Abstract The aerodynamic performance and pressure ratio of modern aircraft engine compressor are degrading due to unsteady flow distorted loads complex phenomena in the blade tip. This paper attempts improve axial transonic stage under conditions by incorporating combination groove casing, tip injection technique with surface roughness effects. is performed for Mach number ranging between 0.8 1...

To explain the effect of tip leakage flow on performance an axial-flow transonic compressor, compressors with different rotor clearances were studied numerically. The results show that as clearance increases, intensity is increased, shock wave position moved backward, and interaction between vortex intensified, while boundary layer weakened. Most all, stall mechanisms varying are different. mai...

This paper focuses on the conceptual design of a two spool compressor for the NASA Large Civil Tilt Rotor engine, which has a design-point pressure ratio goal of 30:1 and an inlet weight flow of 30.0 lbm/sec. The compressor notional design requirements of pressure ratio and low-pressure compressor (LPC) and high pressure ratio compressor (HPC) work split were based on a previous engine system s...

This paper presents a comprehensive validation and verification study of turbomachinery Reynolds-averaged Navier-Stokes flow solvers on the transonic axial compressor TUDa-GLR-OpenStage. Two commercial namely Ansys CFX Numeca FineTurbo are adopted to provide benchmark solutions, which can be used for other RANS in future. Based these solvers, five sets grids, two advection schemes (i.e., centra...

Abstract The stability limit of a tip-stalling axial compressor is sensitive to the magnitude near-casing blockage. In transonic compressors, presence passage shock could be major cause for Identification and elimination this blockage key improving compressor. article, using numerical simulation, within rotor, NASA Rotor 37, quantified parameter. For smooth casing, at conditions near stall maxi...