超音速圧縮機翼列に関する実験

抄録

Blade passages affect much on 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 for five different passages was carried out at the entrance Mach number of 1.55, using Freon 113. Real channel flows were radically different from those with ideal fluid because of the shock-boundary layer interaction, that is, having poor pressure rises through shocks and different starting conditions from theoretical ones. Results at the maximum back pressure are that the Mach numbers ahead of the normal shocks in the divergent portions of the passages and losses in the free streams coincide approximately with those obtained from one-dimentional analysis for an ideal flow, but the locations of the normal shocks and pressure rises through them depend upon the divergent angles and widths of the passages. And separation was observed for the passage with its divergent angle of 10°.