ENGINEERING RESEARCH INSTITUTE UNIVERSITY OF MICHIGAN ANN ARBOR FINAL REPORT SUBJECT: THE MEASUREMENT OF AMBIENT PRESSURE AND TEMPERATURE OF THE UPPER ATMOSPHERE By W. G. DOW Project DirLctor N. W. SPENCER Project Engineer "The research reported in this document has been made possible through support and sponsorship extended by the Geophysics Research Division of the Air Force Cambridge Research Center, under Contract No. AF 19(122)-55. It is published for technical information only, and does not necessarily represent recommendations or conclusions of the sponsoring agency' Project M824 AIR FORCE CAMBRIDGE RESEARCH CENTER, U. S. AIR FORCE CONTRACT NO. AF 19(122)-55 August 1953

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ABSTRACT The general objective of the work under subject contract is to obtain measurements of the ambient pressure and temperature of the upper atmosphere using high-altitude rockets as research vehicles. A new method which permits an essentially point-by-point determination of ambient temperature from measured pressure values has been developed and utilized experimentally. Resulting upper-atmosphere data are presented. The equipment necessary to implement the theoretical method has been developed. The alphatron ionization gage constitutes the basic pressure-sensitive device, and provides the basis for a pressure measurement system that will measure gas pressures in the range 760 mm Hg to 10-3 mm Hg. Associated equipment utilizing a gyroscope to allow determination of missile angle of attack, andmany accessory pieces of equipment, less spectacular but necessary to the experiment, have been developed. This report reviews that portion of the research which was conducted under subject contract. iii

TABLE OF CONTENTS Page LIST OF FIGURES v LIST OF TABLES vii 1. INTRODUCTION 1 2. SUMMARY OF RESEARCH 2 3. DESCRIPTION OF PRIMARY INSTRUMENTATION 4 5.1. Alphatron System 4 3.2. Gyroscope and Camera Installation 7 4. SUMMARY OF ACCESSORY EQUIPMENT 18 5. AIR DENSITY MEASUREMENT BY UHF GAS-DISCHARGE BREAKDOWN 19 6. SUMMARY BY ROCKET FLIGHT 20 6.1. V-2 No. 47, June 1949 20 6.2. Aerobee USAF-3, December 1949 20 6.3. Aerobee USAF 6, June 1950 22 6.4. Aerobee USAF 10, December 1950 26 6.5. USAF 18, September 1951 33 6.6. USAF 31, October 1952 38 7. REPORTS AND PAPERS 45 8. PERSONNEL ENGAGED IN CONTRACT WORK 47 9. GENERAL REMARKS AND SUMMARY 49 APPENDIX 51 ABSTRACT 51 I. INTRODUCTION 52 II. THEORETICAL CONSIDERATIONS 53 A. General 553 B. Nonyawing Cone 54 C. Yawing Cone 59 III. INSTRUMENTATION 62 A. Pressure Measurement 62 B. Angle Measurement 65 iv

TABLE OF CONTENTS (cont.) Page IV. EXPERIMENTAL RESULTS 67 V. SELF-CONSISTENCY IN THE RESULTS 71 LIST OF SYMBOLS 75

LIST OF FIGURES Fig. Page 1 Constructional Features of an Alphatron Gage 5 2 Typical Curve of Alphatron Output Current Vs. Pressure 6 3 Schematic Diagram of Single Alphatron System 8 4 Camera Modification, December 1950 10 5 Photograph of Equipment Recovered from V-2 No. 47 21 6 View of Alphatron Instrumentation, June 1950 23 7 View of Alphatron Instrumentation, June 1950 24 8 Gyro and Meter Panel as Viewed by Camera, June 1950 25 9 Curve of Ambient Pressure Vs. Altitude, June 1950 28 10 Curve of Ambient Temperature Vs. Altitude, June 1950 29 11 Overall View of Instrumentation, December 1950 31 12 Portion of Instrumentation, December 1950 32 13 Cone-Wall Alphatron Electrometer Stage, September 1951 36 14 Complete Instrumentation Rack, September 1951 39 15 Functional Sketch Illustrating Location of Alphatron Gages and Arc Discharge Gap on Aerobee Nose Cone, September 1951 40 16 Minimum Local Cone-Wall Density as Obtained from Gap Experiment, September, 1951 41 17 Curve of Ambient Pressure Vs. Altitude, September 1951 42 18 Curve of Ambient Temperature Vs. Altitude, September 1951 43 19 Physical Quantities Appearing in Nonyaw Concial Flow 54 20 Mach Number Vs. Quotient of Cone Pressures 57 21 Relationship between the Unyawed Surface Pressure and the Ambient Pressure 58 22 Coordinates Describing the Yawing Cone 60 vi

LIST OF FIGURES (cont.) Fig. Page 23 Variation of Output Current with Chamber Pressure 63 24 Elementary Block Diagram of Alphatron Pressure Measurement System 64 25 Variation of Alphatron Pressure Measurement System Output 66 26 Actual Cone Pressures, Total Head (Cone Tip) and Surface, Compared with the Ambient Pressure for Aerobee Rocket of June 20, 1950 68 27 Ambient Temperature at Various Altitudes above Alamogordo, New Mexico (June 20, 1950 at 0838 hours) 69 28 Ambient Temperature at Various Altitudes above Alamogordo, New Mexico (September 13, 1951 at 0437 hours) 70 29 Pressures Calculated from Hydrostatic Equation (dP/dh = - Pgm/RT) 72 30 Pressures Calculated from Hydrostatic Equation using Experimental Temperatures of Fig. 10 73 vii

LIST OF TABLES Table Page 1 Typical Gyro Data (December 1950) 12 2 Upper-Atmosphere Data for 20 June 1950 27 3 Upper-Atmosphere Data for 13 September 1951 37 viii

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN THE MEASUREMENT OF AMBIENT PRESSURE AND TEMPERATURE OF THE UPPER ATMOSPHERE 1. INTRODUCTION The Electrical Engineering Department of the University of Michigan has been engaged since early 1946 in the measurement of the ambient pressure and temperature of the upper atmosphere, and, in addition, certain characteristics of the lower ionosphere regions including ion and electron concentrations and electron mean energies. These experiments have been facilitated through the use of V-2 and Aerobee rockets, the field test portion of the work having been conducted exclusively at military bases in New Mexico. The research accomplished prior to 1 June 1949 was carried out under USAF AMC Contract No. W-33-038-ac-14050. For a summary of those activities, the reader is referred to the final report,1 which gives an indication of the instruments used and status of the resulting data. The report also lists, with abstracts, all the reports issued during the period of that contract. The research conducted from 1 June 1949 to 1 October 1952 was carried out under USAF Contract No. AF 19(122)-55, which terminates 30 Septemiber 1953. The period from 1 October 1952 to the termination date has been devoted exclusively to publication and report writing. This report is the final report on the work conducted under Contract AF 19(122)-55. "Pressure and Temperature Measurements in the Upper Atmosphere", by W. G. Dow and N. W Spencer, USAF AMC Contract No. W-33-038-ac-14050, May 11, 1950. 1 -

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN 2. SUMMARY OF RESEARCH The research performed under this contract has been devoted almost exclusively to the measurement of ambient atmospheric pressure and temperature. In general, a single temperature-measurement method developed by project personnel from fundamental aerodynamic considerations has been employed. The Appendix to this report presents details of the method, which may be summarized as follows: The ratio of the pressure (impact pressure) existing at the apex of a slightly truncated cone, moving axially at a supersonic velocity, to the pressure existing at a point along an element of the same cone may be interpreted as a Mach number. A mathematical consideration of this Mach number and the velocity of the cone (rocket) will yield the ambient temperature. Certain accessory considerations are basic to this computation. The angle of attack, that is, the angle between the axis of the cone and the air stream velocity vector, must be known. Implicit in a determination of this angle and the resulting adjusted pressure values is the assumption that the ambient wind is negligibly small as compared with the stream velocity, that is, the missile (cone) velocity. The instrumentation developed by project personnel to implement the theoretical considerations and to apply the method experimentally uses, as the basic sensing elements, alphatrons and a modified attitude gyroscope. The overall instrumentation system employing these devices has undergone almost continuous development in order to improve the accuracy of the resulting data, facilitate operation in the field and rocket, and improve the reliability of operation in flight. Each field test conducted under the contract employed equipment of this nature and also accessory instruments, the following list of which is typical: (a) camera for recording gyroscope data, (b) telemeter calibration unit, (c) timing devices, (d) missile acceleration indicators, (e) telemetering signal limiters, 2

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN (f) invertor power supply for gyroscope, (g) Alphatron-chamber temperature indicator, and (h) circuit switching devices.

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN 3. DESCRIPTION OF PRIMARY INSTRUMENTATION This section will present moderately detailed descriptions of the various basic pieces of equipment of which a typical instrumentation is composed. In general, the latest model will be described. In many cases, the discussion will be extracted from the many progress reports written during the course of the work. 3.1. Alphatron System An alphatron gage is an ionization gage wherein the ionization results from collisions between alpha particles, emanating from a radium foil source, and gas molecules. The charged particles resulting from this process are collected by electrically polarized elements of the gage and thus constitute a small current when an external circuit is completed. Figure 1 illustrates the constructional features of a typical Alphatron chamber, and Fig. 2 presents an actual curve of output current plotted as a function of pressure (temperature assumed constant). It is apparent that the gage characteristic is linear over a large range of pressure. At higher pressures the departure from linearity indicates recombination before particles are collected. At the lower pressures, the current due to direct collection of alpha particles and the resulting secondary emission becomes more significant, which ultimately serves to limit the usefulness of the device at very low pressures. In application, the total range of pressure measured by project equipment is from a ground level pressure of approximately 760 mm Hg to about 10-3 mm Hg. In order to obtain the desired definition, this total range is divided arbitrarily in many so-called subranges, usually seven to eight in number. The various subranges in the equipment are manifested as different values of alphatron load resistance. The voltage range developed by the alphatron current under conditions of changing pressure in the load resistance is equivalent in magnitude to the desired range of output information signal. However, due to the very small current and hence the very large value of resistance required, it is necessary to provide an impedance match between the high resistance (which may be as high as 5 x 1012 ohms) and the input resistance of the telemeter, usually of the order of 5 x 104 ohms. 4

COLLECTOR EL ECTRODE L RADIUM SOURCE Fig. 1. Constructional Features of an Alphatron Gage.

9,@afnssacId'SA uGlTmD cncIdino uo[0uclTIV Jo aAnLD TIeOTdId'O'T ALPHATRON CURRENT IN AMPERES -N -1 0 - 0. o C CX) mo 0 m 0 30), o

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN A 100 per cent negative feedback amplifier provides the requisite impedance match. The first stage of the amplifier employs an electrometer tube (low leakage in shunt with high resistances) and the following stages employ conventional d-c amplifier techniques. A filament-voltage-variation compensating stage is included. Independent of the amplifier but associated with the system is a switching circuit whose function is to insert the particular value of alphatron load resistance appropriate to a given chamber pressure range. The circuit in effect acts to keep the output signal voltage on scale as the chamber pressure progresses from atmospheric pressure to lower values, during flight use. A Ledex rotary solenoid serves to accomplish the actual resistance switching on activation by the switching circuit. Several figures are included in this report to illustrate specific instrumentations. The external features of alphatron systems are shown. Figure 3 is a schematic diagram of a single alphatron unit. 3.2. Gyroscope and Camera Installation A Sperry model F4A gyroscope has been modified for use in determining missile (cone) angular position during flight. The modifications deemed necessary are generally to adopt the unit to operation under the free-fall conditions of flight and to change the scale markings to facilitate data recording by the missile-borne camera. The data obtained from the gyroscope may be utilized to determine the zenith (z) angle of the missile as shown by the following discussion: The F4A gyro has two degrees of freedom. In the event that an Aerobee were to fly with no spin, it would be possible to mount the gyro so that pure pitch and yaw angles would be read. Such a mounting would require that the axis of the gyro rotor be along the figure axis of the missile (or axis of least moment) initially and, furthermore, that the gyro pitch axis be perpendicular to the plane of the trajectory. For the purpose of this discussion, pitch angle is defined as the angle in the plane of the trajectory and yaw angle as the angle in a plane perpendicular to the plane of the trajectory and the earth. It is not to be inferred, however, that, in normal use of the F4A gyro in a missile, the two angles read are exactly pitch and yaw, for the measured angles vary with missile roll. From the indicated pitch and yaw angles, the zenith angle as referred to the gyro rotor axis, in this example, the tower axis, may readily be computed.

