THE UN I V E R S I T Y OF M I C H I G A N COLLEGE OF ENGINEERING Department of Electrical Engineering Space Physics Research Laboratory Sounding Rocket Flight Report NASA 18.104 AND NASA 18.105 THERMOSPHERE PROBE EXPERIMENTS Prepared on behalf of the project by H. J. Grassl ORA Project 027700 under contract with: NATIONAL AERONAUTICS AND SPACE ADMINISTRATION GODDARD SPACE FLIGHT CENTER CONTRACT NO. NAS 5-21038 GREENBELT, MARYLAND administered through: OFFICE OF RESEARCH ADMINISTRATION ANN ARBOR February 1971

TABLE OF CONTENTS Page LIST OF ILLUSTRATIONS iv 1. INTRODUCTION 1 2. GENERAL FLIGHT INFORMATION 2 3. LAUNCH VEHICLE 4 4. NOSE CONE 7 5. THE THERMOSPHERE PROBE (TP) 5.1. Omegatron 10 5.2. Electron Temperature and Density Probe 17 5.3. Support Measurements and Instrumentation 5.3.1. Aspect determination system 20 5.3.2. Telemetry 23 5.3.3. Housekeeping monitors 24 6. ANALYSIS OF DATA 25 6.1. Trajectory and Aspect 25 6.2. Ambient N2 Density 28 6.3. Temperature 38 6.4. Geophysical Indices 38 7. REFERENCES 43 iii

LIST OF ILLUSTRATIONS Table Page I. Table of Events 3 II. Omegatron Data 12 III. N2 Ambient Density Data 28 Figure 1. Nike-Tomahawk with thermosphere probe payload. 5 2. Nike-Tomahawk dimensions. 6 3. Thermosphere probe instrumentation design. 8 4. Assembly drawing, 8-in. nose cone. 9 5. Thermosphere probe system block diagram. 11 6. Omegatron II. 14 7. Final calibration of the NASA 18. 104 omegatron. 15 8. Final calibration of the NASA 18. 105 omegatron. 16 9. Electron temperature and density probe. 18 10. ETDP system timing and output format. 19 11. NASA 18.104 minimum angle of attack vs. altitude. 21 12. NASA 18. 105 minimum angle of attack vs. altitude. 22 13. NASA 18. 104 sequence of events. 26 14. NASA 18.105 sequence of events. 27 15. NASA 18. 104 omegatron current vs. flight time. 30 16. NASA 18. 105 omegatron current vs. flight time. 31 iv

LIST OF ILLUSTRATIONS (Concluded) Figure Page 17. K (So,a) vs. altitude for NASA 18.104. 32 18. K (So,a) vs. altitude for NASA 18. 105. 33 19. NASA 18.104 ambient N2 density vs. altitude. 34 20. NASA 18.105 ambient N2 density vs. altitude. 35 21. NASA 18. 104 neutral particle temperature vs. altitude. 39 22. NASA 18. 105 neutral particle temperature vs. altitude. 40 25. Solar flux at 10.7 cm wavelength. 41 24. Three-hour geomagnetic activity index (a ). 42 p

1. INTRODUCTION The results of NASA 18. 104 and NASA 18. 105, identical Nike-Tomahawk sounding rockets launched during the solar eclipse of March 1970, are presented and discussed in this report. The payload for each flight, 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 Planetary Atmospheres. 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. Each payload included a "second generation" omegatron mass analyzer, an ion spectrometer, an electron temperature probe (Spencer, and Carignan, 1962), and a solar position sensor. This complement of instruments permitted the determination of the molecular nitrogen density and temperature and the charged particle density and temperature in the altitude range of approximately 150 to 290 km over Wallops Island, Virginia, during 40% (NASA 18.104) and 80% (NASA 18.105) solar obscuration. A general description of the payload kinematics, orientation analysis, and the technique for the reduction and analysis of the data is given by Taeusch, Carignan, Niemann, and Nagy (1965) and Carter (1968). The f(s) curve fitting technique for reduction of the omegatron data is not described here but will be presented in a future report, currently in preparation. The orientation analysis and the reduction of the nitrogen data were performed at SPRL, and the results are included in this report. The ion spectrometer data and the electron temperature probe data were reduced at GSFC, and are not discussed here. 1

