THE UNIVERSITY OF MICHIGAN COLLEGE OF ENGINEERING Department of Electrical Engineering Space Physics Research Laboratory Sounding Rocket Flight Report NASA 18.50 THERMOSPHERE PROBE EXPERIMENT Prepared on behalf of the project by R. W. Simmons, M. F. Carter, and D. R. Taeusch ORA Project 07065 under contract with: NATIONAL AERONAUTICS AND SPACE ADMINISTRATION GODDARD SPACE FLIGHT CENTER CONTRACT NO. NAS 5-9115 GREENBELT, MARYLAND administered through: OFFICE OF RESEARCH ADMINISTRATION ANN ARBOR May 1968

TABLE OF CONTENTS Page ACKNOWLEDGMENTS iv LIST OF TABLES v LIST OF FIGURES vi 1. INTRODUCTION 1 2. GENERAL FLIGHT INFORMATION 2 3. LAUNCH VEHICLE 4 4. NOSE CONE 9 5. THE THERMOSPHERE PROBE (TP) 13 5.1. Omegatrons 13 5.2. Electrostatic Probe (ESP) 21 5.5. Support Measurements and Instrumentation 24 5.3.1. Aspect determination system 24 5.3.2. Telemetry 24 5.33.. Housekeeping monitors 26 6. ENGINEERING RESULTS 27 7. ANALYSIS OF DATA 28 7.1. Trajectory and Aspect 28 7.2. Ambient N2 Density 28 7.3. Temperature 31 7.4. Geophysical Indices 37 8. REFERENCES 40 APPENDIX. DETERMINATION OF THE TOTAL PAYLOAD MOMENTS OF INERTIA 41 iii

ACKNOWLEDGMENTS Over one hundred employees of the Space Physics Research Laboratory of The University of Michigan contributed to the success of the NASA 18.50 Thermosphere Probe Experiment. Some of the personnel with specific responsibilities are listed below: Carignan, G. R. Laboratory Director Caldwell, J. R. Electronics Engineer Campbell, B. J. Design Draftsman Crosby, D. F. Electron Temperature Probe Engineer Freed, P. L. Head Technician Grim, G. K. Electronics Engineer Kartlick, W. G. Omegatron Technician Kimble, R. G. Telemetry Technician Maurer, J. C. Payload Engineer McCormick, D. L. Machinist Niemann, H. B. Neutral Particle Scientist Pate, R. W. Omegatron Engineer Poole, G. T. Head Programmer iv

LIST OF TABLES Table Page I. Table of Events 3 II. Omegatron Data 18 III. N2 Ambient Density Data 36 v

LIST OF FIGURES Figure Page 1. Nike-Tomahawk with thermosphere probe payload.5 2. Nike-Tomahawk with thermosphere probe payload. 6 3. Nike-Tomahawk with thermosphere probe payload. 7 4. Nike-Tomahawk dimensions. 8 5. Thermosphere probe instrumentation design. 10 6. Assembly drawing, 8-in. nose cone. 11 7. Thermosphere probe assembly. 12 8. Thermosphere probe system block diagram. 14 9. Omegatron I. 15 10. Omegatron II. 16 11. OM I final calibration. 17 12. Electrostatic probe mounting configuration. 22 13. Electrostatic probe system timing and output format. 23 14. Minimum angle of attack vs. altitude (OM I). 25 15. Trajectory program output format. 29 16. Sequence of events. 30 17. OM I current vs. flight time. 32 18. K(Sola) vs. altitude for OM I. 33 19. Ambient N2 density vs. altitude (OM I). 34 20. Neutral particle temperature vs. altitude (OM I). 35 21. Solar flux at 10.7 cm wavelength. 38 22. Three hour geomagnetic activity index (ap) 39 vi

1. INTRODUCTION The results of the launching of NASA 18.50, a Nike-Tomahawk sounding rocket, are presented and discussed in this report. The payload, a Thermosphere Probe (TP), described by Spencer, Brace, Carignan, Taeusch, and Niemann (1965), was jointly developed by the Space Physics Research Laboratory (SPRL) of The University of Michigan and the Goddard Space Flight Center (GSFC), Laboratory for Atmospheric and Biological Sciences (LABS). The TP is an ejectable instrument package designed for the purpose of studying the variability of the earth's atmospheric parameters in the altitude region between 120 and 350 km. The NASA 18.50 payload included two omegatron mass analyzers (Niemann and Kennedy, 1966), an electron temperature probe (Spencer, Brace, and Carignan, 1962), and a solar position sensor. This complement of instruments permitted the determination of the daytime molecular nitrogen density and temperature and the electron density and temperature in the altitude range of approximately 140 to 284 km over Wallops Island, Virginia. A general description of the payload kinematics, orientation analysis, and the techniques for the reduction and analysis of the data is given by Taeusch, Carignan, Niemann, and Nagy (1965) and Carter (1968). The orientation analysis and the reduction of the nitrogen data were performed at SPRL, and the results are included in this report. The electron temperature probe data were reduced at GSFC, and are not discussed here. The NASA 18.50 payload described herein was specifically designed and implemented to serve as an engineering flight test vehicle for a second generation omegatron system, as well as to provide a direct comparison of this new system with the previously reported standard omegatron (OM I). A description of the Omegatron II system (OM II) and discussion of the engineering test results are included in Section 6 of the present report. 1