4.5 V. 60V. 60V. a 9..*R C, I'. 150K 10K 470 K.OlnV 100K 45VV I~~~~IALPHATRON GAGE vt 45+ RK V 5702/CK605CX ~ 5702 50 -_5702 K 30V. ALPHA 5 5800'VX41A I.V ~~~~~~~~~~~~SOURCE j-j~~~~~+45v_ 8 ~~ ~ ~ ~ ~ ~ ~ ~ < -0V4Ao5 OUTPUT IO/t RI,, R24 R47 OK VF1.2KY R~~a~ I~ I RI )R ~Rs R4 Rs RI R? I.SV I.S SIA SH ZERO ADJUST MOTOR V7~~~~~~' ZERO ADJUST +6V. ~ 1 ~ ~~~~~~~~~~~~~~~~~~~~~~MOTOR CONTROL lo o l+6V.o lonl.nlon I00/1. 10011. I00a1'1. 10O0.f)t. 100A 100A LEDEX COILS HI LIMIT -I6V. SWITCH c5 ~~~~~~-CC CC)1 1 — 2)Id /I.sv. RY LO LIMIT =?ON 0 I SWITCH 50IK v V 470 K 6 47 K CZ 4)Af - Ijf AMPERITE 250K IHTII R Z3 R21 470K TO ALL HEATERS t IC + 24V + 6V. GND Fig. 3. Schematic Diagram of Single Alphatron System.

-- ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN If the picture is now complicated by missile spin, it is still possible to determine the zenith (z) angle. Consider the following: a missile spinning at a one-second rate, constant angle z, no precession. The gyro would indicate two sinusoids of one-second period, one as pitch angle and the other as yaw, appearing ninety degrees out of time phase. The maximum instantaneous value of either sine curve would be exactly the z angle. If precession about the gyro rotor axis is now allowed, there would be no change in the indications. Therefore, a so-called "cone of ambiguity" is defined whose half angle is the indicated maximum or z angle. It is not possible, therefore, to determine the missile's position on the precession cone. Note that in order to determine the z angle, missile spin is required. In our normal installation, the gyro is mounted as stated above, except that the pitch axis, of course, does not necessarily remain perpendicular to the trajectory during flight. This does not alter the fact that the maximum values of the sine variations of the angles read as pitch and yaw are the z angle. In actual, rather than hypothetical, flight, the axis of the precession cone does not remain parallel to the rotor (or tower) axis, but follows, more or less, the trajectory. This assists in the analysis, for it greatly reduces the ambiguity due to the precession cone by superimposing an envelope on the sine variation, suggestive of an amplitude-modulated radio-frequency wave, thereby altering the gyro indications. The period of the amplitude variations is the precessional period. The period of the sine variations is very nearly the spin period. To determine the z angle, then, from the gyro data, the instantaneous indicated angle readings are used to determine an angle that would be the indicated maximum of a sine variation if the maximum had obtained at the time in question. The unknown in the equation is the spin position, which is readily obtained from the raw gyro data. The angle thus obtained is the zenith angle. This angle information, combined with data concerning the trajectory of missile center of gravity, permits a determination of the angle of attack of the missile. The gyroscope data is recorded in flight on film by a 16-mm motionpicture camera. Early flights utilized a standard GSAP gun camera, which was later discarded due to its unreliability. Later flights used a type B-2 navigation camera modified to the extent of providing take-up spool armor and a more reliable driving motor. A photograph of this modified camera is included here as Fig. 4. 9

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-- ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN Representative gyro data are likewise included in this report. The angles specified are those occurring between the missile's (Aerobee) longitudinal axis and the gyro rotor axis. The independent variable is the frame number, which may be converted approximately to seconds from launching instant by dividing by the frame rate of 7.2 frames per second. See Table 1. 11

TABLE 1 DATA FOR DECEBEO_, 1950 Frame Angle Frame Angle Frame Angle Frame Angle Frame Angle 0 0 50 4.72 109 9.1 158 9.4 207 11.4 12 0 51 4.22 110 9.1 159 9.9 208 11.5 2 0 52 4 1ll 7.8 160 10.5 209 10.7 4 0 53 4.30 112 7.1 161 10 210 10.4 5 0 54 3.60 113 6.1 162 10.4 211 9.4 6 0 55 3.49 114 6.0 163 11.2 212 8.9 7 0 56 3.04 115 5.6 164 9.1 213 9.1 8 0 57 3.36 116 5.0 165 8.1 214 9.9 9 0 58 3.60 117 5.1 166 8.1 215 10.7 10 0 59 4.32 118 5.1 167 6.9 216 11.4 11 0 60 5.53 119 6.o 168 6.2 217 12.1 12 0 61 5.9 120 5.9 169 7.0 218 12.7 13 0 62 6.13 121 6.9 170 6.7 219 12.2 14 0 63 6.37 122 7.7 171 7.3 220 12.3 15 0.2 64 7.18 123 7.8 172 8.9 221 12.47 16 0.7 65 8.51 124 8.2 173 9.8 222 12.00 17 0.9 66 8.05 125 9.7 174 10.4 223 10.5 18 +1.0 67 8.59 126 9.6 175 10.7 224 8.54 19 2.05 68 8.07 127 9.6 176 10.9 225 9.00 20 2.05 69 6.84 128 9.0 177 ll.0 226 8.97 21 4.12 70 5.83 129 8.5 178 10.4 227 8.84 22 4.15 71 5.9 130 7.3 179 10.9 228 9.48 23 4.0 72 5.14 131 7.4 180 10.4 229 10.18 24 5.0 73 4.46 132 6.3 181 10.0 230 11.5 25 5.36 74 3.67 133 5.8 182 8.6 231 12.1 26 4.17 75 3.0 134 5.7 183 8.7 232 12.0 27 3.16 76 3.60 135 6.o 184 7.6 233 11.29 28 2.6 77 4.72 136 5.8 185 9.1 234 11.05 29 2.235 78 5.93 137 6.o 186 8.7 235 8.54 30 2.32 79 6 138 7.1 187 9.1 236 8.04 31 1.49 80 6.7 139 8.0 188 10.5 237 8.4 32 1 81 7.78 140 8.0 189 10.7 238 9.33 33 1.14 82 8.53 141 8.8 19o 12.1 239 9.49 34 1.80 83 8.oo 142 9.8 191 12.0 240 7.61 35 9.87 84 7.99 143 10.3 192 11.1 241 9.8 36 3.165 85 7.2 144 10. 0 193 11.7 242 10.2 37 3.83 86 7.03 145 9.5 194 11.0 243 11.94 38 4.0 87 6.09 146 9.1 195 10.7 244 10.79 39 4.12 88 6.o 147 8.7 196 lo.o 245 10.68 40 4.29 89 4.5 148 7.8 197 9.8 246 9.02 41 4.85 100 3.2 149 7.3 198 9.7 247 10.24 42 5.52 101 2.92 150 6.3 199 9.1 248 9.55 43 5.8 102 3.9 151 5.0 200 9.8 249 9.19 44 6.41 103 5.1 152 5.2 201 9.95 250 8.52 45 7.09 104 6.1 153 5.4 202 11.2 251 9.45 46 7 105 6.7 154 6.1 203 11.2 252 8.96 47 6.19 106 6.7 155 7.6 204 12.5 253 10 48 5.64 107 8.1 156 8.3 205 12.7 254 11.3 49 5.7 108 8.7 157 9.1 206 12.2 255 12.6 12

TABLE 1 (cont.) Frame Angle Frame Angle Frame Angle Frame Angle Frame Angle 256 12.38 307 9.5 358 11.33 409 21.18 460 20.55 257 11.4 308 10.79 359 7.0 410 21.14 461 18.1 258 11.31 309 12.29 360 5 411 20.45 462 16 259 9.22 310 11.88 361 6.70 412 16.07 463 14.4 260 8.74 311 11.13 362 9.25 413 13 464 10.19 261 7.63 312 11.03 363 11.13 414 9.82 465 10.13 262 7.00 313 8.56 364 13.3 415 7.0 466 9.52 263 8.13 314 8.06 365 15.48 416 5.81 467 8.97 264 8.24 315 5.39 366 15.18 417 5.1 468 8.25 265 10 316 7.1 367 14.65 418 6 469 7.43 266 11.33 317 9.48 368 12.95 419 7.03 470 4.66 267 11.23 318 11.4 369 11.85 420 7.58 471 2.83 268 10.5 319 13.85 370 10.05 421 9.15 472 2.83 269 10.48 320 13.79 371 8.o 422 10.43 473 4.93 270 9.8 321 13.1 372 5.02 423 13.35 474 7.96 271 8.5 322 11.29 373 5.45 424 17.65 475 10.0 272 7.15 323 7.3 374 8.54 425 20.8 476 10.37 273 8.25 324 5.0 375 11.79 426 21.0 477 10.53 274 8.05 325 6.33 376 16.1 427 22.4 478 10.33 275 8.72 326 8.72 377 17.42 428 22.58 479 11.32 276 10 327 12.02 378 19.6 429 21.3 480 12.95 277 12.1 328 14.9 379 19.6 430 20.75 481 15.48 278 11.58 329 14.48 380 16.3 431 19.4 482 17.7 279 11.0 330 13.02 381 13.17 432 18.6 483 20.28 280 11.1 331 11.59 382 9.35 433 19.53 484 21.0 281 9.95 332 7.09 383 4.18 434 16.72 1485 22.13 282 8.5 333 4.24 384 3.21 435 15.5 47S6 21.32 283 7.42 334 5.5 385 5.81 436 13.2 487 20~72 284 7.28 335 10.8 386 8.36 437 9.8 488 20.9 285 8.04 336 14.6 387 10.33 438 7.75 489 21.47 286 9.0 337 16.62 388 12.05 439 3.16 490 22.55 287 10.2 338 17 389 13.65 440 3.35 491 23.88 288 11.72 339 14.9 390 15.03 441 4.24 492 24.9 289 11.48 340 12 391 16.98 442 7.06 493 25 290 10.5 341 7.5 392 19.38 443 7.63 494 23.8 291 11.23 342 3.04 393 20.4 444 8.13 495 22.88 292 8.91 343 5.5 394 18.65 445 8.93 496 19.88 293 8.05 344 10.58 395 17 446 9.99 497 20.25 294 7.2 345 15.25 396 13.4 447 12.53 498 18.87 295 7.15 346 17.65 397 9.06 448 16 499 18.12 296 8.56 347 17.77 398 5.38 449 18.88 500 19.57 297 10.05 348 16.70 399 1.41 450 20.52 501 19.4 298 11.48 349 13.17 400oo 5.1 451 21.3 502 19 299 12.49 350 7.78 401 8.25 452 20.3 503 16.45 300 11.28 351 2.24 402 10.75 453 20.75 504 14.05 301 10.73 352 5.66 403 12.94 454 19.46 505 11.68 302 9.66 353 11.0 404 13.3 455 20 506 9.84 303 9.03 354 16.5 405 15.48 456 20.1 507 8.91 304 7.25 355 17.8 406 18 457 21.05 508 8.24 305 7 -29 356 17.55 407 17.4 458 21.77 509 6.76 306 9.17 357 15.11 408 20.15 459 21.24 510 8.54 13

