THE UN IV ER SIT Y OF MI CHI GAN COLLEGE OF ENGINEERING Department of Electrical Engineering Space Physics Research Laboratory Sounding Rocket Instrumentation and Flight Report NASA 18. 78 GA MODEL A PLANETARY MASS SPECTROMETER TEST FLIGHT Prepared on behalf of the project by D. F. rrsby D. L. Jones ORA Project 02681 under contract with: NATIONAL AERONAUTICS AND SPACE ADMINISTRATION GODDARD SPACE FLIGHT CENTER CONTRACT NO. NAS- 11128 GREENBELT, MARYLAND administered through: OFFICE OF RESEARCH ADMINISTRATION ANN ARBOR January 1970

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TABLE OF CONTENTS Page ACKNOWLEDGMENTS v LIST OF FIGURES vi 1. INTRODUCTI ON 1 2. GENERAL FLIGHT INFORMATION 2 3. LAUNCH VEHICLE 3 4. PAYLOAD 6 4.1. Nose Cone and Inlet System 12 4.2. Spectrometer Section 13 4.2.1. Mass spectrometer electronics and quadrupole analyzer tube 15 4. 2. 2. Breakoff device 15 4.2.3. Linear actuator assembly 15 4.2.4. Pressure sensor 15 4. 2. 5. Temperature sensor 15 4.3. Telemetry and Control Section 23 4.3.1. Magnetometer deck 23 4.3, 2. Temperature and filament switch deck 23 4.3 3. Control deck 31 4.3.4. Commutator deck 37 4.3.5. Battery deck 41 4.3.6. Subcarrier oscillator and transmitter deck 41 4.3.7. Thrust axis accelerometer 41 4.4. Pyrotechnic Firing Circuits 41 5. DATA 50 5. 1. Trajectory and Aspect 50 5.2. Temperature 50 5.3. Pressure 50 6. REFERENCES 60 APPENDIX: MODEL TESTS BY GAS DYNAMICS LABORATORIES 57 A. 1. Introduction 57 A.2. Wind Tunnel Tests 3 111

TABLE OF CONTENTS (Concluded) Page A. 3. Model Instrumentation 63 A.4. Test Program 64 A. 5. Test Results 66 iv

ACKNOWLEDGMENTS The Model A planetary mass spectrometer test flight was conducted under Contract NAS5-11128 as a cooperative undertaking of Goddard Space Flight Center's Laboratory for Atmospheric and Biological Sciences and the Space Physics Research Laboratory of The University of Michigan. Obviously a complete listing of those contributing to the success of the flight would be too lengthy to include here; however, personnel with specific responsibilities are listed below. Goddard Space Flight Center J. E. Cooley Project Scientist H. Dewey Mechanical Engineer D. W. Grimes Project Manager D. N. Harpold Experimenter N. W. Spencer Project Director Space Physics Research Laboratory (The University of Michigan) G. R. Carignan Laboratory Director J. R. Cutler Electrical Engineer C. E. Hubler Payload Technician R. G. Kimble Telemetry Technician R. W. Simmons Data Processing Manager Gas Dynamics Laboratories (The University of Michigan) D. E. Geister D. R. Glass v

LIST OF FIGURES Figure Page 1. Rocket elevated for launch. 4 2. Rocket and payload. 3. Payload configuration. 7 4. Payload assembly. 9 5. Payload block diagram. 11 6. Nose cone. 14 7. Breakoff device assembly. 16 8. Linear actuator assembly. 17 9. Conax linear actuator. 19 10. Pressure sensor. 19 11. Pressure sensor calibration. 20 12. Temperature sensor. 21 13. Temperature sensor calibration. 22 14. Magnetometer deck. 24 15. Magnetometer deck interface. 25 16. Magnetometer calibration. 26 17. Temperature sensor and filament switch deck. 27 18. Temperature sensor deck interface. 28 19. Temperature sensor electronics. 29 20. Automatic filament switching circuit. 50 21. Control deck. 52 vi

LIST OF FIGURES (Continued) Figure Page 22. Control deck interface. 33 23. Control deck index wiring. 34 24. Control deck circuits. 35 25. Control functions and monitors. 36 26. Commutator deck. 38 27. Commutator deck interface and reference supply circuit. 39 28. Commutator. 40 29. Battery deck. 42 30. Battery deck interface. 43 31. Subcarrier oscillator and transmitter deck. 44 32. Subcarrier oscillator and transmitter deck interface. 45 33. Subcarrier oscillator block diagram. 46 34. Subcarrier oscillator calibration circuits. 47 35. Accelerometer and interface. 48 36. Pyrotechnic firing circuits. 49 37. Trajectory. 51 38. Temperature vs. altitude. (a) Upleg. (b) Downleg. 52 39. Pressure vs. altitude. (a) Upleg. (b) Downleg. 54 40. Model for Mach 8 tunnel tests. 59 41. Wind tunnel model, nose cone. 60o 42. Wind tunnel model adapter. 61 43. Wind tunnel model, capillary section. 62 vii

LIST OF FIGURES (Concluded) Figure Page 44. Pressure tap locations. 63 45. Bench model of sampling flow inlet passage. 65 46. Predicted cavity pressure vs. altitude. 67 47. Schlieren photograph of flow field around model in an M = 8. 03 stream. 68 viii

lo INTRODUCTION The Model A Planetary Test Flight was designed as the first in a series of test flights to qualify a quadrupole mass spectrometer for high-pressure neutral constituent measurements on future planetary exploratory missions. The spectrometer, which was designed and built by the Laboratory for Atmospheric and Biological Sciences at Goddard Space Flight Center (GSFC), was the first test of a sterilized mass spectrometer electronics system in a flight environment. The spectrometer employed a unique pressure reduction device at its inlet orifice to permit measurements to be made at higher pressures than those measured by previous earth atmosphere devices. The measurement region for this mission was chosen to be 30 to 60 km on the basis that the ambient pressure profile in this portion of the earth's atmosphere corresponds to a region in the Martian atmosphere from 0 to 25 km. The nose cone design incorporated an atmospheric sample inlet system which provided a tolerable pressure and temperature profile at the mass spectrometer inlet orifice for this region. The payload contained temperature and pressure sensors mounted within the nose cone for the purpose of verifying the inlet system design. In addition to the control and telemetry circuitry, the payload also contained a magnetometer to provide aspect information, a thrust axis accelerometer to monitor rocket performance, and a pyrotechnically activated breakoff device to open the spectrometer to the atmosphere at the desired point in the trajectory. The present report describes the payload instrumentation and the sample inlet system design in detail and provides flight trajectory, temperature, and pressure data. The mass spectrometer data are being processed by GSFC and are not presented here. 1

2. GENERAL FLIGHT INFORMATION The general flight information for NASA 18.78 GA is listed below. The table gives the flight times and altitudes of significant events which occurred during the flight. These parameters were obtained from the flight records and radar trajectory information. Launch Date: 21 August 1969 Launch Time: 14:09 GMT; 10:09 AM, EDT Location: Wallops Island, Virginia Latitude: 370~50'14. 915" N Longitude: 79~29'01. 693" W NASA 18. 78 GA TABLE OF EVENTS Event Flight Time Altitude (sec) (km) Lift-off 0 0 Nike Burnout 3.5 1.7 Tomahawk Ignition 11.6 6.2 Auto Filament Switch Enable 19.0 (est.) 14.0 (est.) Tomahawk Burnout 21.0 17.7 Mass Spectrometer Inlet Opening 24.5 24.8 Enter Measurement Region 27.2 30.0 Exit Measurement Region 43.4 60.0 Apogee 236.0 227.4 L. O. S. 460.0 2

3. LAUNCH VEHICLE The NASA 18.78 GA launch vehicle was a Nike-Tomahawk two-stage, solid propellant, fin-stabilized, unguided sounding rocket. The first stage was a standard Nike (M5) rocket motor with a nominal 3.5 sec burning time. The second stage was a Thiokol Tomahawk (TE 416) rocket motor with a zero delay pyrogen igniter and a nominal burning time of 9 sec. The Nike used Aerolab type fins canted 12 min to produce a nominal 1.2 rps roll rate at burnout. The Tomahawk used Astro-Met type fins canted 20 min to produce a nominal 6 rps roll rate at burnout. At Nike burnout, the two stages drag-separated and the second stage Tomahawk coasted until T+12 sec, at which time the second stage igniter fired by means of an on board timer and battery pack located in the firing and despin module (FDM). The Tomahawk and payload were not despun, and the coning angle and angle of attack were thereby minimized. The launch vehicle, illustrated in Figures 1 and 2, performed satisfactorily and boosted the payload to an apogee of 227.4 km, 236.0 sec after lift-off. 5

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Rocket elevated for launch. 4

350.750 145.20 141.42 -- 64.12 59.936 \16.466~ /DI~ t9.00 DIA. FDM L 16.466 DIA. ANTENNA SECTION 36'610 PAYLOAD LINEAR ACTUATORS NIKE MOTOR TOMAHAWK MOTOR PAYLOAD VEHICLE WEIGHT PAYLOAD WEIGHT 18.78GA BOOSTER 1325 LBS. FDM 11.8 LBS. TOMAHAWK 537 ANTENNA SECTION 17. 7 1862 LBS. TOTAL EXPERIMENTS a CONTROL 177.4 206.9 LBS. TOTAL N IKE-TOMAHAWK Figure 2. Rocket and payload.

