The control of secondary flow in compressor cascades

Corner separation at the NACA 65-K48 cascade . Wake measuring equipment at the cascade wind tunner

The losses in axial compressors are primarily caused by tip clearance and secondary flow losses. Due to the high positive pressure gradient in flow direction the mean flow is decelerated and the sidewall boundary layers are thickened. In each compressor stage the sidewall boundary layer thickness is increasing and thereby the risk of flow separation is also increasing.

The secondary flow at the sidewalls (casing, hub) is causing the corner separation on the suction side of the vane which is responsible for losses and the decrease of the achievable pressure ratio in a compressor stage.

By the use of methods of the high lift aerodynamics of airplanes it is possible to eliminate or at least to reduce the flow separation. Vortex generators, boundary layer fences, slots in the vanes as well as trailing edge modifications are utilized to influence the flow. A NACA 65-K48 profile in cascade arrangement is used to investigate the potential of these methods at peak efficiency for their application in compressors.

A special high speed cascade wind tunnel was built in order to investigate the corner separation at realistic Mach and Reynolds number (M = 0.7, Re = 0.6 x 106) and investigate the increase in efficiency by influencing the secondary flow. The sidewall boundary layer thickness is continuously adjustable. Stagger and inflow angle can be adjusted within wide range and the complete cascade can be replaced within a few minutes. Probe arrays and computerised data acquisition allow the assessment of many configurations within a short period of time. Oil-streak pattern can visualize the flow near the vane surface to obtain more information about the average flow field including secondary flow phenomena.