TABLE 1 (cont.) Frame Angle Frame Angle Frame Angle Frame Angle Frame Angle 511 7.28 562 24 612 21.68 663 34.35 714 25.6 512 5.00 563 22.5 613 21.15 664 35.85 715 25.0 513 3.16 564 19.8 614 20.81 665 37.55 716 24.73 514 1.5 565 18.6 615 21 666 37.6 717 27 515 3 566 16.25 616 20.2 667 35.3 718 27.28 516 4.61 567 14.96 617 21.5 668 35.1 719 28.35 517 7 568 16.07 618 24 669 32.2 720 26.5 518 8.04 569 16.69 619 26 670 35 721 24.3 519 7.3 570 15.62 620 28.55 671 29.1 722 23.6 520 7.62 571 16 621 26.95 672 34.3 723 22.0 521 8.25 572 13.89 622 25.35 673 35.85 724 21.2 522 9.38 573 11.87 623 25.39 674 37.15 725 22.25 523 11.18 574 9.47 624 24.3 675 36.63 726 23.05 524 13.52 575 7.6 625 25 676 35.3 727 24.0 525 15.18 576 5.8 626 24.68 677 34.1 728 24.6 526 17.28 577 7.75 627 28.08 678 33.05 729 21.68 527 16.7 578 7.16 628 29.35 679 35 730 21.17 528 17.03 579 7.6 629 31 680 33.3 731 18.58 529 16.48 580 7 630 31 681 34.3 732 18.6 530 18.24 581 5 631 28.7 682 35.05 733 18.4 531 18.62 582 3.16 632 28.37 683 36.65 734 17.5 532 20.5 583 1 633 27.42 684 35.55 735 20.05 533 22.9 584 1 634 29 685 33.25 736 21.4 534 24.1 585 1.8 635 27.32 686 33.1 737 20.73 535 25.13 586 3 636 30.1 687 31.65 738 19.52 536 24.75 587 3.61 637 32.8 688 33 739 18.11 537 24.08 588 3 638 32.3 689 30.6 740 16.26 538 22.85 589 2.06 639 11.85 690 32.3 741 14.83 539 23.75 3.17 64o 31.35 691 34.3 742 15.12 540 23.3 590 5.38 641 30.65 692 34.3 743 14.07 541 25.58 591 7.1 642 29.8 693 32.85 744 17 542 27.42 592 9.65 643 30/30 694 32.07 745 16.05 543 29.1 593 10 644 31.9 695 31.1 746 16.48 544 30 594 11.33 645 32.3 696 27.55 747 16.08 545 28.15 595 10.28 646 34.3 697 30 748 14.7 546 27.5 596 10.2 647 34.6 698 29.8 749 13.03 547 26.2 597 10.25 648 33.8 699 31.1 750 11.93 548 25.0 598 10.42 649 33.07 700 31.6 751 11.0 549 24.5 599 11.12 650 33.1 701 30.95 752 10.93 550 26.6 6oo00 14.15 651 31.7 702 30.9 753 12 551 27.1 601 15.45 652 32 703 29.1 754 13.65 552 29.12 602 17.1 653 26.7 704 28.32 755 12.6 553 30 603 17.07 654 33.3 705 27.4 756 12.6 554 26.35 6o4 16.85 655 35.1 706 29 757 11.32 555 26.8 605 16.7 656 33 707 27.24 758 10.5 556 23.3 606 16.18 657 34.75 708 30.1 759 8.98 557 22.23 607 17.35 658 33-35 709 30.1 760 9.07 558 21.3 608 17.8 659 34.55 710 31.15 761 8.16 559 23.8 609 21 660 33 711 29.05 762 9.06 560 24.5 610 22.75 661 34 712 27.49 763 10.23 561 24.32 611 22.0 662 33.25 713 26.4 764 11.12 14

TABLE 1 (cont.) Frame Angle Frame Angle Frame Angle Frame Angle Frame Angle 765 lo0 816 21.5 877 40.8 919 61.6 970 69.o 766 9.94 817 22.03 878 40.55 920 56.3 971 78 767 8.54 818 23.78 879 43.65 921 60.7 972 79 768 7.76 819 26.0 880 41.4 922 58.6 973 70.68 769 6.38 820 27.1 881 46.15 923 60 974 74.25 770 6.32 821 25.8 882 44.6 924 57.45 975 71.7 771 6.7 822 25.0 883 48 925 63.0 976 76 772 8.1 823 23.8 884 45.4 926 64.68 977 72.6 773 9.85 824 23.45 885 45.5 927 66.o 9.78 73.3 774 10.42 825 24.1 886 43.15 928 67.9 979 72.4 775 10.3 826 25.39 887 43.0 929 60o.5 980 80 776 9.85 827 27.65 888 46.o 930 59.55 981 82 777 8.53 828 29.38 889 46.2 931 60.95 982 73.6 778 7.34 829 30 890 48.9 932 65.0 983 74.6 779 7.27 830 25.05 891 47.7 933 58.7 984 70.7 780 6.37 831 28.62 892 50 934 66.05 985 80 781 8.25 832 28.3 893 47 935 62.9 986 75.4 782 9.83 833 29.88 894 48.6 936 68 987 77.15 783 11.58 834 29.5 895 46.6 937 65.9 988 75.14 784 12 835 29.78 896 47.5 938 63.35 989 84 785 11.6 836 32.5 897 48.6 939 59.35 990 83 786 10.72 837 33.6 898 48.2 940 72.13 991 75.85 787 9.95 838 32.0 899 52.05 941 69.15 992 75.15 788 9.38 839 32.4 900 50.7 942 61.5 993 80.09 789 lo0.0 840 32.05 901 54 943 68.3 994 82 790 10.22 841 30.65 902 48.7 944 67.1 995 77.32 791 12.15 842 31.4 903 52-7 945 72 996 80.04 792 14.0 843 32.4 904 48 946 66.35 997 80.02 793 15.2 844.32.8 905 52 947 66.2 998 88 794 14.45 845 35.2 906 50.1 948 61.1 999 86 795 13.8 846 36.6 907 51.8 949 70 1000 79.35 796 13.31 847 36.55 908 53.2 950 67.4 1001 78.78 797 12.7 848 35.8 909 56.o 951 65.2 1002 82 798 12.65 849 36.1 901 57 952 68.3 1003 85 799 14.09 850 35.25 902 50.75 953 70.5 o1004 82.1 800 17.05 851 38 903 56 954 74 1005 82.04 801 17.4 852 35.65 904 50.6 955 67.4 1006 85.12 802 18.4 853 38.2 905 57 956 70.55 1007 900 803 19.55 854 38.9 906 53.2 957 65.4 1008 900 804 18.1 855 42.15 907 55.95 958 72 1009 84.2 805 17.4 856 41.0 908 55.35 959 70Q35 1010 82.33 806 16.42 857 37.8 909 59.0 960 67.1 loll 86 807 17.28 858 38.5 910 60.2 961 69.55 1012 89 808 17.78 859 42.6 911 53.8 962 73.2 1013 88 809 20.05 870 41 912 58.1 963 77 1014 83.02 810 22.05 871 39.25 913 59.8 964 69.4 1015 900 811 22.35 872 41.5 914 58 965 73.05 lo16 88 812 21.67 873 41.8 915 58.3 966 68.6 1017 89 813 21.23 874 45.35 916 59.25 967 75 1018 90 814 21.0 875 42.2 917 56.2 968 71.5 1019 85.03 815 20.08 876 41.6 918 62.2 969 69.8 1020 90 15

TABLE 1 (cont.) Frame Angle Frame Angle Frame Angle Frame Angle Frame Angle 1021 89 1072 73.58 1123 52.7 1174 46.o 1225 53.1 1022 89 1073 79.1 1124 58 1175 47.2 1226 55.6 1023 89 1074 81 1125 53.4 1176 48.0 1227 51.6 1024 85.17 1075 73.98 1126 60.2 1177 48.4 1228 57.0 1025 82.49 1076 73.02 1127 57.2 1178 48.6 1229 54.1 1026 89 1077 68.78 1128 54.6 1179 45.9 1230 60.4 1027 89 1078 75.00 1129 52.6 1180 48.o 1231 56.2 1028 90 1079 71.48 1130 53.2 1181 43.4 1232 56.7 1029 90 1080 82.05 1131 53.4 1182 46.5 1233 51.8 1030 89 1081 66.40 1132 49.4 1183 45.8 1234 58.4 1031 90 1082 75.09 1133 51.6 1184 49.0 1235 57.2 1032 90 1083 75.96 1134 51.8 1185 43.7 1236 56.3 1033 90 1084 67.3 1135 57.0 1186 57.6 1237 58.7 1034 900 1085 70.4 1136 51.8 1187 55.7 1238 61.1 1035 85.53 1086 67.9 1137 52.7 1188 46.3 1239 61.9 1036 90 1087 71.0 1138 49.6 1189 46.8 1240 57.7 1037 90 1088 62.9 1139 52.6 l190 45.8 1241 61.0 1038 89 1089 70.03 1140 51.5 1191 48.4 1242 55.5 1039 87.03 1090 64.3 1141 48.9 1192 45.3 1243 60.1 1040 88.02 1091 73.95 1142 52.0 1193 50.2 1244 58.1 1041 87 1092 69.45 1143 51.8 1194 47.6 1245 61.5 1042 900 1093 63.85 1144 54.3 1195 50.0 1246 59.8 1043 87 1094 62.05 1145 50.1 1196 47.8 1247 65.4 1044 83.7 1095 67.3 1146 51.0 1197 49.0 1248 66.3 1045 83.72 1096 67.15 1147 45.3 1198 46.3 1249 58.4 1046 89 1097 60.2 1148 50.2 1199 48.3 1250 62.6 1047 89 1098 67.3 1149 47.2 1200 48.6 1251 56.2 1048 86 1099 82.4 1150 50.4 1201 48.7 1252 65.3 1049 82.07 lloo0 1151 48.o 1202 50.4 1253 60.3 1050 81.32 1101 63.4 1152 49.8 1203 45.2 1254 66.o 1051 84.o 1102 69.2 1153 50.9 1204 53.2 1255 61.3 1052 85.O 1103 56.7 1154 48.3 1205 49.0 1256 69.4 1053 79.65 1104 62.9 1155 48.o 1206 51 1257 58.8 1054 82.53 1105 79.4 1156 45.7 1207 45.4 1258 62.1 1055 84 1106 58.2 1157 45.1 1208 50.2 1259 62.6 1056 85 1107 59.5 1158 44.7 1209 48.1 1260 67.0 1057 84 1108 61.4 1159 51.0 1210 52.4 1261 67.3 1058 78.3 l109 65.2 1160 47.4 1211 50.0 1262 62.1 1059 75.03 1110 57.7 1161 50.1 1212 54.8 1263 69.1 1060 83 1111ll 68.1 1162 45.o 1213 54.8 1264 66.5 1061 81 1112 55.9 1163 46.9 1214 49.7 1265 73.3 1062 73.25 1113 76.0 1164 46.0 1215 52.4 1266 69.8 1063 79.28 1114 55.0 1165 47.0 1216 50.2 1267 66.o 1064 80.1 1115 72.4 1166 48.o 1217 52.7 1268 62.6 1065 82 1116 55.6 1167 45.5 1218 47.0 1269 73.1 1066 77.32 1117 60.3 1168 48.8 1219 55.0 1270 71.3 1067 78.03 1118 60.6 1169 47.3 1220 50.4 1271 65.9 1068 69.47 1119 50.1 1170 49.7 1221 58.2 1272 70.9 l069 78 1120 56.6 1171 46.3 1222 54.7 1273 74.2 1070 77.1 1121 54.6 1172 43.2 1223 53.4 1274 77.0 1071 69.57 1122 61.2 1173 47.2 1224 52.4 1275 71.3 16

