Results of supersonic tunnel tests made to obtain the characteristics of conventional control surfaces are summarized. A symmetrical double-wedge airfoil of 7 per cent thickness, with 15 and 30 per cent trailing-edge flaps, was used as a model. Pressure measurements and Schlieren photographs of the airfoil were taken in a two-dimensional channel at Mach Numbers 1.38, 1.48, and 1.58. The pressure data were reduced to lift coefficient, drag coefficient, lift to drag ratio, section moment coefficient, flap moment coefficient, and center of pressure points for every test angle of attack. Comparisons of the experimental data were made with the linearized and shock-expansion theories of references 1, 2, and 3. Agreement with Busemann's secondorder linear theory and the shock-expansion theory is good except where such unpredictable phenomena as the following exist: (1) the shock detachment from the leading edge of the airfoil, (2) the forward movement of the shock emanating at the hinge point of the flap, and (3) separation on the upper flap surface to a limiting case at which the top trailing-edge shock moves to the hinge point. The tests were especially significant in determining the effects of these irregularities on airfoil characteristics. Since a presentation of all the material would be somewhat lengthy, only representative results are given in this paper.