Measurements were made of the flow around a 10-percent-thick, double symmetrical, two-dimensional wedge at a Mach number of 1.30 and of a 14.2-percent-thick wedge at Mach numbers of 1.30 and 1.41 for various angles of attack up to 5 degrees. Results were thus obtained in the vicinity of the theoretically interesting region between shock attachment and the lower limit for completely supersonic flow over the surface of the airfoil. Pressure and Mach number distributions, lift and drag coefficients, center of lift, and pitching moment are presented for the angles of attack used. By means of the transonic similarity laws, the results are compared with each other, with small-disturbance theory, and with shock-expansion theory wherever possible. The data show that pressure distributions on wedges of different thickness and Mach number are similar at the same values of transonic similarity parameter and reduced angle of attack for angles of attack as large as the thickness ratio; that the lift-curve slope is approximately independent of the angle of attack for an angle-of-attack range from -2 degrees to 2 degrees; and that, for the airfoils tested at Mach numbers greater than the attachment value, the center-of-pressure location is nearly independent of the angle of attack, the variation being to plus or minus 3 percent chord for the angles of attack used in this investigation. For the airfoil tested at a Mach number slightly less than the shock-attachment value, the center-of-pressure location was only roughly independent of the angle of attack, the variation of this location being to plus or minus 6 percent chord.