TABLE 1 (cont.) Frame Angle Frame Angle Frame Angle Frame Angle Frame Angle 1276 73 1289 74.7 1302 83.2 1315 1328 1277 69.3 1290 73.3 1303 81.1 1316 1329 1278 75.2 1291 84.o 1304 85.4 1317 85.1 1330 1279 72.2 1292 84.o 1305 86.6 1318 1331 87.0 1280 68.9 1293 77.1 1306 85.0 1319 1332 82.7 1281 68.6 1294 79.2 1307 68.o 1320 1333 89.o 1282 79.0 1295 80.1 1308 89.4 1321 1334 1283 80.0 1296 84.2 1309 1322 49.0 1335 88.1 1284 73.1 1297 84.0 1310 1323 1336 83.0 1285 76.1 1298 78.3 1311 1324 1337 70.5 1286 73.6 1299 80.5 1312 1325 1338 1287 78.2 1300 79.2 1313 1326 51.9 1339 1288 77.2 1301 1314 1327 70.7 1340 84.2 17

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN 4. SUMMARY OF ACCESSORY EQUIPMENT As noted previously, many accessory devices are necessary in the instrumentation in addition to the basic measuring instruments. (1) Telemeter input-voltage limiters are used to prevent the application of any voltage that lies outside the limits of 0 to +5 volts d-c to the telemeter. (2) Panel meters are provided to present certain data, some of which are also recorded by telemetering, to the gyro camera. Some cross checking of data-recording accuracies is thus possible. (3) An acceleration switch is used to indicate missile take-off, end of boost phase, and end of sustaining-motor phase. The switch trips when the acceleration exceeds 2-3 g's, closing the circuit to a small light which is viewed by the gyro camera. (4) Two clocks, one normal spring-wound and the other motor-driven, are included to establish a time base for the camera film. (5) Thermistors are included in some instances for the determination of various temperatures, chiefly alphatron-chamber temperature. (6) Mercury cell battery packs are utilized to provide constant, known voltages for in-flight telemeter calibration. (7) A vibrator-type inventor serves to provide a source of llOv 400w voltage for gyroscope operation. Its input and output are filtered. (8) Rotary solenois selector switches are utilized for all switching operations, including disconnection of pull-off control leads immediately preceding missile take-off. (9) Binary coded neon light systems are used to present alphatronsystem range information tothe camera. 18

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN 5. AIR DENSITY MEASUREMENT BY UHF GAS-DISCHARGE BREAKDOWN During the course of the contract, an investigation was initiated into the feasibility of determining ambient air density by means of a measurement of the RF field strength necessary to cause a gas-discharge breakdown of the air at a point several inches in front of the nose of an Aerobee missile. Laboratory experiments were conducted which resulted in causing the desired breakdown in an evacuated chamber. Although it appeared to project personnel that such an experiment was feasible and that useful results could be obtained, contract support for the work was withdrawn. The work was therefore not continued under this contract. 19

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN 6. SUMMARY BY ROCIKT FLIGHT This section will summarize the work under the contract by reference to the rocket flights in which the research group participated. 6.1. V-2 No. 47, June 1949 A single thermionic ionization gage with newly developed circuitry was utilized in this experiment. The primary purpose of the instrumentation was to provide a field test of techniques that were to be used in later Aerobee flights. The experiment was considered successful and the equipment was recovered intact. No pressure or temperature data were anticipated nor obtained from this test. A photograph of the recovered equipment is included in this report as Fig. 5, 6.2. Aerobee USAF-3, December 1949 The following basic equipment was installed for this flight: (1) Four thermionic ionization gages, each provided with independent automatic control of filament emission to permit reliable and stable operation (10-2 mm Hg to approximately 10-6 mm Hg). (2) Two alphatron-type ionization gages each supplied with an automatic range-seeking circuit for presenting an appropriate voltage to the telemeter in accord with the particular pressure being measured at any given instant. These gages were intended to measure the pressure from the ground level to approximately 10-3 mm Hg, slightly overlapping the thermionic gages. (3) A Sperry attitude gyro to provide information regarding the aspect of the rocket during the useful portion of its flight. Certain difficulties in the missile fuel-control system caused a loss of power a few seconds subsequent to launching. This resulted in excessive missile yaw, causing ejection of the nosepiece and instrumentation, which fell to the ground and was demolished. The experiment was, of course, unsuccessful. 20

BROKEN TERMINAL7 BRACKET TORN LOOSE t BROKEN SEAL Fig. 5. Photograph of Equipment Recovered from V-2 No. 47.

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN 6.3. Aerobee USAF 6, June 1950 The purpose of this flight was to provide a test of missile control functions, and at the same time obtain upper-atmosphere data. Therefore, a portion of the instrumentation (not supplied by this project) was to be used to monitor the operation of the normal equipment carried by the Aerobee, including beacon, cut-off system, ejection system, and related equipment. The balance of the instrumentation, for which this project was responsible, was to be used for upper-atmosphere experiments. It was decided to include two alphatron ionization gages and the aspect determination system using the Sperry attitude gyro. The alphatron gages were mounted so that one would measure ram pressure and the other cone wall pressure, an arrangement suitable for ambient-temperature calculation. The forward portion of the warhead was to be used for the upper-air experiments, so it was decided that the special magnesium nose section this project had constructed for previous warheads would be employed. Accordingly, two alphatrons and their associated circuits were installed in the magnesium cone section (see photographs in Figs. 6 and 7). The circuits used were very similar to those used in the past, with some modifications. For example, the range-switching circuit was slightly altered for more stable operation and also for the incorporation of a time delay. It was felt that due to the spinning of the missile, the instantaneous pressure in the conewall alphatron could vary sufficiently to cause the system to change range in the roll period, under conditions of high yaw. Such a variation and consequent rapid switching would render the data useless. A time delay of approximately two seconds was therefore introduced to delay switching. The aspect determination equipment consisted of a Sperry attitude gyro, type F4A, and a project-modified gun camera used for photographing the gyro sphere. In addition tothe gyroscope, the camera photographed four panel meters and a clock. Two of these meters were Simpson 2-inch square O-l-ma meters, and the other two were Marion "ruggedized" O-l-ma meters. See Fig. 8. One Marion and one Simpson were connected in parallel and indicated the range position of the ram alphatron. These data were transmitted by telemetering as well, the object of the parallel connection being to give some indication of the operation of each meter as compared to telemetering, and also some information relative to the general usefulness of the camera - panel-meter system of data recording. 22

:~~~ D 1: AL PHATRON Fig. 6. View of Alphatron Instrumentation, June 1950. 23

Fig. 7- View of Alphatron Instrumentation, June 1950. 24

25l:00~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~; ff~ ~ ~ ~ ~~i:ID- ----:~_~~~~i__~~~ AS:__:::::j~~:-_::::i.::~(i::i_: —: l: A::~~~~~~ ~ ~ ~~ ~::~ X~:;:~:; 0:ff::00f0:ff f~fr0f::::::::~:~::J —-:-::; Fig. 8. Gyro and Meter Panel as Viewed by Camera, June 1950.

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN The third panel meter recorded the output of the ram pressure unit, which data were likewise displayed by telemeter. The missile was fired at approximately 0840, Tuesday, June 20, and was very successful. Maximum altitude was fifty-seven miles, and most of the instrumentation was recovered intact. This project's equipment operated satisfactorily from take-off until warhead separation at about 160 seconds (the peak of the trajectory), at which time both alphatron units failed. This, however, is approximately the time (zenith) at which the pressure data ceases to be useful, and so it is not considered that there was any loss of information. The shock of warhead blowoff was apparently very severe and caused the failure, the nature of which was not apparent from the recovered equipment. Unfortunately, the telemeter commutator also stopped at warhead blowoff, so that certain operational inforimation that was presented to the telemeter was not available for analysis. The gyro, camera, clock, and panel meters were undamaged by the nosecone blowoff and operated satisfactorily throughout the flight. All the film was used, resulting in a full ten minutes of gyro and meter information. The film supply was exhausted about one minute before impact. Upper-atmosphere pressure and temperature data resulting from this flight are tabulated in Table 2 and plotted in Figs. 9 and 10. It should be noted that the computation of many additional points is possible from the recorded data. In general, however, due to the magnitude of the computation, intervals of 1-2 kmn, are chosen, as this defines the curve adequately. 6.4. Aerobee USAF 10, December 1950 In the early stages of this project's participation in the Air Force Aerobee program, two nearly identical warheads were instrumented for a double firing scheduled in December 1949, one for a daytime firing and the other for a nighttime firing. The first of the two firings was unsuccessful because of a flight failure (USAF 3 as noted previously). The second warhead was used for the December 1950 firing. The instrumentation consisted of two alphatron-type ionization gages for cone-wall pressure measurement from ground level to approximately fifty miles; four thermionic ionization gages for cone-wall pressure measurement from about twenty miles to rocket ceiling; a single thermionic ion gage mounted in the forward portion of the tank section of the missile for measuring the surface pressure at this point; an attitude gyro, Sperry type F4A, with camera for zenith-angle determination throughout the flight; and a single phototube mounted in the rocket skin to assist in determining roll position of the missile. -26

TABLE 2 UPPER ATMOSPHERE DATA FOR JUNE 20, 1950 Measured Derived* Alt. Pressure g m2 Drived log Pmp. log A.km mb aedynes/cm Tem. dynes/cm2 -log10lo Pd 30 12.5 4.095 234 4.095 0.00 32 9.33 3.969 235 3.970 -0.001 34 7.01 3.846 238 3.847 -0.001 366 3.730 250 3.730 0.00 38 4.13 3.616 260 3.617 -0.001 40 3.21 3.507 265 3.507 0.00 42 2.49 3.397 265 3.397 0.00 44 1.94 3.288 268 3.288 0.00 46 1.51 3.178 268 3.179 -0.001 48 1.18 3.070 268 3.070 0.00 50 0.914 2.961 267 2.961 0.00 52 0.709 2.850 265 2.851 0.001 54 0.548 2.739 261 2.739 0.00 56 0.421 2.625 256 2.625 0.00 58 0.321 2.507 249 2.508 -0.001 60 0.242 2.386 241 2.387 -0.001 62 0.182 2.262 231 2.261 0.001 64 0.133 2.123 218 2.128 -0.005 66 0.0964 1.983 202 1.984 -0.001 68 0.0681 1.834 194 1.834 0.00 70 0.0479 1.681 189 1.681 0.00 72 0.0336 1.526 187 1.526 0.00 *The measured temperature values are utilized in an integrating computation employing the hydrostatic equation to obtain the "derived" pressure values. The derived pressure values may then be compared with the actual measured values. This procedure allows an overall check of the data analysis. The derived pressures compare very favorably with the measured values as shown by the last column of figures. 27

I0 39 ~ M8 6 5 - 4 3 2 W 10 Z) 8 75 2 30 40 50 60 70 80 90 ALTITUDE (KM) Fig. 9. Curve of Ambient Pressure Vs. Altitude, June 1950. 28

270 260 _ 260 ROCKET PANEL ADOPTED VALUES 0 CL 220 w 220 21 200 190 180.. 17530.0 35.0 40.0 45.0 50.0 55.0 60.0 65.0 70.0 75.0 ALTITUDE (K.M.) Fig. 10. Curve of Ambient Temperature Vs. Altitude, June 1950. 29

