Compressor bleed self-recirculating system

A gas turbine engine comprising:

• a compressor section, a combustor and a turbine section, in serial flow communication;
• and
• the 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;
• and
• a 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;
• and
• an 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.

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;
• and
• 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;
• and
• recirculating 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.