2. GENERAL FLIGHT INFORMATION The general flight information for NASA 18.50 is listed below. Table I gives the flight times and altitudes of significant events which occurred during the flight. Some of these were estimated and are so marked. The others were obtained from the telemetry records and radar trajectory information. Launch Date: 18 September 1967 Launch Time: 14:10:00.090 EST, 19:10:00.090 GMT Location: Wallops Island, Virginia Longitude: 75029'W Latitude: 37~50'N Apogee Parameters: Altitude: 284.624 km Horizontal Velocity: 522.26 m/sec Flight Time: 263.92 sec TP Motion: Tumble Period: 5.557 sec lRoll Rate: -17.16 deg/sec 2

TABLE I TABLE OF EVENTS (NASA 18.50) Flight Time Altitude Event (sec) (km) Lift Off 0 0 1st Stage Burnout 3.3 (est.) 1.3 (est.) 2nd Stage Ignition 12.0 6.7 2nd Stage Burnout 21.0 19.2 Despin 43.0 (est.) 65.9 (est.) TP Ejection 45.6 69.1 Omegatron Breakoff 81.6 136.2 Omegatron Filaments On, M28 82.5 137.6 Peak Altitude 263.9 284.6 L.O.S. 500.0 5

3. LAUNCH VEHICLE The NASA 18.50 launch vehicle was a two-stage, solid propellant, NikeTomahawk combination. The first stage, a Hercules M5E1 Nike motor, had an average thrust of 49,000 lb and burned for approximately 3.6 sec. The Nike booster, plus adapter, was 145.2 in. long and 16.5 in. in diameter. Its weight unburned was approximately 1325 lb. The sustainer stage, Thiokol's TE416 Tomahawk motor, provided an average thrust of 11,000 lb and burned for about 9 sec. The Tomahawk, 141.1 in. long and 9 in. in diameter, weighed 530 lb unburned. The TP payload, which was 90.2 in. long and weighed 181 lb, including despin and adapter modules, made the total vehicle 376.5 in. long with a gross lift-off weight of 2036 lb. The vehicle is illustrated in Figures 1, 2, 3, and 4. The launch vehicle performed flawlessly and reached a summit altitude of 284.6 km at 263.92 sec of flight time. 4

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ROCKET NO. 18:50 8.. ODI A. PAYLOAD 90..156 ^l^^ i, ^-FIRING & DESPIN SECOND STAGE TOMAHAWK 141.1 376&5 9" DIA. a ^ [r —-p FIRST STAGE NIKE BOOSTER 1 145.2 ll ORDNANCE ITEMS 16.5 DIA. (NOSE CONE OPENING PRIMERS O BREAKOFF LINEAR ACTUATORS 3 DESPIN INITIATION PRIMERS. I — -- y l SECOND STAGE IGNITER 59.936- NIKE BOOSTER IGNITER Figure 4. Nike-Tomahawk dimensions. 8

4. NOSE CONE A diagram of the NASA 18.50 payload, including the nose cone, the despin mechanism, and the adapter section is shown in Figure 5. An assembly drawing of the 8-in. nose cone is given in Figure 6. The payload is programmed to despin at 65 km altitude (43 sec after launch), and the ejection begins at 70 km (45 sec after launch). The ejection system is designed for a tumble period of 4 sec by using a 2.2 lb Neg'ator* force and by limiting the travel of the plunger to 0.75 in. (Carter, 1968). The breakoff devices of both OM I and OM II are removed at 156 km (82 sec after launch), and the filaments of the omegatrons are turned on approximately 2 sec later. *Negtator is a trade name. 9

ROCKET NO. NASA 18.50 GA LAUNCH RANGE NO. G2-3157 TYPE OF ROCKET NIKE-TOMAHAWK DATE OF LAUNCH 18 SEPT 67 (DAY 261) - - - LOCATION WALLOPS ISLAND 9. 250 TIME 19:10:00.09 Z.\ I5 INC ALTITUDE 284KM=176MI RESULTS ALL SYSTEMS NORMAL, GOOD DATA' 28.562 MISC NOTES: CONE I)TWO PROBES:STAINLESS, PLATINUM CON E NOSE CONE TIMER -iL T I L. - BREAKOFF H''_ 1.781 THERMOSPHERE 4OMU - PROBE OM.ADAP-t _"" OM. AMP ff -- 1REGULATOR DECK- - I - 90.156 ____________ _______ TIPOSC DECK ---- -: 1 1.1 NAME WEIGHT C.G. FROM OSC CONTROL DECK- 78406 | NAME WEIGHT | LOG1 C0G 7806 331 SPACE —---— 111 NOSE CONE 103 LB. ___ CONTROL DECK - PROBE 78 LB. COMM. DECK-: _ l COMBINED 181 L B. 52.31" SCO DECK t _ | BATT DECK —----, 1 XMITTER DECK - | ELECTRON TEMP 38.562 PROBE ----, 3.500 ASPECT SENSOR __ i AUX DECK —-—: I OSC DECK, _i_ _..; REGULATOR DECK- IOU.AMPI --- ~ 12.688 =.W OZ ~~~~ ~ ~ ~OM ADAPTER I m 1 1! I I Rll" I IBREAKOFF I- -— SW -.0 1 >~ 7.500 "4 "8 z i =0 2.000 —- 9 10

