The demand for more efficient turbocharger and aviation centrifugal compressors operating at higher pressure ratios and specific speeds with extended flow ranges is focusing research efforts on the inducer and diffuser transonic flow fields. At pressure ratios above 5.0 and specific speeds of unity inducer tip relative Mach numbers exceeding 1.4 can be encountered, precipitating both increased shock losses and diminished stall margin. The results of compressor rig testing on a research 6.8 inch tip (173mm) diameter single stage centrifugal compressor operating with inducer tip relative Mach number up to 1.5 are presented. The test results reveal high efficiency combined with extended flow range. This was achieved through improved impeller stability with shroud bleed, thereby permitting the diffuser to operate stably on its positive slope recovery characteristic.
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