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN The alphatron system used was similar to previous installations except that the range-changing circuit was modified to include a switching time delay as noted previously. The purpose of this delay was to prevent a range change when the output momentarily goes off-scale, as might be the case with a large angle of attack. Under a condition of large angle of attack one can expect that the instantaneous pressure at the gage port may change by as much as a factor of ten or twenty as the missile rolls, with consequent amplifier output swing. Without the time delay, the system would be forced to switch to another range momentarily and then return to the original range. This switching sequence is unnecessary and undesirable, for the useful data exist in the normal range. Hence the time delay system was incorporated. A portion of the ionization gage system employed four thermionic ion gages, mounted for cone-wall pressure measurement similar to the alphatrons. Each gage with its associated circuit was independent of the other gages, so that a possible failure in one would not result in failure of all. The emission current of each gage was controlled by a special servo unit detailed in early progress reports. A voltage indication of the grid and plate current of each gage was presented to the telemeter, as was the output and range indication of each alphatron unit. In addition to photographing the gyro, the camera photographed six panel meters, a light actuated by an acceleration switch, two neon lights, and a clock. The output and range of one alphatron unit and the grid and plate currents of two of the ionization gages were applied to the six panel meters as well as to the telemeter as an alternate means of recording the data. This flight outwardly was very successful. The instrumentation apparently operated properly as intended. However, the gyro information indicated very large angles of attack and this was verified by the pressure data, which exhibited very large variations about the expected mean. These variations were great enough (factor of 10-20) to preclude a normal data-reduction procedure. Accordingly the reduction of data from this firing was postponed to a later date. As of the time of writing of this report, the work has not been resumed, due to pressure of other data reduction from later firings. Photographs of the instrumentation of this flight are included as Figs. 11 and 12 of this report. The only item recovered from this flight was the missile camera film. 50

c~ Lr~ on cH,N~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~a ~ C) a) q.......................~~~~~+ ----------— ~~~l) t mg~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~C

Fig. 12. Portion of Instrumentation, December 195U. 52

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN 6._2. USAF 18, September 1951 Aerobee USAF 18 was instrumented as follows: 1. Three alphatron ionization gages for pressure measurement, two mounted for indication of cone-wall pressure, and the third mounted for impact (rampressure) measurement. 2. A special controlled air gap, as a portion of the nose cone, for measurement of breakdown potential. This measurement allows computation of air density. 3. A gyroscope for measurement of zenith angle of the missile, from which it is possible to compute the angle of attack of the missile with the air stream. 4. A type B-2 aircraft 16-mm motion-picture camera for recording gyro data. 5. Four panel meters photographed by the camera for additional data recording. 6. An acceleration switch used to indicate time of take-off, end of boost, and end of burning of the missile-sustaining motor. 7. Equipment comprised of two fine wires mounted in the air stream off the surface of the nose cone, to obtain data leading to information regarding stream velocity profile in the boundary layer. 6.,5.1. Alphatron System. A newly designed alphatron gage was utilized in the equipment for this test, and other modifications to previous alphatron systems were effected. Since these changes represent advances in the technique of alphatron use, some discussion of them is included in this report. The modifications effected included, in addition to the new gage unit, an increase in the number of subranges from five to seven, a redesign of the "zeroing" system to allow recording of the zero reading during flight, and slight changes in the switching circuit in regard to time delays. In previous systems used by the project in Aerobee firings, the amplifier zero reading was obtained (and correction made) only immediately prior to flight by simulating a known value of pressure (in this case, zero pressure) and recording the corresponding value of output signal. To accomplish this, it was necessary to disconnect the range-switching circuit (which inserts an appropriate value of alphatron load resistance consistent with the pressure being 35

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN measured) from the amplifier and connect it to the circuit ground reference. This operation caused the switching circuit to insert the lowest resistance, zero ohms, which simulated zero pressure. The time involved merely to observe the zero reading was so long (of the order of ten seconds) that the possibility of recording the zero reading during flight was precluded. The system was altered, therefore, to allow observance during flight, the time required being reduced to approximately one second. A small auxiliary relay was installed so that when it was energized, a d-c voltage was impressed across the alphatron load resistance. The amplifier then sees a voltage as if some unique value of pressure were impressed upon the gage, and responds accordingly. The voltage used is of such magnitude that the amplifier output indicates approximately at midscale of the normal 05-volt range, this indication then being considered as the zero reading of the system. Note that the zero reading may be obtained regardless of the value of input resistance that may be in use (except zero ohms), for the battery voltage is of course unaffected by the high resistance that it may shunt. Therefore, it is possible to obtain a zero reading during flight without disturbing the amplifier substantially, for the switching circuit does not know that any change has occurred. This system was utilized for the three alphatron circuits included in the instrumentation of the September Aerobee. A motor-driven cam switch actuated the zero relays for approximately one second every thirty seconds, allowing the recording of any zero shift and hence correction in the analysis of the data. (Any shift is merely added algebraically to a given reading.) In regard to the new style of alphatron gage used for this instrumentation, as compared with the units used previously, the volume occupied is reduced by a factor of ten, the activity of the radium source is much greater, and the overall sensitivity of the unit is greater by a factor of approximately five. The advantage of the reduced volume is obvious; however, the increased activity of the source poses a moderate problem from the standpoint of the health hazard involved. The AEC standard dosage allows 10 milliroentgens (mr) per hour for 8 hours, with a limit of 3 consecutive days, or a total of 300 mr per week. The new type of alphatron contains 3 mg of radium in equilibrium with decay pro- - ducts. The measured radiation rate (gamma) at the surface of the gage is 400 mr/hr. At a distance of approximately 15 inches, the rate is 10 mr/hr., the 8hr. tolerance value. It is obvious, therefore, that extreme care must be exercised in the handling of gages to avoid personnel injury. The increased sensitivity, however, is a considerable advantage, for it allows a reduction in the magnitude of the load resistances required, from 500,000 megohms to 100,000 megohms (for the lowest pressure range desired), which assists greatly in solving construction problems in regard to leakage which may be complicated by humidity effects, for example.

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN The new type of gage will not respond to as low a pressure as the previous model, in spite of the increased sensitivity, for the increased activity of the source produces a greater "dark current", a relatively small current produced in the load resistance as a result of particles emitted by the radium colliding with, and producing secondary emission from, the electrodes of the gage. The construction of the amplifier input stage (electrometer stage) was revised for this flight to allow pressurization. A question has arisen on previous flights regarding the possibility of condensation in this portion of the system, resulting in calibration errors. It was felt, therefore, that it would be desirable to pressurize this portion of the instrumentation for at least one flight to answer this question. A photograph of the completed unit without pressurizing can appears in Fig. 13. The upper-atmosphere data resulting from use of this equipment appears in tabular form in Table 3, and in graphical form in Figs. 17 and 18. 6.2.2. Boundary-Layer Experiment. An exploratory experiment designed to indicate mass-flow conditions in the boundary layer of air surrounding the missile during flight was prepared for the Aerobee launched in September 1951. Two fixed wires were mounted adjacent to the surface of the nose of the missile, approximately 2 feet from the tip. One wire was located approximately 1/16 inch from the surface and the other about 1 inch from the surface. The wires used were tungsten, very small in diameter (0.0005 inch) and approximately 1/4 inch long. The inner wire was heated by the passage of a constant current, and the voltage drop was measured, providing an indication of the temperature of the wire. An amplifier was included which allowed recording of the information by telemeter. The outer wire was not heated, but its continuity was recorded, likewise by telemetering. Upon examination of the results, it was learned that the inner heated wire remained intact approximately 30 seconds after launching, while the outer wire lasted 40 seconds. This is longer than had been anticipated, for it was expected that the wireds would probably not survive the boost phase of the flight. The experiment was considered successful, as it indicated probable success for a future experiment developed from this fundamental device. 6.5.3. Paschen-Law Density Experiment. An independent experiment developed for cone-surface stream density determination was included in the instrumentation of this flight. 355

:!......ATR....A i~~~~~~i?!?i? ~~~~~~~~ ~ ~ ~ $frR (DIN ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~~~~MQNE.I EE Fig. 13. Cone Wall Alphatron Electrometer Stage, September 1951.

TABLE 3 UPPER ATMOSPHERE DATA FOR SEPTEMBER 13, 1951 Measured loglo P loglo P log lo Pm Alt. Pressure dynes/cm2 Temp. dynes/cm2 -log10o Pd km mb OK 30 12.4 4.093 226 4.092 +0.001 32 9.24 3.966 226 3.965 0.001 34 6.83 3.834 228 3.834 o 36 5.21 3.717 251 3.718 -0.001 38 4.07 3.610 272 3.610.000 40 3.20 3.505 277 3.504 +.001 42 2.49 3.397 277 3.398 0.001 44 1.95 3.291 276 3.293 -0.002 46 1.54 3.187 275 3.187 0.000 48 1.20 3.080 273 3.080.000 50 0.935 2.971 269 2.971.000 52 0.727 2.861 264 2.860 +0.001 54 0.561 2.749 259 2.748 +0.001 56 0.431 2.634 253 2.633 +0.001 58 0.329 2.517 246 2.515 +0.002 60 0.247 2.392 236 2.390 +0.002 37

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN A specially designed air gap was incorporated as part of the nose- cone design, Fig. 15, and provision was made for causing an electric arc across this gap. The voltage applied to cause the arc had a sawtooth-shaped variation with time, and a sensing device was utilized to ascertain the voltage at which the arc formed. A consideration of the voltage at breakdown through use of Paschen's Law allows determination of the density of the arc's path. The experiment was successful, operating as anticipated throughout the flight. The resulting data have been analyzed, and are presented here in Fig. 16. A forthcoming technical report is in preparation which will describe this experiment in detail. 6.5.4. Gyroscope and Camera Installation. The gyroscope and camera installation in USAF 18 were similar to that in the December 1950 Aerobee USAF 10. Satisfactory missile-attitude data were obtained. A photograph of the complete instrumentation of USAF 18 is presented in Fig. 14. 6.6. USAF 31, October 1952 The instrumentation for Aerobee 31 was prepared during the period of this contract. The field use of the equipment, however, was accomplished under a continuing contract.2 The instrumentation was similar to earlier units constructed at the University of Michigan, the objective being the measurement of atmospheric pressure and temperature. The equipment included: (a) five alphatrons (b) one gyroscope (c) two cameras (d) accessory equipment associated with a, b, and c. 6.6.1. Pressure Measurement. Of the five alphatrons, two were mounted to indicate impact pressure and the remaining three to indicate conewall pressure. One impact and two cone-wall units constituted the basic pressuremeasuring devices necessary to satisfy the requirements of the temperaturemeasurement method. Each of the actual alphatron chambers of these three units 2Contract No. AF 19(604)-545.

