A preliminary theoretical and experimental investigation has been made of the supersonic aerodynamic characteristics of blunt-trailing-edge airfoils with finite trailing-edge airfoils. Calculations of the drag of a family of airfoils with finite trailing-edge thickness are presented for various values of the base pressure. Theoretical expressions for the lift, pitching moment, and maximum lift-drag ratio are developed using the Buseman second-order theory for two-dimensional supersonic flow. In order to compare the theoretical estimates with experimental data, measurements were taken on the lift and drag on wings of various airfoil sections at Mach numbers of 1.5 and 2.0 and at Reynolds numbers varying from 0.2 to 1.2 million. Rectangular plan forms with an aspect ratio of 4 and a thickness ratio of either 10 or 9.1 percent were used throughout the experiments.

The pressres acting on the base of blunt-trailing-edge airfoils have been measured at Mach numbers of 1.25, 1.5, 2.0, and 3.1 and at Reynolds numbers from 0.2 to 3.8 million. Data are presented for 29 profiles both with laminar and with turbulent boundary layers approaching the trailing edges of the wings. The base pressure is found to be a function primarily of Mach number and the ratio of the boundary layer thickeness at the trailing edge to the trailing-edge thickness.

Comparisons are made between a theory for calculating stresses about rectangular cutouts in circular cylinders of semimonocoque construction published in NACA TN 3200 and previously published NACA experimental data. The comparisons include stresses in the stringers and shear stresses in the center of the shear panels in the neighborhood of the cutout. The theory takes into account the bending flexibility of the rings in the structure, and this factor is found to be important in the calculation of stresses about cutouts. In general, when the ring flexibility is considered, good agreement is exhibited between the calculated and experimental results.