2. GENERAL FLIGHT INFORMATION The general flight information for NASA 18.104 and NASA 18.105 is listed below. Table I gives the flight times and altitudes of significant events which occurred during the flights. Some of these were estimated and are so marked. The others were obtained from the telemetry records and radar trajectory information. Flight: NASA 18. 104 NASA 18. 105 Launch Date: 07 March 1970 07 March 1970 Launch Time: 18:00:00. 079 GMT 18:27:00. 073 GMT Location: Wallops Island, Virginia Wallops Island, Virginia Latitude:'37 50' 14.915"N Latitude: 3750' 14. 915"N Longitude: 75029'01.693'W Longitude: 75029'01.693"W Apogee Parameters: Altitude: 289.32 km 290.21 km Horizontal Velocity: 277.2 m/sec 343.8 m/sec Flight Time: 265. 5 sec 266.0 sec TP Motion: Tumble Period: 11.158 sec 2.451 sec Roll Rate: 47.14 deg/sec -147 deg/sec 2

TABLE I TABLE OF EVENTS NASA 18. 104 NASA 18.105 Event Flight Time Altitude Flight Time Altitude (sec) (km) (sec) (km) Lift-off 0 0 0 0 1st Stage Burnout 3.6 1.4 3.9 1.6 2nd Stage Ignition 12.0 6.6 12.1 6.7 2nd Stage Burnout 21.8 20.2 21.5 19.5 Despin ---- --- 43.6 (est.) 66.9 (est.) TP Ejection 45.8 71.3 45.6 70. 9 Omegatron Breakoff 78.3 131.5 78.7 132.3 Omegatron Filament On 80.8 135.7 79.9 134.3 Peak Altitude 265.5 289.32 266.0 290.21 L.O.S. 502.0 498.0 3

3. LAUNCH VEHICLE The launch vehicles for NASA 18.104 and NASA 18.105 were two-stage, solid propellant Nike-Tomahawk combinations. The first stage of each vehicle, 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, 141.4 in. long and 9 in. in diameter, weighed 530 lb unburned. Each TP payload, which was 90. 5 in. long and weighed 175 lb including despin and adapter modules, made the total vehicle 377.1 in. long with a gross lift-off weight of 2030 lb. The vehicle is illustrated in Figures 1 and 2. Both launch vehicles performed flawlessly. NASA 18.104 reached a summit altitude of 289.32 km at 265.5 sec of flight time, and NASA 18.105 reached a summit altitude of 290.21 at 266.0 sec of flight time. 4

~~~~~ 3 Figure 1. Nike-Tomahawk with thermosphere probe payload. 5

ROCKET NO 18.104 8 18.105 8.0 DIA. PAYLOAD 90.531 ("9- 0., 531I -FIRING a DESPIN SECOND STAGE TOMAHAWK 141.1 376.831 9 DI A. 36.6 - kFRFIRST STAGE NIKE BOOSTER 145.2 ORDNANCE ITEMS 16.5 DIA. O NOSE CONE OPENING PRIMERS {O BREAKOFF LINEAR ACTUATORS O DESPIN INITIATION PRIMERS - -- ^ (.4|9 SECOND STAGE IGNITER -— 59936 - (- NIKE BOOSTER IGNITER Figure 2. Nike-Tomahawk dimensions. 6

4. NOSE CONE A diagram of the payload for both NASA 18. 104 and NASA 18. 105 including the nose cone, the despin mechanism, and the adapter section is shown in Figure 3. An assembly drawing of the 8-in. nose cone is given in Figure 4. The despin mechanism on NASA 18. 104 apparently did not work. Ejection began at 71 km (46 sec after launch), and the resulting tumble period was 11.158 sec. The omegatron breakoff device was removed at 132 km (78 sec after launch), and the omegatron filament was turned on approximately 2 sec later. The NASA 18.105 payload was despun at 67 km (44 sec after launch), and the ejection began at 71 km (46 sec after launch). The resulting tumble period of the payload was 2. 451 sec. The omegatron breakoff device was removed at 152 km (79 sec after launch), and the omegatron filament was turned on approximately 2 sec later. 7