H 0)...................77777m777.~~~~~~~~~~~~~~~~~~~..................... ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ 0

ANODE / i'S-S-''';-. -2s'i-'S GAP SECTION V AI NSULATOR PATH OF I BREAKDOWN —- _ ____ CONDUCTOR IMPACT ALPHATRON CONE WALL CATHODE ALPHATRONS Fig. 15. Functional Sketch Illustrating Location ofr Alphatron Gages and Arc Discharge Gap on Aerobee Nose Cone, September 1931. 40

8.0 9.0 7.0 6.0 3.0 4.0 3.0 2.0 o 0 0 0.8 0.7 0.8_ 0.6.t ~Z~ 05 -' - DERIVED FROM ALPHATRON PRESSURE MEASUREMENTS 040 O - EXPERIMENTAL POINT FROM BR EAKDOWN EXPERIMENT 0.3 G) 0.2L 0.1 I 50 35 40 45 50 55 60 65 ALTITUDE (KILOMETERS) Fig. 16. Minimum Local Cone-Wall Density as Obtained from Gap Experiment, September 1951. 41

Id 9 * 5 ~~~~~~~10 ~ ~ s 8 E 6 w 5 WI 3S 0. 2 10..... 6 5 4 2 10 - 30 35 40 45 50 55 60 ALTITUDE - K M Fig. 17. Curve of Ambient Pressure Vs. Altitude, September 1951. 42

280 -- 270 260 250 L I I 1/ROCKET PANEL ADOPTED VALUES C: 240 -.230 220 210 200 190 180 17530.0 35.0 40.0 45.0 50.0 55.0 60.0 65.0 70.0 75.0 ALTITUDE (KM) Fig. 18. Curve of Ambient Temperature Vs. Altitude, September 1951. 43

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN - was provided with a valve, whose function was to seal the unit and hence maintain a pre-established vacuum in the chamber until some predetermined point in the flight. That is, each chamber was pumped to and maintained at some low pressure (10-2 mm Hg) many hours before flight use, following which the valve was closed, thus sealing the chamber. The valves, which were motor-driven, were then activated at the desired point of the flight, exposing the gages to the atmosphere. This procedure, which establishes a "pressure history" for the gage, contributes to the accuracy and reliability of the device. The reasons justifing this operation are not too well understood at the time of this writing. They may be associated with the release of radon gas from the radium ionizing source that is employed in each alphatron, or they may be due to the presence of water vapor. Laboratory studies are underway to assist in our understanding of the phenomena. The two additional alphatrons that were employed without valves were used as controls for comparison with valved units. The data from these units were considered as supplementary to the data from the valved units. 6.6.2. Gyro and Camera Installation. The gyroscope installation was identical with that of earlier equipment. An additional camera was included, however, to lend assurance that the gyro information would be recorded, although failure of the main camera was not expected nor did it occur. 6.6.3. Results of USAF 31 Firing. Inasmuch as the flight use of the USAF 31 equipment took place under a following contract, a discussion of the launching is not presented in this report. In the interest Of completeness, however, it is appropriate to state that the flight was successful. The resulting data are being analyzed at the time of writing. 44

-- ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN 7- REPPORS AND PAPERS During the course of the contract, several papers and reports were undertaken. Fifteen quarterly progress reports were written which presented in some detail the course of the research. A scientific report was issued detaiiiling the theoretical work and early experimental aspects of ionospheric probe work. It is designated as follows: "Dynamic Probe Measurements in the Ionosphere" by Hok, Spencer, Reifman, and Dow, August 1951. A supplement to this report was also issued. A technical paper bearing the same title was published in the June, 1953, issue of the Journal of Geophysical Research. A paper detailing the project temperature-measurement method was written and has been submitted for publication in the Journal of Applied Physics. It is designated as follows: "Rocket Measurement of Upper-Atmosphere Ambient Temperature and Pressure in the 30-75-Km Region" by Sicinski,. Spencer, and Dow. Three papers were written and presented at technical society meetings: (1) "Rocket Atmospheric Pressure Measurement System", by Spencer and Schulte, presented by Spencer at the I.R.E. National Conference on Airborne Electronics, May, 1951, Dayton, Ohio. (2) "Rocket Measurement of Upper-Atmosphere Density by Paschen's Law," by Smith and Early, presented by Smith at the Conference on Gaseous Electronics; September, 1952, Princeton, New Jersey. (3) "Rocket-Borne Electronic Devices in Upper-Atmosphere Research" by Spencer, presented before I.R.E. Detroit Branch Meeting, March, 1953, Ann Arbor, Michigan. Numbers 1 and 2 above are currently in preparation for submission as technical papers to the Review of Scientific Instruments. The writer has an invitation from the I.R.E. to prepare (3) above for submission for publication as a technical paper in the Proceedings of the

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN I.R.E. An additional paper on the subject of the use of a single gyroscope in a rocket is likewise in preparation for submission to an appropriate technical journal. 46

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN 8. PERSONNEL ENGAGED IN CONTRACT WORK Professor W. G. Dow Supervisor Gunnar Hok Research Engineer W. G. Kartlick Research Technician H. F. Schulte Research Engineer H. S. Sicinski Research Physicist N. W. Spencer Research Engineer The above list constitutes the heart of the working group involved in contract activities. In addition, many other individuals, in most cases students hired on a part-time basis, contributed substantially to the project. Their efforts were applied to computing procedures, equipment construction, laboratory testing, etc. Professor W. G. Dow has served as supervisor of the work under this contract. He has provided very necessary and desirable guidance and counsel in regard to the general objectives, as well as direct participation in many project activities. Mr. Gunnar Hok has acted as technical consultant to the project and has, in particular, contributed directly by his theoretical studies of the bipolar probe. Mr. H. F. Schulte has been intimately associated with instrumentation activities, and has, in particular been responsible for the detailed development of the alphatron system. Mr. W. G. Kartlick has served as the only full-time technician on the project. His careful attention to instrumentation details has been a contributing factor to the success of the instrumentation. Mr. H. S. Sicinski has served as a research physicist. His responsibilities have included the theoretical aspects of project work as well as data-analysis efforts. He has, during the course of the contract, developed the theoretical method used for temperature measurement, as well as being responsible for the initial development work on use of the gyro. 47

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN The writer has served as project engineer for the work, coordinating the overall effort. 48

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN 9. GENERAL REMARKS AND SUMMARY The reader, unfamiliar with the detailed procedure involved in rocket measurement of ambient pressure and temperature, summarized in this report, may conclude that such measurements are readily accomplished. It is not the intent of the writer to suggest that the procedure involved is excessively complicated or unduly lengthy. On the contrary, it is believed to be a straightforward, realistic approach to temperature measurement. In fact is is the only method known to project personnel that yields accurate, essentially point-by-point temperature values. However, as with any scientific investigation where accurate, reliable data is desired, care is required in the design and preparation of the equipment used as well as in consideration of the theoretical aspects of the method employed. Data analysis, too, requires careful attention to details, to insure confidence in the results. In particular, upon reviewing the work performed under this contract, it is the opinion of project personnel that the following may be considered as accomplishments: a) Reliable temperature, pressure, and density information on the atmosphere has been obtained. This constitutes the major stated objective of the contract. b) A theoretical study of certain aspects of the ionosphere has been completed and recorded. Data obtained under an earlier contract have been analyzed. c) A new and experimentally proven method has been developed for the measurement of atmospheric temperature from a rocket. d) A new and partially experimentally verified method involving a so-called bipolar probe has been developed for exploration of the ionospheric regions from a rocket. e) A pressure-measurement system employing an alphatron ionization gage has been developed to the point that it is a reliable tool for use in a rocket. It is capable of providing accuracies of 1 part in 100 over a pressure range of several decades, in particular 760 mm Hg to 103lO mm Hg. A redesign of the gage chamber would allow greater range.......t" 49

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN During the course of the contract about fifteen complete alphatron units were constructed and used. No significant failure was encountered., each flight use, however, indicating desirable improvements such as increasing the number of subranges and providing switching delay as discussed earlier in this report., some filament supply regulation, and means for zero indication during flight. f) A missile-aspect system employing a single gyroscope and associated camera has been developed and used successfully. The method provides continuous missile-attitude information. g) An alternative pressure-measurement system employing thermionic ionization gages has been developed. The troublesome problem of emission change with pressure has been solved with the development of a servo mechanism which maintains constant cathode emission over several decades of pressure change. h) A method based on Paschen's Law regarding electric breakdown of air on a missile surface has been developed. On the basis of a single rocket experiment and moderate laboratory testing, it is believed to be a promising means of determining air density on the surface of a rocket. It is not considered that development of this method is completed; it was abandoned under the pressure of other work. i) A method was conceived and the development initiated for a means of direct air-density measurement through use of an RF breakdown of air at a point ahead of a missile. This work was discontinued because support was withdrawn. A device of this nature could produce density measurements in undisturbed air, ahead of a missile, and hence would not be subject to aerodynamic sources of error. j) Numerous accessory devices have been developed to augment fundamental instrumentation. Such devices, including telemetering limiters, motor-driven clocks, calibrating devices, pressuregage valve systems, etc., are essential to an instrumentation and consequently require as much care in development as the more spectacular equipment items. k) A quantity of technical writing was initiated during the contract term, and some papers were presented at technical society meetings. Publication of several papers is scheduled under a continuing contract. 50

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN APPENDIX This appendix is a reprint of a technical paper submitted to the Journal of Applied Physics. The subject matter of the paper includes primarily a discussion of the theoretical temperature-measurement method used under this contract. Included also is a brief description of appropriate instrumentation. Some repetition of material presented in the body of this report is unavoidable. ROCKET MEASUREMENTS OF UPPER-ATMOSPHERE AMBIENT TEMPERATURE AND PRESSURE IN THE 30 TO 75- KILOMETER REGION H. S. Sincinski, N. W. Spencer, and W. G. Dow* ABSTRACT A method for determining ambient temperature and ambient pressure in the upper atmosphere is described, using the properties of a supersonic flow field surrounding a right circular cone. The underlying fundamentals stem from basic aerodynamic principles as combined with the developments of the aerodynamics of supersonic cones by G. I. Taylor, J. W. Maccoll, and A. H. Stone. The experiment provides the necessary cone pressures, velocities and Eulerian angles, so that a Mach number characterizing the ambient space conditions may be computed. A description is given of the requisite experimental equipment and related techniques. Experimental data from two rocket-borne equipments are presented with the resulting calculated pressures and temperatures as experienced over New Mexico to approximately 70 kilometers. *Department of Electrical Engineering, University of Michigan, Ann Arbor, Michigan. 51

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN I. INTRODUCTION The research reported in this paper has been sponsored by the Geophysical Research Directorate of Air Force Cambridge Research Center, Air Research and Development Command, under Contract Nos. AF 19(122)-55 and W33-058 ac-14050. The University of Michigan Department of Electrical Engineering has been engaged for the past several years in the measurement of the arbient temperature and pressure of the upper atmosphere. These measurements have been carried out in high-altitude rockets, in particular the "V-2" and the "Aerobee." During the period in which V-2 rockets were employed, temperature measurements were implemented by application of the "barometric equation" to a measured curve of ambient pressure versus altitude. Although curves of ambient temperature versus altitude were obtained,3 the computational procedure for obtaining the temperature is primarily one of differentiation, and hence yields only very approximate values of temperature. In an effort to improve the quality of the measurements, a more exact method has been developed which overcomes certain disadvantages of the earlier procedure. iEssentially point-by-point values are obtained in a manner that does not require an averaging process. That is, each temperature point on the curve is determined directly from the experimental pressure data independently of other points and gives the temperature at a particular location in space, as contrasted with values obtained from the barometric equation, which represent an average over a rather considerable altitude interval. The experimental data required for a temperature computation by the new method include in general: a ratio of the nose-cone tip (impact) pressure to the pressure at some point on the cone wall, the instantaneous angle between the rocket's longitudinal axis and a space-fixed reference system, and the magnitude of the missile velocity vector in the same reference system.4 3Rpt. No. 2, Upper Air Research Program, Engineering Research Institute, University of Michigan, July 1948. 4Another temperature-measurement method, similar in that correspondingly fundamental pressure measurements are employed has been utilized by the Naval Research Laboratory. See Haven, Koll, and La Gow, J. Of Geophysical Research V. 57, 59-72 (MIrch 1952). 52

