Gas turbine engine compressor impeller cooling air sinks

Patent metadata

Field	Value
Publication number	US-10830144-B2
Application number	US-201615259883-A
Country	US
Kind code	B2
Filing date	Sep 8, 2016
Priority date	Sep 8, 2016
Publication date	Nov 10, 2020
Grant date	Nov 10, 2020

Abstract

A gas turbine engine includes devices, systems, and methods for providing bleed air from the compressor impeller to the turbine for cooling and/or other use. The bleed air may include compressor cooling air that is routed through the diffuser and external to an outer bypass duct and/or internally to a forward wheel cavity of the turbine.

First claim

What is claimed is:

1. A gas turbine engine comprising
an engine core defining a rotating axis, the engine core includes a compressor having an impeller arranged to rotate about the axis to compress air with an impeller tip and a diffuser for collecting compressed air from the impeller tip, a combustor fluidly connected to receive compressed air from the diffuser for combustion, and a turbine fluidly connected to receive exhaust products from the combustor; the impeller, the diffuser, the combustor, and the turbine collectively defining a core flow path;
an outer shell disposed about the engine core to house the engine core therein; and
a bleed circuit fluidly connected between the compressor and the turbine for communicating a stream of impeller air from the impeller to the turbine, the bleed circuit including a bleed inlet and a transport section, the bleed inlet arranged at the impeller tip and configured to bleed the stream of impeller air out from the core flow path and to communicate the stream of impeller air to the transport section, and the transport section in fluid communication with a deposit junction of the turbine,
wherein the bleed inlet is formed to include a radial section extending radially inwardly at least partially within a clearance between the impeller and a stationary wall located aft of the impeller and configured to receive the stream of impeller air from the impeller tip and an axial section extending axially aft through the stationary wall for communication with a transport section, wherein each of the radial and axial sections are arranged near the impeller tip,
wherein the bleed circuit further includes an inlet passage that receives the stream of impeller air from the axial section, at least a portion of the inlet passage is located aft of the impeller, the inlet passage conducts the stream of impeller air axially forward and radially outward where the inlet passage penetrates the outer shell and connects with the transport section that extends axially relative to the axis outside of the outer shell to a location near the turbine.

2. The gas turbine engine of claim 1 , wherein the deposit junction includes a vane cooling path of a second stage vane of the turbine and the stream of impeller air passes through the vane cooling path to cool the second stage vane.

3. The gas turbine engine of claim 1 , wherein the section connects with a plenum of the second stage vane of the turbine.

4. The gas turbine engine of claim 3 , wherein the at least a portion of the inlet passage is located in the combustor.

5. The gas turbine engine of claim 4 , wherein the inlet passage penetrates the outer shell axially forward of diffuser vanes included in the diffuser.

6. The gas turbine engine of claim 1 , wherein the inlet passage is defined through at least one blade of the diffuser and in communication with the bleed inlet.

7. A gas turbine engine comprising
an engine core defining a rotating axis, the engine core including a compressor configured to provide compressed air, a combustor fluidly connected to receive the compressed air from the compressor for combustion, and a turbine fluidly connected to receive exhaust products from the combustor; the compressor, the combustor, and the turbine collectively defining a core flow path, and the compressor including an impeller having an impeller tip and a diffuser for collecting the compressed air from the impeller tip of the impeller; and
a bleed circuit fluidly connected between the compressor and the turbine for communicating a stream of air from the impeller of the compressor,
wherein the combustor includes an inner combustor casing and a transport wall, the inner combustor casing defining a portion of a combustion housing, and the transport wall positioned radially outward of the inner combustor casing to locate the inner combustor casing between the transport wall and the impeller,
the bleed circuit including a bleed inlet, a transport section, and a deposit junction, the bleed inlet arranged at the impeller tip of the impeller upstream of the diffuser and configured to bleed the stream of air out from the core flow path and to communicate the stream of air to the deposit junction, the transport section connects with the bleed inlet and the deposit junction to communicate the stream of air between the inner combustor casing and the transport wall from the bleed inlet to the deposit junction so that the stream of air is conducted entirely radially outward of the inner combustor casing between the bleed inlet and the deposit junction, and the deposit junction communicates the stream of air back into the core flow path in the turbine,
wherein the transport wall includes a radially extending section near the bleed inlet and an axially extending section that extends aft from the radially extending section and forms a portion of the deposit junction.

8. The gas turbine engine of claim 7 , wherein the deposit junction includes a forward wheel cavity of the turbine and the stream of air enters the forward wheel cavity for purging.

9. The gas turbine engine of claim 8 , wherein the stream of air pressurizes the forward wheel cavity and leaks into the core flow path at a location between a first stage vane and a first stage blade of the turbine.

10. The gas turbine engine of claim 9 , wherein the compressor includes the diffuser having a number of blades arranged to collect compressed air from the compressor and the bleed circuit comprises at least one inlet passage defined through at least one of the number of blades of the diffuser and in communication with the bleed inlet to receive the stream of air.

11. The gas turbine engine of claim 7 , wherein the bleed circuit comprises at least one inlet passage defined through at least one blade of the diffuser and in communication with the bleed inlet to receive the stream of air.

12. A method of operating a gas turbine engine, the method comprising
flowing an engine core flow through each of a compressor, a combustor, and a turbine fluidly, the compressor including an impeller and a diffuser located directly downstream of the impeller,
bleeding a stream of air in a bleed circuit from a tip of the impeller of the compressor out from the core flow path between the tip of the impeller and the diffuser through a bleed inlet of the bleed circuit, the bleed inlet formed to include a radial section extending radially inward at least partially within a clearance between the impeller and a stationary wall and an axial section extending axially aft through the stationary wall for communication with a transport section, each of the radial and axial sections arranged near the tip of the impeller,
conducting the stream of air axially aft in the transport section such that the stream of air is located entirely radially outward of an inner combustor casing of the combustor while being conducted through the transport section, and
depositing the stream of air to a deposit junction of the turbine such that the stream of air in the deposit junction is located radially outside of the inner combustor casing.

13. The method of claim 12 , wherein the deposit junction includes a vane cooling path of a second stage vane of the turbine and the stream of air passes through the vane cooling path to cool the second stage vane.

14. The gas turbine engine of claim 12 , wherein bleeding includes flowing the stream of air through at least one inlet passage defined through at least one of a number of blades of the diffuser of the compressor.

15. The gas turbine engine of claim 14 , wherein bleeding includes flowing the stream of air through at least o

Assignees

• Rolls Royce Nam Tech Inc

Inventors

• Lambert Tony A
• Kapadia Behram V
• Hall Ron

Classifications

• F02C7/18 Primary

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

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F04D29/444

  Bladed diffusers · CPC title

• F02C6/08

  the gas being bled from the gas-turbine compressor · CPC title

• F05D2240/12

  Fluid guiding means, e.g. vanes · CPC title