| Abstract:To analyze the effect of rotor tip clearance values and tip clearance shape modifications on compressor performance,
numerical simulations were conducted on a typical axial flow transonic compressor NASA Stage 35 with different tip clearance values. The
results show that with the increase of tip clearance and leakage flow intensity, the position of the detached shock wave moves forward, the
position of the shock wave acting on the suction surface of the rotor blade moves backward, leading edge spillage phenomenon is weakened,
the interaction between the tip leakage vortex and the shock wave is strengthened, the vorticity in the tip clearance area increases, and the
vortex core near the leading edge moves downstream, but the interaction between the boundary layer and the shock wave is weakened; The
stall mechanisms of compressor with different tip clearance values were different, the stall of the compressor with small tip clearance is
caused by the joint action of the of the tip leakage vortex and the rotor blade suction surface boundary layer separation, while the stall of the
compressor with large tip clearance is caused by tip leakage vortex; Within the range investigated, when the tip clearance is 0.50τ, the
compressor has the best performance, with peak efficiency and stability margin increased by 0.16% and 2.39%, respectively; After the
optimized uniform tip clearance is modified to parallel to divergent tip clearance and sinusoidal tip clearance, the compressor performance
is improved, the detached shock wave moves backward, the interaction between tip leakage vortex and shock wave is weakened, and the
boundary layer separation on the rotor blade suction surface is weakened, and the tip clearance shape modification can reduce blade mate?
rial usage and engine mass. |