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5. THE THERMOSPHERE PROBE (TP) The TP used for the NASA 18.50 payload was a cylinder 38.6 in. long and 7.25 in. in diameter which weighed 78 lb. The major instrumentation for this payload included two omegatron mass analyzers and an electron temperature probe. Supporting instrumentation included a solar aspect sensor for use in determining the attitude of the TP. Figure 7 is an assembly drawing of the Thermosphere Probe. The diagram in Figure 5 shows the location of instrumentation and supporting electronics in the nose cone. Figure 8 is the system block diagram. 5.1. OMEGATRONS The omegatrons used in the payload were of two designs. The first, designated as OM I, was described by Niemann and Kennedy (1966). The second, referred to as OM II, incorporated several innovations in structure, function, and circuitry. Further discussion of the new design is contained in Section 6 of the present report. Table II lists the operating parameters of the gauges and associated electronics. The characteristics of the linear electrometer amplifier current detectors, used to monitor the omegatron output currents, are also listed. The breakoff configuration and omegatron envelope for each omegatron are shown in Figures 9 and 10. The calibration of the NASA 18.50 OM I, performed at SPRL during August 1967, is shown in Figure 11. 13

UMBIUCAL CABLE CONNECTIONS po2 po3 po4 __po1 VECTOR SCO ---------— OMTL OUT/S S -O KHZ +28PW - OM3 SYSTEM OMZ 2ND DER. _ II|ALL REGULATORS, REFERENCES, I ________ OM IST DER. - 40 KHZ I I POWER CONVERTORS, BIAS SUPPLIES, ------ OMg I OIUT/D _I5.4 KHZ I I I. IOSCILLATORS, & CALIBRATORS. _ _______OME RNG - j 3.9 KHZ I ledex control CONTROL LEDEX 3 KHZ I —----------- | / +28PW —* OMI SYSTEM COMMUTATO 7.5 me nitor voltage monitors POWER CONVERTORS, REGULATORS, ------------ ^ BIAS SUPPLIES, OSCILLATORS, a *18 REF, THERM CALIBRATE VOLTAGES amp control CALIBRATOR AND THERMISTOR et. oerbIcontrol +28PWE _ IcontRroml filament control I O U 14.5 KHZ POWER CONTROL ESP-D 30KHZ ext. power — ELECTROSTATIC PROBE l l | SOlAR ASPECT | ASPECT O| KHZ,_____________ ESP-F _ _ 22 KHZ +28PW J T 2P.2 INTERNAL L EDEX a RELAY KFF i ROWER C ONTROL | P! 20%SY + 2lPW 01PO SILVERCELLS --------- — * SOLAR ASPECT ASPECT 10.5 KHZ APPROX. 28V +28PW TIMER + ACTUATORS OMI t FILAMENTS Figure 8. Thermosphere probe system block diagram.

FEED THRU PINCH OFF TUBE BREAKOFF ASSEMBLY iI ^ffll TiEPri -^ —' ACTUATOR CERAMIC ------ RlWL-ENVELOPE I< | Ij^"^ —---— ANTECHAMBER VACUUM SEAL OMEGATRON ---------- ^= —B —--— ~~-MAGNET OMEGATRON ADAPTER~ ~ VACUUM SEAL PROBE HOUSING ELECTROMETER AMPLIFIER DECK EMISSION REGULATOR a BIAS CONTROL ~_ ^WSBJ) T ~ ~ OSCILLATOR DECK AUXILIARY DECK_ Figure 9. Omegatron I. 15

PINCH OFF TUBE BREAKOFF ASSEMBLY ORIFICE l~. M ^-ACTUATOR ANTECHAMBER VACUUM SEAL ENVELOPE - OMEGATRON - G MAGNE OMEGATRON ADAPTER --- - -- VACUUM SEAL PROBE HOUSING A ~ w ELECTROMETER AMPLIFIER DECK EMISSION REGULATOR —- a BIAS CONTROL --. ( ) —--- OSCILLATOR DECK OSCILLATOR CONTROL DECK |(Dl I >| ^ —---- --— LOGIC DECK Figure 10. Omegatron II. 16

ic10 < 1: Iz hi,_ - II ________ 0 t^~~./*I~~ I I I I ~~NAS 18.50,0M I -4 S SEN ITIVITY: 1.34 x 10N PAIRT/CC N2 RUN OF 9- 3-67 OMEGATRON VS. EA7 106 107 108 109 100 loll0" 10 NUMBER DENSITY (PART/CC) 1/9/68 Figure 11. OM I final calibration.

TABLE II OMEGATRON DATA (NASA 18.50) Omegatron Gauge Parameters OM I OM II Beam Current: 1.85 pA 1.93 pA Electron Collector Bias: 75.50 V 76.65 V Filament Bias: -86.38 V -93.05 V Cage Bias: - 0.17 V - 0.20 V Top Bias: - 0.59 V - 0.60 V RF Amplitude, Mass 28: 4.00 Vp.p 3.95 Vp.p RF Frequency, Mass 28: 143.58 kHz 144.62 kHz Monitors Filament OFF: 0.11 V 0.12 V ON: 3.39 V 3.38 V Beam OFF: 0.66 V 0.66 V ON: 3.68 V 3.68 V Thermistor Pressure (zero pressure) Filament OFF: 2.55 V 2.81 V Filament ON: 2.39 V 2.68 V Bias: 3.96 V 4.07 V RF: 3.26 V 3.37 V Calibration Sensitivity: 2.40 x 10-5 A/torr 2.05 x 10-5 A/torr Maximum Linear Pressure (5%): 8 x 10-6 torr 5 x 10-6 torr 18

TABLE II (Continued) Electrometer Amplifier OM I RRange Range Indicator Range Resistor Mass 28 ZPV 1 0.0 V 9.119 x l09 2 5.014 v 2 0.7 v 2.479 x 1010 n 5.008 v 3 1.4 v 6.738 x 10ol Q 5.007 v 4 2.1 v 1.832 x 10ll Q 5.000 v 5 2.8 v 4.979 x 101l 0 4.990 v 6 3.5 V 1.353 x 1012 0 4.989 v 7 4.2 V 3.679 x 1012 X 4.960 v 8 4.9 v 1.000 x 1013 Q 4.977 v 19

