, are necessary to complete the experiment. 2.11 Tracking Primary tracking information is obtained by DOVAP,4 a continuous wavetracking system which utilizes the doppler frequency shift to measure rocket velocie velocity profile may then be integrated over the entire flight path to yield the trajectory. A type UDT/B DOVAP* transponder is included as an integral part of the payload, and either a multistation DOVAP system consisting of a transmitter and at least three receiving stations or a Single Station (SSD), which utilizes interferometer techniques, may be used with the Pitot-Static Probe experiment, A detailed analysis and description of the systems may be found in the references and will not be repeated here.4,5,6 The SSD system offers a particular advantage to the experiment in that the ground support equipment is sufficiently portable to allow its use at remote locations thus maintaining the flexibility of the experiment with respect to possible launch sites. *ITT Labs, 2

IMPACT GAGE ORIFICEAMBIENT GAGE ORIFICE WIND GAGE ORIFICE TELEMETRY SECTION TELEMETRY ANTENNA PULL AWAY DOVAP DOVAP ANTENNA APACHE NIKE PITOT STATIC PROBE ROCKET CONFIGURATION NIKE ~ APACHE SCALE 3/" I' - 0 Fig. 2. 5

DATA FROM NASA 14.22 LAUNCHED 4 FEBRUARY 1964 AT iso - ASCENSION ISLAND SOUTH ATLANTIC OCEAN 140 - 120 E- N — OF PITOT MEASUREMENT (120 KU ) 100 ENO OF UISMBIENT PRESSURE MEASUREMENT (5 KM ) Y I \ / START OF PITOT MEAIURuEM(NT ( 0 KM ) 40 - to -S — TART OF ANMIENT PRIESSURE MEASUREMENT ( 20 KM) -— APACHE lURNOUT (@ 26 SEC / - -AP^CHI I1NITION Q 22 SEC - - NIKE BURNOUT * ).S *EC 0 40 0 120 I 0 200 240 200 320 30 TIME FROM LAUNCH - SECONDS TYPICAL FLIGHT PROFILE FOR THE PITOTSTATIC PROBE EXPERIMENT Fig. 3 4

2.12 Aspect The rocket angle of attack figures prominently in the data reduction of the primary probe measurement and is, therefore, considered a requisite of the experiment. An Adcole Digital Solar Aspect System is built into the payload. The system is composed of a model 2355 shift register and a model 135B aspect sensor that also includes an earth telescope which senses the presence of earth with a 1~ field of view (see Appendix 8.5). The sun angle is determined by sunlight passing through a slit in a quartz block which is screened by a grey coded reticle that allows or inhibits illumination of six photocells. This information is processed in the form of "ones" or "zeros" corresponding to an illumination or nonilluminated condition and then stored in a shift register for telemeter read-out. The angle of incidence of the sun with respect to the vehicle axis determines which cells are illuminated, and 27 or 128 unique combinations of "ones" or "zeros" represent 1~ resolution over a 128~ field of view. A serial read-out of the coded information is used for telemeter. The presence of an earth cell output causes a dc shift in the record that is easily distinquishable. The accuracy of sun data can be read to +0.25~ at transition. As a supplement to earth cell data, magnetometers are also used in the aspect measurement. Magnetic sensors mounted in the rocket determine the vehicle's position with respect to the earth's magnetic field vector. 2.2 DENSATRON SYSTEM The Densatron (Fig. 4), which is capable of measuring pressures found in the region between 30 km to 120 km, is the foundation from which the PitotStatic Probe experiment evolved. A radioactive ionization gage (Fig. 6A), that produces currents in proportion to neutral particle density, is combined with a multirange differential dc operational amplifier to form the Densatron System. A schematic diagram of the system is shown in Fig. 5. A detailed description of the ionization gage can be found in the references and will not be repeated here.2,7 Basically, the gage measures gas density which may then be interpreted in terms of pressure when the gas temperature is known. Figure 6B show a typical gage pressure-current characteristic. The electronic circuits used in the Densatron are composed entirely of solid-state components except for the amplifier input stage, where electrometer tubes are required to obtain the high input impedance. necessary to measure low currents. A thermistor mounted in the ionization chamber measures the gage wall temperature which, it is assumed, is also the gas temperature that must be known for data reduction. A second thermistor measures the temperature of the electronics section. 5

N.~~~~~~~~~~~~~~~~~~~~~~~~~~~~ I Fig. 4. The ]Densatron.

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O-8 RECOMBINATION cn I'9 loCt w ~~ ~ io" IC-) cr (0 |0-9 - W TYPICAL CURRENT - -' PRESSURE CHARACTERISTICS 0 o TRITIUM IONIZATION GAGE W 10-12 o3 RESIDU CURRENT 00' or - 13 10-4 10- 10 2 10io 100 101 102 PRESSURE -- mmHg Fig. 6B

Development of the system occurred over a period of years by various personnel of the Space Physics Research Laboratory. Further development of the system is currently in progress to extend the measurement range, and improve the accuracy and response of the instrumentO A description of the various components of the system follows. 2.21 Amplifier The amplifier block diagram is presented in Fig. 7. Using standard operational amplifier theory,8 an analysis of the amplifier circuit reveals an output voltage which is linearly related to the input current according to the relation EB - Eout i (1) gage Rf where Rf is the feedback resistance of the amplifier and EB is a dc bias voltage (usually about 6 v) added to the feedback loop. By rearranging terms, the output characteristic then becomes, out = B igage (2) From this relation, it can be seen that, given a constant EB, the sensitivity of the amplifier is dependant upon Rfo Since any particular gage pressure-current characteristic (Fig. 6B) will encompass many orders of magnitude, the resolution and accuracy of the measurement is increased considerably by changing the sensitivity of the amplifier in different pressure ranges. This is accomplished by having a number of different Hi-megohm resistors (Rf) available, and employing a selector circuit to insert the proper resistor into the feedback loop, A bidirectional Ledex stepping motor, that drives a specially constructed Ledex wafer switch, provides the necessary mechanical configuration. The selector circuit senses the amplifier output voltage and provides switching pulses to the Ledex motor that will either increase the sensitivity as the output increases beyond 5 v, or decrease the sensitivity as the output approaches 0 v, Five current ranges, each covering approximately one order of magnitude from 10 amperes are generally used in the Densatron. A voltage divider circuit wired to an additional wafer switch on the Ledex indicates the switch position; hence the range of the amplifier, 10

Rf EB 11 HIGH GAIN I DC AMPLIFIE I GAGE(-) r EOUT Fig. 7 11

2o22 Range Selector Circuit The range switching circuit is a form of voltage comparator that provides output pulses when predetermined voltages are sensed at the input. The amplifier output is coupled to the inputs of the sensing circuits through an isolation resistance (R6 in Figo 8) that allows external control of the switching circuit irrespective of the amplifier output level, The operation of the circuit can then be tested,before flight, Three functions are possible with the range control input. The amplifier may be either up-ranged or down-ranged by applying, respectively, a signal above a positive 5 v or a signal below 0 v. Sometimes, it is desirable to hold the amplifier in a particular range which may be done conveniently by applying any low-source impedance voltage lying between the switching pointso Switching action is controlled in either direction by the sensing amplifiers which command the appropriate drive amplifier to operate, The drive amplifiers are identical and are a form of monostable flip-flop which are used to "dump'" the charge from a 450 MFD capacitor through the Ledex coilo The capacitor provides a source of low impedance, high pulse energy needed to switch the Ledexo The pulse width may be adjusted through the monostable for best performance and efficiencyo The input stage to the monostable is "gated'" off during capacitor recharge, so that no switching can occur until sufficient energy is available in the capacitor to reliably switch the Ledexo By keeping the peak charging current to 100 ma or less, a maximum repetition rate of one switch every 3 sec is obtained, Under the conditions stated, the Ledex usually requires a 20 msec pulse that causes the capacitor to discharge to about 10 v, 2o23 Power Supply The many supply and bias voltfages required to operate the Densatron are derived from a dc-to-dc convertor included as part of the instrument A single, 1 w power source between 25 and 35 v will operate the unit, The circuit (see Fig. 5), is a standard square core oscillator operating at approximately 6 kCo Primary regulation is accomplished by zener regulators D5 and D6 and commutating diodes D2 and D3, which c lamp the amplitude of the square wave at a level determined by the voltage drop across the diodes and the regulab.orso The isolation and drift characteristics are improved in the amplifier, by using secondary regulation for the electrometer tube filament supply, and also the positive supply to the rest of the amplifier, 2, INSTRUMFNT T.QION SECTION The operational characteristics of the experiment are primarily defined in the instrumentat;ion sectiono All of the supplementary circuitry and instru12

BI-DIRECTIONAL LEDEX MOTOR WAFFER SWITCH DRIVE DRIVE RI AMPLIFIER AMPLIFIER INCREASE DECREASE R2 SENSITIVITY SENSITIVITY 1 I I I R3 - R4- 5 VOLT DI D2 O VOLT SENSING SENSIN R5 _____ AMPLIFIE'' AMPLIFIE ONIZATION DENSATRON R6DG U CURRGAGE AMPLIFIER CURRENE A VOLTAGE INPUT RANGE SELECTOR CIRCUIT FUNCTIONAL DIAGRAM Fig. 8

mentation needed to mold the payload into a completely independant measurement probe are located here, A block diagram of the system is shown in Fig. 9. The payload operates from a nominal 28-v power source at approximately 400 ma load. All internal instrumentation is controlled by Ledex stepping switches through the pull away system, while the Densatron and other critical circuits are monitored in the same mannero A PAM/FM/FM telemeter system provides real time data transmission with 0.5 w RF power at 231MHz. In-flight calibration is included for all except digitized subcarrier channels, and an electronics timer tests for Densatron amplifier drift after the data portion of the flight, The main data channel is also connected to the DOVAP telemeter input thus providing a back-up telemeter for the experiment. Considerable effort was directed towards the mechanical as well as electrical development of this section in order to keep a close communication of basic objectives and design results, The circuits or components are mounted on a series of 5-3/8 in, diameter cylindrical shaped discs of variable thickness that are placed one on top the other forming a column similar to Fig, 10 and Figg 1lo The column is, in turn, fastened to the payload using four No. 10-52 threaded rods that pass through each deck and screw into a mounting platform, During final assembly, nuts are placed on the threaded rod and made to bear against the top deck causing the entire column to be held slightly in compression. The resulting structure offers a high degree of protection from shock and vibration as well as providing accessibility to any part of the system. Each individual deck is pressure foamed with Ecco-foam* to further insure the protection of components. All of the decks are electrically joined to the system using Cannon D series connectors, This modular design, so described, results in an extremely versatile payload that is easily tested and maintained, The electrical characteristics can be conveniently modified to satisfy varying demands of the system. For example, the data capacity of the telemeter system may be more than doubled without changing either the physical size or appearance of the section. The individual decks found in the instrumentation section are described in more detail below, and the column deck arrangements are shown in Fig. 12. 2 5l Control Deck Two Ledex programmers, mounted in the control deck Fig. 14, are remotely controlled through the umbilical pull-away system to energize internal instrumentation and select the proper functions to be monitored. The energizing *EEmerson and Cummings, Inc. 14

RAM DENSATRON LEDEX PULL QUADRALOOP GAGE ELECTRONICS PROGRAMME AWAY ANTENNAS CALIBRATION TIMER HOUSEKEEPING DOVAP PHASING OFF-200 SEC I | CIRCUITS 1 TELEMETER HARNESS ON -0.5 SEC AMBIENT DENSATRON DATAMETRICS EVECTOR VECTORR AMBIENT DENSATRON I I II TYPE 989 | TYPE TS-54 TYPE TA-58 | TYPE TP 501 GAGE ELECTRONICS COMMUTATOR SCO MIXER TRANSMITTER 2. RPS 70 KC AMPLIFIER 231.4 MC CALIBRATION ^~~~f;~~~~ VOLTAGES VECTOR MAGNETIC -- MAGNETOMETE _ COMMUTATOR TYPE TS-54 SENSORS ELECTRONICS 2.5 RPS - SCO 1.~.so,,. |, -- __._c,.,,o,,, __s_,10.5 KC ADCOLE ADCOLE VECTOR MODEL 135 1 1 MODEL 235 1 TYPE TS-54 ASPECT EYE SHIFT SCO ASPECT EYE REGISTER 22 KCA INSTRUMENTATION BLOCK DIAGRAM Fig. 9

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SHIFT REGISTER l-J"^^~~~~~~~ _. /~ TRANSMITTERVC0 DECK ASSEMBLY -~.NIE'T -F MI...-,,CHONTROL DECK LASCOMMUTATOR DECK BATTERY DECK ENGINEER G RUPERT DRAFTSMAN E.B. RHODEHAMEL SPACE PHYSICS RESEARCH LABORATORY PITOT STATIC PROBE TRANSMITTER DECK ASSEMBLY DEPARTMENT OF ELECTRICAL ENGINEERING INSTRUMENTATION ASSEMBLY UNIVERSITY OF MICHIGAN 7-12-66 ANN ARBOR, MICHIGAN B- E 434 LAST USED R C D L Fig. 12

I IMPACT II IDENSATRON MONITOR GAGE - I OUT -a CONTROL (AMBIENT |,L IDIl _ LEOEX CONTROL GAGE 2' —08 RANGE SWITCH IK POSITION WIND GAGE E I \ I F USED): -__A I^ PACK |I LEDEX YSI K THERMISTOR 44008 Fig. 13A DENSATRONSWITCH POSITION CALIBRATE O REG & TIMER GAGE EXT ",, ^. ePOWER ALL EXT TELEMETER POWER EXTERNAL TLET -- < ^^ = POWER THROUGH a ASPECT - INTERNAL ULL AWAY I —I ^POWER FLY BATTERY ^^- CHARGE PACK D1 MOTO_____iRJ I LEDEX CONTROL I IMOTOR AND MONITOR I CIRCUITS ARE COMMDN --- TO POWER BUS IN....-. RANGE POSITIONS SHOWN IK IK IK IK IK IK I Fig. 13B 19

rM)~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~y' 0 Fig. 14.jControl deck-front view'I~~~~~~~~~~~~~~~ I O~....... i ~i~i:iilli:::ic~:::lil'ii~;::s~i- li:i:ci:i:::::-:i::::i:::i:~::ili!~i:ii'iii ~ ~::::: ii~~~~~~~~~~~iiiiiiiii~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~~~ ~~~~~~~~~~~~:~- ~~iill~:~~;- ~;I- i~!i~ ~ ~ ~ ~ ~ ~ ~~~~~~~~~~~: i~!i-:::::::::i-::-i::i:i:i:::i:iiiii: i.:i~ii:i;iii~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~~~~~~~~~~~~~~~~~~~~~~~::i:::::::::~~iii:~;iiii::::~:::::j ~:::i:::: ~:i~ij:~ii:::i~ ~i:i'ii~i!!iii l;;i:':':~:':::::::'::::'-:::::::-::i'iiiiii::::::::::'':::i::ii:aii4ii~:::::::::::::: ii~~~~~~~~~~~~~~~~~~~~~~~~~ii~~~~~~~~~~~~~~~~~~~~~~~~~i!!i~ ~ ~ ~ ~ ~ ~ ~ ~~ ~~~~~:-:1:iiii-i'ii:i1;:j:::::iil~:r~:i: ~:ii::iii~~::~:-::i::,:: i!11!~iiii 1. on ro d ck -fil vew

Ledex, commonly referred to as the telemeter Ledex, uses a specially designed six-position wafer switch to sequentially turn on the payload with provisions for external power operation and battery charging. The remaining Ledex, referred to as the gage Ledex, uses two standard three-pole, four-position wafer switches to select leads from any one Densatron for ground monitoring through the pull away systemo A separate position on each Ledex is used to disconnect all payload voltages from the umbilical during flight, The functional aspects of the control circuitry is better understood from the block diagrams of Fig. 13A and Fig. 13B, also Tables I and II. Each Ledex has a single control lead for switching the Ledex and also determining switch position, Blocking doides (D! and D2 in Fig. 13A) are so arranged that a negative voltage pulse will actuate the Ledex motor, while a positive constant current source passing through a resistor network on a wafer switch, yields a voltage which then defines the switch position, For example, in position one of the gage Ledex, the 2 ma current source passes through four 1K resistors in series to ground, thus causing a voltage drop of 8 v. In position two, only three resistors are in series which then indicates 6 v. TABLE I GAGE LEDEX PROGRAM FORMAT Switch Switch Position Designation Function Position gnionFunctionMonitor Voltage 1 Gage 1 Control and monitor 8 ram gage 2 Gage 2 Control and monitor 6 ambient gage 3 Gage 3 Control and monitor 4 wind gage (if used) 4 Fly Removes all internal 2 voltages from pull awayo Measures transmitter mounting plate temperature 21

TABLE II TELEMETER LEDEX PROGRAM FORMAT Position Switch.Swih Designation Function Monitor Monitor Position Vo _____________________Voltage 1 Off All circuits dis- 12 connected from pull away, 2 Gage external Applies external: External 10 power to all gages, Voltage at calibration regular, payload and calibration timer. 5 All external Applies external External 8 power to all remain- voltage at ing circuits. payload 4 All internal All circuits switched Battery 6 to internal battery voltage power. 5 Fly All voltages removed 4 from pull away. All circuits energized from battery. 6 Charge Connects battery to Battery 2 pull away for charg- voltage ingo Additional space in the control deck is used for the calibration timer and the calibrate voltage regulator. A test point jack for the 5-v calibration voltage and its adjustment pot is easily accessible at the so-called "rear" of the deck, The control deck further serves as a center for all integration wiring. The Densatrons connect directly to the control deck via a 57-pin cannon connector at "rear" of the deck. The front side contains another 37-pin plug which connects to the main telemeter section wiring harness, and a 15-pin plug which goes directly to the umbilical plug. 22

