Effect of Rotor Tip and Stator Hub Clearance on Aero Performance in Axial Flow Compressors for Aero Gas Turbine Engines

Gas turbine engines are widely used as power plants. Absence of reciprocating and rubbing parts in gas turbine engine helps in lower consumption of lubrication oil and provides higher power to weight ratio with high reliability, apart from reduction in balancing and vibration problems. Turbine inlet temperature and efficiency of components are two main factors which are responsible for gas turbine engine performance. Compressor is one of the main component which directly influences the efficiency of the gas turbine engine. The tip clearance causes leakage of flow which interacts with the main passage flow and the wall boundary layer region causing performance deterioration in compressors. The efficiency and compressor operating envelop gets impacted with the tip leakage flow. In this study, a steady state analysis has been carried out to understand the combined effect of radial clearance on rotor and the stator hub on the aerodynamic performance of intermediate stage axial flow compressor through three dimensional viscous analysis using ANSYS CFX-19.2. The analysis was carried out for 30 different tip and hub clearance configurations varying from 0 to 6% of rotor tip axial chord. Out of 30 configurations only 11 combination of hub and tip clearance were selected meeting the design requirement of surge margin. The study concluded that out of 11 configurations, the combination having 1.5% rotor tip clearance and 3.0% stator hub clearance resulted into higher efficiency of 3%, higher pressure ratio of 4% and surge margin improvement from 15 to 18.3% a gain of about 22%. It is also observed that the stator performance is better with hub clearance for the configuration under investigation.