TABLE II (Concluded) Electrometer Amplifier OM II Range Range Indicator Range Resistor Mass 28 ZPV Bias 1-1 0.85 V 0.91 x 1012 Q 0.034 V + 3.037 V 1-2 1.01 V 0.91 x 1012 Q 0.034 V - 0.985 V 1-3 1.19 v 0.91 x 1012 Q 0.034 V - 4.949 V 1-4 1.38 V 0.91 x 1012 o 0.034 V - 8.902 V 1-5 1.55 V 0.91 x 1012 Q 0.034 V -12.868 V 1-6 1.74 V 0.91 x 1012 Q 0.034 V -16.890 V 1-7 1.89 V 0.91 x 1012 Q 0.034 V -20.643 V 2-1 2.41 V 1.56 x 101l Q 0.050 V + 3.037 V 2-2 2.56 V 1.56 x 1011 n 0.050 V - 0.985 V 2-3 2.76 V 1.56 x o10" 0.050 V - 4.949 v 2-4 2.95 V 1.56 x 101 nQ 0.050 V - 8.902 V 2-5 3.12 V 1.56 x 101" 0.050 V -12.868 V 2-6 3.30 V 1.56 x 101 Q 0.050 V -16.890 V 2-7 3.48 V 1.56 x 0lol 0.050 V -20.643 V 3-1 3.92 V 3.22 x 1010~ 0.050 V + 3.037 V 3-2 4.07 V 3.22 x 1010 ~ 0.050 V - 0.985 V 3-3 4.25 V 3.22 x 1010 Q 0.050 V - 4.949 V 3-4 4.43 V 3.22 x 11~ 0.050 V - 8.902 V 3-5 4.61 V 3.22 x 101~ 0.050 V -12.868 V 3-6 4.79 V 3.22 x 10~ i 0.050 V -16.890 V 3-7 4.97 V 3.22 x 101~ Q 0.050 V -20.643 V Miscellaneous OM I OM II +28 power current all on: 340 mA 420 mA Preflight gauge pressure (N2): 8 x 10-6torr 1 x 10-2 torr Magnetic field strength: 2800 gauss 2800 gauss 20

5.2. ELECTROSTATIC PROBE (ESP) The ESP system consists of two cylindrical Langmuir probes immersed in the plasma and an electronics package which measures the current collected by the probes. The cylindrical Langmuir probes and the mounting configuration are shown in Figure 12. Probe 1, the lower probe, is made of stainless steel, while probe 2 is made of platinum. The electronics unit consists of a dc-dc converter, a ramp voltage generator, a three range current detector, range switching relays, and associated logic circuitry. There are two outputs, a data channel, and a probe and AV monitor channel. The data channel output is a voltage proportional to the collected probe current, and the probe and AV monitor output is a signal showing whether the system is measuring in its high or low AV mode and which probe is connected to the current detector. System timing and output format are shown in Figure 13. The following are the specifications of the Electrostatic Probe system for NASA 18.50: (a) Input Power 1.5 W at 28 V (b) Sensitivity Range 1 20 pA full scale (full scale output is deRange 2 2 pA full scalesI fined as 4.5 V) Range 3 0.2 pA full scale (c) Ramp Voltage (AV) Magnitude Slope High AV -2.9 V to +4.8 V 30.8 V/sec Low AV -1 V to +1.8 V 11 V/sec Period 250 msec High-Low AV alternate every 4.5 sec (d) Output Voltage -0.6 V to + 5.8 V Resistance 2000 Q Bias Level +0.5 V (e) Calibration Sequence (Synchronized with AV) Occurrence of calibration every 54 sec Duration of calibration 1.5 sec (f) Timing (see Figure 13) High-Low AV alternate every 18 sweeps (4.5 sec) Probes switch with every AV sweep (0.25 sec) Detector ranges change sequentially every 2 AV sweeps (0.5 sec) 21

DIRECTION OF FLIGHT PROBE NO. 2 PLATINUM TP CENTER SECTION 9.000 Figure 12. Electrostatic probe mounting configuration. __ r ~ \ 022-022 <'\ J85.875 %00, 0 O.D.-.065 PROBE No. I. L^^^^^^^ y^ STAINLESS STEEL Figure 12. Electrostatic probe mounting configuration. 22

I I +6V — -3 V-.-]-[I,, ~ -250 MS. HIAV 4.5 SEC' - -LO AV 4.5 SEC 3 0.20.a DETECTOR 3 -.2-a I 2 22a RANGE I 200 _a ro 6 V- PROBE I LEVEL IJiJIJJJI IJJ MONITOR 1!VJJJIVSVSVSVSJJJ PROBOE 2 LEVEL MONITOR ev I 6V CALI BRATION I I IO CALIBRATION OCCURS ONCE EVERY 54 SEC AT BEGINNING OF HI AV SWEEP I \^ <-CALIBRATE COMMON MODE CHECK Figure 13. Electrostatic probe system timing and output format.

