Experimental Study of Supersonic Axial-Flow Compressor Blading

Abstract

Blade passages much affect the performance of the supersonic axial-flow compressor blading of internal shock type. To investigate their effects and obtain basic data for the design of a supersonic compressor having a good efficiency, a single channel investigation of five different channels was carried out at an inlet Mach number of 1.55, using Freon-113. Actual channel flows were radically different from those with an ideal fluid because of the shock-boundary layer interaction, that is, poor pressure rises through shocks and different starting conditions from theoretical ones. Results at the spill point are that the Mach numbers before the normal shocks in the divergent parts of channels and the losses in the main streams coincide approximately with those obtained from the one-dimensional analysis for an ideal flow, but the positions of the normal shocks and pressure rises through them depend upon the divergence angles and widths of channels. And separation was observed for the channel with a divergence angle of 10 deg.