A study of the end wall boundary layer in an axial compressor blade row

Abstract

Turbulent end wall boundary layers in the hub region of axial compressor blade rows have been studied in the Gas Turbine Laboratory. Both analytical and experimental work are presented. An attempt was made to predict the growth of the boundary layer momentum thickness as it passes through a cascade of blades using the momentum integral equation and several assumptions This calculation was compared with measurements taken along an assumed free streamline in a stationary cascade of blades supplied with an artificially skewed inlet boundary layer. From this comparison, the relative influence of individual terms in the momentum integral equation is deduced. The effects of inlet skewing and cross flow in the boundary layer in the region between blades is described. Two methods of representing main and cross flow velocity profiles are compared with experimental data.