15~ INC 9.250 4 --- FRANGIBLE RING 28.562 CONE ASS'Y rr n | r! *r- --- |NOSE CONE TIMER 8.00 DIA. _ I1.781 RE LJ --- BREAKOFF (CERAMIC) 1 —-.78., T H THERMOSPHERE PROBE T! I- a. J ------ OMEGATRON OMEG ADAPTOR OMEG. AMP ~1~~.'C — ~ -- MP. DECK 22.796 _ -- -- OSC. LOOP DECK r- --- -- OSC. LOGIC DECK _ _ ~_" ------ REG. DECK ~. _I —---- POWER SUPPLY DECK 90.531,_ ___ DATA DECK _-~- ~ CONTROL DECK COMM. DECK 39.812 S.C.O. DECK E.SRP DECK * ----- XMITTER AND r —-. —--- RBATTERY DECK 45 469 3 LSOLAR bENbOR (ADCOLEJ 3.50 45.469 13.141 ION SPEC. SECTION 1-I 1 1_ ~ L^ " — HOUSING TUBE 3.895 -- UMBILICAL PLUG 2.00 -- BASE 5.125 1 -- AFT. SECTION 2.750 _ RADAR BEACON f- TOMAHAWK IGNITION AND ~~6.~~~625 ~DESPIN CIRCUITS 6.625 DESPIN K-9.00 DIA. - Figure 3. Thermosphere probe instrumentation design. 8

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UMBILICAL CABLE CONNECTIONS VECTOR MO SCOS VECTOR MMO SCO'S po_ 02 po3 __po4 __ po1 a CALIBRATOR /,' ---------------------------------------------— ~yJ2~ -------- ^.?~~ 70 KHZ +28PW OM3L SYSTEM ALL REGULATORS, REFERENCES, POWER CONVERTORS, BIAS SUPPLIES, _____________OUT/D54_____ _ 5 KHZ 8 OSCILLATORS RNG SH ledex ---— J control IRPS IRIG 30SEGMENT COMMUTATOR I ------ * ------ I ^ ^^ -------------------— 1 ^^. ft t i 9^ a l/*AI COMMUTAT) CONTROL LEDEX I N\V 0,1,2,3,4,5 CAL- -- 7.35KHZ +28PW HOUSEKEECPN IBRATE VOLTAGES m nitor voltage monitors ION SPECTROMETER SYSTEM IS DATA 24 KHZ amp control IS ANALO 4o KHZ IS ANALOG CH 30 KHZ,control 22 KHZ c r filament control 145KHZ 3.0 KHZ ------- ^^>>^^. ~~~~~~~~~~~~~~~~~~~~VECTOR POWER CONTROL ESP_ D2 ext power ELECTROSTATIC PROBE 525 KHZ (2 PROBES) 240. mhz +Z8PW LEDEX 8 RELAY ^J 791R9 HR-I + 28 PW YARDNEY ANTENNA SILVERCELLS ASPECT APPROX. —8V ~PWSPRL/ADCOLE SOLAR ASPECT 10.5 KHZ APPROX. 28V +2iPW +28PW BACKUP + 28PW POWER L RAYMOND oG" WOEIGHT *CALIBRATION O,5 0lOms,EVERY TIMER 20 SECONDS ON CHANNELS SHOWN ON OM FILAMENTS BREAKOFF DEVICE Figure 5. Thermosphere probe system block diagram.

TABLE II OMEGATRON DATA (NASA 18.104) Calibration Normalized N2 Sensitivity: 1.64 x 10 5A/torr Electrometer Amplifier OUT/S Range Range Indicator Resistor Gain Bias 12 1-1 0.54 V 1.000 x 1012 -1.98 + 1.42 V 1-2 0.83 V 1.000 x 10 -1.98 - 2.56 V 12 1-3 1.13 V 1.000 x 10 -0.99 - 3.03 V 1-4 1.43 V 1.000 x 1012 -0.99 - 7.00 V 12 1-5 1-73 V 1.000 x 10 -0.99 -10.97 V 1-6 2.03 V 1.000 x 10 -0.99 -14.94 V 1-7 2.33 V 1.000 x 10 -0.99 -18.89 V 2-1 2.81 V 6.664 x 1010 -1.98 + 1.42 V 2-2 3.13 V 6.664 x 10 -1.98 - 2.56 V 1010 2-3 3.45 V 6.664 x 10O -0.99 - 3.03 V 2-4 3.76 V 6.664 x 10 -0.99 - 7.00 V 2-5 4.07 V 6.664 x 1010 -0.99 -10.97 V 2-6 010 2-6 4.39 V 6.664 x 10 -0.99 -14.94 V 2-7 4.70 V 6.664 x 010 -0.99 -18.89 V OUT/D Range Range Indicator Resistor Gain Bias 12 1 1.000 x 101 -0.2486 - 0. o059 V 2 6.664 x 10 -0.2486 - 0.0059 V 12

