Wind-tunnel investigation on aerodynamic characteristics of transonic wings

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 1.48×10⁶, 1.53×10⁶ and 1.57×10⁶, 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.