This document contains reproductions of the technical papers presented at the NACA - University Conference on Aerodynamics held at the Langley Aeronautical Laboratory on June 21, 22, and 23, 1948. The conference was held in recognition of the difficulties, imposed by security restrictions, in keeping abreast of the rapid advances in aerodynamics. The papers were prepared to review the status of a number of fields of interest, to summarize the more important wartime advances that are no longer classified, and to orient reference material for further study.

Memorandum presenting equations for obtaining the wall coordinates of two-dimensional supersonic nozzles. The equations are based on the application of the method of characteristics to irrotational flow of perfect gases in channels. A brief discussion of characteristics as applied to nozzle design is given to assist in understanding and applying the nozzle-design method of the report.

Report presenting testing of seven models that from two different wing series: a series of taper ratio 0.5 and differing aspect ratio and a series of aspect ratio 4 and differing taper ratio. Measurements of lift, drag, and pitching moment were made at Mach number 1.53. The aerodynamic parameters generally varied with aspect ratio and taper ratio in the manner indicated by the linear theory.

Report discussing the effect of half-span and full-span split flaps through a deflection range on the low-speed, longitudinal characteristics of a sweptback wing with a round-nose, extensible leading-edge flap. Information about the effect of split-flap deflection on lift and stalling characteristics, effect of split flaps on gliding characteristics, and effect of leading-edge roughness on longitudinal stability is presented.

"Six numerical examples are presented for steady, two-dimensional, compressible, nonviscous flow in centrifugal compressors with thin straight blades, the center lines of which generate the surface of a right circular cone when rotated about the axis of the compressor. A seventh example is presented for incompressible flow. The solutions were obtained in a region of the compressors, including the impeller tip, that was considered to be unaffected by the diffuser vanes or by the impeller-inlet configuration" (p. 141).