An Effect of Tip Clearance on Aero Performance in Axial Flow Compressors for Aero Gas Turbine Engines

Gas turbine engines have dominated aviation industry, because they are highly efficient and reliable. Axial flow compressor is a critical part which plays a crucial role in efficiently working of an aero gas turbine engine. The entry and exit of the working fluid is in axial direction with respect to rotational axis in axial flow compressor. Higher mass flow swallowing capacityand higher efficiency for a small frontal area are unique features of axial flow compressors making an ideal option for the gas turbines used in jet engines., One of the major source of performance deterioration/enhancement in compressors is because of tip clearancebetween the rotor blade tip and casing. The tip clearance causes leakage of flow. The leaked flow interacts with the core flow and the wall boundary layer and control aerothermodynamics behavior of flow in the tip regions. This tip leakage flow also has an impact on efficiency and operating envelope of the compressor. In this work, a steady state analysis has been carried out to understand the influence of tip clearance on the aerodynamic performance of transonic single stage wide chord blade axial flow compressor stage through three dimensional viscous analysis using ANSYS CFX. The analysis was carried out for different tip clearance configuration varying from 0% to 1.2% of rotor tip axial chord. The characteristics plots like corrected massflow rate versus pressure ratio andcorrected massflow rate versus efficiency were plotted for various tip clearances at design speed. The generated database of tip clearance in this work will predictbeforehand the extent of deterioration due to manufacturing deviation of the components on aerodynamic performance prior to engine testing in the test bed.