Numerical Study on Centrifugal Compressors for Small Aero Engines. Validation of a Radial Impeller and Diffuser Design for Minimal Frontal Area

Abstract

Most modern aircraft engines employ multistage axial compressor configurations. However, in small aero engines, axial compressors face substantial challenges. As mass flow is conserved, blade height decreases with each stage, and in small compressors this reduction becomes significant due to end wall boundary layer growth, large tip gap ratios, and pronounced secondary flow effects. An alternative is the use of an axi-centrifugal compressor configuration, where the high pressure axial stages are replaced by a centrifugal compressor. While centrifugal compressors provide higher pressure ratios within a shorter axial length, they introduce challenges related to engine frontal area. The frontal area scales with the compressor pressure ratio and is constrained by the overall engine geometry. The objective of this thesis is to numerically investigate centrifugal compressor design concepts aimed at achieving high aerodynamic performance within a confined space. The NASA High Efficiency Centrifugal Compressor impeller is replicated and validated. Subsequently, four diffuser designs are developed and evaluated using Computational Fluid Dynamics. Vaned diffusers are analysed and compared using standard performance parameters, and the influence of splitter vanes is examined. The results indicate more stable performance for the wedge diffuser compared to the cascade design. Splitter vanes extend and stabilize the operating range of both designs, but at the cost of increased total pressure losses. Due to wake-related issues in the wedge diffuser, the cascade combined with splitter vanes is found to be beneficial. While the goal of achieving significantly higher performance is only partially met, the study provides valuable insights into diffuser design trade-offs for small aero engine applications.