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An apparatus and method is provided for improving efficiency of a transonic gas turbine engine compressor by bleeding off a shockwave-induced boundary layer from the gas flow passage of the compressor using an array of bleed holes having a downstream edge aligned with a foot of an oblique shock wave

An apparatus and method is provided for improving efficiency of a transonic gas turbine engine compressor by bleeding off a shockwave-induced boundary layer from the gas flow passage of the compressor using an array of bleed holes having a downstream edge aligned with a foot of an oblique shock wave which originates on the leading edge of an adjacent transonic rotor blade tip.

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The invention claimed is: 1. A transonic gas turbine engine compressor comprising: a rotor having a central axis of rotation and a plurality of blades extending into a gas flaw passage through said compressor, each of said blades having a blade tip and a leading edge defined between opposed pressur

The invention claimed is: 1. A transonic gas turbine engine compressor comprising: a rotor having a central axis of rotation and a plurality of blades extending into a gas flaw passage through said compressor, each of said blades having a blade tip and a leading edge defined between opposed pressure and suction surfaces, said rotor being rotatable about said axis of rotation at a speed such that gas flow adjacent said blade tips becomes supersonic, creating oblique shock waves originating at said leading edge and terminating at a shock foot on said suction surface of an adjacent blade; an outer shroud surrounding said rotor, said outer shroud defining a radially outer boundary of said gas flow passage; a plurality of bleed holes extending through at least a portion of said outer shroud adjacent said blade tips to provide gas flow communication between said gas flow passage and an outer bleed passage, said plurality of bleed holes defining a bleed bole array, and said bleed hole array defining a downstream edge thereof substantially aligned in a gas flow direction with said shock foot and extending upstream thereof, said bleed holes in said array being selected in size, number and location to bleed at least a portion of a shockwave-induced boundary layer from said gas flow passage adjacent said outer shroud. 2. The compressor as defined in claim 1, wherein said bleed hole array defines an upstream edge thereof substantially aligned with said leading edge of said blade on which said oblique shock originates. 3. The compressor as defined in claim 1, wherein said bleed hole array is disposed upstream of each said shock foot on said adjacent blade and downstream of said leading edge of said blade on which said oblique shock originates. 4. The compressor as defined in claim 1, wherein said compressor is an axial compressor. 5. The compressor as defined in claim 1, wherein said bleed hole array extends circumferentially about said outer shroud. 6. The compressor as defined in claim 5, wherein said bleed passage is an annular cavity surrounding said outer shroud. 7. The compressor as defined in claim 1, wherein said selection of bleed holes is such that less than 5% of total gas flow through said gas flow passage is bled off by said bleed hole array. 8. The compressor as defined in claim 7, wherein said bleed holes are arranged and sized such that about 1% of said total gas flow is bled off by said bleed holes. 9. The compressor as defined in claim 1, wherein said outer shroud defines a bleed exhaust opening therein, said bleed exhaust opening providing gas flow communication between said outer bleed passage and said gas flow passage such that gas flow, bled into said outer bleed passage via said plurality of bleed holes, is exhausted back into said gas flow passage. 10. The compressor as defined in claim 9, wherein said bleed exhaust opening is disposed upstream from said rotor. 11. The compressor as defined in claim 1, wherein said bleed hole array defines a band substantially parallel to a plane perpendicular to said gas flow direction. 12. The compressor as defined in claim 1, wherein said bleed hole array defines a band disposed at an angle relative to a plane perpendicular to said gas flow direction. 13. The compressor as defined in claim 1, wherein said bleed hole array defines at least three rows of said bleed holes between said downstream edge and an upstream edge of the bleed hole array. 14. A method of bleeding a shockwave-induced boundary layer from a transonic gas turbine engine compressor comprising: providing a rotor having a plurality of blades extending into a gas flow passage of the compressor and an outer shroud surrounding said rotor; rotating said rotor such that gas flow adjacent tips of said blades becomes supersonic, creating an oblique shockwave originating at a leading edge of each said tip and terminating at a shock foot on a suction surface of a respective adjacent blade; providing a plurality of bleed holes in said outer shroud in a predetermined region corresponding to a boundary layer induced by said oblique shock wave; arranging said plurality of bleed holes in a bleed hole array; positioning said bleed hole array such that a downstream edge thereof is substantially aligned, in a gas flow direction within said gas flow passage, with said shock foot of each said oblique shockwave and bleeding said shockwave-induced boundary layer out of said gas flow passage. 15. The method as defined in claim 14, further comprising passively bleeding said shockwave-induced boundary layer. 16. The method as defined in claim 14, further comprising selecting a number and size of said bleed holes such that about 1% of total gas flow through said gas flow passage is bled off by said bleed holes. 17. The method as defined in claim 14, further comprising exhausting gas bled via said bleed holes back into said gas flow passage. 18. The method as defined in claim 17, wherein said gas is reintroduced into said gas flow passage upstream from said rotor.