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN The required pressure measurements are accomplished in the missile through the use of "alphatron" ionization gages, which are utilized in equipment that has been developed by this research group. The data obtained are telemetered from the rocket to ground stations, where they are recorded. The fundamental information required for angle computation is similarly determined in the missile through the use of a single gyroscope. In this case, the data are recorded in the rocket on film, which is later recovered when the missile reaches the ground. Velocity information is obtained by triangulation, employing groundbased instrumentation which tracks the rocket during flight. Temperature measurements have been made on several rocket flights utilizing the new method. The following sections of this paper present the data resulting from two such flights of Aerobee rockets, a discussion of the theoretical basis for the measurements, and a description of the particular equipment developed to obtain the basic data. II. THEORETICAL CONSIDERATIONS A. General Temperature is a typical "intensive" magnitude5 quantity, which, for its determination, rmust be correlated with phenomena measured "extensively." Although the usual laboratory thermometric systems can measure temperature in extensive terms, these techniques cannot be directly extrapolated to supersonic missiles for upper-atmosphere ambient-temperature measurements without producing questionable results. The chief difficulty arises with the formation of a boundary layer about the instrument, which perturbs the temperature experienced. A more promising datum is pressure, which, unlike the temperature, is very nearly constant throughout any boundary-layer section, being nearly equal to the value Just outside the boundary layer. The pressure datum thus "neglects" the boundary layer, approximating an inviscid flow. For the practical case of a cone with a semi-vertex angle of 7.5~, very good argeement exists6 between the inviscid theory and experimental data from viscid tests in the range of Mach numbers and yaw angles experienced. 5For definition of extensive and intensive properties, see Fowle, F. E., Smithsonian Physical Tables. 6Cronvich and Bird, "Pressure Distribution Tests for Basic Conical Flow Research", Ordnance Aerophysics Laboratory, Daingerfield, Texas. 53

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN The principal limitation to applying inviscid theory occurs for large yaw angles (Ee -0o.8 es) where boundary-layer separation occurs.7 B. Nonyawing Cone The problem of supersonic flow around a cone has been successfully analysed by G. I. Taylor and J. W. ~Maccoll,8 with the subsequent embodiment of their results in tabular form by Z. Kopal.9 These results are applicable to cone pressure measurements in order to compute ambient upper-air conditions for given conical geometry and the characteristic Mach numbers. Although surface pressure measurements alone do not constitute sufficient information to deduce the characteristic Mach number, knowledge of the total head pressure will, when taken concurrently with the surface pressures, define the characteristic Mach number. A schematic representation of the experiment and the physical quantities appearing is shown in Fig. 19. Quantities with subscript "1" Pi / Pw P VI I 2 Ti P2 Fig. 19. Physical quantities appearing in nonyaw conical flow. Pressures measured by the experiments described appear inside of the outline of the cone. (For physical significance of symbols, see list in paper.) 7Moore, Franklin K., "laminar Boundary Layer on a Cone in Supersonic Flow at Large Angles of Attack," NACA-TN-2844. 8Taylor, G. I., and Maccoll, J. W., Proc. Royal Soc. London Ser. A, 159, 279311 (1933); Maccoll, J. W., Proc. Royal Soc. London Ser. A, I, 459-472 (1937). 9Kopal, Z., M. I. T. Tech. Report No. 1, 1947, Dept. of Elec. Eng'g., Center of Analysis. 54

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN are the upstream conditions or so-called "ambient values". The dotted area about the cone vertex defines a subsonic flow region existing because the cone is truncated to permit total head measurement. This deviation from purely conical geometry affects the flow locally; however, a wind-tunnel analysis demonstrates that the conical flow regime is established ahead of the cone surface pressure measuring port. The computation of the ambient conditions proceeds through a combination of the Taylor-Maccoll relations and the Rayleigh total-head expression. P' and P (cone-tip and cone-wall pressures) and cone velocity V relative to the ambient air are measured experimentally. The theory presented leads to a relationship (Fig. 20) between Po /s and the Mach number M1, thus permitting a determination of the Mach number from the experimental data. The dependence of Mach number and cone-wall pressure Ps on the ambient pressure Pi appears in the course of this determination. The ambient temperature is determinable from the familiar equation (6) relationship between Mach number and velocity relative to ambient air. It is convenient to initiate the theoretical analysis by stating the ratio Pl/A of the measured pressures in the following identity:* P' P' Pi P R (1) Pi P P P S W O S The theoretical treatment consists of expressing each of the right hand factors in terms of the Mach number and the known ratio of specific heats, thus leading to the Fig. 20 relationship. To accomplish this for the first factor, energy considerations permit expressing the ratio of pressures across the normal shock wave in terms of the Mach number and the ratio of specific heats as follows:* P f2rM-( -_1/(r-l ) rM2]( -) Pi _2M, - (y - 17 + 1) M.- 1) ( _ L) + U 2 (2) Similar evaluation of the remaining three factors on the right of equation (1) requires use of the theory of the conical regime. Taylor and Maccoll10 determined the pressure ratio across the shock wave using a lengthy graphical procedure. Z. Kopalll derived an explicit expression for this ratio using purely algebraic procedures; he also prepared tables providing values of the local velocities (radial velocity Uw, tangential velocity Vw, and sonic *See list of symbols at end of paper. lOTaylor, G. I., and Maccoll, J. W., op. cit., pp. 288-292. liKopal, Z. op. cit., p. xii. 55