TABLE II (Concluded) (NASA 18.105) Calibration Normalized N Sensitivity: 1.68 x 10 5 A/torr Electrometer Amplifier OUT/S Range Range Indicator Resistor Gain Bias 12 1-1 0.55 V 1.000 x 1012 -1.99 + 1.39 V 1-2 0.85 V 1.000 x 10 -1.99 - 2.61 V 1-3 1.14 V 1.000 x 102 -0.99 - 3.07 V 12 1-4 1.44 v 1.000 x 102 -. 99 - 7.05 v 1-5 1.74 v 1.000 x 1012 -0.99 -11.04 V 12 1-6 2.04 v 1.000 x 10 -0.99 -15.03 v 1-7 2.34 V 1.000 x 10 -0.99 -18.99 V 10 2-1 2.82 V 6.664 x 101 -1.99 + 1.39 V 2-2 3.14 V 6.664 x 101 -1.99 - 2.61 V 10 2-3 3.46 v 6.664 x 10 -0.99 - 3.07 V 10 2-4 5.77 V 6.664 x 010 -0.99 - 7.05 V 2-5 4.08 V 6.664 x 101 -0.99 -11.04 V 2-6 4.40 v 6.664 x 10 -0.99 -15.03 V 2-7 4.71 V 6.664 x 10 -0.99 -18.99 v OUT/D Range Range Indicator Resistor Gain Bias 12 1 _1.000 x 10 -0.2487 - 0.0075 V 2 6.664 x 10 -0.2487 - 0.0075 V 13

FEED THRU PINCH OFF TUBE BREAKOFF ASSEMBLY ORIFICE ORIFICE ACTUATOR ANTECHAMBER _ --- VACUUM SEAL ENVELOPE OMEGATRON g MAGNET OMEGATRON ADAPTER ----- VACUUM SEAL PROBE HOUSING ELECTROMETER AMPLIFIER DECK EMISSION REGULATOR R BIAS CONTROL ___ __ B0QS9 1i -SOSCILLATOR DECK OSCILLATOR _ CONTROL DECK Lu 6. - LOGIC DECK Figure 6. Omegatron II. 14

Ilo,.'z. II Oi ~ F I I || 2.07x10 | PART/CC MP 0 -1 Z 10 - 2 1 9 ll 102 1013 i d NORMALIZED N2 SENS11IIVITY - NUMBER DENSITY (PART/CC) Figure 7. Final calibration of the NASA 18. 104 omegatron. Figure 7. Final calibration of the NASA 18. 104 omegatron.

_- I" I I IIH 1 i I [ [ I' 1 [ [ ]' I I I l' 1 fill, I I I I III I I I I fi ll I I - 0 I0 F z -- r^1 3o 10 01 0 _NASA l5 C 0,\ NORMALIZ D N2 SENSI IVITY (3 |_ 2.02 xl0 PART/CC MP w -1 00 - __ — Zio44~~~~ — J. 2I.IU_ |i 1 I "1111 | ~I| || l 1 11 i i - i1111111 10 L i i LLLIL~_ J.~..II0 1(J 108 109 do0 101012I 1013 NUMBER DENSITY (PART/CC) Figure 8. Final calibration of the NASA 18.105 omegatron.

5.2. ELECTRON TEMPERATURE AND DENSITY PROBE The electron temperature and density probe consists of two cylindrical Langmuir probes placed in the plasma, and an electronics unit which measures the current collected by the probes as they are swept through a series of ramp voltages. A typical Langmuir probe is shown in Figure 9. For both flights probe 1 is stainless steel and probe 2 is rhodium-plated stainless steel. Each electronics unit consisted of a dc-dc converter, the AV ramp generator, a three range current detector, and associated logic and control circuits. Timing and sequencing of the various functions are shown in Figure 10. The pertinent system parameters follow. NASA 18.104 NASA 18.105 (a) Input Power 2.2 W at 28 V 2.2 W at 28 V (b) Sensitivity Range 1 4.0 pA full scale (5 V) 4.0 pA full scale (5 V) Range 2 0.4 pA full scale (5 V) 0.4 pA full scale (5 V) Range 3 0.05 pA full scale (5 V) 0.05 pA full scale (5 V) (c) Ramp Voltage (AV) High AV 80.0 V/sec 80.0 V/sec Low AV 24.1 V/sec 24.0 V/sec Period 124.7 msec 125.1 msec (d) Output Voltage -0.59 V to +5.65 V -0.60 V to +6.01 V Resistance 2700 0 2700 ) Bias Level 1.01 V 1.02 V (3) System Calibration Calibration occurs every 31.5 sec for NASA 18.104 and 27.0 sec for NASA 18.105 for a duration of 750 msec. 17

-- 125m SEC ~ 1.5 f 1.5 NOMINAL FLAG VOLTAGES h* —----- z2.250 SEC.AV FLAG -- II —------------------------— I 2 AV FLAG HI AV SLOPE=80V/SEC LO AV SLOPE 24 V/SEC OV 0.2 UA K4 ON DETECTOR 2 UA __ _ --- RANGE 20 UA K3 ON________.750 M.SEC — ~~~I-'~ ~CALIBRATION OCCURS ONCE EVERY 31.5 SEC AT BEGINNING OF HI AV MEASUREMENT 00 CALIBRATION 125ms (SIG. OUT) h —--- 750 —---- "I AV - LO AV CALIBRATION AND COM MODE CHECK 4.5 V 5.0 V (J6- 3) I. 5V ~~bv uowiro r r -5 AV MONITOR fv -1.5 V (33 Ka IMPEDANCE) -5.0V Figure 9. Electron temperature and density probe.

