Gas turbine engine with high speed and temperature spool cooling system

Patent metadata

Field	Value
Publication number	US-9677475-B2
Application number	US-201514597442-A
Country	US
Kind code	B2
Filing date	Jan 15, 2015
Priority date	Jan 15, 2015
Publication date	Jun 13, 2017
Grant date	Jun 13, 2017

Abstract

A gas turbine engine includes a turbine section that includes a turbine rotor arranged in a plenum. A compressor section includes a compressor rotor assembly that has spaced apart inner and outer portions that provide an axially extending cooling channel. Compressor blades extend radially outward from the outer portion which provides an inner core flow path. A rotor spoke is configured to receive a first cooling flow and fluidly connect the outer portion to the cooling channel. The compressor rotor assembly has a coolant exit that is in fluid communication with the cooling channel. The compressor rotor assembly is configured to communicate the first cooling flow to the turbine rotor. A bleed source is configured to provide a second cooling flow. A combustor section includes an injector in fluid communication with the bleed source. The tangential onboard injector is configured to communicate the second cooling flow to the turbine rotor.

First claim

What is claimed is:

1. An engine comprising:
a turbine section includes a turbine rotor arranged in a plenum;
a compressor section includes a compressor rotor assembly having spaced apart inner and outer portions that provide a cooling channel that extends axially, compressor blades extending radially outward from the outer portion which provides an inner core flow path, and a rotor spoke configured to receive a first cooling flow and fluidly connecting the outer portion to the cooling channel, the compressor rotor assembly has a coolant exit in fluid communication with the cooling channel, the compressor rotor assembly configured to communicate the first cooling flow to the turbine rotor;
a bleed source configured to provide a second cooling flow;
a combustor section includes a tangential onboard injector in fluid communication with the bleed source, the tangential onboard injector configured to communicate the second cooling flow to the turbine rotor; and
wherein the compressor rotor assembly has a rotor disk that includes a rotor outer peripheral surface, and wherein the compressor blades are supported on a platform that has a blade inner surface that faces the rotor outer peripheral surface, and wherein the cooling channel is defined between the rotor outer peripheral surface and the blade inner surface.

2. The engine according to claim 1 , wherein the compressor blades are bonded to the inner portion.

3. The engine according to claim 1 , wherein an inlet opening is arranged upstream from and adjacent to the compressor blades, the inlet opening configured to supply the first cooling flow to the cooling channel.

4. The engine according to claim 1 , wherein the compressor section provides the bleed source.

5. The engine according to claim 1 , wherein the compressor rotor assembly includes a hub arranged adjacent to the combustor section, and wherein the coolant exit is provided in the hub.

6. The engine according to claim 1 , wherein the combustor section includes an exit guide vane arranged axially between the compressor rotor assembly and a combustor, the exit guide vane configured to communicate the bleed air radially inward from the bleed source toward the tangential onboard injector.

7. The engine according to claim 1 , wherein the compressor blades are integrally formed as one piece with the rotor disk.

8. The engine according to claim 1 , wherein the turbine rotor is arranged in a plenum configured to receive the first and second cooling flows.

9. The engine according to claim 8 , wherein the plenum is arranged radially inward of a core flow path, wherein the plenum is defined by a forward wall, a shaft, an aft wall, and an inner diameter wall of the core flow path, and wherein the first and second cooling flows are isolated from the core flow path.

10. The engine according to claim 9 , wherein the tangential onboard injector extends into the plenum to deliver at least the second cooling flow to the turbine rotor.

11. The engine according to claim 9 , wherein the shaft, to which the compressor rotor assembly is mounted, communicates the first flow to the plenum.

12. A method of cooling a gas turbine engine turbine section, the method comprising steps of:
providing a first cooling flow through an opening in an inner core flow path of a compressor rotor assembly of a compressor section, the compressor rotor assembly having spaced apart inner and outer portions that provide a cooling channel that extends axially, wherein the compressor rotor assembly has a rotor disk that includes a rotor outer peripheral surface, and wherein a blade is supported on a platform that has a blade inner surface that faces the rotor outer peripheral surface, the cooling channel is defined between the rotor outer peripheral surface and the blade inner surface;
providing a second cooling flow to an injector; and
passing the first and second cooling flows to a turbine rotor to cool the turbine rotor.

13. The method according to claim 12 , wherein the cooling channels are beneath the inner core flow path, and wherein a spoke extends from the inner core flow path into a cooling channel adjacent to the blade.

14. The method according to claim 13 , wherein the compressor rotor assembly includes a coolant exit, wherein passing the first and second cooling flows further includes communicating the first cooling flow from the coolant exit to the turbine rotor.

15. The method according to claim 12 , wherein providing the second cooling flow further includes includes supplying bleed air from a bleed source.

16. The method according to claim 15 , wherein the turbine rotor is arranged in a plenum, and passing the first and second cooling flows further includes communicating the first and second cooling flows to the plenum.

17. The method according to claim 16 , comprising a step of preventing air from an adjacent plenum fore of said plenum from entering said plenum by at least disposing a first seal at a joint between a fore wall of the plenum and a structure of a first rotor in said plenum.

18. The method according to claim 16 , comprising a step of preventing air from an adjacent plenum aft of said plenum from entering said plenum by at least disposing a second seal at a joint between an aft wall of the plenum and a structure of a second rotor in said plenum.

Assignees

• United Technologies Corp

Inventors

• Merry Brian D
• Suciu Gabriel L

Classifications

• F02C9/18 Primary

  by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} · CPC title

• F05D2250/322

  tangential · CPC title

• F02C7/18

  the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} · CPC title

• F01D5/084

  the fluid circulating at the periphery of a multistage rotor, e.g. of drum type · CPC title

• F02C3/10

  with another turbine driving an output shaft but not driving the compressor · CPC title