4. PAYLOAD Figure 3 shows the payload configuration for NASA 18.78 GA. Figure 4 is an assembly drawing of the payload excluding the antenna section and the fire and despin module which were furnished by GSFC. The nose cone section, quadrupole mass spectrometer section, and the control and telemetry section are discussed in this part of the report. Figure 5 is a block diagram of the complete payload. At T+24.5 sec the timer provided a signal to fire the redundant Conax linear actuators (-8 on Figure 4) which fractured the ceramic of the breakoff unit (-4 of Figure 4), thus exposing the spectrometer inlet orifice to the atmosphere. The temperature sensor (-10R of Figure 4) and the pressure sensor (-9R of Figure 4) provided temperature and pressure data within the sample inlet system nose cone cavity from lift-off to loss of telemetry signal. 6

ORIFICE 30 300 NOSE TIP 5I. 0 TEMPERATURE SENSOR PRESSURE SENSOR I' BREAK OFF -j L i QUADRUPOLE TUBE I i. I QUADRUPOLE ELECTRONICS QUADRUPOLE MASS SPECTROMETER I I 250 tEMR & FIL. SW. - I i_. i[11i, MAGNETIC ASPECT CONTROL DECK COMMUTATOR ---. CONTROL & TM I: =: BATTERY DECK 11. IO PULL AWAY _ -— =I SCO, TRANSMITTER = ACCELEROMETER ANTENNA 5.0 TOMAHAWK 9 FIRING CIRCUITS F D M Z L NAME WEIGHT C.G. FROM TIP ROCKET NO. 18.78 GA PROBE 177.4 hbs LAUNCH RANGE NO. WALLOPS IS., VA. FDM 11.8 Ibs TYPE OF ROCKET NIKE-TOMAHAWK ANTENNA 17.7 lbs DATE OF LAUNCH 8-21- 69 TOTAL 206.91bs 26.655" LOCATION WALLOPS ISLAND MISC. NOTES: TIME 10:09EDT 14:09 MT ALTITUDE 227.4 KM RESULTS ROCKET AND PAYLOAD PERFORMED AS EXPECTED. Figure 3. Payload configuration. 7

UMBILICAL CABLE CONNECTIONS AUTO FILAMENT BREAKOFF CHANGEOVER +28 PW'-o IRIG 30SEGMENT SWITCH DE V I C D E 2.5RPS COMMUTATOR LEDEX CALIBRATE CONTROL VOLTAGES, THERMISTORS, CONTROL LOGIC - 28 PWHOUSEKEEPING 0(INFO BKHZ) TEMASS SPECTROMETER CONTROL HV St FIL M 0 FL A INHIB MS OUT Fo —5 V ~~70.0 KHZ OWECFILAMENT CONTROL NR " —RAYMOND TITORSC(8W.. 5 C (INFO 1050 H) ~~~GAUBATTERY CHG.AO POWE CONRO COMMANDTTE TPOWMER CONTROLM. ASPECT 0-5 V 10.5 KHE 3. H (INFO 160 HZ) EX POWER POWER CONTROL 19 HR-I I 1IIIIACCEELE ROMETET R CA L 54 KHZ YARDNEY 0-50G I I r I CH18 SILVERCELLS (B.W. 810 HZ. (INFO 80 HZ) 28V' NOSECONE BACKUPTEMP a PRESS +28 PW SO's POWER T21 + 15 SEC TIMED FIL CONTROL Lw RAYMOND T4 T19 SEC. ACTUATED BREAKOFF ~~~~~~~~~~~~~~~~~~~~~~~~~ TRANM ITER j~~~~~~ II OMN +2MGT24.5 SEC T 30 RAYMOND TIMER Figure 5. Payload block diagram.

4.1. NOSE CONE AND INLET SYSTEM The design of the nose cone and sample inlet system was based on the principle of providing an environment in the earth's atmosphere which would be similar to that encountered by a high velocity entry into the atmosphere of Mars, The design was limited by the performance characteristics of the NikeTomahawk launch vehicle in that the high velocity (15,000-20,000 ft/sec) and resulting high stagnation conditions of a Martian entry could not be simulated. Also, since the Tomahawk stage did not include an attitude control system, the mission was dependent upon an upleg measurement.' As it turned out, however, the payload and attached Tomahawk stage stabilized very soon after encourtering the aerodynamic drag region on the downleg, and consequently some usable data were obtained on the reentry portion of the trajectory. The measurement region was chosen to be 30 to 60 km on the basis that the ambient pressure of this portion of the earth's atmosphere is close to that in the expected Martian atmosphere from 0 to 25 km, The design objectives of the nose cone were to transport a sample of the atmosphere to the mass spectrometer inlet orifice as rapidly as possible and at a temperature and pressure that would not affect the pressure reduction device at the inlet orifice0 The pressure reduction device is a sintered stainless steel leak that provides molecular flow into the mass spectrometer, if the pressure external to the leak is not too high. The leak used in this experiment required that the pressure in the sample chamber be below 100 mmHg, The leak conductance is also dependent, to a lesser extent, on its temperature. The nose cone and the inlet system were designed to maintain the maximum pressure in the sample chamber below 100 mmHg, and to reduce the temperature of the incoming gas to a level such that the total heat input to the leak did not raise its temperature more than 50'F. (More than a 50~F rise results in a change in leak conductance.) On the basis of three requirements, solid stainless steel was selected for the nose cone: (1) capability to withstand the high stagnation temperatures without appreciable chemical reaction with the incoming gas sample, (2) capability to act as a heat sink to the gas sample, (3) capability to insulate the incoming gas sample from aerodynamic heating of the conewallo Stainless steel is the least reactive of the readily available metals capable of withstanding the high stagnation temperatures. Copper, or a metal of similar heat conductivity, would have been the obvious choice for a heat sink but these metals are all very reactive and oxidize readily at high temperatures. The conductivity of stainless steel is high enough to reduce the gas temperature to the required level. The critical flow section of the inlet system was recessed (Figure 6) 2.5 in. to reduce the amount of aerodynamic heating input to a mini12

mum. Because of the large mass and relatively low conductivity of the metal surrounding this section and because the time duration in the aerodynamic drag region was quite short, the external heat contribution from conewall heating was very low. During flight, the gas temperature was monitored by using a platinum wire temperature sensor built by Rosemount Engineering Corporation. This sensor was mounted near one of the exhaust ports to insure that it would be in the flow stream where the time response would be a maximum, Figures 38(a) and 58(b) show theoretical versus measured altitude profiles of temperature. The pressure reduction was accomplished by using a critical flow section (Figure 6) which limits the flow rate by choking the flow and large expansion volume of the sample chamber. The exhaust ports were designed to be large enough with respect to the small diameter of the choking section so that the only restriction to exiting flow would be the external conewall pressure. A pressure transducer built by Spartan Southwest Engineering was used to measure the static pressure in the chamber, The transducer was mounted so that the pressure was monitored near the top of the chamber where purely static conditions most likely prevail. Figures 39(a) and 39(b) shows theoretical versus measured altitude profiles of pressure. To verify the calculations of pressure and temperature, wind tunnel tests were run in which a 1/5 scale model of the nose cone was used (see Appendix). Another objective of the tests was to determine whether there would be a, positive flow rate throughout the measurement region, Since the pressure measured in the chamber was always higher than the conewall pressure, a positive flow rate was indicated, The gas temperature coming into the sample chamber was measured and indicated that the temperature in the sample chamber was about 20% of the free-stream stagnation temperature. This agrees closely with the theoretical temperatures in Figures 38(a) and 38(b). Good data were received from both temperature and pressure sensors throughout the measurement region and close agreement was shown. between the theoretical and the laboratory data, 4.2. SPECTROMETER SECTION The spectrometer section assembly, shown in Figure 4, contains the mass spectrometer electronics (-2R), the quadrupole analyzer tube (-6R), the breakoff device (-4), the linear actuator assemblies (-8), the pressure sensor (-6), and the temperature sensor (-10R). Each of the above-mentioned components is discussed below. 15

375 /A..040.500 -0 NO. 5 0 - a. (70 Mq7r W/V SWq.3 -00.4 - 0es) O.JIL I 2 ~ AI.i.500 elql0. / _/-_7 0/ 0F. 4.=,e(D ooo) 00/LO./05/3000, Figu.o6. Nos) I co ne TH/S _U, F&}C E?. H, -l CON CEN T.I C FIN/NI.-/3%/3 H. 76T Z 3. WlTI4 NO N/CK<, OENrs, SCATt-/'E$ ETC. ('FOB SEAL IV/THO'O IPm,'NG ~'Z -.56.) 14HLES EllOUtS RCEO Figure No. o se cone.