bV -I 125m SEC 1- 1.5 APP. - 1.5 1| —, -2~~2.250 SEC. -! +2V AV FLAG HI AV SLOPE=8OV/SEC LO AV SLOPE=24V/SEC OV.05 UA K4 ON DETECTOR 0.4UA _ RANGE 41 A K3 ON.750 M.SEC CALIRATION OCCURS ONCE EVERY I.5SEC AT BEGINNING OF HI AV MEASUREMENT (ONCE EVERY 27.0 SEC FOR 18.105) CALIeRATION 1251ms (IG. OUT) - I - s 750 HI AV - - L LO AV CA,IBRATION AU, CO t MODE CHECK (J6-3) 1.6 V p P2/ P P2/ P P2/ P P2 PI P2/ P P2 PI P2 PI P2 PI P2 PI P2 PI P2 PI P2 PROSE AND AV O V- j J J J__ / I 1.2 V MONITOR CALIBRATION APPLIES DURING CALIBRATION PERIOD ONLY COMMON MODE CHECK Figure 10. ETDP system timing and output format.

5.3.1. Aspect Determination System The NASA 18.104 and NASA 18.105 TPs used Adcole sensors identical to those used on previous flights. However, each flight used a single-eye (120 deg field of view) system rather than the triple-eye (360 deg field of view) system of previous flights. Also, the accompanying electronics were adjusted at SPRL to account for the solar obscuration encountered. The 120 deg field of view proved sufficient to determine the NASA 18.104 aspect. 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, determined to an estimated accuracy of +5 deg, is plotted versus altitude in Figure 11. Due to an electronics problem the NASA 18.105 sensor failed to function during the flight. However, information obtained from the ion spectrometer (which "saw" the sun) was sufficient to establish the initial parameters necessary for standard reduction of the TP attitude. The resulting minimum angle of attack, determined to an estimated accuracy of +5 deg, is plotted versus altitude in Figure 12. 20

320 NASA 18.104 07 MARCH 1970 18:00 GMT 300- WALLOPS IS., VA. 280 260 Z240 I II 140 200_AG OF ATC -2 140 ANGLE OF ATTACK (DEG) 21

320 NASA 18.105 07 MARCH 1970 18:27 GMT 300 WALLOPS IS., VA. 280 260 - -- F 220 200 I UPLEG DOWNLEG 180 - 4 160 140 0 10 20 30 40 50 60 70 80 90 ANGLE OF ATTACK (DEG) Figure 12. NASA 18. 105 minimum angle of attack vs. altitude. 22

5.3.2. Telemetry For each flight, the payload data were transmitted in real time by twelve channel PAM/FM/FM telemetry system at 240.2 MHz with a nominal output of 2.5 W. The telemetry system used twelve subcarrier channels, as outlined below. NASA 18.104 NASA 18.105 Transmitter: TRPT-251-1RBO TRPT-250RAO (Serial No. 842) (Serial No. 2516) Power Amplifier: TRFP-2V TRFP-2V (Serial No. 144) (Serial No. 450) Mixer Amplifier: MMA-12 IMA-12 (Serial No. 11898) (Serial No. 652) Subcarrier Channels: MMO-11 MMO-11.Serial No.. IRIG - Serial N Center Low Pass NASA NASA Function Band 18.104 18.105requency Filter Used 18.104 18.105 20 8356 7510 124 kHz IS Digital 2500 Hz CA 18 16666 16673 70 kHz OM OUT/S 450 Hz CA 17 19232 20386 52.5 kHz ESP/D 790 Hz CD 16 7831 404 40 kHz IS Analog 1 600 Hz CA 15 606 15015 30 kHz IS Analog 2 450 Hz CA 14 15163 9052 22 kHz IS Analog 3 350 Hz CA 13 15270 15278 14.5 kHz IS Ramp 220 Hz CA 12 18742 12583 10.5 kHz Aspect 330 Hz CA 11 15902 15997 7.35kHz Commutator 120 Hz CD 10 7808 15428 5.4 kHz OM OUT/D 80 Hz CD 9 15777 15728 3.9 kHz OM Range 60 Hz CD 8 674 15667 3.0 kHz IS V Monitor 45 Hz CD s Instrumentation power requirements for each flight totaled approximately 35 W, supplied by a Yardney HR-1 Silvercell battery pack of a nominal 28 V output. 25

