Influence of Endwall Contouring on the Transonic Flow in a Compressor Blade

Abstract

Endwall contouring is successfully applied to compressor rotors to reduce boundary layer loading and to control endwall flow. Over-speeds resulting from large relative thickness and curvature of the rotor root section are compensated by the increase in open flow area which is generated by the concave hub shape. In transonic flow this area increase promotes higher Mach numbers and has a considerable impact on the shock system. To investigate endwall contouring experimentally at engine like flow conditions a novel cascade technique is introduced: A contracting endwall is designed in a way to enable homogenious flow conditions in Mach number and flow angle at the inlet plane of the cascade. For a given blade shape now several endwall contours may be investigated. Experiments are performed at the High Speed Cascade Windtunnel of the University of the Armed Forces, Munich for a linear and a concave endwall for a given blade section. Inlet Mach number level is around Ma1 = 0.9 at typical turning and profile thickness. The results show an increase in pre-shock Mach number and a change in shock pattern from an oblique shock for the linear contour to a normal shock for the concave one. Endwall contouring is demonstrated not only to influence the flow in the vicinity of the endwall but to extend up to a considerable distance in spanwise direction.