4.2.1. Mass Spectrometer Electronics and Quadrupole Analyzer Tube The mass spectrometer system (Kerne, Deskevich, and Elder, 1968; Consultants and Designers, Inc., 1968), consisting of the electronics and the analyzer tube,was supplied by GSFC and is described only briefly here. The mass spectrometer system is designed to measure the neutral atmospheric constituents with masses between 10 and 50 amu. The spectrometer is continuously tunable and scans through one complete cycle every 2 seco The spectrometer output, an analog voltage proportional to the relative abundance of the mass number to which it is tuned, is supplied to the payload telemetry unit. Several housekeeping voltages are also monitored to assure proper spectrometer flight operationo 4.2,2, Breakoff Device The breakoff device assembly drawing is shown in Figure 7, This device provided a seal for the mass spectrometer inlet orifice until the spectrometer was in the desired region of the atmosphere. When the desired altitude was reached (at T+24o5 sec), two linear actuators were fired, which fractured the ceramic at the scored line about its circumference. The upper portion of the breakoff device is captured by a spring-loaded retaining device and the spectrometer was opened to the atmosphere. 4.2o3. Linear Actuator Assembly The linear actuator assembly is shown in Figure 8 and the actuator itself is shown in Figure 9, The actuator housing was a safety precaution which insured that any gases escaping from the actuator itself during its activation- would not contaminate the mass spectrometer measurement~ Two actuator assemblies were used for the sake of reliability. 4,2,4~ Pressure Sensor The pressure sensor and its interface to tthe payload are shpown in Figure 10. The final calibration curve for this 0-1 psia sensor is shown in Figure 11, The pressure sensor was used to provide verification data own the sample inlet system of the nose cone, 4.2.5. Temperature Sensor The temperature sensor is shown in Figure 12 and its final calibration tabulation in Figure 13. The temperature sensor was used to provide data te verify the design parameters of the nose cone sample inlet system. 15

I-'.250 / 8-024-00B - 4 HA T A SS~MBLY I B-024-008 -:3 CEIAMI l / 8- 0e4-00 7 - 2 BSAS' / 1 - 024-006 -/ FtLAVE URT NO. AME | I | rTHO ~~~~~~C,> NACD I * ^__wXD9Y SPACE PHYSICS RESEARCH LABORATORY.Y:: f /U. 5 /Z~DEPARTMENT OF ELECTRICAL ENGINEERING THE UNIVERSITY OF MICHIGAN B/ EAK' OFF U/V/T SS' ANN ARBOR, MICHIGAN A A PROJCT 93MOTHRWIE SMC IEDTLERN E41w l DMDIM. "D.0 t -,am... - t C- 024- 005o Dli. WDI.000eo dic I Mly. Figure 7. Breakoff device assembly.

4100/FlED -3S' - -9 FEED T7H L. 2 SrqoNaq4e -aS FLAT HD.CC2. M4.40AC)2~3/ STN. ST. A/ IF2ZSV31 -7S - IO"L/NC P /I X//6 V/_ON 77-54 5 P 57,qNDAED -,S eOUND sO.-D.5CA5r4- 40A I? tEASS 2 S TANDIWD -3S FLATAID. SCN. 14-40NC-2'.37S N2'SS 2 STatvO'De,. 1I4S SOCH. O CADCPPSC P.ECS~3PNC,.S00 PFI2/<2/ZED 2 S 7,NDQ,-D - 3 SOC. 40D. CPSC2 B4 4ONCI. 5004 P0eE2'/22D / _ L -25 SJ/XKBi/OWS(M T~D;."DO2& C"YC: 3B/ ST/V. ST STANDa9eD -/5 SOC. HO2 CAP 5(2 "- 3?NC-2'.375 PA2KEN/ZE / 0-0e4-0/0 -7 SPACC. I 8-024-0226 -c'- TA I/Nfe 2 A- 04-0/7 -S CONTACT A -0D04 -0/d6- 4 /NSLJL0TO, _ _ / 8- 024-015 - NDAZTF / 0- 004-0/1 - 2 VLLJNOCE / C-24-0/4-, 31-Y T PAT N(O. H A ht SPACE PHYSICS RESEARCH LABORATORY 0G0' A,. S Z - DEPARTMENT OF ELECTRICAL ENGINEERING rcr87L THE UNIVERSITY OF MICHIGAN ACTLJATOE SEOLSO HOUS7N0 0555 ANN ARBOR. MICHIGAN ANS F 40 Figure 8. LinearOT actuator aCIssembly.ERANCE DIY. D1D~~no A.oo"" it. D I No. C Figure 8. Linear actuator assembly.

8(M&I-) LONG E6LC L'A SLCA 5PECIFICATiO f\5: - FIeCZ LP6t $WG, SOLEE VIN.)O \j I. ACTU ATOC APP~LcD F0eCE =Z0O1~l&J AT B~E6INNING OF STADKEP 2. PetC'LF 6PDGE. Wle!22 $MsislWCE~d -.2 3 3. CONTINUIl'! TEST =C.OI AMP 4. MAA. PCTTIVL NO-FICE - O5 AMP S. MiWJ. POSITIlV FIPL -IAMP, eCCOMM. FI-!hG CUeCNT e 2 AMPS 7. OPEIZATING TI ME- ~ A"-PS 0.00? 5EC~. S. OPElATIN~' RANGE ~-(0OF TO 1IbQP 3. ENVIF-OKNMEMJTAL' ESAUEC$EA LEVEL TO Ij~ r j#~ IO. V4C.IGkT C.ST O APPLY THIN SEALIN4G COAT. i 00:9 OF UPOXY TO TH2-AD APPEO.. I INCH DEEP. 4 H 1.25 DIA- ko z - 7 — I, ANOD12L ALUL-i)HUM PA2TS PUE IwIL- A- A TYPE T F-J 1,25 D1A~ —-— Ct 5~~~~~~~~~~~~~~~~~~~5 L-IOG U _ _ _ _ ___ _ _ _ _ _ MIKI. TeAVEL ____IA,~ —~ —I 1_ Il GO-Z" ODIA! DeILL (.154 DIA) THC4U EIRUf3B STAMP MARKiNG TO INCLUDE 3 HOLOS SPACUD AS SHOCN CONAI F PIN, 5JP D MFG. DATE. ON.337 014. 5. C, NIAKINL PtOr~iCrTE? WI-TH A COAT Of vhp-NiSP PEZ MIL-V- IT!. T Y E J. Figure 9. Conax linear actuator.

.30 REF. _ 1.400 MAX —-.300 ~.030.600~o020 2 3% t~~~ —- - ~.500~020 ELECTRICAL IND. HEADER PRESSURE PORT /8 NPT *909FR/40W —HS-3E OPEN TO NOSECONE VOLUME 1.000 ~.020 PRESSURE SENSOR: SPARTON SOUTHWEST PRESSURE TRANDUCER 401-G-6-50 SN 7387 3 II NOMINAL Iprl piF S. 5 REF SUPPLY INCR EASING 2 CONTROL PRESSURE CK TEMP D, IRIG 2.5RPS COMMUTATOR SEGI I ~~I (TO 22KHZ SCO) L LEDEX J 7 O16 _ND SG16 GND OP N cm COMMUTATOR H REFERENCE DECK OP INTERFACE DCM-37 S SENSOR EXCITATION 5.OOOV SENSOR RESISTANCE 5000. SENSOR LINEARITY ~2.5%F.S. Figure 10. Pressure sensor.

SN - 7887 SUPPLY VOLTAGE: 5.00VOLTS 50 0 INCREASING PRESSURE CALIBRATION 6/3/69 8 DECREASING PRESSURE CALIBRATION 6/3/69 E 30 I w C) 20.100 0 0 1 2 3 4 5 OUTPUT (VOLTS) Figure 11. Pressure sensor calibration. 20

1.5 (TYP) U I --. 040 DI. (MAX.) I SENSOR COLOR WHITE.20(MAX) OO1 DIA. PLATINUMi; w (F~( BOTH ENDS) H SCHEMATIC DIAGRAM Figure 12. Temperature sensor. 21

ROSEILOUNT ENGINEERING COMPANY TEST REPORT MODEL 146CY SERIAL 3891 LATr 12 31 68 QUALITY CONTROL APPROVED ACTUAL CALIBRATION POINTS'iEMP K RESISTANCE 273.1500 496.89610 373.1665 690.27000 ALPHA IS.00389100 DELTA IS 1.50501 BETA IS.1100'I'MIt K RESISTANCE 300.00 549.38005 31).00 568.81997 320.00 588.20160 330.00 607.62504 340.00 626.79035 350).00 645.99742 360.00 665.14631 370.00 684.23706 380.00 703.26952 390.00 722.24384 400.00 741.15993 410.00 760.01783 420.00 778.81755 430.00 797.55900 440.00 816.24237 450.00 834.86753 460.00 853.43440 470.00 871.94313 480.00 890.39368 490.00 908.78899 500.00 927.12017 5 10.00 945.39810 520.00 963.61386 530.00 981.77382 540. 0 0 999.87475 550.00 1017.91780 560.00 1035.90290 570.00 1053.82960 580.00 1071.69820 590.00 1089.50860 600.00 1107.26070 610.00 1124.95460 620.00 1142.59050 630.00 1160.16800 40 0() 1177.68740 v6''i) 0. 0() 1! ]. 195. 14860 6.0 00 1212 55160 -/'i.0(0 1229.89650 66C5.00 1247.18330 690.00 1264.41140 700.00 1281.58170 710.00 1298.69370 /20.00 1315.74740'30.00 1332.74310 740 00 1349.68050 5.00 1366.55980 i30 0 1 3. 3 38080 77(.00 1400.14360 780.00 1416.84820 790.00 1433.49470 800.00 1450.88290 810.00 1466.61300 820.00 1483.08480 830.00 1499.49850 840.00 1515.85480 850.00 532.15120 86.00 1548.39030 870.00 1]564.57110 880.00 1580.69380 890.00 1596.75830 900.00 1612.76450 910.00 1628.71260 920.00 1644.60260 930.00 1660.43420 940.00 1676.20770 950.00 1691.92300 960.00 1707.58010 970.00 1723.17900 980.00 1738.71890 990.00 1754.20220 1000.00 1769.62660 Figure 13. Temperature sensor calibration. 22