The pull away system cable assignments are listed in Table III below, TABLE III PULL AWAY SYSTEM CABLE ASSIGNMENTS Connector Connector Pin Function Pin Function Number Number 1 Signal ground 7 Densatron range control 2 Power ground 8 Calibrate timer control 3 Telemeter Ledex 9 External power input control and range 4 Gage Ledex control 10 Power voltage monitor and range 5 Densatron output 11 Amplifier temperature monitor monitor 6 Densatron range 12 Gage temperature monitor monitor 25,32 Battery Supply Module All of the instrumentation in the experiment is powered by a single battery pack specifically designed to fit in the base of the instrumentation section, The pack consists of 19 Yardney HR-1 Silvercels connected in series to yield a nominal supply voltage of 28,5 v. The module will supply a 1 amp load for greater than 1 hr and can be recharged at a 100-ma rate. The cells are arranged on a 5-3/8 in. diameter deck, 3-3/8ino, high with a 2-5/8-in, diameter hole in the centero (The transmitter is fitted into the center hole during assembly ) Construction of the pack is started by placing ventilation ports, made of heat shrinkable tubing over the vent caps in the cellso The cells are then arranged on an aluminum deck, placed in a mold, and scotchcast* in a tubular *Minnesota Mining & Manufacturing Co, 25

shape to a height just above the cell terminals. A cannon plug bracket is then bonded to one of the forward deck supports. After wiring the battery leads to a cannon DEM-9S connector (Fig. 16) the whole assembly is returned to the mold where it is Ecco-foamed to its final configuration shown in Fig. 15. 2.33 Shift Register The Adcole shift register "conditions," data received from the solar aspect eye. Information received in parallel form is "stored," and "read-out" in serial form by the shift register. Whenever used (on all daylight firings) the shift register is located at the top end of the instrumentation column as shown in Fig. 10. The unit is first mounted on a 1/8-in. thick aluminum deck which then serves as the top plate of the column. 2.34 Magnetometer Deck The magnetometer deck assembly, pictured in Fig. 17A, uses two Schoenstedt Engineering Model RAM-5C magnetometers. The sensors are mounted, one along the rocket axis, the other perpendicular to the first in a horizontal plane. The deck is constructed completely with the nonmagnetic material as are other surrounding structures to guard against field distortion and magnetic attenuation. Electronically, the units operate from a nominal 28 v source at 11 ma current, are internally regulated, and provide an output linearly related to the magnitude of the magnetic input according to the following relation: E = 2.40 +.004 H cos 0 where E is the output in volts, H is the magnetic field in millioersteds, and I is the angle between the magnetic field vector, and the sensor positive magnetic axis. The measurement range is +600 millioersteds resulting in a normal output swing from 0 to 4.8 v which is compatible with telemeter requirements. Typical calibration data for this magnetometer type are found in the Appendix. The deck wiring is shown in Fig. 18. The deck is used for all nighttime launchings, and is physically located at the top of the instrumentation column in place of the shift register (Fig. 11). For daytime launchings, a single axis magnetometer is mounted directly on the shift register as shown in Fig. 17B. 24

Z:::::::i~::iii:::::::::::il:iii:::::i:~~i I:'il:::::: ~ t:~ ~i l~i: li:: i i~ii::::::::::j:-:::-::ii:~~i'i.::il':,iBii;:iei:i-iiiiiji::-::i:::::::::::::i::i::I:::-:i::::~i'i:c:::Fig. 15. Battery module -front view.:::::-i::~::ii~iiiiib~i?.:-:;ii~~i.iii-i::::::iic~l'i~9-i~:ia'i_:ii4''::`i:~isic:: ~ii~liiii::i~i.::il:

OND. 2 3 19 HR I I 4 YARDNEY 1 3 SILVERCELLS I I 3 J.- _ 28PWR CANNON DEM-98 BATTERY SUPPLY MODULE WIRING Fig. 16 26

F-JF-J O-q ciC+ PI (D 0 0 C+ FJ

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S LU SCHOENSTEOT YTEL__ TYPE RAM-SC C vlo MAGNETOMETER O Re | A | C | D | | CANNON DEM-9P lOOK 2 l lOOK 00K " 0 —______'"1 ______ J _____ __ ___~ | |VERTICAL 3-r-_M_ —ll___ SENSOR S \10 6 -7I I IOOK __~___y.. __T~ ~HORIZONTAL |A |8| C |o|DF| SENSOR BLU SCHOENSTEOT T EL TYPE RAM-SC E vlo MAGNETOMETER F ore MAGNETOMETER DECK WIRING Fig. 18

2.35 Special Circuits Other circuitry, associated with the instrumentation section, were developed to satisfy specific requirements of the experiment. The operation of these circuits are discussed below. 2.351 Thermistor Supply The temperature measurements needed in the experiment were greatly simplified by the availability of thermistors with 1* resistance/temperature characteristics, Thermistors have a large enough resistance change with temperature, that a voltage divider circuit with the thermistor in one leg, offers sufficient signal conditioning for telemeter, The voltage/temperature curve of Fig. 19 results from the circuit of Fig. 20. Most of the temperature measurements performed in the payload are for near ambient conditions, therefore a YSI-type 44008 thermistor is used most often which offers adequate resolution up to 60~C. Higher temperatures, up to 150~C, are measured using the same basic circuit and a thermistor probe selected to fit the desired temperature range. Most of the nose cone heating measurements graphed in the Appendix were performed in this manner. The appropriate thermistor probe required for certain temperature ranges are listed in the table below. Nominal Minimum YSI* Ro at 25~C Temperature Sensitivity, Part No>______________. ____Range, ~C mv/0C 44008 30K 10-60 50 44011 100K 45-90 42 44014 300K 60-120 35 44015 1 meg 85-150 30 *Yellow Springs Instrumento The maximum temperature error expected due to the thermistor along is less than 0.5~C within the nominal temperature ranges specified. Telemeter may in the worst case add another 0.5~C error thus making a steady state RMS temperature error of about 0O7~C. 2 352 Calibration Regulator Calibration voltages for telemeter are derived from the voltage regulator circuit of Fig. 21. This circuit is temperature compensated to within 0.05o 50

Ra500K\ THERMISTOR. \R, L=0O CIRCUIT CHARACTERISTICS \~\ Es = 8VOLTS Rs = 36.5 K \"^ ~~ \\ ~ YSI 30K THERMISTOR - RLIOOK \\ -40 -20 0 20 40 60 80 100 120 140 160 TEMR C~ Fig. 19 31

SC.9 K ~~~~OUTPUT 36.5 K Pos. — ^A —-tO TELEMETER REGULATED OR METER READOUT a*VOLTS YSI I + PART NO. 44008. SO~~~~3K RESISTANCE AT 2S~C NEG. —0 - THERMISTOR SUPPLY Fig. 20 52

8.012 aR IK 3V INPUT OUTPUT R 332 2t~~~^5V- 31 I <V DI*|ADJUST 2V TEMPERATURE 2 | COM PENSATION O IV I NT K RIO 332 -A 4V ) OUTPUT ^ 01 ^ -------' f _)>E E R9, 332 0 2v TEMPERATURE D2 NETWORK RIO 332 CALIBRATION REGULATOR Fig. 21 55

from -20~C to +60~C, Supply voltage regulation is better than 0.01% and expected load changes effect the output less than 0.01%. Five calibration points are obtained with a voltage divider using 1-v steps. Overall accuracy of these voltages are within +3 mv under all conditions of load, temperature, and input supply variations. This circuit also provides an 8-v source for the thermistor circuits. The design of the circuit follows standard patterns except for the temperature compensation method employedo The base of Q3 is held at a voltage equal to the reference element D2 plus the VBE of Q3o The voltage divider, consisting of R4, the output level adjust and the temperature compensation network, then determines the output voltageo Q1 is the control element, which drops the supply voltage to the required output while QX is added to provide more loop gain to the amplifier, needed to obtain high stability. D1 is a preregulator for the amplifier, which compensates for large input voltage swings. While it is true that the circuit can be made temperature stable by biasing the reference amplifier (Q3 and D2) properly for zero drift, the characteristic may change considerably with circuit component changes because the point of zero drift is highly dependent upon maintaining precise bias conditions. For this reason, the reference amplifier is heavily biased so that changes in other components or conditions will not alter the bias conditions appreciably, which could lead to a change in the temperature characteristics, The compensation network is easily determined by substituting a decade resistance box in place of compensation network and finding a curve of resistances versus temperature needed to keep the output constant. A curve obtained in this manner is shown in Fig. 22. The derivative of this curve taken at two or three points will yield a total resistance change or AR required for compensation, The compensation network can thus be selected by comparison of this data with the computed AR for various resistor-sensitor combinations. The chosen network can then be inserted into the circuit and the curve of temperature versus output obtained. The step may be repeated if better compensation is desired; however, other factors limit the degree of compensation possible with this methodo The temperature response of a typical regulator is shown in Figo 235 An important advantage to this method is that the shape of the curve remains essentailly the same over large periods of time. 2o353 Calibrate Timer The Densatron amplifiers are calibrated periodically before and during flight primarily as a precaution against unforeseen or unnatural amplifier behavior. The amplifiers have very stable drift characteristics and normally require only a small correction factor for data reduction or none at all. A free running unijunction transistor relaxation oscillator (Fig. 24) performs the basic timing functiono The pulsed oscillator output triggers 34

VOLTAGE REFERENCE CIRCUIT COMPENSATION NETWORK CURVE 2040 2030'.t 2020 ~ I.. wf M fR T 2010 2 u. ~- / 1934 -22C13 2000 * Z 1965 -4 19408/ 1980 7 1990 1990 16 -1990 0 10 20 30 40 50 60 /~Fig.2000 24 351980 - 2013 36 2022 45 1970 " ~ / 2036 60 1960 1950./ 1940 - ] -20 -10 0 10 20 30 40 50 60 TEMPERATURE - DEGREES CENTIGRADE Fig. 22 55

+.002 VOLTAGE REGULATOR +.002 _ TEMPERATURE DRIFT h+.001 0 I8.0000 8.0000 |- - i -—'l-' -30 -20 -10 0 10G 30 40 50 70 TEMPERATURE - DEGREES CENTIGRADE -.001 [T 0 -21-C 7.9971 - 6 7.9987 7 7.9995 16 7.9998 -.002. / 25 8.0000 34 8.0002 45 8.0002 60 7.9998 -.003 Fig. 23 36

25-531 VOLTS 620 - R2 RELAYI 39K 22V J820 470 -5VUr E <R6 4.7K 1.2K EXTERNAL 12219SA CONTROL INPUT lO 180 R5 6 N45 CALIBRATE TIMER Fig. 24 357

a transistor switch (Q2) that in turn drives a 1/2 w relay which then activates the calibration circuits in each Densatron. Under normal operation, the off-time is adjusted for 200 sec. An external control lead can reset or start the timing function which in effect sets the first calibrate time after launch, thus insuring uninterrupted data recovery. The on-time is made as short as is necessary, usually about 0.5 sec. The off-time is controlled primarily by the charging of Cl through R3 and by the intrinsic stand-off ratio (h) of the unijunction.. The instantaneous voltage (Vc) on Cl is: = V /1- C s when the effects of R4, R5, and R6 are ignored. The unijunction will fire at its peak point voltage (Vp) which is approximately VP = (2) where n is the intrinsic stand-off ratio and V is supply voltage across Dlo The time required to achieve this level is found by substituting Eq. (1) into Eq. (2) which then yields: t/ ~ =v" - 3C1 (5) V = V 1-c which can also be written 1 R3C (4) (l-n) = c then solving for t, the off-time: Thff n The value of In (1/1 - T) will usually lie between 0.7 and 1.4 which means the time constant R must nearly equal the time required. The peak point current quite often limits the value of R3 to 1 meg or less so that capacitence required must be 200 SiF or larger to obtain the desired off time. It should be noted here that only high-quality, low-leakage capacitors can be used to obtain long off times.

The on-time is controlled by the discharge circuit of Cl which includes R4 and R5 and the unijunction characteristics. R4 may be used to adjust this time which, in practice, usually cannot be made longer than 0.7 sec due to other factors involved. The transistor switch is held in cut-off, during the off-time with a diode, D2, that keeps a slight reverse bias on the base of Q2. The diode D3 provides a current path for the relay current when the transistor is turned off thus avoiding high-voltage transients. By proper selection of R5 and R6 the off-time can be made stable within +2% over the temperature range of -20~C to +80~C. The on-time will vary considerably more by temperature changes than the off-time but is less critical in this application. 2.3554 Pedestal Inverter Supply Since there are no negative voltages available in the instrumentation section, the inverter circuit shown in Fig. 25 is used to provide the negative pedestal voltage needed in the IRIG commutator format. For simplicity, the circuit is represented in three parts: the first part is a Shockley diode relaxation oscillator composed of Rl, D1, Cl, and R2; the second part is a voltage doubler circuit with D2, D3, C2, and C3; while the third part is a voltage regulator composed of D4. Component selection is relatively uncritical except in the oscillator section where the circuit could "lock-up" due to the use of wrong combinations of components. It is best to use a high holding-current Shockley in this application so that the selection of other components are less critical. The oscillator frequency lies between 300 and 400 cps. Typical circuit performance curves are presented in Fig. 26 and Fig. 27. 2.4 TELEMETER SYSTEM Telemeter system design is the selection of many interrelated parameters in a manner devised solely to satisfy the demands or needs of a particular experiment or measurement. Each system is therefore unique in character and make-up. Pitot-Static Probe telemetry are tailored to meet the experiment's requirements, within the scope of primary payload objectives, while at the same time maintaining compatibility with existing telemetry equipment. For the most part, IRIG standards are used throughout the design except where modifications are more desirable and data quality are not effected. Flexibility is maintained wherever possible, so that each payload may be used to gather the most useful information —both engineering and scientific. System compon39

OSCILLATOR VOLTAGE DOUBLER REGULATOR 33K 4E20M Io |f F RI2 I AA I I!! J IOOK INPUT I if lr ) lOOK INPUT.Z L 10f IOK 4, 25-31 VOLTS CI RI g P? 1 PEDESTAL INVERTER SUPPLY Fig. 25

-.6T -1.1.93 -IA. 1.5 -2.12 2.2 - 2.22 0 -2.0 o0 ~~~t~ 1.*+~ *~ l3.3 -2.28 0 > 4.7 -2.32 - -2.0 - 6. - 2.34 ^ \ 10 1-2.36 0 ) 2-2.3 0 0 -2.4 - -.$....,,.... I,,,. IK 2K 3K 4K 5K 6K 7K 6K 6K IOK IlK 12K LOAD RESISTANCE OHMS PEDESTAL SUPPLY CIRCUIT LOAD CHARACTERISTICS Fig. 26 *40 T ~ V. I C AEo ~O.~ *,o30+~ ~;-20 +13 0 S30 ~~~~~~> +~4*25 0 +20 +65 -II + - — 28 * go -1 z W 0 Z -20 -10 0 +10 +20 +s0 *50 6o0 +70 + 0 +O 0 +100 4 - 10 I-TEMPERATURE C' *o -30o 4 -40- - PEDESTAL SUPPLY CIRCUIT TEMPERATURE CHARACTERISTICS Fig. 27 41

ents are modularized in a manner similar to other instrumentation components to simplify payload check-out and testing. Incoming data to the telemeter is nominally in the form of a variable dc voltage of 0 to +5 v magnitude. Slowly changing signals, such as engineering data, are multiplexed by a mechanically driven commutator resulting in a PAM pulse train that is, in turn, FM modulated in a SCO (Subcarrier Oscillator). The SCO outputs are paralleled forming a complex signal that is then passed through a mixer amplifier to the transmitter where frequency modulation is again performed. The expression PAM/FM/FM is thus defined by the data modulation conditioning processes prior to RF transmission. A floating power ground system is used in union with a separate signal ground common to the chassis, thus reducing data degradation due to ground loops and cross-talk noise. Other major considerations in the design, covered in more detail below, include: The RF link; data modulation processes and formats; and the system components used to accomplish the first two. 2.41 RF Link The design of the transmitting system for the experiment is restricted to a large extent by the size of the vehicle and by the available power source. The weight of a transmitter including power source varies in direct proportion to the RF power output, which means a 3 db RF power increase will double the weight required. Added weight to the payload will detract from the vehicle's altitude capability, which in the case of the Pitot-Static Probe configuration amounts to a loss of approximately 1 km/lb. The transmitter must be capable, however, of generating sufficient radio frequency power to allow reliable ground reception of the data. With these facts in mind, the minimum RF power requirements for the experiment are examined in detail below. The design of the RF link for the experiment is dependent upon many nebulous factors, some uncontrollable, that add to the complexity of the problem. An example being that ground station parameters are not usually known since the experiment will be launched in a variety of locations using existing facilities. A practical or beginning approach to the problem is to assume all "worst case" conditions that will take into account the unknowns and then include a reasonable margin of safety in the design. The first step in the analysis is to establish a relation between bthe transmitted power and the signal power needed at the receiver. This can be done 42

somewhat intuitively by summing all of the factors that influence the signal, It is convenient to think of the radiation as occurring between isotropic antennas, and referencing all terms in decibels (db) in order to simplify calculations, The following relation is established by tracing the signal route from transmitter to receiver: PT + + TL + G = S/Ndb Nt (6) where: PT is transmitted power required; GT is the transmitting antenna gain over an isotropic reference; TL equals the free space transmission loss; GR refers to receiving antenna gain over an isotropic reference; S/Ndb is the signal to noise ratio desired at input to the receiver; and Nt is the total noise power referred to receiver inputO The transmission loss is determined using the formula for free space 9,10 transmission, which is written: r/Pt = ArAt/d2 (7) where P and Pt are, respectively, the received and transmitted powers; A and At are, respectively, the effective areas of the receiver and transmitter antennas, which for an isotropic antenna is 2/4.; d is the length of transmission and \ is the radio wave length, The relation may be rewritten for isotropic antennas as: (X )2/d2 (8) which after conversion to decibel form is written: loss = lO lO /Pt 20 log -/ A log -20 4 -20 log d (9) Assuming a transmission distance of 150 miles (241 km) at a frequency of 231,4 me (a = 1.3 m,) the free space attenuation is approximately -130 db. 45

