Combustor Exit Temperature Distortion Effects on Heat Transfer and Aerodynamics Within a Rotating Turbine Blade Passage

Abstract

A numerical simulation of temperature distortion at inlet to a rotating turbine rotor has been performed. A 3D Navier-Stokes code the 2D boundary layer code - STAN5, have been used. The results show that the hot gas is transported to the pressure surface of the blade and that hot gas also migrates to the blade pressure side tip. At locations greater than 50–60% axial chord the hot gas enters the tip-gap and emerges over the suction side. These computations are in agreement with previous experimental results. The secondary flows within the turbine rotor are enhanced by the introduction of inlet radial temperature distortion and this is in accord with previous analytical work. It is shown that the heat flux near the tip region on the pressure side of the blade can be increased by up to 76% due to the re-distribution of the inlet temperature distortion.