4.3. TELEMETRY AND CONTROL SECTION In the telemetry and control section, illustrated in Figure 4, the following components are contained (the number following the component is the component designation number on Figure 4): magnetometer deck (-14), temperature and filament switch deck (-13), control deck (-12), commutator deck (-11), battery deck (-10), subcarrier oscillators and transmitter deck (-9), and the thrust axis accelerometer (-1R). The telemetry and control section provided all payload control and timing functions, battery power, telemetry signal conditioning, and the umbilical connection to the ground support console. 4.3.lo Magnetometer Deck The main purpose of the magnetometer was to provide roll rate data, The magnetometer deck is shown in Figure 14, the interface to the payload in Figure 15, and the calibration table in Figure 16. 4.3.2. Temperature and Filament Switch Deck Figure 17 shows the physical configuration of the electronics which perform the automatic filament switching function and the temperature sensor signal conditioning, and Figure 18 shows the temperature sensor-electronics interface. The temperature sensor electronics, which was designed by GFSC, accepts the temperature sensor input, and then provides as input to the telemetry system a O to 5 V signal proportional to the sample inlet system temperature. Figure 19 is the temperature sensor circuit diagram~ The filament switch portion of this deck provides two functions. First, it allows selection of either filament in the spectrometer ion source through the ground support console; and second, in case of a filament A (preferred) failure during flight, filament B is automatically switched in. Filament A functioned throughout this flight and a switch to filament B was not required. The filament switching circuitry is shown in Figure 20. 23

iiii~....~~~~~~~~~~~~~~~~~~~~~ai~~i!, ~ ill~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~:i~ii —i 4- ii-iiii- — i ii~:iii-E i~iii.-iiE~iii—::: —'~~~~~~~~~~~~~~~~~~~::.!::... Fi::I ~i-~i- igr 14. Magnetometer:: i~" dck.::ii~_-::::::jl/::

SENSOR CABLE J1 J2(HARNESS CONNECTOR AS) (MUST BE AT LEAST 8 IN. LONG) ASPECT ELECTRONICS SENSOR (4 IN. LONG BY I IN. HIGH) 2 B -F --— Ji YEL ASPECT IJ~9.~ 125K TYPE RAM-5C MAGNETIC Pa r) ASPECT SENSOR ALK SIG GND SN 2528 A~It +28 PWR CDB8-19 ELECTRONICS J2 PAYLOAD HARNESS CONNECTOR AS (AMPHENOL 126-223) NOTE: R1,R2, I/lOW. MOUNTED DIRECTLY MECH. DWG. C-024-025 ON CONNECTOR AS. Figure 15. Magnetometer deck interface.

HELIFLUX CALIBRATION DATA MAGNETIC ASPECT SENSOR TYPE RAM-S5C FIELD IN OUTPUT SIGNAL MILLIGAUSS IN VOLTS D C 600 SERIAL NQ g 25",, 550 _16? 500 4 / 450 _/L. 400 350 300? _. 250 200 150 _. o 100 50 0. 2 ___ (BIAS LEVEL) -50 -100'(E>S 4 (':/?DIRECTION OF MAGNETIC FIELD FOR VOLTAGE SIGNALS ABOVE BIAS LEVEL -300.. NOTE: CALIBRATION MADE WITH A 1OOK 350 aOHM RESISTOR FROM SIGNAL -350 <....:'- OUTPUT TO NEGATIVE TERMINAL l.m OF BATTERY SOURCE, AND A lOOK -400 OHM RESISTOR FROM BIAS OUTPUT TO NEGATIVE TERMINAL OF BATTERY -450 SOURCE -50 SCHONSTEDT INSTRUMENT COMPANY -550'.... / SILVER SPRING, MARYLAND -600 _dC)~_ CALIBRATED BY CALIBRATION MADE WITH BATTERY SUPPLY OF L 8 V OLTS DATY __2 -ALB 11262 Figure 16. Magnetometer calibration. 26

I,I~~~~~~~~~~~ iFigure 17. Temperature sensor and filament switch deck.

TEMPERATURE AND T E i:: RATLUE SENSOR ELECTRONICS I,X ~ ~FILAMENT SWITCH DECK (G'r SUPPLIED) ~~L_,~~~~ -)DEM-9P I\ \/ S ji kC*. C. r N.C. ~ CONNECTOR: TEMP TEiP- SIG TEMP SHL 1+28PW 24-D 34V PLATINUM CDB-16 WIRE 25 — 40 ma SENSOR WTEM -RE T TEMP DB TO2.5 RPS RIG COMMUTATOR O-V SEGMENT II (22KC SCO) SIG GND XT-5 SEALED II II CONTROL SECTION [~U ~ FEEDTHPU'S I MARS QP o ON FLANGE I (SEE B-E1344 MODEL A INTERFACE L.EE BCANNON NOTE:SHIELD TO BE GROUNDED DCM-37 ONLY INSIDE TEMPERATURE TEMP & FIL SW DECK PLUG TEMP SENSOR ELECTRONICS. XT-5 S N TEMP' TEMP SIG wT TEMP RET ILK TEMP SH T RED.+ 28 PW CBTEMP 1000 -3000K OUTPUT O-5V FS. POWER +28Vdc ~5V FLOAD 500K.D MINIMUM PHYSICAL SEE SPRL DWG. C-024-024 SENSOR: ROSEMONT 146CY SPECIAL SN 3891 Figure 18. Temperature sensor deck interface.

+PWLi~~~~~~~~~~~~~~~~~~~~~~~~~~ f~~+12V + PW ( + -z17 10 n(1/2 W N37820 IOK5% IN3612 5 % IN3157 IW~%0,5% (8.4V.REF)2K 5% 658%t 2 _ 7 35V 1N3612 POWER INPUT I(N3612 RF 1.1K LM20 11 1 24 - 34V MI %F M2 25 — i40 ma +.1/50 J/5 N12 (12v, 1w) 4 74u f $5V i.1132 vw. 1.Oltuf 5% 22pf PW GNO 1N3612 333K E 1N3612 15v - N2 (f:-.IO KHZ PWU GND ( 0.1% 0.1% 0.1% 6. 2224 V 5K,.I% 2OK 5K 1 % 5% KZN,132~~~~~~~~~~~~~~~~~~32K 4.3 2 K -~~~~~~7V (OFIET %,MF CALIBRATION FOR SENSOR SERIAL NO 3891 TEMP TEMP TEMP 50K~~~~~~~~~~~~~~~~SELCTE * C+ 12SPECIAL F\)~~~~~~~~~ ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~ 5/oOO 300 27 549.38 5 2K A 3 0.% 5.262 6 5% (6.2V REF ) 65 + 7 5 K (THEO.) LH201 ICM IIMF 2 2K ~/POT LH201 K I ~~~~~~~~~~~~ 2 I%5.12-,uf "FIJLMITE"I 600 327 1107.6 2.7164 PRO NASA GSFC DW. GC-1140351 0 2 9 0F BIOK 0.5Tma Figure 9 Tepeatr seso elctoncs 011% ~~~~ ~ ~~~~~IOOK T L_________ DEM-9P 1 I I 20K T POT CALIBRATION FOR SENSOR SERIA\L NO, 3891 PLUG - FIL T~P TM EP5 T(-K ) T(OC) Rs (n) VTM 0.1f~~ /O~ % - ~~~~~~~~~~TO SENSOR: ROSEMONT L _ 300 27 549.38 5.000 2 *146 CY SPECI AL FROM NASA GSFC DWG. GC-1140351, 600 ~ ~ ~ ~ ~ ~ ~ ~ ~~~~~~~I IN87A 11.. /6 3 41 1000 727 1769.63 0.015 REF: P.C. BOARD, GT-1140348'. Figure 19. Temperature sensor electronics.

TIMER SW* 3 QT3 NOT RUN NOT RUN SW!NH!BIT 420 R RUN OFF ON +20FIL FIL CONTROL RELAY + 20 FIL + 28 PW PART OF CONTROL DECK r IK r- -|- - - R6 1 50K + 33/35 SW INHIBIT W/OR N +-20V UNTIL RS RAYMOND TIMER R3 1N645 AUTO FIL B MON IS RUN lOOK 02.01 1N645 22K 220K C( R4Im RIFI 3.5V NOMINAL; O 02 SEL RIO 0 IOK I I ~I I I J II I 35AMPS N"U'NA' - I I PBB DIODES FO6 FIL A ILK 2 SIG GND VLK E SPRL 68030, B-E1359-I- B-~L N GSE I FIL ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~~~~~~~~ IFIL,E, SELECT RETURN I FILA A- SHORT ATOGETB. + FOR FIL B FIL - FOR FIL A DAM15P.P15 PC BOARDDCM 37) SPRL 68030, B-E1359 Figure 20. Automatic filament switching circuit.