The noise power present at the input of a receiver is essentially Johnson noise and is found by the relation: N = KTB (10) where N is the noise power in watts; k is Boltzman's constant; T is the noise temperature, ~K; and B is the receiver bandwidth, cps. The total system noise must include the effect of noise generated in the receiver itself,-'the amount of which is a function of the receivers "noise figure." The noise figure is defined as the S/N ratio at the input of the receiver, divided by the S/N ratio at the output and expressed in decibels. The total system noise referred to the input of the receiver is then: Nt = 10 log1 N + NF (11) which is approximately Nt, =1 7 dbw + NF at a temperature of 25~C using a receiver bandwidth of 500kc. Referring now to Eqo (6), and using the following assumptions: Input power to receiver: /Ndb = db Receiver noise figure: NF = 10 db (5-db is reasonable) Transmitter Antenna Gain: GT = 0 db (isotropic) Receiver Antenna Gain: GR = 8 db (using six turn helicals). The required transmitter power is calculated as Pt = -12 dbw or 60 mw. This figure then represents the minimum RF power required for reliable data reception under the conditions statedo The conclusion of the design then rests with the selection of a reasonable safety factor. Again, many parameters become involved that are not and probably cannot be determined. Additional noise may enter the picture, antenna response fluctuations may occur due to the spinning vehicle, the system may experience some deterioration due to battery-voltage decline or other unknown factors, Earlier Pitot-Static flights used a nominal 2-w transmitter which certainly proved more than adequate. Subsequent flights have used a 500 mw transmitter which has proven to be a near optimum choice. Decreasing the size further will not effect the payload size or weight nor will it decrease the cost noticeably. 44

It should be noted here also that higher transmitter powers than are actually needed during flight are beneficial during ground check-out of the payload where "line of sight" transmission between the rocket and ground station may not exist —a condition certainly expected at some remote launching sites. An examination of past Pitot-Static flights indicates that RF signal reception is good except during the Latter phase of the flights where drop outs do occur, usually due to tumblingo 2, 42 Data Formats Data inputs to telemeter are either multiplexed (time shared) or sent directly to subcarrier oscillators, From the standpoint of maximum information content in minimum bandwidth, PAM is considered the best modulation method to useo10 Multiplexing is possible, and therefore desirable whenever interruption of the data does not destroy its content. Direct channels are used for continuous or analog data inputs, which for the most part include fast changing or unpredictable datao A block diagram of the Pitot-Static Probe data distribution through telemeter are shown in Fig, 28 and described belowio PAM data formats are found in Table IV and V. Normal system requirements are for two or three SCO's to be used with a single two-channel commutatoro The IRIG section of the commutator (J2) supplies twenty-eight data channels of which five are used for calibration voltages. The frame sync pulse carries the 5-v calibration. Seven channels are assigned to each Densatron output and the remaining nine channels carry engineering data. The PAM commutator output is sent to the DOVAP telemeter, in addition to modulating a SCO in the regular telemeter system; thus providing two separate RF links for a major share of the experiment's data. The remaining commutator section (Jl) has 30 data channels. Two are used for 0- and 5-v calibration voltages; the rest are divided between two magnetometers. The output usually modulates a channel 10 (l0o5 kc) SC0. The remaining SCO channel is used to transmit solar aspect information, The shift register output is a digital representation of the solar aspect data that also includes earth cell information. A command eye, associated with the optics, "tells" the shift register to "store" the instant it "sees" the sun, and to "read-out" a short time later. The earth cell output shifts the dce level of the pulsed sun information during the time the earth telescope views the earth~ The resultant aspect data, being unpredictable in time, requires exclusive use of a SCO-usually an IRIG channel A (22 kc)> 45

VERTICAL MAGNETOMETERHORIZONTAL 10. KC MAGNETOMETER COMMAAM O-5V DC PAM/FM OURALOO COMlUTA'.A O:8V DC $CO TOR Q ANTENNAS CALIBRATION T-84 VOLTS /4/ - PAM M/ /FM RAM GAGE DATA v AMBIENT GOA GE -2 -O C DATA COMMUTA PAM O-5V C PAM/FM TEMPERATURE DATA TS-54 CALIBRATION VOLTS 22 KC SOLAR 0-5V DC FM IXER TRANSMITTER ASPECT SCO -AMPLIFIER SYSTEM TA- 58 TRPT- 501 TS- 54 TELEMETRY SYSTEM DATA FLOW DIAGRAM Fig. 28

TABLE IV PAM CHANNEL ASSIGNMENT Channel Number 1 Ram gage output 2 Ambient gage output 3 Ram gage amplifier range monitor 4 Ambient gage amplifier range monitor 5 Nose cone temperature 6 Ram gage out 7 Ambient gage out 8 Ram gage electronics temperature 9 Ambient gage electronics temperature 10 Ram gage output 11 Ambient gage output 12 Ram gage temperature 13 Ambient gage temperature 14 Ram gage output 15 Ambient gage output 16 Nose cone temperature 17 Battery voltage monitor 18 Ram gage output 19 O-v calibration 20 Ambient gage output 21 1-v calibration 22 Ram gage output 23 2-v calibration 24 Ambient gage output 25 3-v calibration 26 Ram gage output 27 4-v calibration 28 Ambient gage output 29 5-v calibration 50 5-v calibration 47!

TABLE V MAGNETOMETER CHANNEL ASSIGNMENTS Channel Function Number 1 O-v calibration 2 5-v calibration 3 - 16 Vertical sensor magnetometer 17 - 30 Horizontal sensor magnetometer 48

A representation of the data is displayed in Fig. 29, which is a photographed portion of a real time paper recording for NASA 14.251o Time codes and the ground receiver AGO characteristic curve showing variations in received signal strength are included with subcarrier discriminator output data. Basically, the data formats are as shown in the figure; however, modifications are periodically implemented to optimize the data recovery scheme consistant with updating the instrument. 2o43 System Component Description 2o431 Transmitter Deck A Vector Manufacturing Model TRPT-501 transmitter, shown fastened to its mounting plate in Figo 30, was chosen for telemeter because of its small size, stability, rugged construction, and reliability. It may be operated from an unregulated power source and is relatively uneffected by environmental changeso The transmitter will drive a 50-ohm antenna with a nominal 0.5 w of RF power at 10% efficiency. Complete specifications are listed in the Appendix. The transmitter deck is assembled in the system at the base of the instrumentation column (Fig. 12) where the transmitter actually protrudes up into the center of the battery pack (Fig. 31). Except for the RF output cable, which passes through the foamed section of the battery pack, the lead connections are made at the bottom of deck through a nine-pin series DEM cannon connector (Fig. 32). Additionally, signal and power ground systems are tied to each other, and to the transmitter chassis through the connector, thus establishing the instrumentation's true ground point. 2o432 Commutator Deck A two-section Datametrics Commutator, Model 989 (Figo 33 and Fig. 34), performs the telemeter multiplexing. One section, designated Jl, has 30 MMB (make before break) contracts while the other section (J2) uses a standard IRIG format with 28-data channels of 60% duty cycle. A pedestal voltage (negative 0.7 v) inserted between each data channel provides channel syncronization for automatic decommutationo The frame rate is 2,5 rps, and the unit requires 1 w at 28 v to operate. 2,433 SCO Deck The combination of a Vector Model TA-58 mixer amplifier with Vector ModeL TS-54 SCO's complete the telemeter system. The units are mounted on a deck as shown in Figo 36. The experiment normally used three SCO's; however, the deck can accommodate up to five SCO's in the event future data requirements 49

TIME CODE l 11 i 2 11i kIir i i Ii l Ji i i l l Jii i ii ii ill i ii ii l ii iI i ii i i li i i ii i jiii iii i iiii iii illliii i lj uiu u uun i i iiu!1.,uuu i i, ii iii i RECEIVER AGC NASA 14.251 16HRS. 53MIN. 12.5SEC. OMT VOLTS FLIGHT TIME 72.5 SEC. ALTITUDE 82.1 KM VERTICAL SENSOR MAGNETOME AN EL HORIZONTAL SENSOR 0 VOLTS __ -STATIC TRACE 1 4 VOLTS i ii i ii ii iii i~~~~. iiiii" i ir^ ^^'- ^l ^ VOLTS, *I"..J=......'. " "'*.. " " " " " "" AMBIENT GAGE OUTPUT. 2 VOLTS -I VOLT COMMAUTATED CHANNEL,Y410~~ IJb^~ Yqr0 Oro-un0 VOLTS NEATIVE PEDESTAL VOLTAOE STATIC TRACE,'..,.,,'.,..~R.. TI ME'C'~DE... —-............A -- vol 111 II I I i i*ii' t' "'*'' 1't1111iiill iin1* i ii i i''''''' " II!'** Fig. 29 5o C, 4 r C r rP AiC~~~~~~~~~~~Bli~~~f............................

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RF OUT SEALECTRO 2 NO - 3002 P_.. MODULATION IN_ VECTOR 6 TRPT- 501 TRANSMITTER SI -- 44008 8 +28 PWR CANNON DEM -9P TRANSMITTER DECK WIRING Fig. 32 55

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CANNON DC M- P l 2 | 3 4 I I llt l3I 4 1j I ll2 I3 4 5 3 | Stlt24|5l26l27l2l |2930l31 |321 3 1 33 5 ST 60 6 6 6 6 6 6 6 6 6 6 6 6 o 6 o o 0 0 6 0 0 6 o 6 6 0 0 5a 10 Is 20 25 30 C OUTP OUND CANNON DCM-S7P 26 CASE GROUND VOC.Ol0jIf FILTER,,o,,voc. l COMMUTATOR WIRING - DATAMETRICS TYPE 989 Fig. 34

CONTINENTAL CONNECTOR MM5-22P A 8 — O0 +28 PWR MIXER C OUT AMPLIFIER IN E -- I OUT SIAS OSCILLATO A 0 28 PWR +2t PWR B ~sco T OUT SCO c OUT D IN__ _________ _ SEALECTRO DL i u PART NO. 3102 A I0 + 28 PWR (8 O7I O I A) r- 2 E 2 OUT SCO C OUT 3 IN E 5 6 SCO CC OUT - IN + 2F PW - (OPTIONAL) D 2IN *6P 10 A 0 —— 13 28 PWR B 14 OUT SCO C I-U — >SD- 15 (OPTIONAL) D IN E +28 PWR SCO C OUT D IN SCO DECK WIRING Fig. 55

,~~~~~~~~~~~~~~_~:ii~::i:~iiiii::;(~il/4:(::::i~i~~i:-~~::-~:i::: ~~~~~~~~~~~~~~~~~?''~-~~,~~~-;~ ~ i~~~ —~"~~~:~:-::: a~:::;:.::i:::i~ia:;;:iii3-;ii'~ ~ ~ ~ l~i~sI~~~~~ ~~~~~~~~~~~~~~~~,> ~ ~ ~ ~ ~~~~ ~~~ ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~:i:::::i::::i::'::::::::::: j::~i:j:::::: ~~~~~~~~~~~~~~~~~~~~~~~~~~:i::I-:::::,:::::,i, i:l:,:a:l~~s ~~r:::~-~,-'2~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~i Fig. 56. SCO deck.

increase. The deck wiring is shown in Fig. 35. The TS-54 SCO has an input impedance greater than 500.k a.linearity better than 0.1%, and an open circuit output of 2 v RMS. The mixer amplifier is used for mixing or "summing" the SCO outputs into a single complex signal in addition to increasing the output levels sufficiently for proper transmitter modulation. The TA-58 has an adjustable voltage gain from 0 to 20 and a frequency response within 0.5 db from 20 cps to 100 kc. The gain is adjusted at the lowest required level to minimize distortion. Also located on the SCO deck is the commutator pedestal bias oscillator explained in an earlier section (see Section 2.354). 2.434 Telemetry Antenna The antenna system uses a pair of model 2.005 quadraloop antennas,11 mounted on the skins covering the instrumentation section (Fig. 37A). These antennas are designed specifically for high spin rate vehicles, and with a low profile to keep aerodynamic drag forces low. A phasing harness (Fig. 28) is used to feed the antennas in parallel, 180~ out of phase, which gives good fore and aft coverage. The antenna driving point impedance is selected near 100 ohms; and by using a one-half wavelength feed line in parallel with a one wave length feed line made of 50-ohm transmission cable, the combination will match a 50-ohm transmitter. As might be expected, the impedance is critical with frequency (a typical curve is shown in Fig. 37B) so that transmitters and antennas must be compatible. The radiation pattern, using the right circular component, is that of a toroid whose axis is normal to the vehicle axis. The VSWR is generally 1.5:1 or less in the range of interest. 2.5 GROUND CONTROL SYSTEM Payload development was paralleled by the design of a ground control system that is needed to provide remote payload control and monitor functions. Divided into three sections, the entire system includes the control deck mounted in the instrumentation section and explained in an earlier section; the umbilical system: and a specially designed ground control console. For the most part, design objectives are similar to those of flight instrumentation with reliability and portability being major considerations. The main function of the system is, of course, to activate the payload; however, additional features include provisions for external power operation and battery charging. The finished design utilizes a step-by-step turn on procedure that, in effect, automatically checks system operation without relying on operator decisions and without jeopardizing payload circuits. Other considerations in the design 58

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IMPEDANCE / FREQUENCY FOR MODEL t.OOS3 QUADRALOOP TELEMETRY ANTENNAS Fig. 37B 60 (r ~ ~ ~ ~ ~ 3. ~

are covered below. 2.51 Umbilical System A 12-conductor pull-away system was somewhat arbitrarily selected as a compromise between using a large plug and cable capable of accommodating all of the required control functions or using a smaller cable and plug that are less bulky and suited more for pull-away operations. By using Ledex programmers, the lead availability is in effect multiplied, which allows use of the latter system. Because it is desirable to keep the operation of the main payload and DOVAP instrumentation isolated and independent of one another, a separate 8conductor ground control cable is used for DOVAP. Each ground control cable to the vehicle is wired to a Jones plug that mates to a female Jones plug mounted in the instrumentation section. The Jones plugs are particularly useful for this application because they have low contact resistance and they can be easily pulled apart without jamming. Control of payload functions and DOVAP instrumentation is possible through the pull-away cables until actual launch time when motion of the vehicle initiates a so called "fly away" umbilical release. The system is simple to use and requires no special equipment other than a few weights to anchor the cable to the ground behind the launcher. The principal of operation is explained as follows using Fig. 38 reference. Consider first, that the system is in a static condition just prior to release. At this time, the cable is tied securely to the ground at point A and is fixed at point B. The impending movement of the vehicle imparts a force Fl, at point B which acts in the direction of motion. This force may be resolved into two components, F2 acting along the cable axis, and F3 which is directed up and away from the vehicle. Since the cable cannot support bending, the force F3 must be located at point B. Due to the cables restriction, a force F4 exists which is opposite to F2, whereby the effects of each force are cancelled leaving only the resultant force F3. Upon release, point B on the cable is accelerated in the direction of F3 along an arc described by the line BC which is the desired result. In actual practice, the cable is taped at various intervals along the length of the Apache motor housing thereby creating a point "B" at each tape point, resulting in a whipping action that carries the plug and cable up and away from the rocket and exhaust gasses. All indications from past flights are that the system has worked very 61

/, /F 9 APPROXIMATELY EQUAL TO 45~ UMBILICAL FLY AWAY RELEASE SYSTEM Fig. 58 /,,,0..__...._J ~~~~~~~~ / PRXMTLYEULT 5 UMILIA LWYRLAESSE I~~~~~~~~.3 62~~~

wello Cables and plugs are generally left undamaged to the point of being reusableo 2.52 Control Console A. functional block diagram of thie contro.l console is shown in Figo 39 Two programming Ledexes, simillar to those located on the control deck described in an earlier section, are used to select the proper control or monitor circuits inside the console and direct them through the pull-away cable to the payloado Each control console Ledex is synchronized to a "companion Ledex" mounted on the payload control deck which, as explained earlier, then directs the pullaway leads to proper circuits to be controlled or monitored, For example, when the instrument.ation Ledex is synchronized to the payload telemeter Ledex in the OFF position, there are no circuits connected to the pull-away leads either at the payload or at the control console. In the next position, GAGE EXT2 the power supply output from the control console is connected to the appropriate pull-away leadso Then, on the payload side of the pull-away cable, the telemeter Ledex directs the power supply leads to the gage circuits, Each succeeding Ledex position correspondingly selects the proper circuits or meters as requested by the payload telemeter Ledex, A single push button switch simultaneously switches both the control, console Ledex and its companion Ledex in the payload. To reduce the possibility of a Ledex not switching, each Ledex is driven from a separate 500 MFD capacitive power source, Control console Ledex positions are displayed by indicator lamps driven from switch wafers Dl and D5o Payload Ledex positions are determined using the range resistors mounted on wafers D4 and D6, Voltages derived by constant 2 ma current sources flowing through these resistors and similiar circuits on the payload Ledex switch wafers, are compared in error amplifiers, A different voltage greater then 1 v (indicating the Ledexes are out of step) will activate a relay (SW-6 or SW-8) which in turn causes the push button switch (either PBI or PB2, whichever the case may be) to glow red by turning on L2 or LLo Normally, when the Ledexes are in step, the push buttons glow green by the lamps LI and L3o Should the case arise where the Ledexes are out of step, the payload Ledexes can be disconnected by SW-5 (or SW-7) from the stepper switch PB 1 (or PB 2) thereby allowing the control console Ledex to "catch-up" to the payload Ledexo The payload operational turn-on sequence begins with the telemeter Ledex in the "OFF" position while the gage Ledex is usually left in the "FLY" position, When the telemeter Ledex is switched to "GAGE EXT' position, the Densatrons, the calibration timer, and the calibration regulator circuits can be operated on external powero A manually operated switch (SW-l) connects the 65

