IDDES method was applied to investigate the highly unsteady flow in a subsonic compressor stator with very large hub clearance and high incidence angle.The blade loading variation frequency was found close to the rotating instability(RI)frequency f RI&333 Hz observed in the experiment.Detailed analysis of the flow physics shows that the loading variation is caused by the periodic swing of the large scale separated flow on the blade suction side surface.The breakdown of the leakage vortex has no significant dominant frequencies,thus cannot be the cause of RI in this compressor stator as normally believed.Furthermore,the vortex shedding and vortex breakdown due to shear layer instability at the outer edge of the blade suction surface separation region excite high frequency unsteadiness that can form the sources of noise.
Space vehicle in atmosphere travels mostly at supersonic speed and generates a very strong bow shockwave around its blunt nose. Oblique shock and conical separated flow zone generated by a forward disk-tip spike significantly reduce the drag by reducing the high pressure area on the blunt nose. This study employs improved delayed detached eddy simulation to investigate the characteristic flow structures around a spike-tipped blunt nose at Mach number of 3 and Reynolds number（based on the blunt-body diameter） of 2.72x10;. The calculated time-averaged quantities agree well with experimental data. Characteristic frequencies in different flow regions are extracted using fast Fourier transform. It is found that two distinct instability modes exist: oscillation mode and pulsation mode. The former is related to the foreshock/turbulence interaction with nondimensional frequency at around 0.004. The latter corresponds to the interaction between turbulence and shock structures around the blunt nose, with a typical coherent structure shedding frequency at 0.092.
The paper investigates the effect of a single circumferential groove casing treatment(CGCT) on a transonic compressor rotor numerically.In particular,the effect of the groove at different axial locations on the flow field is studied in detail and stall margin improvement is also discussed.The present results show that the groove close to the leading edge plays a crucial role in stabilizing the near stall flow structures and,hence,improves the stall margin.The groove at the mid-chord-section of the blade can help exchange and transfer momentums between different directions,and suppress the flow unsteadiness,leading to increased efficiency in rotor performance and extended operation range.The groove located near the blade trailing edge has limited effects on stall margin improvement and may cause additional penalty in efficiency.Through comparison with the recent work on CGCT,some common flow physics can be observed.
