本研究目標在於校驗目前座落在成功大學歸仁校區航太試驗場之反射式震波風洞。該風洞主要目標為進行馬赫數5以上之極音速外流場及馬赫數2以上之超音速燃燒相關實驗。反射式震波風洞主要由震波管，超音速噴嘴及測試段組成。震波管總長度13公尺，驅動段長度3公尺，被驅動段長度10公尺，內徑18公分。
震波風洞校驗實驗主要可以分為三個部分：第一、主要隔板破裂測試及入射震波馬赫數於震波管內之校驗。震波風洞驅動段起始壓力條件由主要隔板破裂壓力決定，主要隔板破裂壓力值在未來實驗為重要資料，研究中同時探討不同鋼板破裂程度對於入射震波馬赫數之影響。實驗結果發現，隔板破裂越完整，震波發展越完全。由震波管上壓力計所量得之訊號，計算出入射震波馬赫數並與理想氣體理論值及美國NASA之CEA程式數值計算結果作比對，實驗量得之震波後壓力值及震波馬赫數均與理想氣體理論值及美國NASA之CEA數值計算結果相符。第二、超音速噴嘴出口流場均勻度實驗。在測試氣體經過超音速噴嘴後，使用皮托管同時量取噴嘴下游1公分處切面之壓力分布。實驗結果顯示，流場均勻度尚可。第三、偵測驅動氣體到達測試段時間之實驗。實驗中成功發展一新的技術，使用一22度半角的二維楔型體，觀察驅動氣體與被驅動氣體在錐形體尾端背階流場中的滑移線角度改變量，進而預估可使用測試時間長度。
實驗結果同時發現高溫氣體會造成壓電式壓力感測器膨脹，進而影響其量測精度，建議未來研究中，要做好壓力感測器的熱隔絕，同時考慮真實氣體效應之修正。

This thesis aims to conduct the calibrations of a reflected shock tunnel at National Cheng Kung University, Taiwan. Notably, this tunnel is the first ground test facility for hypersonic researches among universities in Taiwan. The reflected shock tunnel is composed of: (1) a 3-meter-long driver tube (2) a 10-meter-long driven tube, with the same diameter of 18 centimeters, (3) a Mach 2 nozzle with a throat diameter of 6.54 cm and a nozzle exit diameter of 8.5 cm or a Mach 5 nozzle with a throat diameter of 4.5 cm and a nozzle exit diameter of 22.5 cm, (4) a 2.05-meter-long test section with 4 windows for flow visualization, and (5) a 2.72-meter-long dump tank.
The calibration works are divided into three parts: (1) Rupture pressure calibrations of the cross-grooved primary diaphragms, made of Stainless Steel 304. The rupture pressure tests of the primary diaphragm provide an important data base for the following experiments. The experiments clearly show that to have a well-developed shock wave, the full opening of the primary diaphragm is necessary. The pressure behind the shock waves and the incident shock Mach numbers are measured with PCB pressure transducers. The experimental results are in good agreement with the theoretical ideal-gas prediction and the NASA CEA numerical simulations. (2) Measurements of the flow uniformity at the exit of the Mach 2 nozzle: A pitot rake is placed at the nozzle exit to measure the total pressure distribution. The total pressure is also used to calculate the freestream Mach number. The freestream uniformity is satisfactory. (3) The driver gas detection to determine the true test time: A new technique to detect the arrival of the driver gas is developed. A supersonic flow over a wedge of 22 half-angle is observed with the shadowgraph technique. The angle of the slip line in the back-step region of the wedge flow changes as the driver gas arrives.
It is also found that the thermal insulation of the PCB sensors needs to be improved to avoid the thermal expansion of the sensors, which is expected to cause the pressure measurement errors in both the shock tube and the Pitot rake. The importance of the real gas effects on the shock tunnel calibration needs to be assessed carefully in the future study.