5.3.3. Housekeeping Monitors Outputs from various monitors throughout the instrumentation provided information bearing on the operations of the electronic components during the flights. These outputs were fed to a thirty-segment commutator which ran at one rps. The commutator assignments were as follows: COMMUTATOR FORMAT FOR NASA 18.104 AND NASA 18.105 Segment Segment Number Assignment 1 OUT/D 2 OUT/S Range 3 RF Frequency 4 RF Amplitude 5 Automatic Frequency Control Lock 6 DC Frequency Control 7 Beam Current 8 Filament Voltage 9 Internal Pressure Monitor 10 Thermistor - Gauge Temperature 11 Thermistor - Amplifier Temperature 12 Thermistor - Transmitter Temperature 15 Battery Voltage Monitor 14 +15 Power Supply Voltage 15 RF Voltage Monitor 16 G15 Monitor 17 Sweep Voltage Monitor 18 oThermistor 1 19 Spectrometer Thermistor 2 20 +10 Power Supply Voltage 21 RF Voltage Monitor 22 +3 Power Supply Voltage 23 Sweep Voltage Monitor 24 0 V Calibration 25 1 V Calibration 26 2 V Calibration 27 3 V Calibration 28 4 V Calibration 29 5 V Calibration 30 5 V Calibration 24

6. 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. 6.1. TRAJECTORY AID ASPECT The position and velocity data used to determine aspect, ambient N density, and ambient temperature as a function of time and altitude were obtained by fitting a smooth theoretical trajectory to the FPQ-6 radar data. The theoretical trajectory is programmed for computer solution similar to that described by Parker (1962). The analysis of minimum angle of attack (cmin) as described by Carter (1968) is also incorporated in the program. The output of the computer furnishes amin, altitude, and velocity as a function of time. Plots of cmin versus altitude have already been given in Figures 11 and ]2. Figures 13 and 14 show the occurrence of significant events during the flights. 25

ALTITUDE (KM) 0 0 0 0 0 0 0 0 ol ~ O ~ O ~ O g O 0 NIKE BURNOUT 3.6 TOMAHAWK IGNITION 12.0 TOMAHAWK BURNOUT 21.8 TP EJECTION 45.8 e3 ~"oALXI'N~T s8, 3,o., hij D - 0 rTrI' I I CD 0ro 0 Ot0 1 PEAK 260.5 n4 z 0 — L.O.S. 502.0 0 O 0 Oj CD 0 (^I 0 3 05 0tV

ALTITUDE (KM) O o 00 0 0 0 ~ ~~~0 0 0 0o n ~~ S -- NIKE BURNOUT 3.9 0 ------ ~ TOMAHAWK IGNITION 12.1 TOMAHAWK BURNOUT 21.5 DESPIN 43.6 TP EJECTION 45.6 O0 O' lOMKAM[ENT7dN7 79.9 CD H I ~ - - D P 0 0 -? z (D C) 01 / L.O.S. 498.0 0 03 O0) 0 0

6.2. 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) a2 17 V cos a. min where n = ambient N number density a 2 An. = maximum minus minimum gauge number density during one tumble, A x AI, where A is the sensitivity of the gauge u, = J2KT./m, most probable thermal speed of particles inside gauge 1 1 T. = gauge wall temperature 1 V = vehicle velocity with respect to the earth = minimum angle of attack for one tumble min K(S,A) = the reciprocal of the normalized transmission probability as defined by Ballance (1967), referred to as the geometry correction factor. AI, the difference between the maximum (peak) omegatron gauge current and the minimum (background) gauge current versus flight time is shown in Figures 15 and 16. 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 28

