穿音速機翼之表面視流觀察研究

本研究以成功大學航太科技中心之穿音速風洞，進行半翼展模型的風洞實驗，利用表面流場觀察的方法以瞭解穿音速機翼與加上翼梢小翼(Winglet)的表面流場在穿音速流的特性，以及不同測試狀況下的改變情形。穿音速機翼所用翼形的設計馬赫數爲0.78，實驗之馬赫數分別爲0.75、0.80及0.85，對應的雷諾數分別爲Re(下標 C)=1.48×10^6、1.53×10^6及1.57×10^6，攻角爲0度、2度、與4度。在邊界層自然轉換的情形下，機翼上表面的層紊轉換爲層流邊界層分離所造成，其後發展爲分離泡或邊界層完全分離，並無後緣分離的現象發生。在邊界層面定轉換的情形下，皆有後緣分離現象發生：由於邊界層轉換區誘生器的影響，使震波前移。

關鍵字

視流觀察；翼梢小翼；邊界層轉換

並列摘要

The purpose of this study is to investigate the aerodynamic characteristics of a finit wing model tested in transonic regime by surface flow visualization. The experiments were conducted in a 600mm×600mm transonic wind tunnel located at the Aerospace Science and Technology Research Center of National Cheng-Kung University. The design Mach number of the transonic wing model is 0.78. The Mach number performed in the wind tunnel are 0.75, 0.80 and 0.85, respectively, and the corresponding Reynolds numbers, based on the mean chord of the wing, are 1.48×10^(-6)、1.53×10^6 and 1.57×10^6, respectively. The angles of attack chosen in the experiments are 0°、2° and 4°, respectively. In the case of boundary-layer free transition for flow over the wing, laminar separation occurred on the suction surface of the wing results in boundary-layer transition. This causes the development of separation bubble or massive flow separation downstream. There is no trailing edge separation seen in this case. In the case of boundary-layer fixed transition, adding the transition strip enlarges the region of trailing edge separation developed on the suction surface on of wing, This is associated with the shock located further upstream.

並列關鍵字

Flow Visualization ； Winglet ； Boundary-Layer Transition

國際替代計量

穿音速機翼之表面視流觀察研究

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