5.3. SUPPORT MEASUREMENTS AND INSTRUMENTATION 5.3.1. Aspect Determination System The NASA 18.50 TP utilized a solar aspect sensor made by Adcole Corporation which is identical to the ones used on previous daytime shots. This system functioned properly throughout the flight. The attitude of the TP was determined by using the method of referencing the solar vector and the velocity vector (Carter, 1968). The resulting minimum angle of attack for OM I, determined to an estimated accuracy of +5 degrees, is plotted vs. altitude in Figure 14..53.2. Telemetry The payload data were transmitted in real time by a ten channel PAM/FM/ FM telemetry system at 240.2 MHz with a nominal output of 2.5 watts. The telemetry system used ten subcarrier channels, as outlined below. Transmitter: Driver TRPT-251RB01, Serial No. 1041 Power Amplifier TRFP-2V, Serial No. 222 Mixer Amplifier TA59A, Serial No. 1033 Subcarrier Channels (SCO Type TS58) IRIG Serial Center Low Pass Band No. Frequency Function Filter Used 18 3345-25 70 kHz OM II OUT/S 500 CD 17 1742-5 52.5 kHz OM II 2nd der 790 CD 16 1729-5 40 kHz OM II 1st der 600 CD 15 1720-5 30 kHz ESP-D 450 CD 14 1717-5 22 kHz ESP-F 330 CD 13 3044-25 14.5 kHz OM I OUT 220 CD 12 3211-25 10.5 kHz Aspect 400 CA 11 1980-25 7.5 kHz Commutator 160 CD 10 1676-5 5.4 kHz OM II OUT/D 80 CD 9 1661-5 3.9 kHz OM II RANGE 60 CD Instrumentation power requirements totaled approximately 40 watts, supplied by a Yardney HR-1 Silvercell battery pack of a nominal 28 V output. 24

300 280 260 240 w 220200 NASA 18.50,OMI 18 SEPTEMBER 1967 19:10 GMT WALLOPS IS.,VA. 180 160 140 120 0 4 8 12 16 20 24 28 32 36 40 44 ANGLE OF ATTACK (DEGREES) 1/9/68 Figure 14. Minimum angle of attack vs. altitude (OM I). 25

5.3.3. Housekeeping Monitors Outputs from various monitors throughout the instrumentation provided information bearing on the operations of the electronic components during flight. These outputs were fed to a thirty-segment commutator which ran at one rps. The commutator assignments were as follows: COMMUTATOR FORMAT FOR 18.50 (OM I G5, OM II G6) Segment Segment Number Assignment 1 Amplifier Range 2 Output 3 Filament Monitor 4 Beam Monitor 5 OM I Bias Monitor 6 RF Monitor 7 Internal Pressure Monitor 8 Thermistor-Gauge Temperature 9 Thermistor-Amplifier Temperature 10 _ Thermistor-Filament Regulator Temperature 11 Thermistor-Transmitter Temperature 12 - Thermistor-Filament Regulator Temperature 13 Thermistor-Amplifier Temperature 14 Thermistor-Gauge Temperature 15 Internal Pressure Monitor 16 OM II RF Monitor 17 Bias Monitor 18 Beam Monitor 19 Filament Monitor 20 OUT/D 21 Comparator Ramp Monitor 22 Open 23 Battery Voltage Monitor 24 0 V Calibration 25 1 V Calibration 26 2 V Calibration 27 3 V Calibration 28 4 V Calibration 29 and 30 5 V Calibration 26

6. ENGINEERING RESULTS The launching of the NASA 18.50 Thermosphere Probe was normal and all systems performed as expected. The prime objective of this flight was to obtain data from a new omegatron system designated the Omegatron II, which was designed to provide high resolution measurements of the density variation within the chamber as the TP tumbled. These data are used to determine the ambient N2 kinetic temperature directly from the measurement of the OM II. The data obtained from the OM II were as anticipated. However, the reduction techniques and the analysis are considerably more complex than those required for the standard omegatron. Consequently, at the present time, we do not have complete results. A report describing the total OM II system and theory is in preparation, and the results from the NASA 18.50 flight will be included. 27

7. ANALYSIS OF DATA The telemetered data were recorded on magnetic tape at the Wallops Island Main Base and the GSFC Station A ground Station facilities. Appropriate paper records were made from the magnetic masters, facilitating "quick look" evaluations. The aspect data were reduced to engineering parameters from paper records. The omegatron and housekeeping data were reduced by computer techniques from the magnetic tapes. 7.1. TRAJECTORY AND ASPECT The position and velocity data used to determine aspect, ambient N2 density, and ambient temperature as a function of time and altitude were obtained by fitting a smooth theoretical trajectory to the FPQ-6 and FPS-16 radar data. The theoretical trajectory is programmed for computer solution similar to that described by Parker (1962). The output format is shown in Figure 15. The analysis of minimum angle of attack (amin) as described by Carter (1968) is also incorporated in the program. The output of the computer furnishes Anin, altitude, and velocity as a function of time. A plot of omin vs. altitude for OM I has already been given in Figure 14. Figure 16 shows the occurrence of significant events during the flight. 7.2. AMBIENT N2 DENSITY The neutral molecular nitrogen density was determined from the measured gauge partial pressure as described by Spencer, et al., (1965, 1966), using the basic relationship: / An u \ n = - K(So a) a 24T V cos a min where n = ambient N2 number density a Ani = maximum minus minimum gauge number density during one tumble, A x AI, where A is the sensitivity of the gauge Ui = 2kTi/m, most probable thermal speed of particles inside gauge Ti = gauge wall temperature V = vehicle velocity with respect to the earth amin = minimum angle of attack for one tumble K(S0,o ) = the reciprocal of the normalized transmission probability as defined by Ballance (1967), referred to as the geometry correction factor. 28

