Flow structure driven end-wall contouring for a highly loaded axial compressor

Reducing stage loss by secondary flow structures is important for improving the efficiency and the stall margin in highly loaded axial compressor stages. A numerical investigation of the NASA 37 transonic axial compressor shows the significant growth of separated flow on the stator suction surface. Simulations show that about two thirds of the stator blade span is covered by separated flow at the trailing edge. The flow structure on the hub surface is affected by a cross-flow from the blade pressure side. In this work, a flow structure driven approach to the design of an end-wall treatment is used to address this flow separation. To decrease the interaction of the cross-flow with the suction surface flow, a guide groove is defined such as to guide the cross-flow driven by the passage pitchwise pressure gradient towards the trailing edge of the adjacent stator blade. The groove shape is defined by five geometry functions. These functions use six independent parameters in total to define the groove shape. A Beta distribution function is used to create a smooth connection between the blade root and the groove. The effect of these parameters on the total pressure loss and on the flow structures is evaluated through RANS simulations. The comparison between the numerical predictions with the groove against corresponding results with an axisymmetric hub, which are validated against experiment, shows a reduction in the stator trailing edge flow separation area. This gives a reduction in the total pressure loss of about 5%. Limit streamlines over the hub indicate that the groove reduces the mass flow feed to the suction side trailing edge region of flow separation. A parametric study on the groove slope and the depth is carried out to maximise this beneficial effect.