TELEMETER AC 110VAC FUNCTIONS LDrIoax I OVAC IrID r0 3 CALIBRATE RELAY, ______X -"I 1.^1 ~-| Dl —----— C31 0 S —- CONTROL ^ OFF,<s- I i..*~." GA 2 I IOPNOF SPL POWER MONITOR L^ZPS ~5> 1 8 sw 2 |2v SUPPLY POSITION eXT. -ON OULAt AND SENSINGO LEADS INDICATORI XT 1 DISCHARGE LAMPS AlU II RM 0: DNSAT:RON FL 5T DOW' RANGE- CONTROL L^-, 151~MT CR I oEETo CHARGE 6 -[1UTIOF] 6~ CUT OFF CHA REOULATED 2 2 V EXTERNAL POWER r- ---------- — 4- BATTERY CHARGE MONITOR WAFDRIVERS INSTRUMNTATION / / D1,D2 3,04 0 - TELEMETER LEDEX |) Pt.1 /- CONTROL / MONITOR DRIVE IERROR, Io, % IK, -- AMP R — 1o lit L _2 IA K I R LED EX Y DK IAI POSITION CONSTANT SW-6 CONSTANT M GE 3 MONITOR CURRENT CURRENT F K I SOURCE J | SOURCE G,i\ -/7 ) H PB-2 Pb-2ER CONTROL MONITOR:P 302 1000/50 1000/50 DRIVE _ WAFERS. o, 06, r - 6.3..ERROR 2 I —-- 1^~1 J r6 |r-AML flER AMPLIFITEPERATUE POSI'TION GAGE 2KANGE ONITOR INDICATOR — GE3 LEOEX 103 CONSTANT SW-9 CONSTANT LAMPS POSITION CURRENT CURRENT GROUND CONTROL C ONSOLERC Fig. 59 * VOLT THERMISTOR RIEGULATOR GAGE/TRANSMITTER TEMPERATURE do AMPLIflER TEMPERATURE OUTPUT MONITOR - -— ^ RANGE MONITOR GROUND CONTROL CONSOLE Fig 539

"external power supply regulator" output to the pull away. Additional power leads, separate from the current carrying leads, are used for "sensing" the voltage at the payload thereby reducing the effect of line voltage drop on the voltage appearing at the payload. Diodes between the sense leads and the output leads at the control console, clamp the maximum available output voltage at a safe level in the event the sensing leads are disconnected. The gage Ledex, operating independently from the telemeter Ledex, can be used to select any one of three Densatrons for monitor or control purposes irrespective of the telemeter Ledex positiono The Densatron output, amplifier range, amplifier temperature, and gage temperature are monitored using meters M2 through My, while a three-position switch, SW-4, is used to select one of three required range control voltageso The calibrate relay control switch (SW-3) is tied in parallel to each Densatron and may be activated in any position including the "FLY" positiono When the telemeter Ledex is switched to "ALL EXTERNAL," the complete payload is operating on external powero The next position is used to switch from external power to internal power —again, all circuits are operating. At this time the internal battery voltage is monitored at the control console, and a RTM (running time meter) is turned on to record the battery discharge time. The next step in the turn-on procedure, puts all circuits in a flight condition and removes all internal payload voltages from the umbilical plugo The "discharge RTM' is still activated. The following position, turns off all of the instrumentation, and puts the payload internal batteries in a position for chargingo The output of a 100 ma constant current source can be manually connected to the batteries by the switch SW-2. When a predetermined voltage level (usually 38 v) is measured by the meter Ml, the meter cut off relay automatically disconnects the charge circuit from the batteries, A separate "charge RTM" records the battery charge time, The control console, pictured in Figo 40, is mechanically built to withstand the rough handling environment often time encountered during shipping and operating the equipment at remote launching siteso The unit is built in two sections which are rack-mounted with shock mounts in the shipping container, By removing the front and back panels of the shipping container, the console is ready for operation by merely plugging in the appropriate cables,

. ~~ ~........................................ _ _.._.... I': N~, F1 U~tS MMhNUTlr, a r k~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~::: j: X _1. 1 1I |||;a,,..E~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~...... Fig. 40. Ground control console-front view. 66

35 MECHANICAL DESIGN The complete nose cone is fabricated in three sections for housing the Densatrons, telemeter instrumentation, and the DOVAP transponder respectively; and hereafter described as the probe section, the instrumentation section, and the DOVAP section. Each section is independently assembled, both electrically and mechanically, before mating to remaining payload sections (see Figs. 41A, 41B, and 41C). The basic structure is described starting at the top or forward end of nose cone, then proceeding along its length until reaching the Apache second stage, The ram pressure orifice and gage chamber are contained in a 3,5-ino diameter hemispherical nose tip (Figo 42A) that fastens directly to a 30-in. long stainless steel tube (Figo 42B), which then connects to a stainless steel center section (Fig. 43) that contains the ambient pressure chamber on the forward side, and either a wind gage chamber, or else no chamber at all on the aft side. (A few later model nose cones were modified to include an additional ambient chamber on the aft sideo) Proceeding further, the center section then mates to a magnesium casting (Fig. 44), that serves as a transition section between the probe section and the larger diameter of the instrumentation sectiono Four Unistrut* columns welded to steel rings (Fig. 45) at each end form the instrumentation structure with a mounting platform included between the columns for the instrumentation deckso A second Unistrut section provides space for the DOVAP transponder and connects to the Apache motor coupling adapter. In the final configuration, a pair of rolled aluminum skins enclose each Unistrut section, The multiple sections of the payload are kept in alignment using dowels and other means explained in detail later, and tight tolerances are maintained between sections to increase strength and rigidity of the structure, The overall length of the payload is 74,2 ino with a nominal weight of 62 lb including DOVAP and all internal instrumentation. The center of gravity is located 37 ino behind the nose tipo A semi-assembled view of the nose cone including a despin module, but without DOVAP, is shown in Fig, 41Bo 3.1 PROBE SECTION ASSEMBLY Figure 47 shows the probe section assembly as well as individual gage as*Unistrut Corporation, Wayne, Michigan. 67

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? —........~~~~~~~~~~~..',~LJ__~~Z~IB o r- -~f-~,o-~p~p~F r........ Is'~~~~~ _Hlr~ J,ri ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~ — ~A —s~,, —--—,,''''DI.. — "~ I__~_.~~J:_ d/!';~"~_.~IF Y~'~_S~-HH --— i — ~~~=D eij~o -:.-~-""L — _... PMPg. ~1C 7o~~~~~~~~~~~~~~~~P

:':::: z:i~i:IzI:~:I::1:... ~~ s::i~ ii::i::;::i: j.... i;;....:;.... i......!:.....:JE:::: i:E::::: i::::::: i::E::i:::::: i::::::::: E:::::::E::::::::::::::E: i:E:E::: i i:: E::::E::::::::::::: i:::::::::::::::i::'::::::::l::iU~:'iiiii:::':i:~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~i... "::::s:::::::::: i::i:i::::::::::::iiEE::::: EE:::::: i::::::::i: ~::::::~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~::~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~-:. i::.::::::....::::::::: E::::~~~~~~~~~':::: E Fig. 42A. Nose tip —front vriew. 71

na ~ Fig. 4-i2B. Nose tube-side view.

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4L. Nose cone transition

0 2~~~~~~~~~~~~~~~ ordl AV*.: 0 X 4 -p 03 LC\ bO Respon ibit at - X —v *? < 0z0;00 00<000550::090 an; i0S00Xt~dA 5 is it Xi~fo aid tS0:z of: X;Sn~l0Z00009009z000004 S> 4 005 0fffffS, 400 i an' Sat; ffi t X0S040 0 t of0X000000000 St......i.....

Fig. 46. DOVAP section with hardware. 76

l ^^^^~h y-RAM GAGE ADAPTER RING HEADER PLATE, — TEFLONI SEAL GAGE BODY ASSEMBLY l A r AME RICIUM 241 RADIOACTIVE SOURCE DENSATRON GAGE MODEL E" iDENSATRON AMPLIFIER 4 FEED THRU POLARIZATION CYLINDER TEFLON SEAL —'0 RiNG GAGE BODY POSITIVE ION5 COLLECTOR~ MdOU4.NTING COLLAR L NR~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~OSE TIP I~-COLLECTOR MOUNT L-'dI RUA R GAGE BODYPROBE SNASSEMBLY H R BACK SEALING PLATE DENSATRON A IMP RTHNIA- COATED IRIDI FILAMENT WIRE RETAINER RING 6 FEED THRU8 DENSATRO GAGE DDENSATRON GAGE UPOEL 0"EADER PLAT.^* -,,,^.,_;- To "...m. U. us.aaaY or M PITOT-STATIC PROBE IAOS PRSE ASSE BLY ____________________INSULATING RN.___INGe*_ -0_. 0 LAD006-184 Fig. 47 MOUNTINGe COLLAR' I I I I IE~i~ i ~~ ~POSITIVE ION COLLECTOR GAGE BODY AM H A W 1 I RING'FEED T'RUS COLLECTOR MOUNT POLARIZ~ATIONI CYLINDDIER -— I I I`F/ AMB)IENT GAGE ADAPTER RING" INSULATING RING -— I I I'd RINGL SEAL RING I AMBDIENT CHAMBER THE Um9rom T'l, M PITOT- STATIC PROBE ANN AR NmR. OSE PW Fig. 47