NASA 18.5) CM 1 LAUNCH TIME (GMT) YEAR 1967 DAY 261 H.OUR 19 MINUTE ir SECOND.*^ INITIAL CONDITIONS r IME 6C.Cr -"SECONCS FRFu LAUNCH ALTITUDE 32.3234.o FT RANGE 87023.0 FT VELCCITY 6322.3 FT/SEC FLIGHT PATH ANGLE 73.33C9 DEGREES UP FROM LCCAL FCRI7CNTAL PLANE _AZIMUTH 119.6q9',^ OEGREES EAST OF LOCAL NCRTH LO NGITUDE -75.220. ^ DEGREES ( AST ) LATITUDE 37 7213 DEGREES (+NURTH) NC. WIND SPECIFIED CGNE CUORPRECTIO N -.40^ MOMENTUM VECTOR INPUT BY SPECIFYING PHI LS = 351.C ANC THEIA LS = 142.C CCMPUTED MOMENTUM VECTCR IN EARTH FIXEO CCCRDINATES IS.93C215.328469 -.16373a MOMENTUM VECTOR INPUT BY SPECIFYINC PHI LS = 355.0 AND THETA LS = 143.C COMPUTED MOMENTUM VECTOP IN EARTH FIXED CCCRIINATES S IS.914231.3714E7 -.161799 PEAK PARAMETERS TIME ALTITUDE F ALPHA V*COS ALPHt PHI V RANGE F VELOCITY F VXFX VZFX AZIMUTH LATITUDE ALTITUDE M - RANGE M VELOCITY M VYFX ELEVATION LONGITUDE 263.92 c93386i 38.18 41n.56 57.75 3'-41C 5 1713,46 1447.03 -29.37- 123.097 37.227 284624 35.5,7 424.81 61.20 131395 522.26 -917.13 -.000 -74.208 Figure 15. Trajectory program output format.

ALTITUDE (KM) 0 0 0 0 0 o 0 0! I I I I O - NIKE BURNOUT 3.3 TOMAHAWK IGNITION 12.0 TOMAHAWK BURNOUT 21.0 ~- ~. - ODESPIN 43.0 TP EJECTION 45.6 ItM |OM BREAKOFF 81.6 H- ~ - OM FILAMENTS ON 82.5 0 CD 0, r e I 0 o m PEAK 263.9 0 0 o 0 CD ct0 z 0 I oc 0 0 L.O.S. 500.0 O 0 1 I I

AI, the difference between the maximum (peak) omegatron gauge current and the minimum (background) gauge current vs. flight time for OM I is shown in Figure 17. The background current is the result of the outgassing of the gauge walls, and the inside density is due to atmospheric particles which have enough translational energy to overtake the payload and enter the gauge. The outgassing component is assumed constant for one tumble and affects both the peak reading and the background reading, and, therefore, does not affect the difference. From calibration data obtained by standard techniques, the inside number density, Ani, is computed for the measured current. By using the measured gauge wall temperature, the most probable thermal speed of the particles inside the gauge, ui, is computed. The uncertainty in this measurement is believed to be about +2% absolute. V, the vehicle velocity with respect to the earth is obtained from the trajectory curve fitting described previously and is believed to be better than +1% absolute. Cos amin is obtained from the aspect analysis described by Carter (1968). Since the uncertainty in cos mnin depends upon %min, for any given uncertainty in 0min, each particular case and altitude range must be considered separately. Figure 14 shows that the minimum angle of attack for the upleg is generally less than ten degrees, so with an assumed maximum uncertainty in Qmin of +5 degrees, the resulting uncertainty in cos 0min is less than +2%. The data for low angle of attack were used as control data. K(So,a), the geometry correction factor vs. altitude is shown in Figure 18. As can be seen, the maximum correction for OM I is about 12%, or K(So,o) =.88 at about 140 km altitude for the upleg data. The correction factor, determined from empirical and theoretical studies, is believed known to better than 2%. The resulting ambient N2 number density, obtained from the measured quantities described above, is shown in Figure 19 and is tabulated in Table III. The uncertainty in the ambient density due to the combined uncertainties in the measured quantities is thought to be 10% relative and 25% absolute. 7.3. TEMPERATURE The ambient N2 temperature profile shown in Figure 20, tabulated in Table III, was obtained by integrating the density profile to obtain the pressure and then by relating the known density and pressure to the temperature through the ideal gas law. The assumption that the gas is in hydrostatic equilibrium and behaves as an ideal gas is implicit. Since the temperature depends only upon the shape of the density profile and not upon its magnitude, it is estimated that the uncertainty in its magnitude is +5% absolute. 31

NASA 18.50,OMI 18 SEPTEMBER 1967 19:10 GMT WALLOPS IS.,VA. lo-lo 101. IC-, *.'. ** (PK - BK G ) Ia. ~~~~~~ - o W.. 10 - ~ * x x xx (BKG) x x x X 1-13 x X 10 -r * x,x x - _ — x x x x Figure 17. OM I current vs. flight time. 32

I I I I I I I T l I 300 280 NASA 18.50, OMI 18 SEPTEMBER 1967 19:10 GMT WALLOPS IS.,VA. 260 240 220 1 — 200 UPLEG / DOWNLEG 180 / 160 140 120-.82.84.86.88.90.92.94.96.98 1.00 1.02 1.04 1.06 1.08 1.10 GEOMETRY CORRECTION FACTOR,K(So,) 1/9/68 Figure 18. K(So,a) vs. altitude for OM I. 33

340- NASA 18.50, OM I 18 SEPTEMBER 1967 320 19:10 GMT WALLOPS IS.,VA. 300 280 260 240 < 220 200 180 160 140 120 l.. I IIIl1 I I III9 I I 1 I I I11 1(C7''''''108 10 10 AMBIENT N2 DENSITY (PART/CC) I/s9/s Figure 19. Ambient N2 density vs. altitude (OM I).

