提升飛行器的空氣動力特性一直都是飛機設計的一大重點，在設計飛行器的機翼時通常會加入高升力裝置，以滿足飛行器在不同的飛行條件所需的升力(如飛行器在起飛或降落過程)。除此之外，前人之研究亦利用控制面改變機翼之幾何外形，在巡航過程中能有較高的升力或俯仰力矩，然而當後緣襟翼之偏轉角(deflection angle)達到一臨界角度時，襟翼上表面可能產生界邊層分離，因此在機翼弧線最佳化的設計上仍需進一步考量。其次，後緣襟翼上表面流場可用簡單的凸角流來模擬。若自由流為超音速，即為一般熟悉的Prandtl-Meyer Expansion。當自由流為次音速，前人之研究指出在接近凸角時，流場因黏滯與非黏滯交互作用的影響產生擴張波，當自由流馬赫數及凸角角度增加時，次音速流場將轉換成穿音速流場，甚至有震波的產生。而當震波達到一定強度時，會進一步導致凸角下游邊界層分離。
本研究依據前人研究的結果，進一步探討雷諾數( 8.04 × 104 - 1.63 × 105)對單一凸角流的影響，由表面壓力結果顯示，在設定的雷諾數範圍內，雷諾數對於穿音速凸角流流場的影響不大。其次，本研究亦分析圓凸角流及雙凸角流的特性，一般而言，在相同的總凸角角度條件下，較低之馬赫數尖峰值將延遲流場由次音速轉換至穿音速。此外，在圓凸角流及雙凸角流的流場中，亦觀察到二次擴張波，此一現象應與兩個lambda shock結構相關。在壓力擾動方面，在較低雷諾數的測試條件下，當邊界層分離時，其壓力擾動值較高，可歸因於較短的分離泡長度。而圓鈍型凸角流及雙凸角流之壓力擾動值則明顯降低。由先前的研究得知壓力擾動值與馬赫數尖峰值成正相關；因此，在圓凸角流及雙凸角流的流場中，所觀察到較低馬赫數尖峰值可用以解釋此時的較低壓力擾動值。
最後，利用hodograph equation 所推導出來的相似參數β (=M^2 η/√(1-M^2 ))適合用於分析可壓縮凸角流。根據在雷諾數效應、圓凸角流及雙凸角流的實驗中，其最小表面壓力(或馬赫數尖峰值)亦與β呈正相關。而壓力擾動值也有類似的現象，但在高馬赫數及大偏轉角之測試條件下，震波震盪的現象亦導致壓力擾動值有一個明顯的躍升。

Improving aircraft performance is one of the major goals in the aviation design. In the past decade, there have been many attempts to improve the aerodynamic efficiency of aircraft. One of the methods is to maximize lift-to-drag ratio in different flight conditions. To achieve this, the wing of airplane must be able to change its configuration, using deflected control surface. Such device like flaps and ailerons could be employed to provide variable camber control within the operational flight envelopes.
Previous investigations have shown that the convex corner is an idealized configuration that be used to model the compressible flows over the deflection control surface at off-design condition. The general characteristics of the flowfield include sudden expansion near the corner followed by recompression. In the research presented in this dissertation, a series of experiments of transonic convex corner flows were performed to investigate the aerodynamic characteristics and to gain a good insight into the interaction of a shock wave and a turbulent boundary layer. First, two flat plates with different lengths were used to develop naturally different incoming boundary layer thicknesses or Reynolds numbers. The Reynolds number ranged from 8.04 × 104 to 1.63 × 105. Compared to subsonic expansion flow, the Reynolds number has an influence on the flow expansion and compression near the corner apex in the transonic flow regime. The subsequent work focused on transonic round convex-corner. Test models with three different radii of curvature from 13.7 to 41.7 times the upstream boundary layer thickness were employed. The total turning angles were 13-, 15- and 17-deg. It was found that the peak Mach number decreased with increasing radius of curvatures for a given turning angle. Finally, experiments were performed wherein a presence of bi-convex corner. The total turning angle  ranged from 10- to 17-deg, in which the first convex-corner angle 1 was 5- or 7-deg. It was known that the length of the first corner L1 also plays an important role. With increasing L1, the less flow expansion is observed, while a shorter L1 resulted in a more significant decrease in the amplitude of (σp/pw)max. Not only the bi-convex corner, but also the round convex-corner was observed dual expansion process in transonic expansion flows. In addition, a delay in the transition from subsonic to transonic flows was found in both test configurations. A similarity parameter β (=M^2 η/√(1-M^2 )) was also employed to correlate the Mpeak, (σp/pw)max and shock-induced boundary-layer separated length. The experimental results showed that the shock structure, boundary-layer separation and peak fluctuating pressure were dependent on freestream Mach number, turning angle and model geometry. The two-threshold method was also used to estimate the zero-cross frequency. The relationship between peak pressure fluctuations and zero-cross frequency was obtain and compared with results from the studies using sharp convex corner.

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