Effect of inlet radial distortion on aerodynamic stability in a multi-stage axial flow compressor

Abstract

Inlet radial distortion caused by boundary layer separation is experimentally studied in a three-stage axial flow compressor. Two lattice rings were designed to generate tip and hub radial distortion, respectively, and distortion intensity was measured using a five-hole probe installed upstream of the first rotor blade row. The unsteady flow field is captured by using a collection of pressure transducers installed on the casing wall with circumferential and chordwise spatial resolution. The stall route measured on the casing wall indicates that the first stage stalls under all inflow conditions, tip radial distortion can modify the scale of stall inception to long-length scale (6–8 blade passages), unlike the uniform inflow and hub radial distortion with typical short-length-scaled inception (3–4 blade passages). Detailed measurements indicate that tip radial distortion can strengthen the tip leakage vortex and increase the tip separation to induce the compressor to stall earlier than uniform inflow. However, hub radial distortion redistributes the mainstream flow toward the blade tip region, thereby weakening the unsteadiness of tip leakage flow, reducing the outlet flow angle and increasing the outlet axial velocity near the blade tip to remove the separation or blockage, and finally extending the stall margin. The aerodynamic measurements downstream different rotor blade rows also indicate that the inlet radial distortion mainly affects the flow field at the first stage, which is obviously weakened at the second stage and completely unaffected at the third stage.