4.533. Control Deck The control deck is shown in Figure 21 and the payload/control deck interface is shown in Figure 22, Figures 23 and 24 show the circuitry involved in the control deck. The control deck contains the Raymond "G" timer, the Ledex rotary stepping switch, the internal/external power control, and the mass spectrometer ion source filament on/off control. The Raymond "G" timer was actuated at lift-off and provided three timing signals: at T+15 sec backup power was supplied to the entire payload; at T+19 sec redundant filament power was supplied and, after a three sec delay to allow for filament stabilization, the automatic filament switching circuit was enabled; and at T+24.5 sec the pyrotechnic activating signal was generated, thus opening the mass spectrometer breakoff device. The Ledex rotary switch provided power to various payload components and supplied various monitor points to the ground support control console depending on which of twelve possible positions it was in. The table of functions and monitors is shown in Figure 25. 31

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m r rri L o, VIOLET <(A O M V -4)LEDEX (3n 0 D7 -IR L BL SI G GND Z Z 0PWGNDNa(9 ASPECT o' PW GND y I DS5-R8 ASPECT (~z n W/BLK 06-2R O YEL L QP DATA LU0 D5-R6 1-~.ANODE D4 -12R TAN C - tANODE D4-12R TAN.... ~M/C -2 D 1-OR I 3 GRN D3-R4 FIL REF D-GRN @ TAN D4-R4 - V AC D6-12R... M/C- 4 cD WoR, V MULT Dl-12F t w P W (t PNK D3-R5 RED p PNK TS 0 I 2 8 n f i T ~ i W/PUR D4-R2 PUMP i mt} zs; n@ GRY D5-R2 m 0 -r- VPUMP INT/EXT I I r;v — PRESS o- 1 03-R7-2 TEMPE - I-:~ Q >rW/PL OR DR-R7 (D GRI PUR S GMPO Zm } o ~ I I ('~i - - w / - - t ORD.RC^A RET. C3 04-RB_ PRESS F-b z + 20 FzL 0 0nOD- 1 AU XT+O R FIL [' 20I 4 —O E 4OFIL MON. - R ~I5 OBRN BLU.BL SG- N D-R4 NV INHIBIT FIL SW PW-SIG GN 1-5R _ PNKP ORD. CMD. a1 ORD. CMD rn 0:0 a PW-F ASEC PWR I v IBRN OR N a COMM PW RED N OP PW I.~ T PW PU - P GrN O~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~r~~~~~~~~~~~~~~~~~~

VOLTAGE MON'" W/IRED Di —II11/ QP,ETC. p (TI P)7 ARF2I BAKP-l XTMR PWR I- CDA_2CDo-15i. BACKUP 5-11RONT ASPECT-C T: B 9-11 I/RI ED Comm 8 C I JCDB -35 ~~II O / ~CDB- 19 I j j O 2RED00 C~ltD TII' T' REDUNDANT 9~ pi TEORD TIE TIE ~~~~~~~'/CDB-I9 ~~~~~~~~~~~~~ ~ ~ ~ liE I O~~~~~~~~~~~< 4 ~~~POINT 93 ~~~~POW TIEWT ~~POINT C C~~~~~~~~~~~~~~'i 9~~~'-r. X~'4'~~' ~" C REAR DUMMY 16K S IK CATHOE 210C DMSW 20 XTMR /C400C R2 PWR 2-F100 QPFIL CONT -2 P\P3 FLCONT-2 WC 17 M/C-3 M/C-4 RNG C PA-4 XTMR PA-10MCb I I 7P RA U L~~~~~~~~~~~~ HGIPMCP\ CA / 35o PI OPOTA 10 I/E-4 Q N.. 0 CDA-36 v3'~I PUMP I I9 ~~~~~~~~~~~~~IIC I/ II D COB-37 II 2ii CDB-13 R19 2 N, C..C. HC.~~.0D4-R6 10:. -. r~N.C-. N.C. P-U 0? DA 34 ~ ~ ~ ~ DA1 "/,/E-4 "L,..~ I~~~~~~~~~~~~~~~~~~~~ c-'~:"'~ u ~~~~~~~~~~~~~~~ -— r, -~o...'.. —. TIE 9PW C POI N".:;~ ~ —" "' M ~- -': _?,AC NT \~ ~ ~ ~ ~~T F..L..,.....CPSAL T 2 REF CA1NDE 0 RDANE INHIBIT CA1 TIMER PW TS Comm T ASPECT TIMER PW CDA-16 CDA-17 D-1AMC DECK I DECK 2 DECK3 M/C- 2 ~~~~~~~~~~~~~~~DB-2 ILMON3M/DECK &3Z 1~~~~~~~~~DCI2 AUTO D CK6I DECK 3 C/C -3 M/D DECK 4 DECK 5 DECK 6 NOTE BENE: DECKS SHOWN IN POSITION I REDUNDANT POWER SWITCHING OP 5-I TRANS II POWER COMMON IS LABELED "C" ( +28) Figure 23. Control deck index wiring.

LEDEX & 2 CONTROL 3 TIMER 4 BATTERY 5 PULLAWAY oN645 N645 DI D2 _ _ __4E30M D3 I - EXT- INT 1+ _ LATCHING KII L NENL IENT EXT EN INTERNAL-EXTERNAL POWER CONTROL 2 CONTROL +20 FIL 3 4 VOLT LATCHING. R=620n a 10.1/50 RI I -OF ON I +-, LATCHING RELAY KI ON OFF L _ _ FILAMENT CIRCUIT IS CONFIRMED OFF BY 10K 5 RESISTANCE TO GROUND. FIL CONTROL MODULE +20R Figure 24. Control deck circuits.

LEDEX POWER MONITOR / CONTROL LEADS OSITION VOLTAGE P, SCO ASPECT IMER F U N C T I 0 N S MONITOR REFCAL Comm XTMR HV 2 3 R=O EMPRESS COM COMMON INHIBIT DUMMY THERM DUMMY THERM I X:| UMBILICAL CONTINUITY OP FIL CONTROL 210C XTMR TEMP 40oc 2 1 I | I I | X I | _ OP CHECK OP FIL CONTROL I PUMP V PUMP P DATA 3 O I P FIL CONTROL I. I PUMP ANODE. OP DATA 4 FIL REF VAC V MULT QP DATA 5 Ts TE I ANODE QP DATA 6 I I I.... FIL SELECT I PUMP I ANODE QP DATA 7 TEMP a PRESS TEMP PRESSURE AUTO B FIL MON QP DATA COMMUTATOR COMMUTATOR QP FIL CONTROL COMMUTATOR ASPECT OP DATA S ASPECT 9 XTMR ON OPFIL CONTROL I I PUMP I ANODE IP DATA 10 I I I PREFLIGHT CHECK OPEN OPEN OPEN OPEN II FL I GHT BAT MON OPEN XTMR TEMP OP DATA R = I00 1 R=IK 12 X BATTERY CHARGE OPEN OPEN CHARGE NOTE I/E CONTROL ACTIVE IN ALL POSITIONS. Figure 25. Control functions and monitors.

4.3 4. Commutator Deck The commutator deck, shown in Figure 26, contains the commutator and the 0 through 5 V precision reference supply voltages. The circuit and the interface diagrams are shown in Figure 27 and the commutator is shown in Figure 28. The commutator cyclically sampled its 30 inputs at the rate of 75 samples per second and supplied these sampled data to the telemetry system. The commutator segment assignments were as follows: Segment No. Segment Assignment 1 mass spectrometer anode current Ia 2 mass spectrometer filament reference fil ref 3 mass spectrometer quadrupole rod voltage Vac 4 mass spectrometer multiplier voltage Vmult 5 mass spectrometer ion source temperature Ts 6 mass spectrometer electronics temperature TE 7 mass spectrometer vac ion pump current I 8 mass spectrometer vac ion pump voltage Vp 9 mass spectrometer +20 V filament monitor 0 V = fil off 10 mass spectrometer automatic B filament 0 V = fil A 11 temperature 12 pressure 13 battery voltage/6 14 transmitter temperature 15 zero reference 16 mass spectrometer anode current Ia 17 mass spectrometer filament reference fil ref 18 mass spectrometer quadrupole rod voltage Vac 19 mass spectrometer multiplier voltage Vmult 20 mass spectrometer ion source temperature Ts 21 mass spectrometer electronics temperature TE 22 mass spectrometer vac ion pump current Ip 23 mass spectrometer vac ion pump voltage Vp 24 0 V reference 25 1 V reference 26 2 V reference 27 3 V reference 28 4 V reference 29 5 V reference and frame sync 30 5 V reference and frame sync P7

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+15R 22 K F R PUR SC_36 A2 2N9 5 6, RD ~~SC - _36_TAN -L (+28) N A T YELFILBMON. A II TEMP; ( 2.999) O +3__ I N 9 6 I 3. 60 +4 5.1 K> 332Qa1 W/ REDY NAC Rl2 W/GRN P f 9 +3 -5 REF.LW/GRY -14 ) TH-XMTR KO f 332 Q 1 f W /GRN _ C B 2 ( R +R13999) 3' fRF R GRY CM-2 CALIBRATE VOLTAGES (8 98 -0 2 G Y so. C.U3 CASEBRATG GROUND R MEASURED BY FLUKE P 332 GWR CMMODEL 825A SN1405 R14 WTCM-5 WHT C3-6 26 JAN 69 (1.0015 CM-6 Yl TAN+1 ~O W/GRN CM-7 ___ ___ __ W/BLK THERMISTOR 28 REFERENCE SUPPLY c iRN. +2 RED3V ORG +5RE.. PC BOARD +4 +5 REF SPRL 68030, B-E1360 +5 + 5REF 470K 15 PINK ~~~~~~+15 +15 R BLU 3 v BAT COMM OUT BLK SIG GND SIG GND BRN NOTES' COMMUTATOR PIN 33 IS COMMUTATOR PW GND CASE GROUND REF PW REF PW COMMUTATOR REFERENCE DWG B-E 1531 COMMUTATOR W/GRN DECK TYPE 884 SN PLUG 28/6 PLUG CM 2.5 RPS DCM- 37S 51K COMM R 16 R17 Figure 27. Commutator deck interface and reference supply circuit.

