GAS TURBINE ENGINE

• US 20120060507A1
• Filed: 09/02/2011
• Published: 03/15/2012
• Est. Priority Date: 09/10/2010
• Status: Active Grant

• Current Assignee: Rolls-Royce PLC (Rolls-Royce Holdings Plc)
• Original Assignee: Rolls-Royce PLC (Rolls-Royce Holdings Plc)
• Inventors: BOLGAR, Crispin D., KING, Jonathan M., SNOWSILL, Guy D., SHEATH, Michael J., DAILEY, Geoffrey M.
• Granted Patent: US 8,973,371 B2
• Time in Patent Office: Days
• US Class Current: 60/782
• CPC Class Codes: F01D 11/001 for sealing space between stator blade and rotor
• CPC Class Codes: F01D 11/04 using sealing fluid, e.g. steam
• CPC Class Codes: F01D 11/06 Control thereof
• 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: F02C 3/13 having variable working fluid interconnections between turbines or compressors or stages of different rotors controlling flow ratio between different flows of multi-flow jet-propulsion plant, e.g. ducted fan F02K3/075
• CPC Class Codes: F02C 6/08 the gas being bled from the gas-turbine compressor
• CPC Class Codes: F02C 7/18 the medium being gaseous, e.g. air F02C7/125 takes precedence

Abstract

A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section. The second flow of compressed air bypasses the combustor and arrives at the turbine section to fowl a cooling flow. A first portion of the cooling flow is routed to a front face of the rotor disc and a second portion of the cooling flow is routed to a rear face of the rotor disc.

Claims

1. -13. (canceled)

14. A gas turbine engine comprising in flow series a compressor section, a combustor and a turbine section, the engine further comprising a turbine section cooling system including:

a first compressed air bleed arrangement which bleeds a first flow of compressed air from a high pressure stage of the compressor section, the first flow of compressed air bypassing the combustor and arriving at the turbine section to form a sealing flow between a row of stator vanes and an adjacent downstream rotor disc and/or a cooling flow for either or both of the stator vanes and the rotor disc; and

a second compressed air bleed arrangement which bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section, the second flow of compressed air bypassing the combustor and arriving at the turbine section to form a cooling flow, a first portion of the cooling flow being routed to a front face of the rotor disc and a second portion of the cooling flow being routed to a rear face of the rotor disc.

15. A gas turbine engine according to claim 14, wherein the turbine section cooling system further includes re-pressurising means which re-pressurises the first and/or second portion of the cooling flow after the second flow of compressed air is bled from the one or more lower pressure stages of the compressor section.

16. A gas turbine engine according to claim 15, wherein a row of turbine blades is mounted to a rim of the rotor disc, at least a part of the re-pressurised first and/or second portion of the cooling flow being subsequently routed into the turbine blades.

17. A gas turbine engine according to claim 15, wherein the re-pressurising means comprises a first centrifugal pump adjacent the front face of the rotor disc for pressurising the first portion of the cooling flow and/or a second centrifugal pump adjacent the rear face of the rotor disc for pressurising the second portion of the cooling flow.

18. A gas turbine engine according to claim 16, wherein a drive arm connects the rotor disc to the high pressure stage of the compressor section, the first centrifugal pump penetrating the drive arm to extend over the front face of the rotor disc both outboard and inboard of the drive arm.

19. A gas turbine engine according to claim 17, wherein the, or each, centrifugal pump comprises:

a cover plate axially-spaced from the respective face of the rotor disc such that a channel producing radially outward delivery of the respective portion of cooling flow is formed between the cover plate and the respective face, and

a row of circumferentially spaced vanes positioned in the channel and co-rotational with the rotor disc to pressurise the respective portion of cooling flow delivered through the channel.

20. A gas turbine engine according to claim 18, wherein the, or each, centrifugal pump comprises:

a cover plate axially-spaced from the respective face of the rotor disc such that a channel producing radially outward delivery of the respective portion of cooling flow is formed between the cover plate and the respective face, and

a row of circumferentially spaced vanes positioned in the channel and co-rotational with the rotor disc to pressurise the respective portion of cooling flow delivered through the channel.

21. A gas turbine engine according to claim 19, wherein a backing plate is mounted to the respective face, the channel being defined between the cover plate and the backing plate.

22. A gas turbine engine according to claim 20, wherein a backing plate is mounted to the respective face, the channel being defined between the cover plate and the backing plate.

23. A gas turbine engine according to claim 14, wherein the second portion of the cooling flow is routed via a central bore of the rotor disc.

24. A gas turbine engine according to claim 14, wherein the second compressed air bleed arrangement comprises passages for the second flow of compressed air extending radially inwards from the one or more lower pressure stages of the compressor section, the passages co-rotating with the one or more lower pressure stages to reduce swirl pressure losses in the second flow of compressed air.

25. A gas turbine engine according to claim 14, wherein the second compressed air bleed arrangement comprises a vortex reducer to reduce the tangential velocity of the second flow of compressed air as it is bled from the one or more lower pressure stages of the compressor section and thereby reduce swirl pressure losses in the second flow of compressed air.

27. A method according to claim 25, further comprising the step of:

re-pressurising the first and/or second portion of the cooling flow after the second flow of compressed air is bled from the one or more lower pressure stages of the compressor section.

26. A method of cooling a turbine section of a gas turbine engine which comprises in flow series a compressor section, a combustor, and the turbine section, the method comprising the steps of:

bleeding a first flow of compressed air from a high pressure stage of the compressor section, the first flow of compressed air bypassing the combustor and arriving at the turbine section to form a sealing flow between a row of stator vanes and an adjacent downstream rotor disc and/or a cooling flow for either or both of the stator vanes and the rotor disc;

bleeding a second flow of compressed air from one or more lower pressure stages of the compressor section, the second flow of compressed air bypassing the combustor and arriving at the turbine section to form a cooling flow;

routing a first portion of the cooling flow to a front face of the rotor disc; and

routing a second portion of the cooling flow to a rear face of the rotor disc.

28. A method according to claim 26, wherein a row of turbine blades is mounted to a rim of the rotor disc, and the method further comprises the step of subsequently routing at least a part of the pressurised first and/or second portion of the cooling flow into the turbine blades.