semblies, Each radioactive ionization gage is an integral part of the Densatron amplifier system, and the two are mounted in the nose cone as a single unit. The probe assembly begins by mounting the appropriate Densatron(s) to the center section (Figo 43)0 The chamber shown has been modified to include two ambient measurements. A further modification is to be made to include a wind chamber (see Section 7ol), Each gage is vacuum sealed from the amplifier section by "0" rings and in turn each chamber gage assembly is VacuumL sealed from the remainder of the payloado The Densatron mated-center-section is next connected to the magnesium structure with the lower Densatron protruding down into the transition sectiono The two parts are secured using screws located externally along the periphery of the magnesium sectiono A close tolerance is needed at this point to insure a rigid junction. Ascrew type coupling is not used here because Densatron electrical cables would be twisted during assemblyo Next, the steel tube is fitted down over the upper Densatron and connects to the center section via a quarter-turn acme threaded coupling. Finally, the ram gage Densatron is dropped into the steel tubing and rests against a lip provided. With an "0' ring attached, the hemispherical nose tip then screws into the tubing, thus completing the probe assemblyo Inside the steel tubing the Densatrons are held in place by nylon supports,s which in turn are supported by steel retainer rings welded to the inside diameter of the steel tubing, Densatron cables pass through the nylon supports and feed directly through the center section to the rear of the magnesium section where they connect to the telemeter section via a 57-pin cannon connector. Before assembly to the payload, each Densatron is wrapped in alternate layers, usually three each, of armalon and aluminum foil which effectively forms a radiation heat shield against the high skin temperatures caused by aerodynamic heating (see Appendix)o 3.2 TELEMETER SECTION ASSEMBLY With the exception of the aspect eye and pull-away plugs, all of the telemeter instrumentation are located on standard cylindrical shaped decks. The decks are assembled into a column such as Fig, 48, and fastened to a steel mounting platform by four 10-32 threaded rods that pass through each decko Additional support for the structure is gained using a bracketed plate (Figo 4b9) that attaches to the top steel ring of the Unistrut section. Two pull-away plugs, one for DOVAP, the other for the main payload, are mounted to a bracket (Figo 45) which in turn screws to brackets attached to two of the Unistrut columnso Located opposite to the pull-away is a mounting plate for the aspect eye which is welded to the remaining two Unistrut columns (Figt 45), 78

~~~~~~~~~~~~~~~~~i~~~~~~~~~~~~~~J:::Phis:::i Fig 48. nstumetaionsecio wih eecroncs-rot vew

* qGeaowq;aodIdns uum-TOo SuT.oqs an;-onJ;s uo-;:4e.ueumxqsuT Jo dOl,'6t * $

The quadraloop telemetry antennas (Fig. 48) are fastened by screws to the aluminum skins surrounding the telemeter section. The completed telemetry section connects directly to the magnesium casting using eight 1/4-ino stainless steel bolts. The upper steel ring of the Unistrut section is snug-fit piloted to the magnesium section, and a keyway is used to maintain alignment and rigidity. 353 DOVAP ASSEMBLY DOVAP instrumentation fits into another Unistrut section, Fig. 46, located to the rear of the payload. The Apache motor coupling adapter is used also for the mounting base for the: DOVAP'can." A hold-down ring (shown in Fig. 46) that fastens to welded brackets in two of the Unistrut columns is all that is needed to hold the can in place as shown in Fig. 52. In the past, the DOVAP antenna coupling networks were fastened to the'can" as shown in Fig. 50 and 51. Brackets for the networks are now provided which are screwed to the Unistrut columns as shown in Fig. 535. Piloting and keying arrangements similar to those employed in the telemeter section are used on both ends of the DOVAP section to help maintain strength and rigidity of the structure. Since DOVAP antennas are- mounted directly on the second-stage Apache, it is desirable to split the payload at this point and consider DOVAP to be a part of the rocket~ Therefore, final assembly of the payload to the rocket involves installing the DOVAP can, joining the two Unistrut sections with eight 1/4-ino bolts, and putting skins around the DOVAP can to finish the operation. The entire sequence is usually accomplished in less than 30 min. 3.4 DESPIN MODULE A yo-yo despin module,8 which will be used in the future payloads for wind measurements (see Section 7.1 and Figs. 54, and 55)^ is designed to fit between the telemeter section and the DOVAP section using the same mounting holes, keys, etco, that are used to mate the two sections. Electrical wiring is independent of the remaining payload, therefore no modification or integration other than mounting are necessary when a despin module is used, 81

coo Fig. 0. DOVAP section mounted on Apache. Fig. 50. DOVAP secti'on mounted on Apache.

N'ASA W-63- 340 Fig. 51. DOVAP transponder installed-side view.

TNASA ~T-63- 338 Fig. 52. DOVAP transponder installed-top view. Fig. 52. DOVAP transponder installed —top view.

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99 ~p~uas~ a:puruda'(;'T.. - - - -. -.:i I.. ". - - i'' -. --.-.::. - "':-i: - 1- 11.....:,,.:..:::....,q:,.1.1 -............~,:;:''!::..:,i....,..,:.,:,. - -. -...':-Ii,.-i:......:... -. —-.. -- -. --:;'.- 1 1,::,.:.:,~~~~~~~~~~.,~~~.",:il.,",",.,..~~~~~~~~~~~~~...:.,~~~~~,:,.. - -~~~~~~~~~~-...' bd -,:.....:i'i:::i:i.ii:x::`:::::-::::.1:!.: -. i:.-..]:.], i~i'..: -.-..':_....._'-....- -!:::... -:..:!:.'.''i'.-!:..: a.-!: l~i -, i.,... -: ii:...'...i:il; i i:: i i.. -.:::..::I:::,:: i"' i,:'::'..1.i....: —'.i —:"',.".'~:*:: ]:i:]':..~i*::.i:.:,::: I::: p::::'~:: -.::: sq.,, "::::,,..-.:::..... -' ";::- I i, -: iI;:':i: -,'il:-i::c i.. I - - -:-::,,-~ii:':,:..ii ~::i.'.:ii]a:..:i~x —."::::,.:l-: Iii: Ill. -:.::: - l...1-1~~... I - 1".. I. -.:.:~~~~~~~~.............-~~~~~~~~~i~::x::-.....i:::::i::i::~::i:; i:: i:- -::::.::-::i,::::::,:::' I::..:....~~~~~~.j~~~::::..i:'-.'::::::::. ~~~~~~~~~~~~~~..................:~::::'::,..l..1li:.... I - -.. -.~:1:::' ~.:':,:'-'',..::....... - I... I ij.,c'I~:-iiiii::::ll..i.i..-:'::,:':i:::',,:'ii i::: ii. - liiil;;i:::i I............. 1. - I'll", -Y. -,.,.; i.:i::'*:::-:-:::":. i:. -...' I........ I I, ll: I,.,.:...~~~~~~~::. ~ ~,~~~~:::',::::.,:::::::....:~~~~~~~~~~~~:~~-:,''.:X...".-.... -.. -........I~~~~~~~~~~~~~::..:: I':I. i:;l',:i:::c.i: ~~~~. ~ ~ -,~~:~~j~~~.-,~~~:::~~~::~~~..1 I....- -..: ~~~~~~~~~~:.:::::'.:::::::w'.X -...... I... ~~~~~'':::: i::::r:........ I - I I I - II:a:::i: II.1 ~ ~ ~ ~ I1,IIII- I:.~~~v".I..I:I.. III......I I. I..... ~~~~':~: k x. i.I..,..;:...~j;:].l,Ii-iI. -ii:.:. p -.:::.,:1.::1:...., I.,..I... I...-iI, II..I.i:::i ~ ~ I...::::::.. ~ ~ ~ I ~~I. I..I ~ ~ ~ I.II. I.. I~~~~ ~~..::..:.::""..:.,:::'.,......~~~~~~~~~~~~I.I...I~~~~...I~~~1,.-...,. ~ ~ ~ ~..I~~~I.:~~.::::::,:::~~~~~-::,..'. I I:: i.. i:. I I::::i,:,~::i::::::: ~~~~..::~~~~::: -,~~~~~,:: ~~~~~~~~:~~~~::': ~~~~~: "-~~~:,.:..I.:. - I,:'.:......:... I..... I:::::i.::::.: - i::ii::ii. i::. i. I:: ~:II..i:::.:::::.., I..:I, I: ~ ~ ~ ~ ~ ~ ~ 1.,,:::;'~~~~~~~~~~~ m~~]~~::~~~ ~~~x,: ~~~:::i::..:~~~~~~.';:.'::~~~~~x::.,.]::.,:,~~~~~~~I ~ ~ I,",. I ~ ~ ~ ~ I.1,II. I.:..:~ ~~~~~.,~~~~.,..,~~~, II ~ ~ ~.I~~I.::,.:::i~.:. I,,.1....... I -. I.,..:::.:.I I I''.'....l.-,,.....:i 5. i::,.::,.:::::::I.~~~~~~~~~~~.:, ~~~~~~.-'..:-...:....."...:: ":..: mi~ ~ ~~ ~~ ~~ ~ ~ ~ ~ ~ ~~~~~~~~~~~~~~~~~~~~~~~c::::::.,.............::.~~~~~~~~~~~~~~~~~~~~~~~~~~~~,.;:::,.::,...:%:X'::.. ].i:~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~]~~~: ~ ~ ~.....,~~~~~~~~~~~~ -','' ~~~~~~~~~~~:.'..:....-.~~~::::.::-~j:..:::::":::."::::.:::::i-' i:..:: 1:I.. =:. —--;,.m.,...::' _ i"::.:..::,.'. I 1.:.,. — — ~~~~~~..::i -. -,'.. II"..,.-..- - -.-' I -.,:.. I.. 1..1.1 1 I I:: -'".:.i: i.-':: I i... I l:i:ii:I::.:,::.:::d: %:..'..::.,:,:, I..II, 1- I.II IIi I I...:"'::.,::..: -::::,,;...... I - " ",'.1 41 1. 11,. 1 1 I.. 1...... ~ - -,.. I 1..,. ~'.. ",':.,.. I ".'..' I ~ ". ~.' ":..::::.::;:::",:::::,::,.:.a. 1.1:....,. v X:.::..:': - -'.",..'.' I...... 1:- -., ~.:: ".. ~ I:::':~::: ~ ~.,::,%.,, I::.~: p::,.,:':':::.,.,:,.:... I - I.,., -.1,"I I...,.1 1. I.. I..,.. 1.',. I,....1. I:': I:1:1:...11... —-'~ 1:......,:',:........1,. -,-11;:-."......,......'......-.I:. -, -,. i. -,~.:::.::

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4o TESTING 4.1 SYSTEM TESTS The objectives of payload testing are (1) to prove design adequacy, and (2) to define operational characteristics and improve the reliability of each payload. The first part is accomplished through simulated methods testing and test launchings. In the case of the Pitot-Static Probe experiment, the basic design was proven by the prototype nose cone launching AA6.340 and since verified by a number of successful launches that followed. The environmental tests, explained in more detail below, performed on NASA 14.19 at Goddard Space Flight Center give further support to the design. In addition, a theoretical study for aeroelastic flight loading of the nose cone was made by Thiokol Chemical Corporation,12 which concluded only moderate loading was to be expected during flight. Nevertheless, static load tests using parameters dictated by the study were conducted on NASA 140285 at Goddard Space Flight Center to determine whether or not structural deficiencies exist. Almost negligible payload distortion was observed during the test. The second objective of determining operational characteristics and reliability, is basically directed towards individual payload testing with emphasis on fabrication and assembly techniques rather than design aspects. For the most part, testing is done at three levels of payload development. First, each circuit is individually performance tested with respect to temperature changes, voltage changes, etc. Then every solder connection is visually inspected before a circuit or connector is considered usable. Next each subassembly, such as the Densatron, or the telemeter system, is throughly checked out before being installed in the nose coneo Before any commercially purchased equipment is used, the units are tested against manufacturer's specificationso The nose cone is periodically inspected during fabrication, and the gage assemblies are vacuum tested for leakso The battery pack is carefully monitored during all charge and discharge operationso The last tests are concerned with the system as a whole and include a vibration test using test levels applicable to the Apache payload (see Section 4,21), Instrumentation tests associated with rocket check out concludes individual payload testingo By the time a particular nose cone is ready to launch, considerable data is available on the performance and history of the components as well as the complete assembly, and a reasonable confidence in its capability to perform the measurement exists o 88

4.2 ENVIRONMENTAL TESTS Formal testing of Pitot-Static Probe payloads originated with NASA 14.19 at Goddard Space Flight Center in the spring of 1962, where vibration and thermovacuum tests were performed and the payload was dynamically balanced, 4o21 Vibration The vibration test consisted of shaking the payload along three mutually perpendicular axes in accordance with specified levels for the Nike-Apache vehicle^ which are; Thrust Axis - 1 g (gravitational constant) sinusoidal survey sweep from 10 cps to 2000 cps followed by a constant maximum velocity of 3 ino/sec. from 10 cps to 145 cps and 7 g from 1l5 cps to 2000 cps. Lateral X axis - same as thrust Lateral Y axis - same as thrust, except no 1 g survey. Anticipated high loading areas of the nose cone, such as the nose tip and the rear of the ambient Densatron gage housing, were instrumented with accelerometers for evaluation, purposeso Results of the tests indicated no structural weakness in the design, and there were no failures of internal instrumentation. Data:'from the test showed a strong resonance with a Q (amplification factor) of approximately 6 in the range between 350 cps and 400 cps along the thrust axis. In the lateral axes, the major structural resonance occurred at approximately 14 cps with a Q of 9o Other minor resonances (Q < 4) appeared in the 150 cps to 200:cps rangeo The same vibration test parameters are used on every Pitot-Static Probe payload as part of a standard check out procedure. Visual and operational examinations after the shake test together with observations during the test determine the payload's launch statuso 4,o22 Thermal Vacuum Test This test examined the Pitot-Static Probe's capability to function in a space environment. The nose cone was placed in a Tenny Thermal Vacuum Chamber where the temperature was stabilized at -15~Co The chamber was then evacuated to a pressure of 10~5 mm Hg and held for 30 min with the payload power turned ono Thermocouples, for temperature measurement, were spaced along the length of the nose cone, and mounted in local heating areas of the instrumentationO 89

A second test at positive 35~C was performed using the same procedure (see the Appendix). Payload operation was normal during both tests, and no large or damaging temperature rises were recorded, therefore inferring the payload may operate safely in a space environment. 4.23 Dynamic Balance Two Pitot-Static nose cones have been examined for dynamic unbalance along the spin axis. The first, NASA 14.19, was tested at GSFC while the second, NASA 14,21 was examined attached to its Apache motor, at the Wallops Island facility (see Figo 56 and Fig. 57)o In both cases, initial unbalance measurements were small or insignificant when compared to the rocket system, as was expected. The payload's basic geometry and symmetrically located internal instrumentation preclude the possibility of a serious unbalance existing, Data taken from fourteen Pitot-Static Probe flights gives an arithmetic mean value of 9,1~ for the included angle of the precession cone. The standard deviation is 5.2~ with a minimum cone of 2o5~ and a maximum cone of 20.3~. 90

NASA W-63-330 Fig. 56. Dynamic balance of nose cone at Wallops Island. 91

:::.: I-::::::.:s" iii i:::':;:::::~: ji'i:;:i:i:':::::.-.:.. :l:'l:ciiiiii:i -::::i:i~'::'i':::::is::i:::: i::::::: -'iiiniiii?~3':... ~:::::::':-i:(:I::::::::::::::::::nli::(si;(i-i-ii::::- iiiliii::::::ji:::::: i:i-iir iiiiib:::i::::-':::" Cii_ ir;::~i:::i:: iii:ii:ii i-Riiiii:iici i'::iBPgiB Tdeisp q i: -~~`cbi?Elsiieur~`l i:i:ii-,i:iiii ii:i:i:ai-ii~:i:iii"ii::':':a:~:i:::::( ii ic' i:iii-:ii i-.,i~;'rii'ii':::::ia::i: lil'iiiciic';iisii'ii-:ci:::::i:i:ii' Iliili-i BFsBsBB ii I ii:;:::j:: -:':::i:::: ::::.-_::::~:i:i::::;:::::::: BOi::::(-:::::::::: ii::i:::i:-:::'i::::::::::::::: i::::::::':::::i: i::_::(::r:li -:I:::::: i::::::;i':::::::::ii:::::::::..:.:::: -_.''iiiia:::i:i::-:::::::::: 3F'. "~`" -::::~:::-:~::::::::~:i:::~:: i: i'l:' iiiiaiii"i:i:i::i:iI:;iiiia~ic:iici, -::':-::-:::::::::::::":':"" ii:l-'iaii'iiiii::l:::::::.:.:.::~::i.;;;;:~:a::::j::r::::,::i:::::i::~:r:-::j:::::j:j:-:I:-:-:j::j::::::: iiii:Bii ii3i::l'i:iilii8i'ij'Zi:illj:::::i::::::::::-::::::::i:::,:,:i:i::::l:'-:ii d:::::C: NA W-63- 327 Fig. 7. Dynamic balance o nose cone and Apache at Wallops Island 92

5. SYSTEM PERFORMANCE The engineering design of the Pitot-Static Probe payload is evaluated with respect to data accuracy and system reliability. Data accuracy is examined strictly from an instrumentation viewpoint with little regard to the absolute accuracy of the measurement. A complete systems error analysis,15 is currently being studied, which will cover all aspects of the experiment, and should be available soon. It is convient to consider aspect data separate from other data channels because in general, instrumentation effects on aspect accuracy can be ignored. 5.1 ASPECT DATA ACCURACY Magnetometer data and solar aspect data are used primarily to determine the rocket angle of attack. Since the solar aspect information is in digital form, telemeter errors are negligible, and the accuracy of the angle of attack measurement is limited to the accuracy of the sensor itself, namely +1~. Stated without proof, the magnetometer system has been used to measure the angle of attack to within 5~. This figure being derived from comparisons of solar aspect data to magnetometer data in previous flights. 5.2 TELEMETERED DATA ACCURACY Instrumentation parameters that effect data accuracy are found principally in the telemeter system and include both frequency response and noise in addition to absolute voltage errors. Although the subjects are interrelated, each will be treated separately so that the contribution of each to the total expected error may be realized. 5.21 Response Multiplexed or commutated data are, by nature, effected by response because of the switching involved. The effect may be limited, however, by using proper data sampling and SCO bandwidths. In this case, the commutator using a thirty channel IRIG format at 2.5 frames/sec, produces a data pulse width of approximately 8 msec. Naturally, the leading edge of the pulse is effected more by switching transcients, whereas, it is possible, given enough time, for the trailing edge to be at the correct level. Automatic decommutation equipment has the capability to 93