5 10 15 20 25 30 35 40 45 50 56 60 uMmTuMM MMTMMTT mnn urm 3.5312 I| I` I I I | I I I I- I I I I -1-1-1 1ITO CASE GMCANNON DC-37P OR EOUIV. 3.250 I-N-cU~tlncD~cQDO) U@F0>g~oK9*990-SZ!EdNEIFIICVXLCeRISR // I.170 DIA. THRUDA.TU3HL J I J2 WIRING IDENTICAL ||*- / ON A 2.812 DIA. B.C..4060_ _Ii.4375.o tIFROM OUTPUT J2 FROM OUTPUT JI W C3 - n n`o o IDE -9P.4062 /1. fl 1 \ X l l 1,5 \00 _ 1.000 K \ CANNON DE-9P - 1.062 ii\ OR EOUIV. ROTOR POSITION ADJUST 0 NOTES: 1. TWO POLES. 2. FRAME RATE:(AT 28:3% VOLTS) SEE TABULATION BLOCK. 3. PHASING; NOMINAL COINCIDENT "ON"TIME OF CORRESPONDING PEDESTAL JU DATA POINTS. 4. MOTOR VOLTAGE: 28 VOLTS AT 2 WATTS MAX. 5. CONTACT RATING 50VOLTS 50 MA, MAX. 6. CONTACT RESISTANCE I LESS THAN Ia. 7. NOISE' LESSTHAN 20,uVOLTS,RMS,O-IO KC BANDWIDTH DURING CONTACT'ON' TIME. 8.LEAKAGE RESISTANCE:GREATER THAN 100 MEGOHMS AT 250 VOLTS, MAX. TYPE NO. FRAME RATE 9. WEIGHT- 11.02 MAX. 10. ENVIRONMENTAL' TYPE 952 -3 * 2.5 RPS, + 5, -15 % A. SHOCK' 150 G'S, II MILLISEC, ANY AXIS. B. VIBRATION 35 G S,20 TO 2000 CPS,ANY AXIS. TYPE 953-3 5 RPS, +5, -15% C. 125 G'S, ANY AXIS. D. OPERATING TEMP. RANGE -55~F TO +1750F. TYPE 951-2 I RPS,+5, -15% E. ALTITUDE UNLIMITED, 10 HRS MAX EXPOSURE TO HARD VACUUM. II. SWITCH FORMAT ( BOTH POLES) FOR AUTOMATIC DECOMMUTATION. A. 28 DATA POINTS -DUTY CYCLE 52% +-5% OF CHANNEL INTERVAL. B. FRAME SYNC SEG: 150% *S5% OF CHANNEL INTERVAL. * (884) 37 PIN CONNECTOR USES SAME C. COLLECTOR ACTION' BBM. D. PEDESTAL VOLT:O- +3V (REL TO MINUS SIDE OF PWR LINE). WIRIN AS (952 - ). E. STABILITY: 1%/ OF SETTING UNDER ALL OPERATING CONDITIONS. Figure 28. Commutator.

4.3.5. Battery Deck The battery deck is shown in Figure 29 and the schematic and payload interface is shown in Figure 30. The battery deck supplies internal power for the entire payload including the Conax linear actuators. 4.3.6. Subcarrier Oscillator and Transmitter Deck The SCO and transmitter deck, shown in Figure 31, contains the transmitter, mixer amplifier, four subcarrier oscillators (SCO's), and the SCO calibration systems. The components for this PAM/FM/FM telemetry system were supplied by the Sounding Rocket Branch of Goddard Space Flight Center and the telemetry system was then assembled, calibrated, and tested by the Space Physics Research Laboratory. The interconnection diagram for this deck is shown in Figure 32, The SCO calibration system was designed to place 0 and 5 V, 50 msec pulses on each of three SCO channels every 15 sec. The commutator SCO was not calibrated in this manner since a six-point calibration was included in commutator segments 24 through 30. The SCO calibration block diagram is shown in Figure 33 and the component circuitry is shown in Figure 34. 4.3.7. Thrust Axis Accelerometer The thrust axis accelerometer, Figure 35, was provided by the Sounding Rocket Branch of Goddard Space Flight Center to monitor the performance of the rocket motors. The accelerometer operated satisfactorily throughout the flight. 4.4. PYROTECHNIC FIRING CIRCUITS The pyrotechnic firing circuits are shown in Figure 36. As can be seen, battery power is connected to the Raymond timer only in Ledex positions 10 (pre-flight check) and 11 (flight)0 After lift-off but before the mass spectrometer inlet opening, the Raymond timer kept a direct short across both Conax linear actuators to protect against premature firing due to transient or spurious radiated signals. At T+24,5 sec the timer removed the short and applied full battery voltage across the redundant Conax linear actuators, Four current limiting resistors were placed in series with the actuators to protect the battery in case an actuator should present a short after firing. In addition, the timer contact was only a momentary closure. The pyrotechnic firing circuit and breakoff device performed as required for this shot, 41

I:~~~~~~~~~~~~~~~~~~~~~~~~~~~~IA qw -:: —~ 4=1~~~~~~~~~~~~~~~~~~~~~~~~~~~j~i 40,~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~:::- -:::::i-:: Fiue2.Batr ek

BA I #OD n XT-1 21 CM - 35 3 CDB-3 19HR-I CDB-4 CELLS el- 819 Lo CDB-36 ""+ 28V P no CDB -37 DEM-9S PLUA + + - t+ - - - + + - + I + -+- I _1 19 HR - I — + - + SIL-RC LLS 81 — + - ~ -4-__+ -I + - +t, MECH DWG. C-024-021 SAME MOLD AS PITOT BATTERY DECK. PLUG BA DEM93 BATTERY PACK Figure 30. Battery deck interface.

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Subcarrier oscillator and transmitter deck.

WINCHESTER CONNECTOR SMRE -I4PJ M Oog~O -J O aEU~ owo~~omo G 9 DBM-28P wco 4 O -90 No ~ POWER CONNECTOR 0XU MATES WITH VIKING vP5/4CE6 NOTE: XL ALL WIRING TO BE WITH *26 AWG. PIUG SC' PIN SCO BANK +28 A SCO PW C II N C. GND VIKING VP5/44615 (PIN VIEW)I -T, TRANS*'3XER AMP I M NC. _JN 11642 2 i 14 1 22.0 R COMM OUT SN 262 WHT I I CAL WM TI' A O~P DATA 3 IS1 70.0 L -."ji WN CAL 2 H N~~~~~~~~~~~~Wl ASPECT 4 12 10.5S 24IIW' MODULATION INPUT 5 10 5.4 K D~~~~~~~~~~~ SN 6745 GnD + 15L rrWT 6 H I 1 ~~~~~~~~~~~~+5 + E TRANSMITTER VECTOR +28 SPRL FET TR -2125 CALIBRATION V1 TR-2125 I I I I I ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~SEE B~E 1330 a 3.5 WATT VECTOR MFG B E 1331. 241.4 MHZ MMM-651-6 SN 975 SI I1 N 191 &LU A SN 975 ~ ~ ~ ~ ~ ~ ~ iL2.L.22~~~~~~~~.. RAY ~~~~~13) XTMR TEMP NC RI -r C2RED SCO MOUNT REFERENCE DWG. B-E1530 IOOpf T CI GRA 12 TH-XTMR RET R2 Ipf C ~~~~~~~~~~~~~T,5 C3 TO PHASING TEE 11LACK AND ANTENNA MOD. RET. allC PWR. GND. BROWN PWR GND GRFF 2203A CONNECTOR GND PRINTED CIRCUITS MOTHER: SPRL 68030 8-EI355 s I N MODULES: SPRL 68031 B-EI6354 I i N TRANSMITTER TRANSMITTER REFERENCE DWG. B-E1529 Figure 32. Subcarrier oscillator and transmitter deck interface.

+I5R OP DATA + 15R X,-,X I I C~ —---— cCS+1 UNI1U5 CTMONOSTABLE J-FET N J-1 FLI P FLOP SWITCH SWITC 50 ms 0-5 REF Y Y C A E a BB A UNIJ3UNCTION C +15 R +5RI I~EC PULSER 15SEC A P BOARDS MONOSTABLE NJ-F FL IP FLOP SWIC S 100 BE 3. E 35 B ~ ~ ~ ~ ~ J-1E PC BOARDS w MOTHER: SPRL 68030,,~ 8-E 1355 NOTES: I.FOR MODULE SCHEMATICS MDLS PL681 SEE DRAWING B~EEI330. 8E15 2.FOR SCO AND TRANSMITTER DECK WIRING SEE DRAWING B-E1342. 3. X IS NORMALLY LOW. 4. X LOW=4 INPUT A ACTIVE Figure 33. Subcarrier oscillator block diagram.