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 amin depends upon amin, for any given uncertainty in amin, each particular case and altitude range must be considered separately. Figures 11 and 12 show that the minimum angle of attack for the upleg is generally less than 20 degrees, so with an assumed maximum uncertainty in aCin of +5 degrees, the resulting uncertainty in cos Cmin is less than +3%. The data for low angle of attack were used as control data. K(So,a), the geometry correction factor versus altitude, is shown in Figures 17 and 18. As can be seen, the maximum correction is about 6%, or K(So,) =.94 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 Figures 19 and 20 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 10o relative and 25% absolute. 29

NASA 18.104 07 MARCH 1970 18:00 GMT _.WALLOPS IS., VA. 100 oliO C3 V. I0\ 10.(PK - BKG) I0 WE | X z X* (PKX 00 X * -. X X 0 100 200 300 400 500 600 FLIGHT TIME (SEc) Figure 15. NASA 18.104 omegatron current vs. flight time. 3o X *x 50

l- l I I I l NASA 18.105 07 MARCH 1970 - * 18:27 GMT _* WALLOPS IS., VA. * ~ I ~: _; (.PEAK- BKG) _ 5 * " * ~ F * * z 9. ( -Z I 4 (PEAK-BKG) * Cz _ * S SS~ LSJ _ 8 x 100~~~~~~~~~~x _ S 10. L;l|*lli ~ Z ~ *; 0 103 Z ~ X LS X X i Xx xxxx.x xx** xx ~' 8 10 z o.. 0 I00 200:00 400 500 600 FLIGHT TIME (SEC) 31,o7 —--------------------------- 0 100 200 30 400 500 60

320 NASA 18.104 07 MARCH 1970 18:00 GMT 300 WALLOPS IS., VA. 280 260 240 220 200 DOWNLEG / UPLEG 180 160 140.94.95.96.97.98.99 1.00 GEOMETRY CORRECTION FACTOR Figure 17. K (S,c) vs. altitude for NASA 18. 104. 32

320 NASA 18.105 07 MARCH 1970 18:27 GMT 300 - WALLOPS IS., VA. 280 260 240, 220 200 UPLEG / DOWNLEG 180 160 140.94.95.96.97.98.99 1.00 GEOMETRY CORRECTION FACTOR Figure 18. K (S,a) vs. altitude for NASA 18. 105. 33

NASA 18.104 340 07 MARCH 1970 18:00 GMT |~~~~~~~~~~~~320 ~WALLOPS IS., VA. 300 280 260 Y240 w fw220 _20 180 160 140 120 __. I, _1_I!!.j 1._ 1 1 11 111I 1 I i I I I IIL 107 108 109 1010 AMBIENT N2 DENSITY (PART/CC) Figure 19. NASA 18. 104 ambient N density vs. altitude. 2

NASA 18.105 07 MARCH 1970 340 18:27 GMT WALLOPS IS.,VA. 320 300 280 260 "240 "220 ~_200 \J1 180 160 140 120 III11,. 1 __1; I I.I Llal I 107 108 109 10'~ AMBIENT N2 DENSITY (PART/CC) Figure 20. NASA 18.105 ambient N density vs. altitude. 2

TABLE III N2 AMBIENT DENSITY DATA NASA 18.104 7 March 1970 18:00 GMT 13:00 EST Wallops Island, Virginia Altitude Temperature Density (km)i (OK) (part/cc) 150 712 5.05 x 101 155 749 2.34 160 785 1.82 165 820 1.43 10 170 853 1. 1 x 10 175 885 9.17 x 10 180 918 7.44 185 945 6.11 190 971 5.05 195 993 4. 21 200 1013 5. 54 205 1029 2.99 210 1044 2. 55 215 1056 2.16 220 1o68 1.85 225 1078 1.59 230 1087 1.36 235 1095 1.18 240 1105 1.02 x 1l0 245 1110 8.80 x 10 250 1116 7.63 255 1122 6.62 260 112P 5.75 265 1133 5. 00 270 1138 4.36 275 1143 3.80 280 1148 3.31 285 1154 2.89 8 289 1157 2.60 x 10 Fit Parameters: T: = 1185 ~K To = 701 oK at 150 km Pb = 5.0 x 10-8 torr a = 2.00 x 10-2 36

TABLE III (Concluded) NASA 18. 105 7 March 1970 18:27 GMT 153:27 EST Wallops Island, Virginia Altitude Temperature Density (km) (~K) (part/cc) 145 676 4. 08 x 1010 150 712 3.09 155 746 2.57 160 781 1.84 165 814 1.45 10 170 846 1.16 x 10 175 881 9.50 x 10 180 914 7.54 185 944 6.16 190 972 5.07 195 996 4.21 200 1018 3.53 205 1056 2.98 210 1052 2.55 215 1066 2.15 220 1078 1.84 225 1088 1.58 230 1097 1.37 235 1105 1.18 240 1113 1.02 x 10 245 1119 8.86 x 10 250 1125 7.70 255 1130 6.69 260 1135 5.83 265 1159 5. 08 270 1145 4.43 275 1146 3.86 280 1149 3.38 285 1151 2.95 8 290 1155 2.58 x 10 Fit Parameters: T 1 = 1168 ~K To = 857 ~K at 170 km Pb =.09 x 10-28 torr a = 2.60 x 10 37