320 300 NASA 18.50,0M I 18 SEPTEMBER 1967 280 _19:10 GMT WALLOPS IS.,VA. 260 _ 240 0 0 - / 5 220 / 200 180 160 140 600 700 800 900 1000 1100 TEMPERATURE (~K) 1/9/68 Figure 20. Neutral particle temperature vs. altitude (OM I). 35

TABLE III N2 AMBIENT DENSITY DATA (NASA 18.50) 18 September 1967 14: 10 EST 19: 10 GMT Wallops Island, Virginia OM I OM I Altitude Temperature Density km ~K part/cc 140 616 4.07 x 1010 145 653 3.00 150 690 2.24 155 726 1.70 160 760 1.31 165 792 1.03 x 101o 170 822 8.20 x 109 175 849 6.60 180 873 5.32 185 895 4.36 190 916 3.60 195 936 2.97 200 955 2.47 205 971 2.04 210 987 1.71 215 998 1.45 220 1010 1.23 225 1018 1.04 x 109 230 1027 8.95 x 108 235 1034 7.65 240 1039 6.57 245 1045 5.63 250 1050 4.86 255 1055 4.16 260 1059 3.6o 265 1063 35 11 270 1067 2.69 275 1070 2.32 280 1074 2.00 284 1076 1.79 x 108

7.4. GEOPHYSICAL INDICES The 10.7 cm solar flux (F10 7) and the geomagnetic activity indices (ap) for the appropriate periods are shown in Figures 21 and 22. 37

220 - FA.7 SOLAR FLUX VS. TIME 210 ~ AUGUST-SEPTEMBER 1967 20190 80 - 110.7 I97 160- g(I\ 1 Figure1 Soa fuat 950 (- \ II-)38 920JULY AUGUST SEPTEMBER 1967 Figure 21. Solar flux at 10.7 cm wavelength. 58

26 0 NASA 18.50 S 24 18 SEPTEMBER 1967 < 19:10 GMT 22 OQp=12 20 I8 16 Op 14 12 10 6 2 0 2 4 6 8 10 12 14 16 18 20 22 24 TIME (GMT) Figure 22. Three hour geomagnetic activity index (ar). 39

8. REFERENCES Ballance, James 0., An Analysis of the Molecular Kinetics of the Thermosphere Probe, George C. Marshall Space Flight Center, NASA Technical Memorandum, NASA TM X-53641, July 31, 1967. Carter, M. F., The Attitude of the Thermosphere Probe, University of Michigan Scientific Report 07065-4-S, April, 1968. Niemann, H. B., and Kennedy, B. C., "An Omegatron Mass Spectrometer for Partial Pressure Measurements in Upper Atmosphere," Review of Scientific Instruments, 37, No. 6, 722, 1966. Parker, L. T., Jr., A Mass Point Trajectory Program for the DCD 1604 Computer, Technical Document Report AFSW-TDR-49, Air Force Special Weapons Center, Kirtland Air Force Base, New Mexico, August, 1962. Spencer, N. W., Brace, L. H., and Carignan, G. R., "Electron Temperature Evidence for Nonthermal Equilibrium in the Ionosphere," Journal of Geophysical Research, 67, 151-175, 1962. Spencer, N. W., Brace, L. H., Carignan, G. R., Taeusch, D. R., and Niemann, H. B., "Electron and Molecular Nitrogen Temperature and Density in the Thermosphere," Journal of Geophysical Research, 70, 2665-2698, 1965. Spencer, N. W., Taeusch, D. R., and Carignan, G. R., N2 Temperature and Density Data for the 150 to 300 Km Region and Their Implications, Goddard Space Flight Center, NASA Technical Note X-620-66-5, December, 1965. Taeusch, D. R., Carignan, G. R., Niemann, H. B., and Nagy, A. F., The Thermosphere Probe Experiment, University of Michigan Rocket Report 07065-1-S, March, 1965. 40

APPENDIX DETERMINATION OF THE TOTAL PAYLOAD MOMENTS OF INERTIA 41

SYSTEMS TEST DEPARTMENT 4 X REPORT NO. TR- 2240 LR NO. 3099 TEST REPORT REPORT NO. TR- 2240 DATE 6 Sep 67 PERFORMED FOR: The University of Michigan 2455 Hayward Ann Arbor, Michigan TEST: Sinusoidal - Random Vibration and Moment of Inertia Determination ITEM: Nike - Tomahawk Payloads, Ser. Nos. 18. 49 and 18. 50 Thermosphere Probe, Ser. No. 18. 50 TEST DATE: 6 Sep 67 PERFORMED AT:' PERFORMED AT Bendix Aerospace Systems Division WORK ORDER NO: 86661 -441 -01-3099 AUTHORIZATION; P.. R-79473 REQUESTED BY: J Marer REPORT SENT TO: J. Maurer PERFORMED BY _____ L. M. Skjei, Section Engineer APPROVED BY:>'' 1-4 R. H. Culpepper, Project Engineer jk 440-6A 42

Bendix Aerospace Systems Division TR 2240 REFERENCE (a) Nike-Tomahawk Vibration Test Specification, dated 22 Apr66, Rev. 6 Dec 66 (copy attached as Appendix A). INTRODUCTION Two Nike-Tomahawk Payloads, Ser. Nos. l8.49 and 18. 50,were vibrated along each of three orthogonal axes according to Reference (a), except as described in Table 1. The purpose of the tests was to determine if the test items could withstand the vibration environment specified. The mass moments of inertia of the Thermosphere Probe 18. 50 were determined experimentally on the trifilar test stand. The purpose of the tests was to determine the mass constants about the roll (X) and pitch (Y) axis. SUMMARY OF RESULTS Visual examination of the test items following the test disclosed no apparent structural degradation as a result of the applied vibration. The University of Michigan personnel, who performed the functional check of the test item during and after vibration, reported satisfactory operation. The moments of inertia of the Thermosphere Probe Assembly S/N 18. 50 are shown below. 2 lb ft sec Roll Axis (X) 0. 0906 Pitch Axis (Y) 3. 1201 METHODS AND DATA Vibration The test items were tested individually. Each test item was attached to a fixture provided by The University of Michigan and mounted on the vibration exciter for vibration in the Z axis and on the slip plate for vibration 45