+15R +15R A X:NORMALLY LOW) A Y (NORMALLY HIGH) C I —..... 1 L____-..... D_ I I j I 560OOK I_ R2 330K R2 2N4852 R2RI 3 10I0R I TT II F,R E 82 IN645 R I I22/1V5 aEI 12N4104 D2I12N4104 I +64_ _ /____o ~CI A47 Q QI IN645 0 2 DI, l I I- "D3 RI 1D 336 K 80K R 6 R 7 FET SWITCH BI N64 -_ I X LOW S INPUT A ACTIVE L........ _.IN645 __ UNIJUNCTION PULSER FROM UJT 15 SECONDS PULSER MONOSTABLE FLIP FLOP PC BOARD ~ 50 ms lOOms SPRL 68031,B-E1354 R 4 680 K 680 K C I.luf.2 )4f Figure 34. Subcarrier oscillator calibration circuits.

WARNING NOTE: RELEASING INTERNAL PRESSURE WILL DAMAGE GIANNINI CONTROLS CORPORATION SENSOR. Do b= LETAIR I LINEAR ACCELOMETER OUTI 24117 WW-50-20 RESISTANCE 2000L RANGE O-50 G's AMPS.010' SERIAL NUMBER 381-4 + G'st NCR ~ ~ ~B A BLK SIG GND COMMUTATOR S REFERENCE DECK PLUG CA +5 SIG REF GND C CPS H F ~~~~~~~~~~PLUG XT SE R.INT +5 REE INCEAIN VIOLET ACCEL 231.4 M68 I~ ~ ~ SNO REITAC 2OOOI Fiur 155 Acelrmee andR interface. G'S 0-5 5.4SESRLIERT A BLK SIG GND SENSOR PLUG ACC ~~~~~~~~~~~~~~~~SCO TRANSMITTER DECK SENSOR INTERFACE 5.OOOV SENSOR RESISTAN)CE 2000.m Figure 35. Accelerometer and interface. SENSOR LINEARITY

PAYLOAD POWER LEDEX POWER CONTROL SWITCH CLOSED IN POSITIONS 10-11 (FLIGHT POSITION IS NUMBER Ii) DI-R5 a D3-F4 - 19 HR-I CELLS APRX. 28V QUADRUPOLE PHOSPHOR-BRONZE CAM NO. 2 OF 3 SECTION. FLANGE CONTACTS RAYMOND TIMER SWITCH HERMETIC CAN f / SEALED FEED THRU MOVES AT24.5 SECONDS CONNECTOR QP AFTER LIFT OFF r —— I CASE GND AT / J TRANS MITTER I L ACTUATOR QUADRUPOLE CONTROL DECK SUADLUPOLE BREAKOFF ACTUATORS CONNECTOR- MN ALL FRE.OA Figure 36R Pyrotechnic firing circuits. Figure 36 Pyrotechnic firing circuits.

5. DATA The telemetered data were recorded on both magnetic and paper tapes at the Wallops Island Main Base and Gcddard Space Flight Center Station A receiving stations. The mass spectrometer data were reduced from paper records by Goddard Space Flight Center personnel and are not discussed here. The temperature and pressure data were reduced by computer techniques from the magneteic tapes. 5. 1. TRAJECTORY AND ASPECT The angle of attack of the payload was assumed to be less than 50 throughout the meaningful portion of the flight. This was based on the facts that the payload was not despun, the dynamic unbalance was very low, and the static stability margin was extremely high. The first 83 sec of trajectory information were obtained from MPS-19 radar data which was fitted and smoothed by computer techniques at Wallops Island, The remaining portion of the trajectory was supplied (also by Wallops Island) in the form of Spandar data. Figure 37 shows the trajectory and the occurrence of significant events during the flight. 5. 2. TEMPERATURE The temperature data were reduced from the decommutated magnetic tapes. Figures 38(a) and 38(b) show the theoretical and measured temperature of the nose cone inlet system versus flight time. 5.3. PRESSURE The pressure data were -reduced from the decommutated magnetic tapes. Figures 39(a) and 39(b) show the theoretical and measured -rose cc'one ii let system temperature versus flight time. 50

1~ ALTITUDE (KM) o. NIKE BURNOUT 3.5 TOMAHAWK IGNITION 11.6 AUTO FILAMENT SWICH NABLE 19.0 MASS IP COETER INLET OPENING 24.5 ENTER MEASUREMENT REGION 27.2 EXIT MEASUREMENT REGION 43.4 0 * )0 *H APOGEE 236.0 H - C/ 0 m o Po 0 I / >z 2,_

600 I 1 I I (A) UPLEG 550 NASA 18.78 WALLOPS IS., VA. 500 14:09:00.000 GMT 21 AUGUST 1960 LONG. 750 29' W 450 LAT. 370 50' N *-MEASURED x - THEORETICAL Li ul [L~a 400' I* 0~~~~~ 350 0. 6 0*~~ APOGEE 300 x 0 0 250 0 20 40 60 80 100 120 140 160 180 200 220 240 TIME FROM LAUNCH (SEC) Figure 58. Temperature vs. altitude.

(B) DOWNLEG 550- NASA 18.78 WALLOPS IS., VA. 14:09:00.000 GMT 21 AUGUST 1969 500- LONG. 75- 29' W LAT. 370 50' N 450 n 400 (LI 350 APOGEE 300......... *. * 250 220 240 260 280 300 320 340 360 380 400 420 440 460 TIME FROM LAUNCH (SEC) Figure 38. (Concluded).

(A) UPLEG 60 NASA 18.78 SWALLOPS IS., VA. 50 14:09:00.000 GMT 21 AUGUST 1969 x LONG. 75W 29'W 40 LAT 37 5dO'N X -MEASURED E x - THEORETICAL LI0 \J1 U) 20 IL 0. 10 X 0 0 *3*e. ** 0*. *e 0 000. ee * ** 0eo * e.@ @.0*0 * **.00's.0 *e. 0 20 40 60 80 100 120 140 160 ISO 200 220 240 TIME FROM LAUNCH(sec) Figure 59. Pressure vs. altitude.

(B) DOWNLEG 60 NASA 18.78 WALLOPS IS., VA. 50 14:09:00.000 GMT 21 AUGUST 1969 LONG 75" 29'W 40 -- LAT 376 50'N E 30 E w Uf) 20 w a. flO APOGEE 0** 0.0so s * * e. *e @ 00 sog 60 0 @6 IL L I I -- 220 240 260 280 300 320 340 360 380 400 420 40 46 TIME FROM LAUNCH (sec) Figure 59. (Concluded).

6. REFERENCES Consultants and Designers, Inc., Final Report for Planetary Quadrupole Mass Spectrometer (Model A) Electronics, Contract NAS 5-9245, prepared for Goddard Space Flight Center, Greenbelt, Maryland, March 1968. Kerne, B., J. Deskevich, and W. Elder, Design Review of the Mass Spectrometer, Operations Research, Inc., Technical Report 445, prepared for Goddard Space Flight Center, Greenbelt, Maryland, 8 January 1968. Simmons, R. W., NASA 18.56 Thermosphere Probe Experiment, The University of Michigan Sounding Rocket Flight Report 07065-11-R, May 1969.

APPENDIX MODEL TESTS BY GAS DYNAMICS LABORATORIES A. 1. INTRODUCTION In order to make predictions regarding the air flow rate, pressure, and temperature in a particular air sampling system, it was desirable that a series of model tests be made at conditions somewhat similar to the most severe conditions anticipated for the flight model tube. The most severe conditions of interest here are approximately the following: Mach number 7.33 Altitude 61,000 ft Stagnation pressure after normal shock 69 psia Stagnation temperature 4610~R It was not practical to simulate these conditions, but it was possible to test a scale model. in the Gas Dynamics Laboratories' hypersonic tunnel without excessive effort. This tunnel was designed to operate at; a, Mach number of 8 with stagnation pressures up to 600 psia and stagnation temperatures up to 10000F. A 1/5 scale model was chosen for the tunnel tests since that was considered to be about the maximum size that could be tested witho;Iut choking the tunnel. Figure 40 is a photograph of this tunnel model. Figures 41, 42, and 43 are drawings of the model. Although the external dimensions (diameter and length) of the model are 1/5 the corresponding dimensions of the flight nose cone, it was not considered appropriate that the inside diameter of the minimum diameter secrtion of the sampling supply tube should be scaled down by 1/5. A 1/5 scaling factor on the diameter would reduce the minimum passage from 0.040 in. diameter to 0.008 in. diameter. It was considered more meaningful to scale the restricted passage area to 1/5 The length of the restricted passage in the model was chosen so that the L/D ratio would be the same for the model as for the full scale unit. The dimensions of the o+her sections of the sampling duct were scaled in a roughly similar manner and are not critical in determining flow conditions. The tunnel model was designed so that the minimum diameter section of the sampling supply duct could be changed easily. Three different inserts were made (see Figure 42), but only the 0.018 in. diameter insert was used in the tests. Further tests did not appear to be warranted at the time.

