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A compressor for a gas turbine engine having a bleed air recirculation system includes a plurality of bleed holes extending through the shroud at a first axial location thereon substantially adjacent the blade tips. The bleed holes have a closed outer perimeter along their complete length. An annula

A compressor for a gas turbine engine having a bleed air recirculation system includes a plurality of bleed holes extending through the shroud at a first axial location thereon substantially adjacent the blade tips. The bleed holes have a closed outer perimeter along their complete length. An annular bleed cavity surrounds the shroud and is in communication with outlet openings of the bleed holes. The bleed holes provide communication between the main gas flow passage and the bleed cavity. The bleed cavity includes exit passages having outlets disposed in the shroud at a second axial location which is upstream of both the first axial location and the leading edge of the blades of the rotor. Bleed air is passively bled from the main gas flow passage via the bleed holes, recirculated through the bleed cavity and re-injected back into the main gas flow passage at the second axial location.

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1. A compressor for a gas turbine engine comprising: a rotor having a hub defining a central axis of rotation and a plurality of blades radially extending from the hub to project into an annular main gas flow passage of said compressor, each of said blades having a remote blade tip and a leading edg

1. A compressor for a gas turbine engine comprising: a rotor having a hub defining a central axis of rotation and a plurality of blades radially extending from the hub to project into an annular main gas flow passage of said compressor, each of said blades having a remote blade tip and a leading edge defined between opposed pressure and suction surfaces, said rotor being rotatable about said axis of rotation;an annular shroud circumferentially surrounding the rotor and having a radially inner surface adjacent to the blade tips, the inner surface of the shroud facing said main gas flow passage; anda bleed air recirculation system including: a plurality of bleed holes linearly extending through the shroud at a first axial location thereon substantially adjacent the blade tips, each of the bleed holes having a closed outer perimeter along a complete length thereof defined between an inlet opening and an outlet opening of the bleed hole, a linear bleed hole axis extending along the complete length between a center of the inlet opening and a center of the outlet opening of each said bleed hole, the inlet opening of each bleed hole being disposed in the radially inner surface of the shroud adjacent the blade tips and the outlet opening being disposed in a radially outer surface of the shroud, the bleed air flowing through the bleed holes from the inlet opening to the outlet opening, the outlet opening of each of the bleed holes being located circumferentially upstream of the inlet opening relative to a direction of rotational flow in the main gas flow passage as driven by a direction of rotation of the rotor, such that the linear bleed hole axis is disposed at an acute angle relative to a radially extending reference axis disposed at the inlet opening of the bleed hole; andan annular bleed cavity surrounding the shroud and disposed in gas flow communication with the outlet openings of the bleed holes, the bleed holes defining inlet passages to the annular bleed cavity and providing gas flow communication between said main gas flow passage and the bleed cavity, the bleed cavity having one or more exit passages having outlets thereof disposed in said shroud at a second axial location thereon, the second axial location being disposed upstream of the first axial location and upstream of the leading edge of the blades of the rotor;wherein, in use, a pressure differential existing between the bleed holes and the bleed cavity exit passage outlets circulates a portion of gas flow in the main gas flow passage through the bleed cavity, said portion being fed into the bleed cavity via the bleed holes, and re-injects said portion in the bleed cavity back into the main gas flow passage at the second axial location upstream of both the bleed holes and the leading edge of the blades of the rotor. 2. The compressor as defined in claim 1, wherein each of the bleed holes has a non-constant cross-sectional area along said length thereof. 3. The compressor as defined in claim 2, wherein the bleed holes are frusto-conical. 4. The compressor as defined in claim 3, wherein the inlet openings of the frusto-conical bleed holes define a first cross-sectional area and the outlet openings thereof define a second cross-sectional area, the second cross-sectional area being greater than the first cross-sectional area to thereby define a radially inwardly tapered bleed hole. 5. The compressor as defined in claim 1, wherein the bleed holes have a substantially circular cross-sectional perimeter. 6. The compressor as defined in claim 1, wherein the inlet opening and the outlet opening of each of the bleed holes define different perimeter shapes. 7. The compressor as defined in claim 1, wherein the inlet opening of the bleed holes has an elliptical perimeter shape. 8. The compressor as defined in claim 1, wherein the bleed hole axis define a bleed path along which the bleed air flows from the inlet end to the outlet end of the bleed holes, the bleed hole axis and therefore the bleed air having both a radial and a circumferential component, the circumferential component being opposite in direction to a circumferential component of the rotational flow in the main gas flow passage. 9. The compressor as defined in claim 1, wherein the bleed holes are in substantial alignment with a direction of tip leakage air flow formed radially between the blade tips of the rotor blades and the inner surface of the shroud. 10. The compressor as defined in claim 1, wherein the acute angle defined between the linear bleed hole axis and the radially extending reference axis is between 25 and 65 degrees. 