6.3. TEMPERATURE The ambient temperatures shown in Figures 21 and 22 and tabulated in Table III -were obtained by integrating the hydrostatic equation using the measured N2 density profile to obtain a partial pressure profile, and by relating the known density and pressure to the temperature through the ideal gas law. In this procedure the assumptions of hydrostatic equilibrium and perfect gas behavior are implicit. It can be shown that the density integral is stable and highly convergent when carried out in the direction of increasing density. The pressure or temperature at the initial (upper) boundary of integration is determined analytically by means of a least squares fitting procedure using a fitting function based on the empirical expression for the temperature profile given by Jacchia (1964), and more particularly by Walker (1965). The procedure is described in detail by Simmons (1969). The fit parameters listed in Table III are the apparent exospheric temperature (TL), the reference temperature at the lower boundary (To), the apparent N2 partial pressure at the upper boundary (Pb), and an estimate of the exponential model shape factor (a). 6.4. GEOPHYSICAL INDICES The 10.7 cm solar flux (F10 7) and the geomagnetic activity indices (ap) for the appropriate periods are shown in Figures 25 and 24. 38

NASA 18.104 07 MARCH 1970 18:00 GMT 300 - WALLOPS IS., VA. 280 260 240 wJ / 220 I-/ 200 180 160 140 700 800 900 1000 11.00 1200 TEMPERATURE (oK) Figure 21. NASA 18.104 neutral particle temperature vs. altitude. 39

NASA 18.105 07 MARCH 1970 18:27 GMT 300 WALLOPS IS., VA. 280 260 240 220 I-y/ 180 160 140 700 800 900 1000 1100 1200 TEMPERATURE (~K) Figure 22. NASA 18. 105 neutral particle temperature vs. altitude. 40

Fio.7 SOLAR FLUX VS. TIME DEC 1969-MARCH 1970 220 20 180 I 160 140 A 120 _ 0a0 z z I I _ _ _ _ _ _ _ _ DEC 1,69 JAN 1,70 FEB 1,70 MAR 1,70 1970 Figure 23. Solar flux at 10. 7 cm wavelength. 41

70 07 MARCH 1970 WALLOPS IS., VA. 60- Ap =42 50 ~~~~\ap =_I\ 0 201- \ -3 \ ap \0 L I I I I I I I I I I I< I I I0 0 2 4 6 8 10 12 14 16 18 20 22 24 2 4 42 z 6 5 0 2 4 6 8 10 12 14 16 18 20 22 24 2 4 MARCH 7, 1970 GMT TIME 42

7. 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 51, 1967. Carter, M. F., The Attitude of the Thermosphere Probe, University of Michigan Scientific Report 07065-4-S, April 1968. Jacchia, L. G., Static Diffusion Models of the Upper Atmosphere with Empirical Temperature Profiles, Research in Space Science, Smithsonian Astrophysical Observatory Special Report No. 170, 1964. 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. Simmons, R. W., NASA 18.49 Thermosphere Probe Experiment, University of Michigan Sounding Rocket Flight Report 07065-9-R, May 1969. 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, 26652698, 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. Walker, J. C.G., "Analytic Representation of Upper Atmosphere Densities Based on Jacchia's Static Diffusion Models," Journal of Atmospheric Sciences, 22, No. 4, 462-463, July 1965. 45

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