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN velocity, a) for any given cone angle as a function of the Mach number. By using values from these tables in his expression Pw (7 _ 1)(c2 _ uW - VW2) (3) Pi 47 v_ - (7 - 1)2 (C2 -, - vW2) the second factor on the right of equation (1) can be evaluated in terms of Mach number and 7. The quantity c, appearing here, is a useful reference velocity, sometimes defined as the maximum velocity attainable by converting all the heat energy of the fluid into uniform motion. In terms of the local sonic velocity a and the velocity V relative to ambient air, the reference velocity c is, defined as [ V (7 - 1) (4) The third factor on the right-hand side of equation (1), the ratio of static pressure behind the shock wave to the stagnation pressure, is found directly from the Bernoulli integral and the assumption of adiabatic flow behind the shock surface.11 This ratio, in terms of the local velocities provided by Kopal's tables, is Pw 2 Lastly, the ratio Ps/Po of cone surface pressure to stagnation pressure behind the shock wave follows from equation (5) on setting the tangential velocity (vw) equal to zero. The results from using these four evaluation procedures in equation (1), expressed in terms of the Mach number for the nonyaw case of a 7.5o halfangle cone, are shown in Fig. 20. With the requisite Mach number known from Fig. 20, the ambient pressure is available from equation (1) after dividing both sides by Po/Pl. The result of this procedure is shown in Fig. 21. Computation of the ambient temperature depends on the definition of the Mach number and the adiabatic sonic velocity relationship. These express the ambient temperature explicitly as ( ( (Rl ) (6) 56

~~~M~~~I3~~~~ ~ t vT[ OU~ -~ II...TMI 9 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 Po/Ps Fig. 20. Mach Number Vs. Quotient of Cone Pressures for Nonyaw Case of a 7.50 Half Angle, Supersonic Cone.

1.8 1.7 1.6 1.5 // /:M,I PD/ 1.3 1.2 1.0 - I 2 3 4 5 Mg Fig. 21. The Relationship between the Unyawed Surface Pressure and the Ambient Pressure as a Function of the Free Stream Mach Number for a Nonyawing, 7.50 Half Angle Supersonic Cone. 58

- ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN The velocity V appearing here is the relative flight speed. In the experiments presented, no provision was made for estimating local wind conditions; consequently the missile velocity relative to the earth was taken as the defining velocity. Tacitly, this statement assumes that the winds present are negligible compared to the relative flight speed. These data have given the Fig. 29 temperatures, through which a reasonably smooth curve could be drawn. The departure of the measured temperature from the smooth curve is of the nature that would result from the presence of a wind field varying in speed and direction. Uniform wind fields, on the other hand, would yield temperatures continuously higher or lower than the actual ambient temperature. C. Yawing Cone Under the conditions experienced by high velocities, missiles are fundamentally yawing bodies; that is, the missile's longitudinal axis does not usually maintain coincidence with the free-stream relative velocity vector. Consequently, the experiment must employ a yawing-cone theory for its analysis. A theory for yawing supersonic cones was developed by A. H. Stonel2 which included second-order yaw effects. Stone's analysis led to Z. Kopal's13 tabulation of the perturbation coefficients for use in the solution of supersonic flow fields about large-yaw cones. Of particular interest is Stone's expression for the cone's surface pressure, since it provides the basis for a pressure experiment on a yawing cone. In terms of the coordinates of Fig. 22, the surface pressure Pa is 00 00 Pa = Ps +e n cos n0 + c2 Pn Cos no... (7) n=O n=O PS is the surface pressure for a zero-yaw angle (e = O), while the perturbation coefficients nn and Pn are available from reference 9. Stone's analysis demonstrated that all the second-order yaw terms except n = O and n = 2 vanish under the boundary conditions on the cone, while the Rankine-Hugoniot conditions reduce all the first-order terms to zero except n = 1. In application of equation (7) the perturbation coefficients as given by Kopal need to undergo a transformation by means of a Taylor expansion for 12Stone, A. H., Jour. of Math. and Phys., 0 XXX 200 (Jan. 1952). l3Kopal, Z., Report No. 3, M.I.T. Department of Elec. Eng'g. Center of Analysis and Kopal, Z., Report No. 5, 1949, M.I.T. Dept. of Elec. Eng'g. Center of Analysis. 59

2 // I \Xx G A.___>...XIS MZ Z 3 2 CONE Fig. 22. Coordinates Describing the Yawing Cone. Fig. 22. Coordinates Describing the Yawing Cone. 6O

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN utilization in the desired reference frame.14 Kopal's tabulation presents these coefficients relative to arguments of the unyawed reference system, where the desired coefficients are perturbed by angle Ae. Thus the pressure at any perturbed position 8 is given by the relation P(G) = P(G + AG) = P(9) + P'(9) AG + 1/2 P"(G)AG2.. (8) where the barred quantities are with respect to the unyawed reference frame and the primes refer to differentiation with respect to 9. With equation (8) we still need an expression for the yawed conical surface with respect to the unyawed reference frame. In the region between the shock wave surface and the yawing cone the variables will be constant over surfaces of a generally conical nature. If it is assumed that these surfaces remain the equationl1 for any of these surfaces is 9s + AGS = es + e cos ~ - (1/2) c2 cot 9s sin2 ~.., (9) where the particular surface is defined when is and AGs are stated. Using equations (8) and (9) in equation (7), then collecting terms to the order C2, and evaluating the derivatives gives the pressure at the yawed solid cone surface as P = Ps f1 c+ 2 cos 2p + 7 I + C2 Po + a)...) I The experimental data provide values for Pa, c, and 0. These quantities along with equation (10) permit computation of Ps, the unyawed pressure. With Ps determined, the experiment reduces to the nonyaw case for which expressions involving the ambient pressure and temperature have already been given. P. are surface pressures measured by suitable gages located in the cone, and e and 0 are computed from a combination of the trajectory locus and data from a missile-borne gyroscope. From the definitions, the yaw-angle computation makes prerequisite an assumption regarding environmental winds. In both experiments presented the wind velocities are assumed to be small compared to the missile velocity. The yaw angle is then determined with the wind vector tangent to the missile trajectory, while the missile aspect is taken from a gyroscope. 14Van Dyke, M. D.; Young, G.; Siska, C., Jour. Aero. Sci., Vol. 18, p. 355, May 1951. 15Stone, A. H., "On Supersonic Flow Past a Slightly Yawing Cone," Jour. of Math. and Phys., A, 73 (1948), Eq. 34. 61

ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN III. INSTRUMENTATION The instrumentation that has been developed and used in Aerobee rockets by this research group to obtain the fundamental data required utilizes an alphatron gage as the basic pressure-sensitive element, and a gyroscope for missile angular-position determination. This section of the paper will describe briefly the manner in which each of these devices is employed. A. Pressure Measurement An alphatron is an ionization gage wherein the ionizing energy is obtained from alpha particles emanating from a small quantity of radium, generally of the order of milligrams. The essential external characteristics of the particular gage chosen for use in this investigation are illustrated in Fig. 23, which presents a typical curve of output current versus chamber pressure. The lowest measurable pressure is determined fundamentally by the "dark current", the value on the curve to which the lower portion of the curve is asymptotic. The upper limit is determined by recombination of ionized particles before the ionization products are collected and measured, as evidenced by the bending of the curve at the higher currents. Two features important from the standpoint of circuit requirements are immediately apparent from the curve: the very small current that constitutes the basic inforuation signal, and the rather large ranges of current and pressure which must be accommodated in an instrument which utilizes the device over its useful range. The small current implies either the use of a relatively small (few megohms) load resistance and very large voltage amplifications, or the use of very high values of resistance, with the consequent problem of impedance matching. The extensive useful range, on the other hand, demands the use of several subranges in order to obtain a reasonable definition in the ultimate pressure data. Figure 24 is an elementary block diagram illustrating the circuit developed16 to meet these requirements. The alphatron current is passed through 16Developed from an original design by J. R. Downing and G. Mellen, Rev. Sci. Instr. 17, 218 (1946). 62

icY UL) Lrl < - 10 M-. 0 0 Cr 4 i0 5121 -J~~2 - 1Q10 106 icY 100 10 102 PRESSURE IN MM OF MERCURY Fig. 25. Variation of Output Current with Chamber Pressure for a Particular Alphatron Pressure Gage.

ROCKET SURFACE PORT ALPHATRON ALPHATRON D.C. LOAD ~ ^.,AMPLIFIER RESISTANCE RANGE CHANGING CIRCUIT OUTPUT TO TELEMETER SYSTEM Fig. 24. Elementary Block Diagram of Alphatron Pressure Measurement System.

-- ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN a resistance of sufficient magnitude to produce a voltage equivalent to the- desired information signal. Because of the very small current, this resistance may be as high as 250,000 megohms. The voltage obtained across this resistance is applied to a 100* negative-feedback d-c amplifier which acts essentially as an impedance-changing device. The first stage of the amplifier employs an electrometer tube in order to provide an input resistance that is large compared with any probable alphatron load resistance. The following sections of the amplifier are, in sequence; a voltage amplifier, a heater voltage regulator, another voltage amplifier, and finally a cathode follower stage. Inside the feedback loop the voltage gain is high, of the order of 4500. However, with feedback, the voltage gain of the system is unity, whereas the current gain is significantly high. Since the output voltage is equivalent to the input voltage (100* feedback) the current gain is numerically equivalent to the ratio of the input load resistor (alphatron load) to the cathode resistor of the cathode follower. The voltage obtained from the cathode follower constitutes the desired data and is accordingly applied to the recording system, in this case the telemetering system. In order to provide for the several subranges, different possible values of alphatron load resistance are provided, one for each subrange. It is the function of the range-changing circuit to select and insert the particular load resistance appropriate to a particular range of chamber pressure. To accomplish this, the amplifier output signal is applied to the range-changing circuit, which uses two thyratrons to control a bidirectional rotary solenoid. If the information signal voltage exceeds a predetermined value (in either direction) the next lower or higher value of resistance is inserted, thus returning the information signal to an "on-scale" value. An automatic range-selection device is, of course, necessary because the equipment operates unattended through the total pressure range encountered during a rocket flight. Figure 25 illustrates the variation of output signal with pressure for a particular subrange. In this case the alphatron load resistance is 5000 megohms. The complete alphatron equipment in a particular rocket includes several nearly identical, independent units similar to that described above, differing perhaps only in regard to choice of pressure subranges required by particular gage locations, for example, cone wall or cone vertex mounting. B. Angle Measurement The gyroscope used for missile angular position measurement is a modified Sperry type F4A unit. The modification was accomplished primarily in

5 4 2.1.2.3.4.5.6 7.8.9 1.0 2 3 4 5 6 7 8910 PRESSURE MM. HG. Fig. 25. Variation of Alphatron Pressure Measurement System Output with Pressure for a Particular Sub-Range.

-- ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN order to allow operation under free-fall conditions. However, in addition, an attachment was developed that enables the establishment of zero position prior to rocket flight. Recording of gyroscope data is accomplished by photography of the gyroscope sphere (gyrostat) position in reference to a missile-fixed coordinate system. The film on which the position is recorded is recovered at the end of the rocket flight. IV. EXPERIMENTAL RESULTS Two experiments based on the above theoretics have been successfully completed. Pressure equipment was instrumented in the Aerobee Sounding Rocket type of missiles for launching by the U. S. Air Force at the Holloman Air Force Base at Alamogordo, New Mexico. The first experiment, on 20 June 1950, carried one impact pressure gage and one cone surface pressure gage. For comparison purposes the two experimental pressures and the resulting computed ambient pressure are shown in Fig. 26. From this first experiment, the ambient pressure data are reliable to 1 part in 35. The seemingly relatively large scatter in the cone surface pressure is a result of the missile's rotation as it assumes increasing yaw angles with altitude. The impact pressure is generally without such cyclic variations, since it remains independent of the yaw angle for values up to about 30~. Although not shown in entirety, maxima in the ratios of both cone surface and impact pressure to ambient pressure occur in the neighborhood of 35 kilometers in altitude. These maxima are presumably the result of the combination of maxima in the Mach number and the missile velocity; as such they must be constructed as having been caused by missile behavior, and not as representing properties of the atmosphere. The temperatures computed from these data are shown in Fig. 27. No comment is offered on this curve other than that the maximum probable error is believed to be +80K to about 60 kilometers and +135K above about 60 kilometers. The second experiment was completed on 13 September 1951. The instrumentation represented considerable improvement over that of 20 June 1950, with two cone surface gages and a greatly increased information-reporting capacity. The increase in sampling information rate resulted in a reduction of the overall probable error of the final temperature data. For Fig. 28 the temperatures up to 50 kilometers have a maximum probable error of +50K while the probable error above 50 kilometers is +70K. The ambient pressure resulting from this flight has an improved accuracy such that it is reliable to 1 part in 65. These data are shown in Fig. 29 as the "experimental-points ". 67

5 -_ A = PI AMBIENT PRESSURE 4 _ o = Po' ACTUAL CONE SURFACE PRESSURE x = Po ACTUAL CONE IMPACT 2 _ PRESSURE xrxxx ALTITUDE IN KILOMETERS PRESSURE IN MILLIBARS eI02....I I 10 I. 9 - --—' E 6 _= k 0 0 __ __1 OC NN 20 30 3540 45 50 55 60 65 70 3 bJ ALTITUDE (KM) 60I 102 1 I0 20 30 35 40 45 50 55 60 65,.ALTITUDE (KM) Fig. 26. Actual Cone Pressures, Total Head (Cone Tip) and Surface. Compared with the Ambient Pressure for Aerobee Rocket of June 68

300! 290 280 270 260'%...~~~ ~ —----- - ROCKET PANEL w ____ADOPTED Ic:::)2~~~ VALUES 240 w.aw 230 22C 21C 200 190 Fig. 27. Ambient Temperature at Various Altitudes above Alamogordo, New Mexico. These temperatures are computed from Aerobee Rocket data of June 20, 1950 at 0858 hours using the assumption of a limited wind field. 69

260 1 - -250_ I CKET PANEL 220 2080 30.0 35.0 40.0 450 50.0 55.0 60.0 65.0 70.0 75.0 ALUES TITUDE (KM) Fig. 28. Ambient Temperature at Various Altitudes above Alamogordo, New Mexico. These temperatures are computed from Aerobee rocket data of September 1, 1951 at 0437 hours using the assumption of 70 30. ] 5.!! 400 40 5. ] O 6. 5. 00 7 ALTITUDE~~~~~~~~~~~~~~~~ "K Fig. 28. Amb~~~~~~~~~~~~~~~~ie Tmeau a aiu liue bv aogr, New Mexico. These temperatures are computed from Aerobee rocket~~~~~~~~~~~~~~~~~~data of Septmber 13, 19~ at 0437 hous using the......... h a limited wind field,~~~~~~~~~~~~~~~~~'~ ~ ~ ~ ~~~7

-- ENGINEERING RESEARCH INSTITUTE ~ UNIVERSITY OF MICHIGAN V. SELF-CONSISTENCY IN THE RESULTS It has been pointed out above that each experimental point, for the temperature curves of Figs. 27 and 28, is evaluated from the experimental data independently of other points. Furthermore, in the nonyaw case, the temperature calculations leading to Figs.. 29 and 30 employ only the ratio of the measured total head pressure and cone surface pressure, not their absolute values. Thus, the absolute values of the pressure measurements are not employed in determining temperatures. It is obviously possible to employ the experimentally determined temperatures in the familiar hydrostatic equationl7 thereby determining the absolute values of the ambient pressures by a method that does not directly use the pressure measurements obtained by the rocket instrumentation. In such use of the hydrostatic equation an inverse-square variation of gravity is assumed, and a mean molecular weight of 28.966 for air is employed. Balloon observation results are employed to provide the absolute value of pressure at 30 kilometers, which serves as a constant of integration. Figures 29 and 30 present, for the two data sets, Figs. 27 and 28 respectively, the absolute values of ambient pressure determined as follows: (a) The values indicated by the circles were obtained by direct point-by-point determination from the data, using Fig. 21 in this case the ambient pressures are obtained, practically speaking, by applying an appropriate correction to the conewall pressures. (b) The values indicated by the crosses were obtained by employing in the hydrostatic equation the Figs. 27 and 28 temperatures obtained point-by-point from the data. The very good agreement between the two sets of points in each case provides a rather satisfying self-consistency check of the system of instrumentation and data reduction employed. The deviations between the two sets of points are less than the experimental errors of the method. Of course, selfconsistency between these two different rocket flights and observations by other methods18 is also of interest. However, seasonal and diurnal variations of 17Mitra, S. K., "The Upper Atmosphere," The R__sl Soc. of Bengal Monograph Series, Vol. V, 1947, See equation 3, p. 5 (dP/P = -mgdh/kl,). l8The Rocket Panel,. Rev., 88, 1027 (1952). 71

2 _ _ _ _ ___2 5 lx z HYDROSTATIC EQN.PTS ~~4 t -- ~~~~~~~oz EXPERIMENTAL PTS. t l I I I I I.I I. I I 2 _ 87 M? LUCr- I I O 10- _(dP/dh 10 Flight. ~7~~~ ~72 4 - -2_._._ 2 ,........~~~~7

10 I_ X 102 x a HYOROSTATIC EQN. PTS. 6 o 0 EXPERIMENTAL PTS. -1 1 I 1 1 r 1 I ~ II4 -33 10. I I I Q:: -,8737 3 3 75

ENGINEERING RESEARCH INSTITUTE * UNIVERSITY OF MICHIGAN the temperature curve certainly exist and must be taken into account in comparing the two sets of results here reported with one another and with results obtained by other methods. We wish to take this opportunity to thank Mr. Ralph E. Phinney for his interest and valuable discussions involving the basic aerodynamics of this problem. 74

LIST OF SYMBOLS Pa Cone surface pressure of yawed cone, along ray at angle 0 from plane of yaw. P1 Ambient pressure PO Stagnation pressure behind conical shock wave P' Stagnation pressure behind normal shock wave 0 P2 Static pressure behind normal shock wave PW Static pressure at shock wave surface on downstream side of a conical shock wave PS Cone surface pressure when yaw angle is zero (e = 0) T1 Ambient temperature M1 Free-stream Mach number M2 Mach number behind normal shock wave V Free-stream velocity vw Tangential particle velocity UtW Radial particle velocity us Cone surface particle velocity Os' ~Cone half angle c See equation (4) e Yaw angle ~TI/s Stone's first-order perturbation coefficient PJ/s Stone's second-order perturbation coefficient P O/P-s Stone's second-order perturbation coefficient Spherical coordinate Spherical coordinate 75

t Angle of rotation about cone's longitudinal axis measured from plane defined by cone's longitudinal axis and the wind velocity vector v a Local sonic velocity h Altitude g Acceleration of gravity m Mean molecular weight R Universal gas constant y Ratio of specific heats 76

UNIVERSITY OF MICHIGAN 3 901 5 02086 64741111111 3 9015 02086 6474