Sensitivity Study of Stagger Angle on the Aerodynamic Performance of Transonic Axial Flow Compressors

Abstract

AbstractGas turbine engines are widely used in the aviation industries because of high power to weight ratio. The gas turbine engine consists of the three main components: compressor, combustion chamber, and turbine. The design of the axial flow compressor depends on various design parameters such as blade stagger angle, blade height, blade chord, rotor tip clearance, and stator tip clearance. Blade stagger angle is a critical parameter which plays vital role in performance of the gas turbine engines. In this study, a steady state analysis has been carried out to understand the effect of rotor and stator blade stagger angle on the aerodynamic performance of single stage axial flow transonic compressor through three dimensional viscous analysis using ANSYS CFX 19.2 software. The analysis was carried out for various stagger angle configuration varying from 24° to 36° for rotor and 06° to 18° for stator. The study concluded that as the rotor stagger angle is increased from 24° to 36°, the mass flow is reduced by 16.5%, peak pressure ratio by 2.8%, and increase peak efficiency by 6.8%. The effect of change of stator stagger angle from 06° to 18° is insignificant.KeywordsAxial flow compressorStagger angleTip clearanceSurge marginComputational Fluid Dynamics (CFD)Pressure ratioEfficiency