[1] Bradshaw, B., “Experimental Fluid Mechanics,” Pergamon Press, Oxford, 1970.
[2] Micol, J. R., “Hypersonic Aerodynamic/Aerothermodynamic Testing Capabilities at Langley Research Center: Aerothermodynamic Facilities Complex,” AIAA Paper 95-2107, June 1995
[3] Stalker, R., “Hypervelocity Aerodynamics with Chemical Nonequilibrium,” Annual Review of Fluid Mechanics, Vol. 21, pp. 37-60, 1989.
[4] Adela, B. Y., “Experimental Investigation of Mixing and Ignition of Transverse Jets in Supersonic Cross Flows,” Ph.D thesis, Stanford University, December, 2000
[5] Holder, D. W., Schultz D. L., “On the Flow in a Reflected Shock Tunnel,” A. R. C. R&M No.3265, 1962
[6] Pennelegion, L. and Gough, P. J., “The Change in Shock-Tunnel Tailoring Mach Number Due to Driver Gas Mixtures of Helium and Nitrogen,” A. R. C. R&M No.3398, 1965
[7] Gaydon, A. G. and Hurle, L. R., “The Shock Tube in High Temperature Chemical Physics,” Reinhold Publishing Corporation, 1963
[8] Smith, C. E., “The Starting Process in a Hypersonic Nozzle,” J. Fluid Mech. Vol. 24, part 4, pp. 625-640, 1966
[9] Huber, P. W., Schexnayder, C. J. and Mc Clintion, C. R., “Criteria for Self Ignition of Supersonic Hydrogen-Air Mixture,” NASA TP-1457, 1979
[10] Yang, C. C., Wen, C. Y., “Reflected shock Tummel Design,” A thesis submitted for degree of Master of Engineering, 2010
[11] Gordon, S. and McBride, B. J., “Computer Program for Calculation of Complex Chemical Equilibrium Compositions and Application I. Analysis,” NASA RP-1311, 1994
[12] Gordon, S. and McBride, B. J., “Computer Program for Calculation of Complex Chemical Equilibrium Compositions and Application II. User Manual and Program Description,” NASA RP-1311, 1996
[13] Holder, D. W. and Schultz, D. L., “On the Flow in a Reflected Shock Tunnel,” A. R. C. R&M NO.3265, 1962
[14] Pennelegion, L. and Gough, P. J., “The Change in Shock-Tunnel Tailoring Mach Number Due to Driver Gas Mixtures of Helium and Nitrogen,” A. R. C. R&M NO.3398, 1965
[15] Gaydon, A. G. and Hurle, L. R., “The Shock Tube in High Temperature Chemical Physics,” Reinhold Publishing Corporation, 1963
[16] Simth, C. E., “The starting process in a hypersonic nozzle,” J. Fluid Mech. Vol. 24, Part 4, pp.625-640, 1996
[17] John, D. A., “Modern Compressible Flow,” Mc Graw Hill, Third Edition, 2004
[18] James, E. A., “Gas Dynamics” Allyn and Bacon, INC, Second Edition, 1985
[19] John, D. A., “Fundamentals of Aerodynamics,” Mc Graw Hill, Fourth Edition, 2005
[20] Mertzkirch, W., “Flow Visualization,” Academic Press Inc., 1987
[21] Stalker, R., “A study of the Free-piston Shock Tunnel,” AIAA Journal, Vol. 5, No. 12, 1967
[22] Bakos, R. and Erdos, J., “Optimizing Pressure Recovery in a Detonation Driven Reflected Shock Tunnel,” Proceedings 21th International Symposium on Shock Waves, July 20-25, 1997
[23] Roy, S. H. and Larry, C. F., “Behavior of Burst Diaphragms in Shock Tubes,” McDonnel Douglas Astronautics Corporation, Huntington Beach, 1975
[24] Bradly, I. N., “Shock Waves in Chemistry and Physics,” John Wiley & Sons Inc., 1962