"The present report contains the results of a series of observations obtained for a wing of symmetrical profile for different angles of yaw. The shock tunnel with 0.4m x 0.4m cross section, of the Institute for Gas Dynamics at L.F.A. - Braunchweig was available for this work. The profile used was a 9 percent, thick hyperbolic profile with maximum thickness 40 percent aft, which was calculated by the method of F. Ringleb (2). Since the conformal transformation of this profile is known, the theoretical pressure distribution could be determined exactly for the case of incompressible flow. Then by the use of the Prandtl rule one may extend the comparison of theory and experiment to the case of higher velocity of incident flow" (p. 1).

Wind-tunnel measurements on projectiles are discussed. Tests at the Gottingen Tunnel are described. The tunnel operates on the Prandtl principle, that is, a brief stationary air stream produced in an evacuated tank by induction of atmospheric air.

This report addresses a method for the approximate calculation of compressible flows about profiles with local regions of supersonic velocity. The flow around a slender profile is treated as an example.

"Two procedures for calculating the lift distribution along the span are given in which a better account is taken of the distribution of circulation over the area than in the Prandtl lifting-line theory. The methods are also applicable to wing sweepback. Calculated results for the two methods were in agreement" (p. 1).

The calculation of infinitesimal conical supersonic flow has been applied first to the simplest examples that have also been calculated in another way. Except for the discovery of a miscalculation in an older report, there was found the expected conformity. The new method of calculation is limited more definitely to the conical case.

"The motion of different bodies immersed in liquid or gaseous media is accompanied by characteristic sound which is excited by the formation of unstable surfaces of separation behind the body, usually disintegrating into a system of discrete vortices (such as the Karman vortex street due to the flow about an infintely long rod, etc.). In the noise from fans, pumps, and similar machinery, vortex noise frequently predominates. The purpose of this work is to elucidate certain questions of the dependence of this sound upon the aerodynamic parameters and the tip speed of the rotating rods or blades" (p. 1).

"In the following, high-speed measurements on a swept-back wing are reported. The curves of lift, moment, and drag have been determined up to Mach numbers of M = 0.87, and they are compared to a rectangular wing. Through measurements of the total-head loss behind the wing and through schlieren pictures, an insight into the formation of the compression shock at high Mach numbers has been obtained" (p. 1).

"The previous measurements on airfoils with hinged nose disclosed a comparatively large low-pressure peak at the bend of the hinged nose; which favored the separation of flow. It was therefore attempted to reduce these low-pressure peaks by reducing the camber of the forward profile and thereby ensure a longer adherence of the flow and a maximum lift increase. The forces were measured on a rectangular wing with double-hinged nose and end plates, the pressure distributions were measured in the center section of the wing" (p. 1).

Force measurements and pressure distribution measurements on the midsection were made on a rectangular wing with slotted droop nose and end plates, on which could be placed a choice of either a plain flap-split flap combination or a slotted flap.