"throw out" unwanted portions of the pulse. In other words, the static portion of the pulse may be measured separately while ignoring the transients. A SCO having a nominal intelligence frequency response of Ilt.k (modulation index of five) will carry eight harmonics of the basic pulse frequency with negligible distortion. 9 Under these conditions over 75% of the pulse width remains at a constant level (assuming the dataare constant) which by actual measurement has less than 0.1% error. 5.22 Noise Sources of noise that can effect data accuracy emanate from the Densatrons as part of the output signal, or they are introduced in the modulated signal. Other noise sources likely to occur in the instrumentation, such as ripple, stray signal pickup, cross talk, and ground loop errors, are all minimized to levels far below those mentioned above. 5.221 Densatron Output Noise Densatron noise outputs vary directly with the range of current being measured; this effect coupled with component variations produce unique noise characteristics in each unit. Table VI lists typical noise voltages found in Densatrons. Besides the converter switching spikes, the principal character of the noise is high frequency, random noise. TABLE VI DENSATRON AMPLIFIER NOISE CHARACTERISTICS Current Noise Level, Range my Peak-to-Peak 10-8 2-5 Random noise 10-9 2-5 Random noise 10-10 2-5 Random noise 10-11 7-10 Random noise plus 1/f noise 10-12 40-60 Primarily 1/f noise. Level may increase as high as 100 mv after amplifier frequency compensation Note 1: Converter switching spikes less than 1 jsec duration, 1 to 2 v magnitude are present on all ranges. Note 2: Maximum sensitivities are 10 amp for ambient gage amplifier and 10-12 amp for ram gage amplifier. 94

The effect of the noise on data accuracy is measurably reduced by the automatic filtering and integration of the signal as it passes through telemeter and ground station circuits. To isolate and assess the actual data error due to noise is extremely difficult. However, experience has shown that noise errors on current ranges greater than 10-11 amp are indiscernable while the 10'12 amp range is probably faced with an error of +0.5%. 5.222 Telemeter System Noise Noise generated in telemeter becomes appreciable as signal to noise (S/N) are reduced. Obviously, then signals must be kept large with respect to noise for optimum performance. There are three possible areas to consider in a FM/FM telemeter system, first the SCO bandwidths, second the RF carrier deviation, and third the RF signal level at the receiver. The SCO bandwidths, according to IRIG standards,l3 are based on a modulation index of five to achieve the highest signal to noise ratio possible, and still provide adequate bandwidth within the subcarrier frequency spectrum. As for RF carrier deviations, IRIG standards are again used which limit the maximum carrier deviation at + 125 kc. Of course, the carrier deviation must be divided between each SCO in the system. By using subcarrier preemphasis,10 with a minimum deviation of 5 kc for the low-frequency channels, S/N ratios are somewhat equal foi each channel and yield comparable results. In practice final deviations are based on the outcome of experimental tests for removing cross modulation products that create unwanted disturbances in the signal. Typical deviation values for a given payload appear below: Required equired Actual Actual Channel Cent. Freq. Calculated De n M. Deviation Mod. Index Deviations 12 10.5 kc 10.9 kc 11 kc 1.045 A 22 kc 33.3 kc 34 kc 1.54 C 40 kc 80.8 kc 80 kc 2.0 The RF power requirements are covered specifically in an earlier section (see Section 2.41) that concludes that a signal strength deficiency does not exist, at least during the measurement portion of the flight. 95

From the standpoint of noise, data accuracy is not appreciably effected from any of the sources mentioned above-a possible exception being in the abnormal condition when low signals are received. 5.23 Voltage Errors Because data acquisition by itself is a complex process which involves large amounts of equipment, accumulative errors can be large even though individual errors are small. Fortunately, major error sources are accounted for in the design by using in-flight voltage calibrations which automatically compensate for circuit and component drifts in both airborne and ground equipment. As stated earlier, the calibration regulator supplies reference voltages for SCO calibrations that are accurate to +0.05% under specified environmental changes. The SCO linearity specifications are +0.25% of design bandwidth, however, in practice most of the units used are linear to +0.1% or better. The accumulation of these errors yields a probable error not greater than +0.12%, which represents also a "worst case" error because, (1) the calibration regulator is usually more accurate than stated for the reason that environmental extremes are not experienced, and (2) linearity errors are greatly reduced because six calibration points are used. Modulation and demodulation of the RF signal will not add appreciable error as long as two conditions are satisfied (1) a strong RF signal is received at the ground station, which is certainly true during the data portion of the flight, and (2) the modulation index is sufficient to capture significant data sidebands. The second condition is guaranteed in the design by using IRIG standards. The data error at this point which is due entirely to the payload is estimated at 0.12%. Further demodulation and signal conditioning through the SPRL Data Conditioning System,l4 will add a theoretical error of 0.5%. Putting all of the errors together, we can conservatively estimate that from the SCO input to the digital output of the ground station a theoretical "worst case" error of 0.56% is possible although not expected. The complete system has been tested in the laboratory with a resultant error of 0.3% which takes into account also, the errors generated by noise and response limitations for which no specific figures were given above. Since a large percentage of the error is in the ground station, it is possible to reduce the error further by careful maintenance and calibration and/or future improvements in ground equipment, 96

5,24 Error Summary Angle of attack measurements are accurate to +1~ using the solar aspect system, and within 5~ with the magnetometer system. The payload has a theoretical telemetered data error of +0.12% which increases to 0.56% in the ground station. The Densatron 1012 amp current range will probably have a RMS error of around 0.75%, because of the high noise level on this rangeo The instrumentation errors, as stated, are characteristic of past payloads, and in general are typical of the errors expected in future payloads until such time as system redesign is necessary. 5.3 RELIABILITY System reliability is no better than the reliability of component parts, and/or the quality of workmanship that characterize the system, In an attempt to establish a high degree of reliability in the Pitot-Static Probe system, considerable emphasis is generated in the following areas of payload development, (1) design, (2) fabrication, (5) inspection, and (4) testing. The design philosophy has been to use simple circuits, redundancy wherever economically feasible, provide adequate safety margins, and otherwise follow standards and specifications as set forth in appropriate NASA documents. 16,17 Fabrication and inspection are supervised by personnel certified by a NASA quality assurance school. High reliability, quality components are used in circuits, and soldering standards and procedures are implemented as outlined in "Quality Requirements for Hand Soldering of Electrical Connections" NASA Quality Publication, 200-4.16 The instrumentation are built to withstand shock and vibration using various potting compounds and cable lacing where necessary. Testing is implemented at various stages of circuit and payload development, and individual performance tests of each circuit and subassembly is mandatoryo Vibration testing and approximately 25 hr of instrumentation tests and check out complete the testing, and provide a final degree of confidence in the payload, Although the program outlined above is designed to provide a reliable payload the results can never really be proven except by actual launchingso Perhaps then, the best definition of system reliability is found in the history of past flightso To date, the Pitot-Static Configuration has 97

been launched twenty-five times without evidence of any payload malfunction. On three occasions, however, pressure data were not obtained because of rocket malfunctions. The first loss of data occurred on NASA 14.29 which had no second stage ignitiono The other two rockets "broke up" due to problems associated with the Apache. In each instance the payload instrumentation performed as expected and data undoubtably would have been acquired had the rockets functioned normally. 5.4 SUMMARY OF SYSTEM PERFORMANCE Data quality has been examined with respect to frequency response, noise, and absolute voltage errors, with the result that a worst case error of +0.56% is possible, however, certainly not expected. This error represents a fixed error due strictly to data conditioning processes in payload instrumentation and ground station equipment. It does not reflect the errors found in either theory, pressure calibrations, or the measurement technique all of which are covered in a separate report.15 Past Pitot-Static flights have proven the payloads capability to operate in a rocket environmento A program of quality control using appropriate NASA publications for guidelines and NASA certified personnel for inspection, has been implemented to enhance the payloads reliability. The results of the program can probably best be summarized by noting that none of the three failures recorded out of twenty-five flights are attributed to payload malfunction. 98

6. LAUNCH OPERATIONS The final launch preparation of Pitot-Static Probe payload begins approximately three weeks prior to a scheduled launcho Each payload is completely assembled, checked out, and otherwise put in a flight condition before shipment to a launch site, Only a few operations, which cannot be done before hand, are required at the launch site, Advantages to this procedure are (1) the payload is assembled under laboratory controlled conditions and (2) range support requirements are minimized. A typical time schedule of events preceding a launching is listed below, 99

TABLE VII LAUNCH OPERATIONS TIME SCHEDULE Launch Time Function L-21 Days Payload fabrication and preliminary test completed. L-19 Days Final Densatron/gage assembly and preparation for calibration. L-16 Days Gage pressure calibrations completed. L-16 Days Densatrons installed in nose cone. L-15 Days Operational tests. L-14 Days Vibration test of complete payload except for DOVAP transponder. L-14 Days DOVAP section shipped to launch site. L-13 Days Operational tests after vibration. L-13 Days Gage pressure correlation tests. L-12 Days Payload shipped to launch site in evacuated vacuum shroud. L- 8 Days Payloads arrive at launch site. L- 7 Days Personnel arrive at launch site. L- 6 Days Payload operational tests including telemeter. Vacuum system assembled. L- 5 Days Vacuum pumping of nose cone begins and continues until launch day. L- 3 Days Magnetometer calibration tests. L- 2 Days Gage pressure correlation tests. L- 1 Day Complete horizontal test with payload and rocket assembled on launcher. L- 3 Hours Final assembly of payload to rocket on launcher. 100

TABLE VII (Concluded) Launch Time Function L- 2 Hours Vertical instrumentation test. L-20 Minutes Payload vacuum shroud removed. L-17 Minutes Start of launcher elevation, L- 8 Minutes DOVAP on external power. L- 5 Minutes Payload on external powero L- 3 Minutes Payload on internal power, L- 2 Minutes DOVAP on internal power. L- 1 Minute Payload in flight conditon, L- 0 Rocket launched. L + 390 Seconds Approximate payload LOS Note 1: The Densatron gage pressure calibrations are performed using the system shown in Fig. 58 and Fig. 59. Note 2: The gage pressure correlation tests are performed with the payload completely assembled and the probe section surrounded by a vacuum shroud (Figo 60). 101

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7o FUTURE CONSIDERATIONS The basic philosophy governing the design of the Pitot-Static Probe system has been to constantly up-grade the experiments capabilities and accuracy without effecting its operational status. The payload, as it is described herein, represents the design as it stands to day which is the result of many incorporated changes since the programs inception. The subjects discussed below are physically realizable and undoubtedly will be great step forward in improving the scientific value of the experiment, 7o 1 WIND MEASUREMENT Standing foremost in scheduled changes for Pitot-Static Probe payloads is the addition of a wind measurement system capable of measuring atmospheric winds in the region between 90 km and 110 km, The system is now under development and should be completed by late 1967o Included in the development is a specially designed hot filament pressure gage and amplifier combination, The measurement will be made using a single pressure port located immediately behind the ambient pressure ports that will give, because of the vehicle spinning, a wind modulated output. Also added to the payload will be a despin module (Figs. 54 and 55) that uses a yo-yo despin techniquel to reduce the vehicle spin rate from a high of 10 rps to about 2;rps, Flight tests of the modules have been made on PitotStatic Probe rockets NASA 14168 and NASA 14,169o Another instrument, associated with the wind measurement and currently in the development stage, is a moon sensor aspect system to be used for nighttime launchings. The expected completion date is late 1966.* 7.2 AMPLIFIER SYSTEM Currently in the design stage is a new amplifier system for the radioactive ionization gages. The amplifier employs a new range switching concept that shortens amplifier recovery time between range changes. Additional current ranges will also be used to improve both resolution and accuracy of the measurement. The amplifier is scheduled for completion in early 1967. *The first lunar sensor was successfully flight tested Feb, 1, 1967, on NASA l4o 316. 106

753 IONIZATION GAGES An attempt to extend the useful measurement range, both in lower and higher altitudes, will be made with a new gage configuration that uses americium as the radioactive source, The first flights with the new gage are scheduled for mid-1966. Overlapping ambient pressure measurements will be performed using an americium gage from 20 km to 55 km and a normal tritium source gage in the region between 40 km and 85 km. It is now expected that the americium source gage will eventually replace the tritium source gage altogether in the experiment. It is also proposed, at this time, to develop a hot filament gage to be used primarily in the wind measuremento The advantages of this type of gage are improved sensitivity and much better electrical and chamber response. Preliminary work has already been completed. 107

8. APPENDIX 8,1 NOSE CONE HEATING Temperature changes within the nose cone caused by aerodynamic heating have been measured by thermistor circuits similar to Fig. 19. The data are graphed in Fig, 62 to Figo 69. The recorded temperatures appear normal for a Nike-Apache payload 19 Early measurements, such as those on AA6,5340 and NASA.4,21, were limited in accuracy because of poor thermistor response and calibrations. Recorded temperatures are probably no better than +10~Co The YSI thermistors used on NASA 14o251 and NASA 14 285 have a guaranteed 1% accuracy besides having a fast response, In these cases, the recorded temperatures are most likely within +2~C except in peak heating periods where response is again a limitationo Measurement resolution decreases with increasing temperatures; therefore, more accurate data can be expected at the lower recorded temperatures, 108

NOSE CONE TEMPERATURES A.A. 6.340 200 LAUNCHED OCT. 171960 FT. CHURCHILL MANITOBA 160 120 H~~~~ 80 THERMSTOR TEMP Co LOI 40J ^^^ >^ (CJ C-SENSOR LOCATED ON A3 AMBIENT GAGE BACKING PLATE 0 B 0 40 80 120 160 200 240 280 320 360 400 TIME SEC. Fig. 62

260 NOSE CONE TEMPERATURES N.A,S.A 14.21 LAUNCHED DEC.7,1963 220 - WALLOPS ISLAND, VIRGINIA TEMP C~ 140 -. THERMISTOR LOCATION I00 - / A a B 100 60 20 L' I I, I.I I I I I 0 40 80 120 160 200 240 280 320 360 RANGE TIME SECONDS Fig. 63

260'NOSE CONE TEMPERATURES NAASA.A 14.21 LAUNCHED DEC. 7,1963 220 WALLOPS ISLAND, VIRGINIA TEMP C~ 180 140 100- // | L WIND M THERMISTOR LOCATION 60 CABLL 20 0 40 80 120 160 200 240 280 320 360 RANGE TIME SECONDS Fig. 64

260 NOSE CONE TEMPERATURES N.A.S.A 14.21 220 LAUNCHED DEC.7,1963 TEMP C WALLOPS ISLAND, VIRGINIA 180 140 100 THERMISTOR LOCATION WIND 60 CABLE 20 F 0 40 80 120 160 200 240 280 320 360 RANGE TIME SECONDS Fig. 65

260 NOSE CONE TEMPERATURES N.A.S.A. 14.21 LAUNCHED DEC. 7,1963 220 WALLOPS ISLAND, VIRGINIA TEMR C~ 180 140 140 THERMISTOR 100ook NOTX LOCATION / 100 -H NOTE: SENSOR J LOCATED I (/ ^ ^^^^^ G I JON TOP PLATE OF 60 I- NSTUMENTATION COLUMN, i, I I 0 40 80 120 160 200 240 280 320 360 RANGE TIME SECONDS Fig. 66

NOSE CONE TEMPERATURES 120 N.A.S.A. 14.251 1 _ LAUNCHED FEB 27, 1966 ASCENSION ISLAND 100 90 80 v 770 TEMP C~ 60 50 40 ( 31 30 20 10 I I I I 0 50 100 150 200 250 300 350 400 TIME SEC Fig. 67

