Supersonic Cascade Tunnel Used to Evaluate Compressor Blade Performance

Abstract

Large improvements in specific fuel consumption and weight have been achieved during the development of the aircraft gas turbine engine in the past decade. The turbo-machinery requirements have become more demanding during this period, resulting in a significant increase in relative Mach number and diffusion loading in fan and compressor stages. Considerable effort was devoted to develop design systems for this type of turbomachinery, in which relative supersonic inlet velocities are realized for rotor and stator blades. One of the major requirements was to develop a design system for the selection of supersonic blade elements which are required for fan and compressor stages employed in turbofan engines. Since subsonic blade selection design systems are based primarily on empirical relationships which are derived from two-dimensional cascade data, a two-dimensional cascade tunnel was developed for: (a) the investigation of supersonic cascade aerodynamics; (b) the development of high performance supersonic blade elements; and (c) the production of basic supersonic cascade data, which is required for the extrapolation of current subsonic design systems. The present paper describes this facility, presents some typical data, and discusses some of the results obtained.