Anti-vortex device for a gas turbine engine compressor

• US 8,453,463 B2
• Filed: 05/27/2009
• Issued: 06/04/2013
• Est. Priority Date: 05/27/2009
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Ciampa, Alessandro, Grewal, Daljit Singh, Caron, Jean-Francois
• Primary Examiner(s): Gartenberg, Ehud
• Assistant Examiner(s): SUTHERLAND, STEVEN M
• Application Number:
  US12/472,720
• Publication Number:
  US 20100300113A1
• Time in Patent Office: 1,469 Days
• Field of Search: 607/82, 607/85, 60/390.8, 608/06, 607/26, 602/69, 415/115, 415/175, 415/144, 416/198.R, 416/203, 416/175
• US Class Current: 60/782
• CPC Class Codes: F01D 5/08 Heating, heat-insulating or cooling means specially adapted for radial flow machines or engines F01D5/04
• CPC Class Codes: F01D 5/081 Cooling fluid being directed on the side of the rotor disc or at the roots of the blades F01D5/087 takes precedence
• CPC Class Codes: F01D 5/082 on the side of the rotor disc
• CPC Class Codes: F01D 5/084 the fluid circulating at the periphery of a multistage rotor, e.g. of drum type
• CPC Class Codes: F01D 5/085 cooling fluid circulating inside the rotor
• CPC Class Codes: F01D 5/087 in the radial passages of the rotor disc
• CPC Class Codes: F04D 29/284 for compressors
• CPC Class Codes: F04D 29/321 for axial flow compressors

Abstract

An anti-vortex device for use in a compressor rotor assembly of a gas turbine engine is described. Spaced-apart radial passageways extend from an axially extending passage provided in a central area of the device to an outer peripheral rim surface thereof. The radial passageways channel air from the primary gaspath about the rotor assembly to the axially extending passage where the air is directed into a central axial passage of the rotor assembly.

Claims

1. A compressor rotor assembly mounted for rotation about a central axis of a gas turbine engine, comprising an anti-vortex device having a body mounted between adjacent rotor discs, the rotor discs having a peripheral rim surface defining an inner boundary of a primary gas path of the engine, the anti-vortex device defining circumferentially spaced-apart radial passageways extending from respective axially extending central passages to an outer peripheral rim surface of the device, the axially extending central passages being defined in a central solid area of the body of the device, the axially extending central passages being independent from each other and communicating with an associated one of said radial passageways, each said radial passageway receiving bleed air from the primary gas path and directing it to an associated one of said axially extending passages, wherein the axially extending central passages extend axially forwardly and rearwardly from the radial passageways.

2. The compressor rotor assembly as claimed in claim 1, wherein the anti-vortex device comprises a solid body, and wherein said radial passageways extend through said solid body in an X-shaped configuration.

3. The compressor rotor assembly as claimed in claim 2, wherein said axially extending central passages being constituted by through bores disposed about a center point of said solid body and spaced to communicate at opposed ends thereof with a central axial passage of the compressor rotor assembly.

4. The compressor rotor assembly as claimed in claim 2, wherein said radial passageways are cone-shaped passageways tapering inwardly from an inlet end thereof at said outer periphery to an outlet end.

5. The compressor rotor assembly as claimed in claim 1, wherein said anti-vortex device has a drum body formed from a solid mass with said radial passageways and axially extending passages being machined from said solid mass, and cavities formed in said drum body between said circumferentially spaced-apart radial passageways.

6. The compressor rotor assembly as claimed in claim 1, wherein the anti-vortex device is spaced radially inwardly of an air bleed gap between said adjacent rotor discs.

7. The compressor rotor assembly as claimed in claim 2, wherein said solid body defining an X-shaped structural web between said axially extending central passages.

8. The compressor rotor assembly as claimed in claim 5, wherein said drum body is further provided with a transverse rod receiving through hole disposed between the radial passageways.

9. The compressor rotor assembly as claimed in claim 6, wherein the rotor discs respectively form part of an impeller and an adjacent compressor rotor, and wherein the anti-vortex device is clamped between opposed faces of the compressor rotor and the impeller.

10. A gas turbine engine comprising a compressor having at least two rotors mounted for joint rotation about a central axis, a combustor and a turbine section;

the compressor has an anti-vortex device secured between said at least two rotors, the anti-vortex device having a solid body portion, circumferentially spaced-apart radial passageways defined in said solid body portion, each said radial passageway extending from an axial passage extending through the solid body portion in a central area thereof to an outer peripheral rim surface of the solid body portion, the solid body portion defining a solid structural web centrally between the axial passages, the outer peripheral rim surface being spaced inwardly of an air bleed gap formed between said at least two rotors and in communication with a gas path of the engine, the anti-vortex device being configured for channeling air from the gas path in non-interference therewith through said air bleed gap and into said radial passageways and said axial passages, said axial passages extending axially forwardly and rearwardly relative to said associated radial passageways for redirecting said air under pressure in two opposite axial directions.

11. The gas turbine engine as claimed in claim 10, wherein said axial passage is in communication with a central axial passage of the gas turbine engine extending into said turbine section, the air under pressure drawn into the axial passage being directed into the central axial passage to provide cooling air for the turbine section.

12. The gas turbine engine as claimed in claim 11, wherein said radial passageways are disposed along two diametrical axes intersecting each other in an X-shaped configuration.

13. The gas turbine engine as claimed in claim 12, wherein said axial passage comprises individual through bores disposed about a center point of said solid body portion, each of said through bores communicating with said central axial passage of the gas turbine engine and with an associated one of said radial passageways.

14. The gas turbine engine as claimed in claim 13, wherein said radial passageways are cone-shaped passageways tapering inwardly from an inlet end thereof at said outer periphery to said outlet end communicating with said axial passage.

15. A method of reducing total pressure drop and the formation of free vortex in a flow of compressed air bled inwardly from a compressor of a gas turbine engine, the method comprising:

i) providing circumferentially spaced-apart radial passageways in an anti-vortex drum mounted to the compressor, the radial passageways extending from a plurality of axial passages extending through the anti-vortex drum in a central solid area of the anti-vortex drum to an outer peripheral rim of the anti-vortex drum;

ii) bleeding compressed air from a gas path of the compressor through said radial passageways and re-directing said compressed air coming from the radial passageways in two axially opposite directions in said axial passages when said compressor hub is rotating; and

iii) directing at least some of the compressed air bled from said axial passages of the anti-vortex drum, via a central axially-extending passage of the compressor, to a turbine section of said gas turbine engine to cool turbine components in said turbine section.

16. The method as claimed in claim 15, wherein said split compressed air flows in opposite directions, said split compressed air flowing in a direction opposite to said air directed to said turbine section being directed to pressurize a buffer seal.

17. The method as claimed in claim 15, wherein step (i) comprises clamping an anti-vortex drum in concentric alignment between two adjacent rotors and spaced inwardly of a peripheral gap formed at an outer periphery of the rotors.