140 130 NOSE CONE TEMPERATURE N.A.SA. 14.285 120- LAUNCHED AUG, 26,1966 WALLOPS ISLAND 110 00 -TEMP, C~ 9080 70 60 50 -. THERMISTOR LOCATED —. 40-/ AT ASPECT SENSOR MOUNTING PLATE 30 2010 - 0 20 40 60 80 100 20 40 60 80 200 20 40 60 80 300 20 40 60 80 400 TIME SEC. Fig. 68

110 NOSE CONE TEMPERATURE 100 TRANSITION SECTION N.A.S.A. 14.289 90- LAUNCHED AUG. 7, 1966 FT, CHURCHILL, MANITOBA 80 TEMP Co 70 60 50 40 H / THERMISTOR LOCATION 30 20 10 0 NOTE: SENSOR OUTPUT WENT TO 0 VOLTS AT 36 SECONDS CAME BACK TO,2 2 VOLTS AT 373 SECONDS INDICATING A ttMPERATURE OF 200C~ I I I I I I I I I I I I I I I I I I I 0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30 32 34 36 38 40 RANGE TIME SECONDS Fig. 69

8,2 ENGINEERING HISTORY OF PITOT-STATIC PROBE LAUNCHINGS 117

Rocket Number: AA6.540 NASA 14.19 NASA 14.20 NASA 14.21 NASA 14.22 Location: Ft. Churchill Wallops Island Wallops Island Wallops Island Ascension Date: 10-17-60 6-6-62 12-1-62 12-7-65 2-4-64 Time: 2104Z 2540Z 2054Z 1545Z 0155Z Result: Successful Successful Successful Successful Successful Peak Altitude: 140 km 124 km 152 km 140 km 158 km Measurements: Ambient pressure Ambient pressure Ambient pressure Ambient pressure Ambient pressure Impact pressure Impact pressure Impact pressure Impact pressure Impact pressure Nose Cone Temp. Wind pressure Wind pressure Nose Cone Temp. Instrumentation: Densatron Model C Densatron Model C Densatron Model C Densatron Model E Densatron Model F Tritium Gages Tritium Gages Tritium Gages Tritium Gages Tritium Gages Bendix TM-System Bendix TM-System Bendix TM-System Vector TM-System Vector TM-System 2 watt 2 watt 2.5 rps IRIG 2.5 rps IRIG Commutator Commutator H O0 Aspect: None Solar with Solar with Solar with 2 Axis earth cell earth cell earth cell Magnetometer Comments: Prototype Pump down Pump down Vacuum Shroud Launched in Nose Cone Apparatus Apparatus Support of Lost DOVAP Track Project FIRE at T+50 Seconds Vacuum S

Rocket Number: NASA 14.23 NASA 14.24 NASA 14.25 NASA 14.26 NASA 14.27 Location: Ascension 8~S Ascension Ship 52~35'S Ship 35~14'S Ship 60o00'S i5~W 78~20'W 74~15'W 78~00'w Date: 4-15-64 4-15-64 4-15-65 4-6-65 4-13-65 Time: 1556Z 0122Z 1600Z 1634Z 1600Z Result: Successful Successful Successful Successful Successful Peak Altitude: 158 km 156 km 139 km 142 km 149 km Measurements: Ambient pressure Ambient pressure Ambient pressure Ambient pressure Ambient pressure Impact pressure Impact pressure Impact pressure Impact pressure Impact pressure Instrumentation: Densatron Model E Densatron Model D Densatron Model F Densatron Models E&F Densatron Model F Tritium Gages Tritium Gages Tritium Gages Tritium Gages Tritium Gages Vector TM-System Vector TM-System Vector TM-System Vector TM-System Vector TM-System 2 watt 2 watt 1/2 watt 1/2 watt 1/2 watt Aspect: Solar with 2 Axis Solar with Solar with Solar with earth Magnetometer earth cell earth cell earth cell \O Comments: Launched in Launched in Vacuum Shroud Vacuum Shroud Vacuum Shroud Support of Support of Project FIRE Project FIRE Vacuum Shroud Vacuum Shroud

Rocket Number: NASA 14.29 NASA 14.47 NASA 14.48 NASA 14.63 NASA 14.64 Location: Ship - Ascension Ascension Ship 44025'S Ship 0000'S Wallops Island 77047TW 84~'w Date: 11-19-64 5-22-65 5-22-65 4-9-65 5-8-65 Time: 18842 02022 14002 20262 1748Z Result: Failed Successful Successful Successful Successful Peak Altitude: 15.8 km 158 km 158 km 147 km 145 km Measurements: Ambient pressure Ambient pressure Ambient pressure Ambient pressure Ambient pressure Impact pressure Impact pressure Impact pressure Impact pressure Impact pressure Instrumentation: Densatron Model F Densatron Model F Densatron Model E&F Densatron Model F Densatron Model F Tritium Gages Tritium Gages Tritium Gages Tritium Gages Tritium Gages Vector TM-System Vector TM-System Vector TM-System Vector TM-System Vector TM-System 1/2 watt 1/2 watt 1/2 watt 1/2 watt 1/2 watt 2.5 rps IRIG 2.5 rps IRIG \-ir~~~~~~~~~~~~~~ ~~~~~~Commutator Commutator 0 Aspect: Solar with 2 Axis Solar with Solar with Solar with earth cell Magnetometer earth cell earth cell earth cell Comments: Apache failed Launched in Launched in Vacuum Shroud Vacuum Shroud to ignite Support of Support of Instrumentation Project FIRE Project FIRE worked Vacuum Shroud Vacuum Shroud Vacuum Shroud No Data

Rocket Number: NASA 14.65 NASA 14.66 NASA 14.67 NASA 14.168 NASA 14.169 Location: Ship 0~52'S Ship 24~5'S Ship 60~000S Ft. Chruchill Ft. Churchill 84~9'w 76~5'w 78~00'w Date: 3-9-65 4-4-65 4-13-65 11-9-65 11-10-65 Time: 0626z 16o6z 0405Z 1840Z 1630Z Result: Successful Sucessful Successful Successful Failed Peak Altitude: 145 km 140 km 151 km N/A N/A Measurements: Ambient pressure Ambient pressure Ambient pressure Ambient pressure Ambient pressure Impact pressure Impact pressure Imapct pressure Impact pressure Impact pressure Instrumentation: Densatron Model F Densatron Model F Densatron Model F Densatron Model F Densatorn Model F Tritium Gages Tritium Gages Tritium Gages Tritium Gages Tritium Gages Vector TM-System Vector TM-System Vector TM-System Vector TM-System Vector TM-System 1/2 watt 1/2 watt 1/2 watt 1/2 watt 1/2 watt 2.5 rps IRIG 2.5 rps IRIG 2.5 rps IRIG 2.5 rps IRIG 2.5 rps IRIG Commutator Commutator Commutator Commutator Commutator ru Thermistor Pressure gage Aspect: 2 Axis Solar with 2 Axis Solar with 2 Axis Magnetometer earth cell Magnetometer earth cell Magnetometer 1 Axis Magnetometer Comments: Vacuum Shroud Vacuum Shroud Vacuum Shroud Despin Apache head-cap Launched with failure NASA 18.03 Despin Vacuum Shroud Launched with NASA 18.02 Vacuum Shroud No Data

Rocket Number: NASA 14.251 NASA 14.252 NASA 14.285 NASA 14.286 NASA 14.289 Location: Ascension Ascension Wallops Island Wallops Island Ft. Churchill Date: 2-27-66 2-27-66 8-26-66 8-26-66 8-7-66 Time 1657Z 1911Z 0423Z 1949Z Result: Successful Failed Successful Successful Successful Peak Altitude: 131.2 km N/A 149 km 150.5 km N/A Measurements: Ambient pressure Ambient pressure Dual Ambient Dual Ambient Dual Ambient Impact pressure Impact pressure pressure pressure pressure Nose Cone Temp. Impact pressure Impact pressure Impact pressure Nose Cone Temp. Instrumentation: Densatron Model F Densatron Model F Densatron Model F Densatron Model F Densatron Model F Tritium Gages Tritium Gages Americium & Americium & Americium & Tritium Gages Tritium Gages Tritium Gages Vector TM-System Vector TM-System Vector TM-System Vector TM-System Vector TM-System 1/2 watt 1/2 watt 1/2 watt 1/2 watt 1/2 watt ~or ~ 2.5 rps IRIG 2.5 rps IRIG 5 rps IRIG 5 rps IRIG 5 rps IRIG rOZ ~Commutator Commutator Commutator Commutator Commutator Aspect: Solar with 2 Axis Solar with 2 Axis 1 Axis earth cell Magnetometer earth cell & Magnetometer Magnetometer Magnetometer Comments: Support of Apollo Apache headcap COSMET-NAM COSMET-NAM New Apache Reentry Exper. No. 26 failure Program program program head-cap design 24 hr after Back-up vehicle Nose Cone Temp. New Apache Nose Cone Temp. reentry for Apollo support Vacuum Shroud Headcap Design Vacuum Shroud Nose Cone Temp. Vacuum Shroud New Apache Vacuum Shroud No Data Headcap Design

Rocket Number: NASA 14.315 NASA 14.316 NASA 14.317 NASA 14.318 NASA 14.319 Location: Ft. Churchill Ft. Churchill Ft. Churchill Ft. Churchill Ft. Churchill Date: 2-1-67 2-1-67 2-1-67 2-1-67 1-31-67 Time: 0109Z 0826Z 0346z 0538z 2317Z Result: Failed Successful Failed Successful Successful Peak Altitude: N/A 160 km (Radar) 161.5 km (Radar) 164.4 km (Radar) N/A Measurements: Dual Impact Dual Impact Dual Impact Dual Impact Dual Impact pressure pressure pressure pressure pressure Ambient pressure Ambient pressure Ambient pressure Ambient pressure Ambient pressure Nose Cone Temp. Nose Cone Temp. Instrumentation: Densatron Model F Densatron Model F Densatron Model F Densatron Model F Densatron Model F Hot Filament & Hot Filament & Hot Filament & Hot Filament & Hot Filament & Americium Gages Americium Gages Americium Gages Americium Gages Americium Gages 2.5 rps Commutator 2.5 rps Commutator 2.5 rps Commutator 2.5 rps Commutator 2.5 rps Commutator Po Aspect: 1 Axis 1 Axis 1 Axis 1 Axis 1 Axis Magnetometer Magnetometer Magnetometer Magnetometer Magnetometer Lunar Sensor Comments: Launched 2nd Launched 5th Launched 3rd Launched 4th Launched 1st in series of 6 in series of 6 in series of 6 in series of 6 in series of 6 Rocket Failed Vacuum Shroud Commutator Vacuum Shroud Vacuum Shroud Vacuum Shroud failure Data may be recoverable

Rocket Number: NASA 14.5322 Location: Fto Churchill Date: 2-1-67 Time: 1158Z Result: Successful Peak Altitude: N/A Measurements: Dual Impact pressure Ambient Pressure Instrumentation: Densatron Model F Hot Filament & Americium Gages 2o5 IRIG Commutator Aspect: 1 Axis Magnetometer Solar Sensor Comments: Launched 6th in series of 6 Vacuum Shroud.12.

8,3 EQUIPMENT SPECIFICATIONS 80' Transmitter 125

TRANSISTORIZED CRYSTAL-CONTROLLED TRANSMITTER model TRPT-501 Model TRPT-501 is an all-silicon, transistorized, phase-modulated telemetry transmitter capable of transmitting the intelligence from any telemetry subcarrier system in the 215 to 260 MHz telemetry band. The unit has a power output of 0.5 watt. It is mechanically rugged and ultra-stable over wide temperature ranges. The transmitter is capable of supplying maximum power output with the use of available transistors in the telemetry band. ELECTRICAL SPECIFICATIONS POWER OUTPUT: 0.5 watt nominal. OUTPUT IMPEDANCE: 50 ohms nominal. FREQUENCY RANGE: 215 to 260 MHz. TUNE UP TOLERANCE: +0.005%. CARRIER DEVIATION: +150 kHz maximum; +125 kHz nominal. TYPE OF MODULATION: Phase. INPUT IMPEDANCE: 5 kilohms minimum. MODULATION CHARACTERISTICS: See modulation chart. MODULATION SENSITIVITY: 0.2 volt rms for a modulation index of 1 measured at 1 kHz. MODULATION FREQUENCY RANGE: 100 Hz to 200 kHz. MODULATION DISTORTION: Less than 1.5% for the following modulation index: Frequency Index 400 Hz to 5.4 kHz 4 7.35 kHz to 22 kHz 2 30 kHz to 70 kHz 1 POWER REQUIREMENTS: +28 volts dc +10% at 150 ma. 127

CONNECTOR PIN IDENTIFICATION (WHEN USED) 100 f —-------— \A- NO CONNECTION 3 80~ [ -- I9 I II0 ~ TOP VIEW C — MODULATION RETURN 6^0 =I_ _ _ _ _ _0=9=2- _1_ D — - MODULATION INPUTU 4O F _ C Rr4 E -28V DC- Q 1/ 5. _ ^ 0 WHITE _ MODULATION rNDUT t ( W. 4 CONTROLS: Eigh scede a mBLACKnS- GROUND O FOR +e28 VOLT y r e i n ENVIRONMENTAL SPECIFICATIONS ThZe unitwill m promn6 ce < charcteristics: 3 6. 5' 5 l^ ^ ^ = -=SIDE V IE W ~3 S < O S6 $~0 D A T R OLE' S3 EAC)| 0 6 F'^^^s^/ MODULATION FREQUENCY ( KC) NOTE> _2 /i il/ - MODULATION FREQUENCY (KC) MECHANICAL SPECIFICATIONS SIZE: Cylindrical: 2.6 inches diameter x 1.5 inches long. Square: 2.6 inches x 2.6 inches x 1.5 inches. WEIGHT: Approximately 9 ounces. MOUNTING: Any flat surface by use of three through bolts. CONNECTORS: See te outline drawing. CONTROLS: Eight screwdriver adjustments accessible, by removing nameplate. ENVIRONMENTAL SPECIFICATIONS The unit will maintain the following performance characteristics: FREQUENCY STABILITY: + 0.005%. DISTORTION: Less than 3%. POWER OUTPUT: 1.5 db. FM NOISE: Less than 2% per channel. when subjected to any of the environments listed below: TEMPERATURE: -20~C to +80~C. HUMIDITY: 90% relative humidity to 50~C. ALTITUDE: Sea level to vacuum. VIBRATION: From 5 to 30 Hz at 0.2 inch double amplitude. From 30 to 2000 Hz at 25 g. Four 15 minute sweeps in each of the major axes. SHOCK: 200 g for 11 milliseconds. 3 shocks in each major axis. ACCELERATION: 100 g for 1 minute in any axis. ORDERING INFORMATION WHEN ORDERING SPECIFY: Model Frequency Shape TRPT-501 128

8,32 Mixer Amplifier 129

SUBMINIATURE TRANSISTORIZED MIXER AMPLIFIER TYPE TA-58 GENERAL DESCRIPTION: The subminiature transistorized mixed amplifier is a high-gain, feed-back audio amplifier. It is used for summing the outputs of subcarrier oscillators. The unit has been designed for use in airborne equipment such as guided missiles and other similar equipment. It meets all applicable military specifications. The unit is characterized by extreme ruggedness. The supporting framework is light but sturdy and transistors and other component parts have been anchored securely to withstand existing environmental conditions such as; shock, vibration, acceleration, temperature, and humidity. Selection of materials, finishes, and relative position of components have been based upon effective heat transfer characteristics and anti-corrosive properties. 131

I. PERFORMANCE CHARACTERISTICS: GAIN: Adjustable to approximately 20 x by means of two resistors mounted beneath the rark chassis. FREQUENCY RESPONSE: Less than ~0.25 db from 100 to 100,i0 cps. INPUT IMPEDANCE: 10 kilohms. OUTPUT IMPEDANCE: Less than 5000 ohms. INPUT SIGNAL: 100 millivolts min. OUTPUT SIGNAL: 2.0 volts rms. HARMONIC DISTORTION: Less than 0.5% at 1.5 volts rms output. NOISE: Equivalent noise input 30 microvolts. STABILITY: ~10% variation in supply voltage will cause no change in unit gain. POWER REQUIREMENTS: + 28 volts dc at 10 ma. II. ENVIRONMENTAL CHARACTERISTICS: ACCELERATION: Capable of withstanding at least 200 g. SHOCK: Capable of withstanding at least 200 g. VIBRATION: Capable of withstanding 25 g in each major axis from 55 to 2000 cps. TEMPERATURE: Operating range from -55~C to + 125~C. III. PHYSICAL CHARACTERISTICS: SIZE: Space requirements have been given utmost consideration resulting in an extremely compact unit. Overall dimensions: 7/8" x 1-1/6" x 1-3/8", excluding connector and potentiometers. The outline drawing is included in the bulletin. WEIGHT: The total weight is approximately 11~ ounces. 132

IV. INSTALLATION AND OPERATION: MOUNTING: Mounting against flat surface with one screw extending through the unit. CONNECTOR: De-Jur connector. TEST POINTS: Two test points available for input and output voltage measurements. V. SPECIAL REQUIREMENTS: All of the above specifications may be modified to suit specific installations and systems. 133

VECTOR MANUFACTURING COMPANY, INC. TECHNICAL DATA DWG. NO. TD-178 UNIT DATE UNI |TA-58 I 7-9-62 VECTOR FOU G.CO. IC. A 7- 9 MIXER AMPLIFIER no. S i OUTPUT - TEST GAIN POINT ADJUST INPUT TEST POINT TOP VIEW CONTINENTAL CONNECTOR MM5-22P 0 0 I 0D a I1~ (s.oltE c A 0 0 0 M 4AX. PIN CONNECTIONS SERIES A, -:.f15 +8A -- NO CONNECTION 5 I B +28V DC 32 I ~ _!____,_ c SIG OUTPUT._J ||||t)|N|o U - D ----- SIG INPUT t L _, E GROUND (SIG, PWR 8 CHASSIS) SIDE VIEW SERIES B MOUNTING SCREW A. CHASSIS GROUND - -0T,262 B ---- +28VDC C --- SIG OUTPUT D SIG INPUT 0.95MAX. E —G ROUND (SIG B PWR) 0.369.570.468 MAX. - l —,CONTINENTAL CONNECTOR MM5-22P BOTTOM VIEW MIXER AMPLIFIER, TA-58, OUTLINE DRAWING AND PIN CONNECTIONS. 134

8 Voltage Controlled Oscillator;1 5

TRANSISTORIZED VOLTAGE CONTROLLED SUBCARRIER OSCILLATOR Type TS-54 GENERAL DESCRIPTION: The TS-54 Subcarrier Oscillator is a transistorized, voltage controlled oscillator designed for telemetry applications. The oscillator converts intelligence in the forms of varying dc voltages into frequency modulated subcarrier signals for further transmission by wire or radio to distant receiving stations. The major functional sections contained in the TS-54 are a voltage regulator, a compensated amplifier, a free-running multivibrator, and an output filter. The oscillator has been designed for use in aircraft, missile, re-entry vehicle, and satellite programs and meets all applicable military specifications, and is available in all standard IRIG channels. The oscillator is also characterized by extreme ruggedness due to the supporting framework being light but sturdy, and the transistors and other components having been anchored securely to withstand environmental conditions such as shock, vibration, and acceleration. Selection of materials, finishes and relative position of components have been based upon affective heat transfer characteristics and anti-corrosive properties, 137

I. PERFORMANCE CHARACTERISTICS: INPUT IMPEDANCE: 1 megohm ~20% or 500 kilohms min. available. INPUT VOLTAGE: 0 to 5 v, ~2.5 v, 0 to 3 v or ~ 1.5 v. OUTPUT IMPEDANCE: 47 kilohms. OUTPUT VOLTAGE: 0.85v p-p with a 10 kilohm load. 0 to 2 volts rms at test point. MODULATION SENSITIVITY: 0 to 5 volts dc for ~7.5% or ~15 % deviation. -2.5 volts dc for ~7.5% or 15 % deviation. 0 to 3 volts dc for ~7.5% deviation. -- 1.5 volts dc for ~7.5 % deviation. Special units available with ~40% deviation. DRIFT: Within ~0.25% of design bandwidth for a period of 8 hours at ambient temperature following a warm-up period of 15 minutes. DISTORTION: At center frequency the total harmonic distortion of output is less than 0.75 %. LINEARITY: Less than ~0.25% of design bandwidth from best straight line. This linearity is maintained at temperatures from -20~C to +80~C. STABILITY: A change in supply voltage of ~10% will vary the center frequency less than ~ 0.5% of design bandwidth at temperatures from -20~C to +80~C. INTELLIGENCE FREQUENCY For modulation index of 5.0 or greater, the intelligence freRESPONSE: quency response is within 0.1 db of the dc response. AMPLITUDE MODULATION: Less than ~5%. SENSITIVITY TO SOURCE Changing the source impedance from zero to infinity varies IMPEDANCE CHANGE: the frequency less than 1.0% of design bandwidth. SENSITIVITY TO OUTPUT 10 to 1 change in the load impedance will cause less than 1% IMPEDANCE CHANGE: change in output frequency. INTELLIGENCE FREQUENCY Harmonics of intelligence frequency are suppressed 40 db COMPONENT IN OUTPUT: below output signal. POWER REQUIREMENTS: 28 volts dc ~10% at 15 ma. II. PHYSICAL CHARACTERISTICS: SIZE: Overall dimensions: 7" D x 1-1/16" W x 138" H, excluding connector. WEIGHT: 1.75 oz. SPECIAL FEATURE: This unit is completely encapsulated. III. ENVIRONMENTAL CHARACTERISTICS: TEMPERATURE: The operating range is from -55~C to +125~C. At any information input, the output frequency is stable within 1 % of design bandwidth (based on best reference) for a temperature change of 0~C to +80~C; likewise, the output frequency is stable within ~2% of dbw (based on best reference) for a temperature change of -20~C to +80~C. Special units are available with stability of ~ 1% of dbw (based on best reference) for a temperature change from -20~C to +80~C. HUMIDITY: When exposed to a relative humidity of 95% for 2 hours at — 50~C center frequency is stable within ~1% of design bandwidth. 1538