The four radial exhaust ports in the model were drilled with a No. 25 drill (0.149 in. diameter). This diameter was 1/5 of the exhaust hole diameter in the flight cone; thus, the model exhaust area was 1/25 of the full scale cone. One of the main objectives of the tunnel tests was to determine whether the exhaust ports in the nose cone were properly sized and positioned to exhaust the flow from the sampling system inlet without restricting that flow rate. The use of undersized exhaust ports made the tunnel test results conservative in that if no restriction occurred with the undersized passages, then the correct passage area would provide even less restriction. For example, at the inlet temperatures tested (from 100 to 300~F), it was found that the actual flow rate (through the 0.018 in. diameter by 0.50 in. long insert) was about 60% of the flow rate calculated on the basis of a short choked orifice. It was also found that the downstream pressure needed to be less than about 1/3 of the inlet stagnation pressure in order that the flow rate would not be dependent on the downstream pressure. In other words, if the spectrograph cavity pressures were maintained at less than 1/3 the upstream stagnation pressure, the exhaust ports would not be restricting the sampling flow rate. Although more extensive tests would have been desirable, it was believed that the test results obtained were adequate to interpret the tunnel test results. 58

Figure 40. Model for Mach 8 tunnel tests. 59

3.5863.111 1. 175.500.125 DIA. N TE.750 DiA 1 300 1.120 15DILI x 1.800 DIA. DIA i _*48 DRILL, ROUND NOTE Q~ LEADING EDGE& 0,1\ ~~~~~~~~~~~~.175 0 NOTE: A 1.460DIA. BOLT CIRCLE DRILL a TAP FOR 6-32 THREAD,.375 DP 4 HOLES EVENLY SPACED. (900) B *25 DRILL NORMAL TO SURFACE OF CONE. 4 HOLES EVENLY SPAC'ED. (900) NOSE CONE MODEL DIMEN - INCHES SCALE - FULL MAT'I. - ALUM 1 REQG Figure 41. Wind tunnel model, nose cone.

DRILL FOR CLEARANCE ON 6-32 CAP-SCREWS 4 HOLES EVENLY SPACED NOTE B DI A. TURN DOWN FOR SMOOTH.30 (FIT IN NOSE CONE..250 NOTE: A. MATCH BOLT CIRCLE PATTERN OF NOSE CONE. NOSE CONE MOUNT B. DRILL & TAP TO MATCH EXISTING FULL SCALE THREAD ON MODEL STING. (3/8-24) MATR. ALUM 2 R EQ'D. Figure 42. Wind tunnel model adapter.

L N o. L 0 (DIA.) I.50.018 2.50.025 3 1.00.025.250(DIA COUNTERSINK ORIFICE MAT'L. STAINLESS DIMEN. INCHES 3 REQ SLIP FIT IN REAR OF CONE (1/4" DRILL HOLE) R0 #48 DRILL.250(-) DIA. No. L 1. 1.10 2. 1.60 Figure 43. Wind tunnel model, capillary section.

A.2. WIND TUNNEL TESTS As mentioned previously, a 1/5 scale model of the nose cone was tested in the Mach 8 facility. A major portion of the air sampling passage was duplicated in the nose cone. The sampling system used in the scale model was designed with the capillary section removable. This feature served two purposes: first, any scale effects on the sampling flow rates could be corrected, and second, the effects of capillary size and material (heat transfer rates) could be changed if the sampling system performance was not adequate. The wind tunnel model was made of aluminum and the capillary sections were made of stainless steel. Although the full scale nose cone sampling system was made of stainless steel, it was felt that the saving in time and machining costs justified an aluminum model. Since the effects of heat transfer on the flow would be controlled by the capillary section, the capillary inserts were made of the same material as the full scale system. The model duplicated the full scale nose cone to a point 2.5 in. back on the cylindrical body. The model terminated there with an adapter plug which mated the nose cone to a tunnel sting. The sting in turn held and positioned the model in the tunnel flow field. A. 3. MODEL INSTRUMENTATION Provisions were made for monitoring seven pressures and one temperature on the model. Figure 44 indicates the points where pressure measurements were taken. Points 1, 2, 4, and 5 were all static pressure measurements on the surface of the cone, while points 6 and 7 were static pressure measurements on the surface of the cylindrical body joining the cone. Point 3 is the pressure in the nose cone cavity. A temperature measurement of the gas flow into the cavity was also made. An iron-constantan thermocouple was used for the temperature measurement. The cavity is the plenum into which the sampling air flows before venting through the cone surface. EXHAUST PORTS TAP NO.3 INSIDE NOSE CONE CAVITY 6, Figure 44. Pressure Cap locations.

The pressure measurements were made via a pressure transducer data acquisition system. The output of the transducer is digitized and punched on paper tape. The temperature measurement was recorded on an x-y recorder. The y channel was run in time-base mode giving a recording of temperature versus time. A.4. TEST PROGRAM Because of time limitations, only a small number of tests were planned. Also, the information that could be gathered from the model was reduced to one particular aspect of the gas sampling system. The major emphasis was placed on whether the gas sampling system would maintain an adequate flow rate over a given altitude range. In the wind tunnel. tests, this constituted the measurement of the cavity pressure level as a function of tunnel total pressure levels. The tunnel operating conditions were such that a Mach 8 flow could be produced with a total pressure variation between 50 psia and 500 psia, the altitude equivalent being approximately 200,000 to 120,000 ft. It was anticipated that the exhaust ports in the model could be easily enlarged but the test results indicated that it would not be necessary. The rate of air flow through the sampling inlet duct is determined by the temperature and pressure (and composition) of the air upstream of the minimum diameter section, provided that the pressure in the downstream spectrograph cavity is below a certain critical value. This critical pressure would be about 1/2 of the upstream stagnation pressure if the restricted (throat) section were only a few throat diameters long, but in the case where the restricted section is many diameters long, the critical pressure must generally be somewhat lower. A series of bench model tests were made to determine, at least approximately, the values of this critical pressure for various inlet conditions. Figure 45 is a drawing of this bench model of the sampling flow inlet duct, This model also was designed to allow for easy changes of the minimum diameter section, but only the 0.018 in. diameter insert was tested. Upstream of this bench model an electrically heated stainless steel tube was used to raise the incoming temperature of the gas. The air flow fed into the bench model was measured by a calibrated capillary tube system. During a test, the flow rate was maintained at a constant value independents of the heat input and the pressure upstream of the test section. The bench tests were not nearly extensive enough to allow accurate extrapolation to the flight model, but certain approximate values were determined. 64

I.375 -.001 _375_-_ -001 MAKE I EACH, STAINLESS STEEL L D _9 -__a/Xmg1 ~ ~~5 7~1 t0.5.018 0.5.025 1.0.0 25 _g- L.. 2.0.040 ALL t.0005 1/8 NPT-FEMALE BREAK THROUGH 1/16 DRILL ( SAME AS OTHER END 1/16 DRILL, 900 FROM TAP 1.250.500 BREAK THROUGH WITH.125 -.375 - 70 DRILL.375 DIA.,169 SEE NOTE I 1.0.55 1.00 -.50 1.45 NOTE: 1. DRILL AND TAP FOR NO. 8-32 RETAINING SCREWS, 15 LOCATE 4, 900 APART; 450 FROM RADIAL HOLES. 2. MAKE I RETAINING ASSEMBLY - MAKE INSERT PER UPPER 1.68 SKETCH. INSERTS SHOULD SLIP EASILY INTO ASSEMBLY. 1AQ SNUG FIT. Figure 45. Bench model of sanpling flow inlet passage.

A.5. TEST RESULTS Three series of tests were run. The total pressure for these runs varied from 60 to 390 psia with a flow Mach number of 8.03. Total temperatures for these tests was 760~F. The tunnel operating conditions are listed below. Since the primary emphasis was on the sampling system flow rate, only the data related to this feature were reduced. A plot of cavity pressure versus equivalent altitude is presented in Figure 46. Also plotted on this figure are the computed values for the pressure on the surface of the cone and the total pressure behind a normal shock. The results indicate that the cavity pressure was slightly higher than the theoretical cone surface pressure, which would necessarily be the case as long as there was flow from the sampling system. Also, the fact that the cavity pressure is roughly 1/6 of the stagnation pressure at the entrance of the sampling system inlet indicates that the flow in the restricted portion of the supply tube was choked. As mentioned previously, a thermocouple was installed to measure the temperature of flow entering the cone cavity. Preliminary indications were that flow temperatures were about 20% of the free stream total temperature. A Schlieren photograph showing the flow field over the test model is also included (Figure 47). This photograph indicates that the low flow rate from the cavity did not significantly disturb the external flow. More extensive tests with the tunnel test model could be made, but the results presented here are believed adequate to meet the limited objective which prompted this work. TEST LOG Tunnel Conditions Equivalent Run Total Total Mach ARDC Std. No, Pressure psia Temp'-F No. Alt. -Ft. Alt. -Ft. 1 62.85 760 8.o3 183,000 2 174.3 760 8o03 155,000 3 391.6 760 8.03 133,000 66

'88' 800o 700 600 500 400 300 \0 ~ TOTAL PRESSURE BEHIND NORMAL SHOCK AT Mz 8.03 - CALCULATED 100 70' 60 E40 30 vc 20 C) 0 6 CONE SURFACE e _ PRESSURE - CALCULATED 4 0 - MEASURED CAVITY PRESSURE 100 120 140 160 180 200 ALTITUDE(103ft) Figure 46. Predicted cavity pressure vs. altitude. 67

Figure 47. Schlieren photograph of flow field around model in an M = 8.03 stream. 68

UNIVERSITY OF MICHIGAN III3 9015 02947 4940I I II 3 0502947 4940