A transonic axial flow compressor rotor

the NASA Rotor 67

was chosen to investigate the triggering mechanism of internal flow instability in transonic axial flow compressor at the 100%

80% and 60% rotating speeds with the help of numerical method. The comparative analysis of numerical results and experimental data showed that the trends of experimental performance curves were finely repeated by numerical results under three design rotating speeds. The fundamental flow mechanism was obtained by the detailed analysis of internal flow field in compressor. As the mass flow rate of compressor reduced at three rotating speeds

the starting position of tip leakage vortex (TLV) moved to the blade leading edge gradually

and tip leakage vortex also turned towards the pressure surface of adjacent blade. The deviated degrees between tip leakage vortex trajectory and compressor rotating shaft for near stall point were 3 degree

6 degree and 9 degree than that for near peak efficiency point at the 100%

80% and 60% rotating speeds

respectively. The blockage resulted from the interaction between tip leakage vortex and shock wave led to the internal flow instability in compressor at the 100% and 80% rotating speeds

and tip leakage vortex broken occurred at the 80% rotating speed. While at the 60% rotating speed

the leading edge spilled flow(LESF) of blade tip caused by tip leakage vortex near adjacent blade was the primary cause of internal flow instability in compressor

while a small scope boundary layer flow separation(BLFS) near the trailing edge of blade suction surface was not the primary cause.