本研究目的為探討微尺度下超音速流場之現象，主要設計超音速噴嘴並在噴嘴後方加上測試段，製造微型超音速風洞，並在測試段內放入結構物，研究在微尺度下，超音速流場經過結構物後之流場變化。本研究之超音速噴嘴設計為出口面積與喉部面積比Ae/A* = 1.68，噴嘴出口處之馬赫數為2，喉部寬度為500 m，流道深度150 m。本研究先利用數值模擬軟體ANSYS CFX模擬在不同流體條件設定下之微型風洞流場現象，發現在層流設定中，因黏滯效應較強，造成震波提前發生。若是在忽略黏滯力設定中，其結果與等熵流場所計算出之結果最為接近。此外，在考慮黏滯效應之模擬結果中，流道近兩側壁面產生邊界層，而忽略黏滯力之設定則無。接著模擬在不同入出口壓力下微型超音速風洞中之流場現象，並發現在水平測試段內之流場因側壁面產生邊界層，導致測試段內之壓力逐漸上升。因此將測試段向外擴張角度，藉由此設計可讓流體之邊界層對流場影響減小。由結果可發現，當漸擴角度3。且入口壓力100 kPa以及出口壓力調整為20 kPa，測試段可維持馬赫數約為1.7之穩定流場，故以此設定做為後續在測試段放入結構物模型討論流場的設計。本研究設計兩種不同結構物放置於微型風洞測試段中，分別為頂角20。及30。底邊長度固定為150 m之三角形。在模擬結果中，發現超音速流場經過尖角後出現壓力逐漸上升之現象，與巨觀尺度下常見壓力驟升之斜震波現象不同，推測原因為在微尺度下黏滯效應影響較劇烈所造成。
後續實驗使用PMMA壓克力 製作微型風洞，主要利用黃光微影製程搭配熱壓印翻模技術可得到表面光滑之流道。本研究利用PSP(Pressure Sensitive Paints)螢光壓力感測塗料分析超音速微型風洞內之壓力分佈，並發現在測試段漸擴角度為3。，入出口壓力分別為100 kPa及30 kPa下，有穩定之超音速流場，馬赫數約為1.4。與模擬結果相比，可發現因表面粗糙度導致流體於噴嘴處提前出現震波，故壓力較模擬結果高。在放置三角形結構物於超音速微型風洞中之實驗，可成功量測結構物後方之加速段，在頂角30。三角形結構之實驗中，沿y = 100 m之軸向壓力由44 kPa降低至32 kPa。然而無法由實驗結果觀察結構物前方壓力逐漸上升之現象，推測原因為表面摩擦效應與結構物對流場產生之壓力上升影響產生抵銷。
綜合以上之成果，本研究成功利用PSP螢光壓力感測塗料，量測在微尺度下超音速流場之壓力分佈，有助於在微尺度下超音速流場之研究。實驗結果發現在微尺度下雷諾數較低，導致黏滯效應影響較劇烈，因此流場現象與巨觀尺度或是以等熵公式計算超音速流場有明顯不同。

The objective of this study is to investigate the physical phenomena in supersonic flow at micro scale. The supersonic micro nozzles have been designed using commercial software ANSYS CFX and the test section was added behind the nozzle outlet, therefore constructed the micro supersonic wind tunnel. Wedge structures with different deflection angles were put into the test section to investigate the change of flow properties in supersonic flow. The area ratio of nozzle throat and exit in the designed nozzle was 1.68 and the designed Mach number was 2. The throat width is 500 m and the depth of the nozzle is 150m. In the simulation, different settings of fluid conditions have been applied to examine the change of flow field. The result shows that the shock wave occurred upstream in laminar flow condition. The simulation result is close to the theoretical result with isentropic assumption if the viscous effect is ignored. Moreover, it can be seen that boundary layer growing from wide wall with viscous flow setting, but not in the inviscid flow setting. In addition, the simulation result shows that the pressure in the test section rise gradually due to the effect of boundary layer. Therefore, divergent angle of test section was designed to reduce the effect of boundary layer and maintain the flow speed. The result shows that there is a steady supersonic flow at Mach number of 1.7 in the test section if the divergent angle of test section is 3 degrees and inlet/outlet pressure is 100 kPa/20 kPa. Two different kinds of wedge structures were designed and placed in the test section, with half-angles of 10 degrees and 15 degrees and the length of base is 150 m. From the simulation results, it can be seen that pressure rise gradually as fluid flow through the structure, which is different than the sudden pressure rise caused by oblique shock wave as expected in macro scale. This is due to much strong viscous effect in micro scale.
For the experiment, the structures of micro wind tunnels were fabricated by photolithography and hot embossing using PMMA material. In this study, PSP (Pressure Sensitive Paints) was utilized to investigate the pressure distribution inside the supersonic micro wind tunnel. The results show that the supersonic flow condition can be kept steady at Mach number 1.4 if the divergent angle of test section is 3 degree and inlet/outlet pressure is 100 kPa/30 kPa. Compared to the simulation results, the pressure is higher than the simulation results due to multiple weak shock waves happened earlier caused by surface roughness. In the experiments of supersonic flow with wedge structures in micro scale, the pressure drop through triangle structure was measured but the pressure rise cannot be identified.
Overall, this study demonstrates that isentropic theory is no longer suitable in micro scale due to small Reynolds number and severe viscus effect. In this study, pressure distributions in supersonic micro wind tunnels have been successfully obtained by PSP technique and the results are used for further investigation and design of supersonic flow in micro scale.

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