A numerical case study of a multistage, highly-loaded, relative supersonic compressor is presented. The purpose of the investigation was to highlight the changing shock structure while throttling the compressor and to give insight into possible compressor instabilities. The computational fluid dynamic (CFD) study was conducted with the NASA code ADPAC, utilizing the mixing-plane assumption for the boundary condition between adjacent, relatively-rotating blade rows. A steady, five-blade-row, numerical simulation using the Baldwin-Lomax turbulence model was performed, creating several constant speed lines. The results show that the shock structure in the downstream rotor isolates the upstream rotor from the exit conditions until the shock detaches from the leading edge. The shock structure in the upstream rotor then moves, changing the conditions for the downstream rotor. This continues as the compressor is throttled until the shock in the upstream rotor detaches from the leading edge. CFD indicates that this event causes a rapid drop in the mass flow rate, creating a mismatch between stage-one and stage-two that results in compressor instability.
Numerical Study of Embedded Supersonic Compressor Stages
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Kempf, SG, Guillot, S, Ng, WF, Wellborn, SR, & Chriss, RM. "Numerical Study of Embedded Supersonic Compressor Stages." Proceedings of the ASME Turbo Expo 2005: Power for Land, Sea, and Air. Volume 6: Turbo Expo 2005, Parts A and B. Reno, Nevada, USA. June 6–9, 2005. pp. 237-246. ASME. https://doi.org/10.1115/GT2005-68611
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