本實驗室在原有之超音速燃燒研究基礎下，使用新建構之馬赫2連管風洞進行超音速燃燒實驗研究。風洞使用下吹式氣源系統，並以鋁礬土球床式蓄熱器與氫氧焰燃燒系統將流場加熱至全溫約1700 K之高焓氣流，接著經過超音速噴嘴將流場加速至馬赫 2進入實驗測試段。
本研究依序驗證風洞之建溫、建壓時間，及穩定區域，並使用新開發之液膜式水冷全壓量測管克服流場高溫，搭配光學位移平台及可拆式壁面，進行侵入式量測取得測試段內流道截面全壓，搭配壁面量測之流場靜壓數據，藉由瑞里皮托管方程式計算取得截面速度與均勻程度，確保流場能達成實驗所需條件、且均勻穩定，能做為日後超音速燃燒之研究設備。
實驗結果顯示，風洞開啟後由預熱至978 K之床式蓄熱器將儲氣槽連續釋放之氣流加溫進入風洞，溫度爬升至設定目標 780 K約需 360 秒，壓力爬升至設定目標230 psi約需 150 秒，接著經由氫氧焰二次燃燒加溫至目標溫度1700 K，溫度爬升時間約14秒。風洞總建溫、建壓時間約在374秒，其溫度、壓力穩定時間可持續50秒以上。測試段流場截面量測部分，全壓及靜壓平均值分別為73 psi及11 psi。流場截面25點之馬赫數在2.15~2.38之間，平均值為2.27，其中截面最大相對誤差為5 %。
從實驗後靜壓數據發現，噴嘴後平均靜壓約為11 psi，低於環境壓力之13.7 psi，屬於過膨脹現象(over-expansion)。由於燃燒段壁面溫度與氣流溫度相差僅66 K，且噴嘴入口全壓與設定值相差不遠，可研判燃燒段內之高溫條件造成流體之動力黏度(dynamic viscosity)顯著增加，加速了噴嘴壁面之邊界層成長。邊界層厚度約為16.74 mm，相當於噴嘴喉部面積與出口面積分別縮小為設計值的44%與58%，此現象改變了原先設計之噴嘴喉部與出口面積比，造成平均馬赫數相較設計值馬赫2 有較大相對誤差(+14%)。

This laboratory, building on its existing expertise in supersonic combustion research,conducted experimental studies using a newly constructed Mach 2 wind tunnel. The wind tunnel employs a blow-down gas source system and uses a bauxite ball-bed heat regenerator in conjunction with a hydrogen-oxygen flame combustion system to heat the flow field to a total temperature of approximately 1700 K, achieving a high-enthalpy gas flow. The flow is then accelerated through a supersonic nozzle to Mach 2 before entering the test section.The study sequentially verified the wind tunnel's temperature and pressure ramp-up times, as well as its stable operating range. A newly developed liquid-film water-cooled total pressure measurement probe was used to overcome the high-temperature flow field. Combined with an optical displacement platform and removable wall surfaces, intrusive measurements of the total pressure across the cross-section of the flow channel in the test section were obtained. By pairing these data with static pressure measurements from the walls, the Mach number and flow uniformity were calculated using the Rayleigh pitot tube equation. This ensured that the flow field met the required experimental conditions and was uniform and stable, making it suitable for future supersonic combustion research.The results showed that after activation, the bauxite ball-bed heat regenerator preheated the gas flow continuously released from the storage tank to 978 K before entering the wind tunnel. The temperature ramped up to the target of 780 K in approximately 360 seconds, while the pressure reached the target of 230 psi in about 150 seconds. Subsequent secondary heating by the hydrogenoxygen flame combustion system raised the temperature to 1700 K in about 14 seconds. The total ramp-up time for temperature and pressure in the wind tunnel was approximately 374 seconds, with stable temperature and pressure conditions lasting over 50 seconds.Regarding the flow field cross-sectional measurements, the average total pressure and static pressure are found to be 73 psi and 11 psi, respectively. The Mach number at 25 measured points within the cross-section ranges from 2.15 to 2.38, with an average value of 2.27, and the maximum relative error in the section is 5%.From the post-experiment static pressure data,it was observed that the average static pressure behind the nozzle was approximately 11 psi, which is lower than the ambient pressure of 13.7 psi, indicating an over-expansion phenomenon. Given that the temperature difference between the combustion section wall and the airflow was only 66 K, and the nozzle inlet total pressure closely matched the set value, it can be inferred that the high-temperature conditions in the combustion section significantly increased the dynamic viscosity of the fluid, accelerating boundary layer growth along the nozzle wall. The boundary layer thickness was estimated to be approximately 16.74 mm, leading to reductions in the nozzle throat and exit areas to 44% and 58% of their design values, respectively. This alteration changed the originally designed throat-to-exit area ratio, resulting in an average Mach number that deviated significantly(+14%) from the designed Mach 2 value.

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