Investigation of a One Stage Axial Flow Supersonic Compressor, Part 2. Blade Section Performance of the Blunt Trailing Edge Rotor R-32

摘要

A high pressure ratio, supersonic axial flow compressor was investigated in a range of relative inlet Mach numbers at tip radius of the rotor from 0.66 to 1.45. The compressor blading was the blunt trailing edge blade. All tests were made by using a Freon-12 + air (20%) atmosphere because of power and speed limitations. Part II describes the blade sections performance, flow angle and velocity distribution and diffusion parameters. All data discussed concern measurements in a plane one axial chord downstream the rotor, which is tested without inlet guide vanes and without stator. (Author)