Bendix Aerospace Systems Division TR 2240 in the X and Y axes. The test setups are shown in Figure 1. An accelerometer, mounted on the fixture, was used to control the vibration input to the test item. For the sine tests, the accelerometer was the detector for a constant displacement/acceleration servo system. The output of this accelerometer was measured with a true rms voltmeter and recorded on an X-Y recorder. These recordings are shown in Figures 2 through 11. The sensitivity of the accelerometer system was checked prior to the test by subjecting the accelerometer to a sinusoidal vibration of 0. 25 in. double amplitude at a frequency of 44 cps, to give an acceleration of 25 g peak. Vibration displacement was measured with an optical wedge, frequency was measured with an electronic counter, and accelerometer system output was monitored with a true rms voltmeter. Calibration dates are shown on the attached list of test equipment. Testing of each test item was performed in accordance with the requirements as shown in Appendix A. The University of Michigan personnel functionally operated the test items during and after each axis of vibration. For the random vibration portion of the test (Tests 3, 6, 8 and 9 of Appendix A) the vibration system was equalized to the required spectrum shape without the test item mounted. Prior to the random vibration test, with the test item mounted the spectrum equalization was touched up at approximately half level. This required from 10 to 35 seconds of vibration. The random spectrum was checked using the ASD-40 spectral density analyzer. Moment of Inertia Determination The test items were mounted on the trifilar pendulum apparatus as shown in Figure 12 and the platform was allowed to oscillate through approximately 1 inch. The period of oscillation of the combined test item and platform was determined. At the conclusion of testing the period of oscillation of the platform alone was determined. 44

Bendix Aerospace Systems Division TR 2240 I ~ I -I test item - combined test item platform alone or and platform wtap 2t - wpap2p 2 2 4 TI L 4 IT2L Where: Wt = Platform plus test item weight a =20 inches L = Filament length, 107. 88 inches w = Platform weight, 21. 05 lbs P t = Period in seconds, combined test item and platform p - Platform period in seconds, 1. 52114 P I - Test item moment of inertia in lb in sec2 Testing was witnessed by The University of Michigan personnel: J. Maurer, P. Freed, B. Halpin, D. Phillips, M. Street, J. Vidolch, D. Haseltine, R. Pate and S. Dietrick who returned test item to The University of Michigan. 45

0~~~~~~~~~~~~~~~~~~~~~~~~> -.,"'~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~..... 0~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~r MM XM. i.......... N, ~~~~~~~~*~~~~~~~~~~~~~*% -~ ~ ~ ~ ~ ~ ~ ~ ~ ~~~~~~~~~- ----- MIN......................................... OF~* 110.l..... RIM.~~~~~~~~~~~~~~~~~ zu~~~~~~~~~~~~~~~~~~~~

evtslmmvd-111 TIR 2240 SYSTEMS TES DEPARTMENT Figure 12 TEST SET Ui Moment of Inertia Determination Pi.tch AxiS (Y) 440-40 A 4( PR1.:r~~~~~~~~~~~~~~~~~~aui.2:r-:~~~~~~~~~~~~~~~~~rziin El -111 1

I 89,d J Systems Division... J Ann Arbor MichiganR 2240 SYSTEMS TEST DEPARTMENT TEST EQUIPMENT Test: S-R Vibration and Moment of Inertia Date Used: 6 Sep 67 Test Item: Nike-Tomahawk Payloads 18.49 and 18. 50, Thermosphere Probe 18. 50 Calibration Item Model No. BxS No. Scale Range Calibration._._._ _.._____....... -_ _ _ _Date Manufacturer Serial No. Accuracy puantity MeasurecLast Next Vibration Systemn PP /175-240 249..Vib o ysem _ — PP/ 175 - 240 —- - - - - - N/A N/A Ling 13 30... True RMS Voltmeter_ _ _2409 __ 1_ 1237_ 0-300Qm _ _ 3/21 9/21 Bruel and Kjaer 72900 2 ___ __67 67 Oscillator... _1018_ ______ 50567 5-2000 _ _ 6/6 12/6 Bruel and Kjaer 70666 Hz 67 67 Logarithmic Converter_ _60B 11220 10 dbJi_ _ _ _ 8/14 11/14 F. T.. Moseelpy _ |840.___...67 67 X Y Recorder_ _2D_ _ _ _ 10010 1 mvLin _ _ 7/15 10/15 F. L. Moseley 110 2% g___67 67 FilterAnalyzer _. _ ASD 40_ _ _ _ 50417 _0-1._ _ 4/10 10/10 Ling 60... 2% lg2/cps |6 7 Accelerometer _ 2221. 6/29 12/29 Endevco FC 87 _67 67 SlipESync _103 AR_ 921 5 5/11 11/11 Chadwick Helmuth 1350 _67 67 Trifilar Stand 5/8 11/8 BSD 13239B. 67 67 ~_________ ______,__~ ____. —— ___

UNIVERSITY OF MICHIGAN 3 9011111 1 111 0324 4360 3 9015 03524 4360