A design strategy for a 6:1 supersonic mixed-flow compressor stage

Abstract

A surge in the small jet engine market due to aero-propulsion purposes generates requirement to develop compact and robust high-performance compressors. We address this necessity through the design of a single-stage high-pressure ratio mixed-flow compressor. Its compactness and reliability demonstrate its ability to replace a multistage axial design in the small aero-engine segment with high-performance envelope. We have perceived that though many design approaches are readily available for centrifugal and/or axial stages, mixed-flow compressor design systems are scarce.In this paper we intend to provide the designer a comprehensive background knowledge of a mixed-flow stage design. A brief historical development of these designs since the 1940s has been provided. It is observed that for a high-pressure ratio demand it necessitates a supersonic rotor exit flow. Hence, tandem stator configurations were investigated in the past to reduce blade loadings for efficient diffusion. However, most of the previous stage designs were inefficient due to inability of the stators to efficiently diffuse this supersonic flow. A tandem design based on Quishi et al. [1] has been implemented to solve this problem.A unique mean-line procedure based on isentropic equations is defined for mixed-flow stage. It is followed by a geometry construction technique based on Bezier curves. Furthermore, a rotor design evaluation study is conducted for 3.5 kg/s mass flow based on the mean-line code and additional computational analysis. Current computational results [2] have shown single-stage mixed-flow compressors designed using this method to generate reasonably high-pressure ratio up to 6:1 with 75.5% efficiency.