ALTITUDE: With constant temperature and at any altitude from sea level up, center frequency is stable within ~0.5% of design bandwidth. VIBRATION: Center frequency is stable within ~0.5 % of design bandwidth when subjected to 30 g rms vibration from 55 to 2000 cps in each major axis. SHOCK: After a shock of 200 g of 10 millisecond duration in the direction of each major axis, center frequency is stable within ~0.5% of design bandwidth. ACCELERATION: Under constant acceleration of 100 g in each direction of each major axis, center frequency is stable to within +1% of dbw. IV. INSTALLATION AND OPERATION: MOUNTING: Against a flat surface with one captive screw extending through the unit. CONNECTION: Continental Type MM5-22P connector is used for input and output signal and power supply connections. CONTROLS: OUTPUT - Provides adjusting the TS-54 output voltage level. LOWER BAND-Provides adjustment of the lower frequency deviation limit. UPPER BAND-Provides adjustment of the upper frequency deviation limit. TEST POINT: Enables monitoring output voltage and frequency. V. SPECIAL REQUIREMENTS AND OPTIONAL FEATURES: All of the above specifications can be modified to suit specific installations and systems. Because of the extremely high stability of the TS-54 oscillator, special units are available without the deviation limit and output potentiometers. These oscillators are adjusted at the factory to the customers specifications. VI. ORDERING INFORMATION: To order specify; Model TS-54 Frequency, Deviation, and Input signal voltage and polarity. Patent Pending Specifications are subject to change without notice. Verification of specifications will be given with each order. 139

VECTOR MANUFACTURING COMPANY, INC. TECHNICAL DATA DWG.NO. TD-126 VECTOR MFG. CO.INC | LO UNIT TS-54 DAT 4-4-62 UPPER BAND! oTNHR. | O —-LOWER BAND ADJUST L T.l ^AN- ADJUST ~UJ U S SUBCARRIER OSCILLATOR v sr MODEL TS-54 OUTPUT, —' V auR ADJUST OUTPUT TEST POINT TOP VIEW CONTINENTAL CONNECTOR MM'5-22P'1 0 0 _I0 8 0 WD I SIDE! ~ ~ ~ <)c c AE I I II N^AX. PIN CONNECTIONS SERIES A 0.419 NI SCREWI~~ ~A —A -- 28V DC 5_j B —— INO CONNECTION -32 L C- SIG OUTPUT - 1111U1 UU-~ oD —-- SIG INPUT 8. L-~ E —- GROUND (SIG PWR a CHASSIS) 6 SIDE VIEW| SERIES B NO. 2-56 MOUNTING SCREW A —-- +28V DC -B262 8 — CHASSIS GROUND /-'o2"~ - _/ C SIG OUTPUT D —--- SIG INPUT 0- 9253 EMX —-—.E GROUND (SIG a PWR) 0.369 I o.5e7 I 0.468 0.46.MAX, " ^ —CONTINENTAL CONNECTOR ~1- I.I I MM5-22P BOTTOM VIEW 14o

8. 4 Magnetometer Calibration 94'

HELIFLUX CALIBRATION DATA MAGNETIC ASPECT SENSOR TYPE RAM- 53 CX FIELD IN OUTPUT SIGNAL MILLIGAUSS IN VOLTS D C 600 4._975_ Y Axis SERIAL NQ 2 550 4.792 500 4.615 450 4.412 400 4.209 350 3.980 300 3.753 250 3.508 200 3.301 150 3.080 100 2,874 50 2.691 0 2.500 (BIAS LEVEL) -50 2.312 -100 2.'131I -150 1.922 -200 1 709 DIRECTION OF MAGNETIC FIELD FOR'* —~~ i~l~fVOLTAGE SIGNALS ABOVE BIAS LEVEL -250 1.502 -300 1.264 NOTE: CALIBRATION MADE WITH A IOOK OHM RESISTOR FROM SIGNAL u-350 1. * 025 OUTPUT TO NEGATIVE TERMINAL OF BATTERY SOURCE, AND A IOOK -400 __181 ] OHM RESISTOR FROM BIAS OUTPUT TO NEGATIVE TERMINAL OF BATTERY -450.588 SOURCE -500.393 SCHONSTEDT INSTRUMENT COMPANY -55n.204 www ~ ~~-550 ~-04-~ ~SILVER SPRING, MARYLAND -600.018 CALIBRATION MADE WITH BATTERY SUPPLY OF 28,0 VOLTS DATE Dec 17, 1963 11262 142

8355 Aspect Sensor and Shift Register.1,.45

SHIFT REGISTER MODEL 235 rhe Model 235 Shift Register works with one Model 135 or Model 131B Aspect Sensor for spinning vehicles. SPECIFICATIONS INPUT: Power: +26 to +40 VDC, 0. 5 watt Solar aspect bits shifted in in parallel upon signal from command eye. OUTPUT: Format: Serial seven bit Gray coded word indicating angle plus end of word bit, 50% duty cycle. Earth output superposed as shift in reference levels. 145

Voltage levels into 10 K ohms: Earth telescope Earth telescope activated not activated Reference 1.25 +.1 v 3.9 +.1 v "zero" 0. 25 +. 1 v 2.9 +. 1 v "one" 2,25 +. 1 v 4.9 +.1 v Shift out rate: 1 to 1000 bits per second selectable by changing capacitor. Internal oscillator may be disconnected and external shift out pulses used with characteristics as follows: amplitude: +5 v pulse width: 5 us rise time: 5 us P rf: 10 Kc SIZE: 2" x 3 1/4" x 4 5/8" WEIGHT: 15 1/2 oz. CONSTRUCTION: Printed circuit boards 5- -1 —---- 7 v1 | I lS 0 6)~ ) 3 BENDIK j TOeP-1.- 835 -- NO.!3 (..85) DIk tMTE.TO0P-IQ- BP: 4 MTG. HOLES \BEmDI< DlTOpP-1+ -185 F[ - r~~~~ |- MATE JTOGP -1- 18 P i~ - - —.3Lt L-_. ^ * J L3/32 / BEONDI ATOEP-I4-l P MATE: TOcP-I4-185 (POWER) 146

ASPECT SENSOR MODEL 135 The Model 135 Aspect Sensor is used on spinning vehicles. The sensor combines a Model 131B sensor with the addition of a telescope which senses the presence of the earth in its field of view. The telescope utilizes a silicon solar cell as the photodetector. SPECIFICATIONS SOLAR ASPECT SENSOR Field of view: +640 Angular resolution: 1 0 Accuracy at transition: +. 250 EARTH TELESCOPE Field of view: 1 +. 250 conical field of view. SIZE: 1 15/16" x Z 7/16" x 1 11/16" WEIGHT: 4 oz. SPIN RATE: 12 rps or less 147

@~ ~ ~ I a\ (^^K""^'^ + ~ II_ *-v^ -i t' _ ~ ~ ~ ~ \ - I r ^ "} -^' | v_ -1 I~~~II - H1~~~~~~~. l1 o6/ -- ~- - \ — \ -11 "I - 3i

8.4 THERMAL VACUUM TEST 149

MEMORANDUM REPORT NO. 322-5-62 May 16, 1962 TO: Test and Evaluation Division Files FROM: Thermal-Vacuum Test Section Thermodynamics Branch SUBJECT: Nike-Apache Neutral Particle Pitot - Static Experiment (NASA - 14.19 UA), Results of Thermal-Vacuum Tests INTRODUCTION Flight Objectives: The flight objectives of the Pitot — Static Experiment are as follows: 1. To measure atmospheric pressure, temperature and density in the region of 30 - 120 KM altitude. 2. Measurement of high altitude atmospheric winds. 3. Test of the interferometer Dovap system as a means of gathering trajectory data for the Pitot - Static probe experiment. 4. Test of the optical aspect system. Environment: Sounding rockets launched during the month of May at Wallops Island are subject to 95% probability of launch temperature extremes of 0~C to 35~C. Since aerodynamic heating is not a factor because of short flight time launch temperature extremes were used for test parameters. 150

Atmospheric pressure during the planned experimental flight varies from sea level ambient to 2.1 x 10-5 mm Hg at an altitude of 120 KM. Payload Personnel: Payload functional tests were conducted by the following University of Michigan Engineers: J. C. Maurer - System Engineer G. F. Rupert - Instrumentation Engineer J. J. Horvath - Project Engineer PROCEDURE The thermal-vacuum tests were performed at GSFC on 5 May and 6 May 1962 in the Stokes 8' x 8' Vacuum Chamber. The following procedure was observed: 1. Initial check-out of experiment in thermal-vacuum chamber at ambient conditions. 2. Evacuate chamber to approximately 40 microns and backfill with dry nitrogen. 3. Reduce chamber temperature to 0~C ~2C and stabilize. 4. Energize experiment and evacuate chamber to less than 2 x 10-5 mm Hg. 5. Payload personnel to perform functional tests coincident with chamber evacuation. 6. This phase of test to continue for not less than 30 minutes from start of evacuation. 7. Return to atmospheric pressure. 8. Repeat steps 3 through 7 using a temperature of 35~C ~2C in step 3. 151

RESULTS Run #1 - Five hours were required to reduce the payload temperature to 0~C and stabilize. Chamber evacuation was started at 1630 hours (5/4/62) and continued for 35 minutes, reaching a final pressure of 1.8 x 10-5 mm Hg. Payload engineers performed functional tests at ten minute intervals during evacuation. All test objectives were met and the payload functioned without incident. Run #2 - Payload temperature was increased to +35~C and stabilized at atmospheric pressure. Chamber evacuation was started at 2348 hours (5/4/62) and continued for 32 minutes reaching a final pressure of 5.6 x 10-5 mm Hg. Payload engineers performed functional tests at ten minute intervals during evacuation. Tests were performed without incident. CONCLUSIONS The Pitot - Static Probe experiment was subjected to thermal-vacuum environmental testing which was in acceptable agreement with proposed flight conditions and met all performance requirements. It was noted on run #2 that the minimum chamber pressure recorded was 5.6 x 10-5 mm Hg, which is in excess of the 2 x 10-5 mm Hg specified under "Procedure." In view of the final low pressure of run #1 (1.8 x 10-5 mm Hg) and the uninterrupted performance of the payload it is felt the pressure differential of run #2 is not of sufficient magnitude to invalidate the test results. 7. 4L 6a W. W. AUER Thermal-Vacuum Test Section WWA:kmf 152

Distribution List: Director's Reading File Office of Technical Services Thermodynamics Branch (3) Thermal-Vacuum Test Section (2) Systems Evaluation Branch (2) GSFC Coordinator (W. S. Smith) GSFC Project Manager (N. W. Spencer) 1553

TABLE I Temperature History of Payload Test @ 0~ +20C (Temperatures in OC) Time 1615 1630 1635 1640 1645 1650 1655 1700 1705 TC #1 1.0 1.5 1.0 0 -0.2 -1.0 -0.5 -0.8 -0.5 TC #2 1.0 1.5 1.0 0 -0.2 -1.0 -0.5 -0.8 -0.5 TC #3 1.0 1.0 1.0 0.2 0 0 0 0 0.5 TC #4 0 0.5 1.0 2.0 2.0 2.0 4.0 4.0 5.0 TC #5 1.0 1.0 0.2 -0.2 -1.0 -1.0 -1.0 -1.0 -1.0 TC #6 1.0 1.0 -1.5 -3.7 -3.8 -3.8 -3.0 -2.8 -2.5 TC #7 0.2 0.5 -2.2 -3.8 -4.0 -4.0 -3.8 -3.2 -3.0 TC #8 1.0 1.0 0.2 -0.5 -0.5 -1.0 -0.5 -0.5 -0.5 TC #9 0 0.5 4.0 11.0 7.2 5.0 7.0 13.0 11.0 TC #10 0 0.5 -3.5 -4.5 -3.8 -3.0 -2.0 -1.5 -1.0 TC #11 (Chamber Wall) 0 0.2 0.2 -0.5 -1.0 -1.0 -1.0 -1.2 -1.0 TC #12 (Chamber 1.5 1.5 -22.5 -11.0 -3.8 -1.5 -0.5 0 0.1 Ambient) Pressure (mm Hg) 640 640 100 14 2.2 0.240 6.5x 7.0x 1.8x 10-3 10-4 10-5 ka 4ro 3 <U O ed Cf << C: * O*r4 N -)0 0 0| 0| H4 0O O' OD ON rU d ri r cn - -%On0 o <-< 0o-~5 rd (a Ir in M o r-1 1ir A 0o 0 og o O ( + O 0 J O kJD Od >i >i> 4 E-4 P4 r- > ( ( C'- M 04 Pp 1P4

Pitot Static Probe Thermal-Vacuum Tesii t I''-:: I::: h-^ [':.\'^'::*:!|::"!:^|::. ^- f Payload Test @ 0~' -2 CO It':i: l l " "I' k7 ~: "' Control' Point Thermocouples- f 4!jr 1* 1 ^ ^.^:.-:l:^:...j_.j —1..0l —~-~~~~~~~~~T #l..Inside Gauge Or ifice!''1...i;:l':..': -7 - -^^ 1::^ l' "^ j^ j7 <:- ^;: *.-;'^-^-T ^'.0 -—', ~~~~~~~~~~~~~TC #2 Outside Gauge' Orif ice.'i"^-!:-^^^...:;"^^.':' "^r'+'n:^'..*"0 —*^ ~~~~~TC #3 Base of Electrometer *ipfer1-"-:^ TC #4 Heat Sink of Power Converter TC #9 Transmitter Base..Pllate _ _.______: - ^ -. Payl oar A. Onri PL Off P/L L: *;-:: ^" * *!-:,.Stabilized __-__ 545 *IJ~~~~~~~~~~~~~~~~~~~~~~~~~~~~ 54 $4 L I L' _ I ~~~~~~~~~~~~~~~~~~~~~~~~.J~~~~~~~- _1545 1600 1615'1630 i 1645 1700 Time 2.: 24 Hour. Clock I' _J: 1~9K3Kt~iI;SL19L~~5 f..,...4

TABLE II Temperature History of Payload Test @ 35~ +2~C (Temperatures in ~C) Time 2330 2345 2350 2355 2400 0005 0010 0015 0020 TC #1 34.5 34.5 35.0 34.0 34.0 34.0 34.0 34.0 34.0 TC #2 34.5 34.5 34.0 34.0 34.0 34.0 34.0 34.0 34.0 TC #3 35.0 35.0 35.0 35.0 35.0 34.8 35.0 35.0 35.0 TC #4 34.5 34.5 35.5 37.0 37.0 37.2 39.5 41.0 40.5 TC #5 35.0 35.0 34.5 34.0 34.0 34.0 34.0 34.0 34.0 TC #6 34.5 34.5 32.5 30.5 30.5 31.0 32.0 32.5 32.5 TC #7 34.5 34.5 32.5 31.0 30.5 31.0 31.5 32.0 32.2 TC #8 34.5 34.7 34.0 34.0 34.0 34.0 34.0 34.0 34.0 TC #9 35.0 34.0 37.0 44.0 43.0 39.8 43.5 50.0 50.8 TC #10 34.5 34.5 32.0 30.5 31.5 32.5 33.2 34.0 34.5 TC #11 (Chamber Wall) 35.5 36.0 36.0 36.0 36.0 36.0 36.0 36.0 36.0 TC #12 (Chamber 34.5 34.5 5.2 22.5 31.2 34.0 34.2 34.5 34.5 Ambient) Pressure (mm Hg) 640 640 120 12 1.2 0.09 9.2x 9.Ox 5.6x 10-3 10-5 10-5 C) mi 010 4-) N 4-i 0 l% ol o'dad'cd-d^' ao'do 4H o - a H od o riQ Co0 -r Ha Co H o o (a n M rt v -" >H O ao E- i cn <N > 4 0 n %-7 P

- ~..'I".':.- I". —::'-[-: 1-:-,~::-'r!,-:" -.... Pitot Static Probe Thermal-Vacuum Test: i:.:.' - — i Payload Test @ 35~ ~2~C -'.- -:.: -' - -.r- --- - Control Point Thermocouples ----- -----— I - -,. -, - 1 - -,- 1 — - -. — J -- j - - j,'i'i'-: -:':; -. i: J. -.,...-::I- I-;-'' 1::~ ~.....:.~,'..'!:."...: _:,2! —''-'.''~:- -1.' - i'-'-.-'-' -:t.....,... i...r —-;' —--— 1.. -_.'.; TC #1 InsidgeGau Orifice —. — -: —---- - | TC #2 Outside Gauge Orifice':I-, —.!::::.. -:'' - | ~ -_ TC #3 Base of Electrometer Amplifier:j - - -.:.'' _.TC #4 Heat Sink of Power Converter I. —-: —-—. -- X.- TC #9 Transmitter Base Plate''; —: —.-: — - -1 -- |CPayload i: - P/L 0 P/L Off P/L OnL _ i- r- -;:..-..Stabilized.:.....:: j..:, t —:i-.: —; -:,.:-::.- - -::'.-. —. " -.' - - 5'0....:....-: j.... —. —-t -'! —. —-' 1 ~ __,.',. *'.::::i::.::";:::'' -::!'::-::. "':. ":, -:. _1.:..::'.: |:':- -' -| —- --':-'.-o-. - ----... —. -- |''' I,.t'...._. ----—'' —-: —!. -:i-':3. ":'_ _.;!:.:j:-:.2.4..: l.' 20'.; 1 - 0... ~:'' —:-.. - -: -: t t, -: - 1:-' u-_v- f-..:-:i- -- 135 o 3 A r__ L___ N r- — I ~ —t _ _....,'.~.........;.........t t'....'-" -.......'....'......-...-........'.........,~~~_,:: I::. -:::,:..:======================================= ==== === = =.::.:::,::...............~-....."....... —' —-.....~ -.1.._ _:..-........_......r?l - - -...-~...1............ —... —

Nike-Apache Pitot-Static Probe (NASA 14.19 UA) Thermo Couple Locations g____...... TC #12 Chamber Ambient:~-.S —---- T.C #5 Nose Surface TC #6 Skin Surface Electronics ---- TC #7 Skin Surface Electronics TC #2 Outside Gauge Orifice TC #1 Inside Gauge Orifice Electronics Surnace -, TC #3 Base of Electrometer Amplifier TC #9 Transmitter Base Plate TC #4 Heat Sink of Power 1\\\_C ^\\\\\Converter DOVAP k TC #10 Frame Surface *i _ ~TC #11 Chamber Wall NOT TO SCALE 158

9. REFERENCES lo Ainsworth, J. Eo, D. F. Fox, and H. Eo LaGow, "Upper Atmosphere Structure Measurement made with the Pitot-Static Tube," Jouro Geophysical Research 66(10), 3191-3211, 1961. 2o Spencer, No W., and R. L. Boggess, "A Radioactive Ionization Gage Pressure Measurement System," Jouro Am. Rocket Soc., January 29, 1959. 5, Horvath, Jo Jo, R. Wo Simmons, and L. Ho Brace, "Theory and Implementation of the Pitot-Static Technique for Upper Atmosphereic Measurements," Univ. of Michigan ORA Report 03554, 04673-1-S, March 1962. 4o Hines, P. Bo, "DOVAP Systems and Data Reduction Methods," New Mexico State University, Physical Science Lab., Univ. Park, No Mo, January 31, 1962o 5. Seddon, Co J., "Preliminary Report on the Single-Station Doppler Interferometer Rocket Tracking Technique," NASA Technical Note D1344, January 19635 6. Spafford, Mo, R. Wiack, and Ro Woodman, "The Rocket Interferometer Tracking (RIT) System," NASA Technical Note D2682, March 1965o 7o El-Moslimany, M. A., "Thoretical and Experimental Investigation of Radioactive Ionization Gauges," The University of Michigan Research Institute 2096, 2406, 2597, 03554-1-S, May 1960. 8, Burr-Brown, "Handbook of Operational Amplifier Applications," Burr-Brown Research Corporation, 19635 9. Peaks, H. Jo, "Some Basic Considerations of Telemetry System Design," NASA Technical Note D355, June 1960o 10. Sliltz, Ho Lo, et al., "Aerospace Telemetry," Volo II, Prentice Hall Inco, Englewood Cliffs, No J., 1966o 11o Keenan, B. M., "Design and Performance of the Models 20035, 2,004, 2.005 and 2~009 Telemetry Quadraloop Antennas," Physical Science Laboratory, New Mexico State University, Scientific Report No. 1, December 1, 1961o 12. Thoikol Chemical Corporation; Astro-Met Division; "Aeroelastic Flight Loads on a Pitot-Static Probe Payload," July 29, 1966o 135o IRIG Document No 106-60, "Telemetry Standards Revised June 1962," August 1962o Defense Documentation Center AD284370o 159

REFERENCES (Concluded) 14. Caldwell, J., "The Space Physics Research Laboratory Data Conditioning System," Univ. of Michigan ORA Report 05776-1-E, January 1966. 15. Horvath, J. J., Report of Pitot-Static Probe system error analysis to be published in 1967. 16. NASA Quality Publication NPC 200-4 "Quality Requirements for Hand Soldering of Electrical Connections," August 1964. 17. Gay Jro, J. A., "Reliable Electrical Connections," Third Edition NASA Technology Handbook SP.5002, December 19635 18o Fedor, J. V., "Theory and Design Curves for A Yo-Yo Despin Mechanism for Satellites," NASA TN D-708, August 1961. 19. Sterhardt, J. A., "Environmental Test of Nike Apache Rocket NASA 14.111 GT," Goddard Space Flight Center technical report No. X-671-65-236, June 1965o 20. Vector Division of United Aircraft Corporation "Transistorized Crystal Controlled Transmitter, Type TRPT-501," Vector Technical Bulletin No. 50335 21. Vector Division of United Aircraft Corporation "Subminiature Transistorized Mixer Amplifier; Type TA-58," Vector Technical Bulletin 2041. 22, Vector Manufacturing "Transistorized Voltage Controlled Subcarrier Oscillator, Type TS-54," Vector Technical Bulletin 1021, October 1962o 235 Adcole Corporation "Aspect Sensor, Model 135," Specification Sheet. 24. Adcole Corporation "Shift Register, Model 235," Specification Sheet. 160

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