11. A gas turbine engine comprising: a compressor section, a combustor and a turbine section, in serial flow communication; andthe compressor section having at least one axial compressor including: an axial rotor having a hub defining a central axis of rotation and a plurality of blades radially extending from the hub to project into an annular gas flow passage of the compressor, each of said blades having a remote blade tip and a leading edge defined between opposed pressure and suction surfaces of the blades, said rotor being rotatable about said axis of rotation;a shroud circumferentially surrounding the rotor and having a radially inner surface adjacent to the blade tips, the inner surface of the shroud facing and radially enclosing said annular gas flow passage; anda plurality of bleed holes extending through the shroud at a first axial location thereon substantially adjacent the blade tips, each of the bleed holes having a closed outer perimeter along a complete length thereof defined between an inlet opening and an outlet opening of the bleed hole, a linear bleed hole axis extending along the complete length between a center of the inlet opening and a center of the outlet opening of each bleed hole, the inlet opening of each bleed hole being disposed in the radially inner surface of the shroud adjacent the blade tips and the outlet opening being disposed in a radially outer surface of the shroud, the bleed air flowing through the bleed holes from the inlet opening to the outlet opening, the outlet opening of each of the bleed holes being located circumferentially upstream of the inlet opening relative to a direction of rotational flow in the main gas flow passage as driven by a direction of rotation of the rotor, such that the linear bleed hole axis is disposed at an acute angle relative to a radially extending reference axis disposed at the inlet opening of the bleed hole; andan annular bleed cavity surrounding the shroud and disposed in gas flow communication with the outlet openings of the bleed holes, the bleed holes defining inlet passages to the annular bleed cavity and providing gas flow communication between said main gas flow passage and the bleed cavity, the bleed cavity having one or more exit passages having outlets thereof disposed in said shroud at a second axial location thereon, the second axial location being disposed upstream of the first axial location and upstream of the leading edge of the blades of the rotor;wherein, in use, a pressure differential existing between the bleed holes and the bleed cavity exit passage outlets draws a portion of gas flow in the main gas flow passage into the bleed cavity via the bleed holes and re-injects said portion in the bleed cavity back into the main gas flow passage at the second axial location upstream of both the bleed holes and the leading edge of the blades of the rotor. 12. The gas turbine engine as defined in claim 11, wherein each of the bleed holes has a non-constant cross-sectional area along said length thereof. 13. The gas turbine engine as defined in claim 12, wherein the bleed holes are frusto-conical. 14. The gas turbine engine as defined in claim 13, wherein the inlet openings of the frusto-conical bleed holes define a first cross-sectional area and the outlet openings thereof define a second cross-sectional area, the second cross-sectional area being greater than the first cross-sectional area to thereby define a radially inwardly tapered bleed hole. 15. The gas turbine engine as defined in claim 11, wherein the bleed holes have a substantially circular cross-sectional perimeter. 16. The gas turbine engine as defined in claim 11, wherein the inlet opening and the outlet opening of each of the bleed holes define different perimeter shapes. 17. The gas turbine engine as defined in claim 11, wherein the inlet opening of the bleed holes has an elliptical perimeter shape. 18. The gas turbine engine as defined in claim 11, wherein the bleed hole axis defining a bleed path along which the bleed air flows from the inlet end to the outlet end of the bleed holes, the bleed hole axis and therefore the bleed air having both a radial and a circumferential component, the circumferential component being opposite in direction to a circumferential component of the rotational flow in the main gas flow passage. 19. The gas turbine engine as defined in claim 11, wherein the acute angle defined between the linear bleed hole axis and the radially extending reference axis is between 25 and 65 degrees. 20. A method of bleeding tip leakage flow from a gas turbine engine compressor comprising: providing a rotor rotatable about an axis of rotation within an outer shroud surrounding said rotor, the rotor having a plurality of radially projecting blades extending into an annular gas flow passage of the compressor, the annular gas flow passage being radially enclosed by an inner surface of the outer shroud, each of said blades having a remote blade tip and a leading edge defined between opposed pressure and suction surfaces, the inner surface of the shroud being adjacent to the blade tips and radially enclosing said annular gas flow passage;rotating said rotor to generate a main compressor flow within the annular gas flow passage in a first rotational direction corresponding to a direction of rotation of the rotor, a tip leakage flow being formed between the blade tips and the inner surface of the shroud, the tip leakage flow being in a direction opposite to the direction of rotation of the rotor;bleeding off at least a portion of the tip leakage flow using a plurality of bleed holes through which bleed air flows, the bleed holes extending through the shroud at a first axial location thereon substantially adjacent the blade tips, each of the bleed holes having a closed outer perimeter along a complete length thereof defined between an inlet opening and an outlet opening of the bleed hole, a linear bleed hole axis extending along the complete length between a center of the inlet opening and a center of the outlet opening of each said bleed hole, the linear bleed hole axis defining a bleed path along which the bleed air flows from the net end to the outlet end of the bleed holes, the inlet opening being disposed in the inner surface of the shroud adjacent the blade tips and the outlet opening being disposed in an outer surface of the shroud, the bleed air flowing through the bleed holes from the inlet opening to the outlet opening along the bleed path, the outlet opening of each of the bleed holes being located circumferentially upstream of the inlet opening relative to a direction of rotational flow in the main gas flow passage as driven by the direction of rotation of the rotor, such that the linear bleed hole axis is disposed at an acute angle relative to a radially extending reference axis disposed at the net opening of the bleed hole; andrecirculating and re-injecting the bleed air back into the annular gas flow passage of the compressor via bleed exit passage outlets located at an axial location disposed upstream of both the bleed holes and the leading edge of the blades of the rotor, wherein the recirculation of the bleed air is at least partially driven by a pressure differential between the bleed holes and the bleed cavity exit passage outlets.