Abstract
An experiment is carried out on a transonic flow over an airfoil, which is partially modified from an arbitrary airfoil. The modification is made to satisfy the shockless condition predicted by the same author [Jour. Japan Soc. Aero. Space Sci. 26 (1978) 191]. Static pressures are measured at 88 points on the airfoil surface, when the Mach number at wind tunnel wall is 0.5∼0.8, the measuring angle of attack is 0°∼4° and the Reynolds number is 4.0∼4.6×10 6 . A shockless transonic flow is observed under a measuring condition close to the design condition. The supersonic region extends from 0.4% to 63% of airfoil chord length and the local maximum Mach number is 1.36. The results support that shockless airfoils can be designed by partial modifications of an arbitrary airfoil.
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