Thermal shield for gas turbine engine diffuser case

Patent metadata

Field	Value
Publication number	US-10837364-B2
Application number	US-201715417257-A
Country	US
Kind code	B2
Filing date	Jan 27, 2017
Priority date	Jan 27, 2017
Publication date	Nov 17, 2020
Grant date	Nov 17, 2020

Abstract

A gas turbine engine comprises a fan delivering air into a bypass duct as bypass air and into a core engine. The core engine includes a high pressure compressor, a combustor, and a turbine section including a fan drive turbine driving a fan rotor through a gear reduction. The high pressure compressor having a downstream most location and air in a chamber radially outward of the combustor being air downstream of the downstream most location in the high pressure compressor. A tap taps compressed air and moves compressed air through a heat exchanger, then returns the compressed air back into a core engine housing and passes the returned air through a conduit radially outwardly of the combustor. The air is passed from the conduit radially inwardly to cool the turbine section.

First claim

The invention claimed is:

1. A gas turbine engine comprising:
a fan delivering air into a bypass duct as bypass air and into a core engine;
said core engine including a high pressure compressor, a combustor, and a turbine section including a fan drive turbine driving a fan rotor through a gear reduction;
said high pressure compressor having a downstream most location, and compressed air in a chamber radially outward of said combustor being air downstream of said downstream most location in said high pressure compressor, and said compressed air passing to said combustor;
a tap for tapping compressed cooling air and moving the compressed cooling air through a heat exchanger, then returning said compressed cooling air back into a core housing and passing the compressed cooling air through a conduit radially inward of said core housing, and radially outwardly of said combustor, and said cooling air passing from said conduit radially inwardly to cool said turbine section;
wherein said conduit is defined between inner and outer core housing walls and extends in an upstream direction to be aligned with a compressor diffuser positioned downstream of said downstream most location in said high pressure compressor;
said inner core housing wall is positioned outward of the chamber that supplies the compressed air to said combustor and defines said chamber with said combustor, such that said compressed air in said chamber contacts said inner core housing wall;
wherein said conduit extends generally circumferentially about a center axis of said gas turbine engine; and
wherein components extend through said inner and outer core housing to provide functions with regard to said combustor and a location through which said components extend is sealed to limit leakage of said compressed cooling air in said conduit radially outwardly.

2. The gas turbine engine as set forth in claim 1 , wherein said components include a fuel injector nozzle.

3. The gas turbine engine as set forth in claim 1 , wherein said inner core housing wall is provided with at least one coating.

4. The gas turbine engine as set forth in claim 3 , wherein said at least one coating includes a thermal barrier coating.

5. The gas turbine engine as set forth in claim 4 , wherein said at least one coating also includes a low emissivity coating.

6. The gas turbine engine as set forth in claim 3 , wherein said at least one coating includes a low emissivity coating.

7. A gas turbine engine comprising:
a fan delivering air into a bypass duct as bypass air and into a core engine;
said core engine including a high pressure compressor, a combustor, and a turbine section;
said high pressure compressor having a downstream most location and compressed air in a chamber radially outward of said combustor being air downstream of said downstream most location in said high pressure compressor, and said compressed air passing to said combustor;
a tap for tapping compressed cooling air and moving the compressed cooling air through a heat exchanger, then returning said compressed cooling air back into a core housing and passing the compressed cooling air through a conduit radially inward of said core housing, and radially outwardly of said combustor, and said cooling air passing from said conduit radially inwardly to cool said turbine section;
wherein said conduit is defined between inner and outer core housing walls and extends in an upstream direction to be aligned with a compressor diffuser positioned downstream of said downstream most location in said high pressure compressor;
said inner core housing wall is positioned outward of the chamber that supplies the compressed air to said combustor and defines said chamber with said combustor, such that air in said chamber contacts said inner core housing wall;
wherein said conduit extends generally circumferentially about a center axis of said gas turbine engine; and
wherein components extend through said inner and outer core housing walls to provide functions with regard to said combustor and a location through which said components extend is sealed to limit leakage of said compressed air in said conduit radially outwardly.

8. The gas turbine engine as set forth in claim 7 , wherein said tap is from a location downstream of said downstream most location in said high pressure compressor.

9. The gas turbine engine as set forth in claim 8 , wherein said heat exchanger is placed in the bypass duct with the bypass air cooling the compressed cooling air in said heat exchanger.

10. The gas turbine engine as set forth in claim 7 , wherein said heat exchanger is placed in the bypass duct with the bypass air cooling the compressed cooling air in said heat exchanger.

11. The gas turbine engine as set forth in claim 7 , wherein said components include a fuel injector nozzle.

12. The gas turbine engine as set forth in claim 7 , wherein said inner core housing wall is provided with at least one coating.

13. The gas turbine engine as set forth in claim 12 , wherein said at least one coating includes a thermal barrier coating.

14. The gas turbine engine as set forth in claim 13 , wherein said at least one coating also includes a low emissivity coating.

15. The gas turbine engine as set forth in claim 12 , wherein said at least one coating includes a low emissivity coating.

16. The gas turbine engine as set forth in claim 7 , wherein said components include a fuel injector nozzle.

17. A gas turbine engine comprising:
a core engine including a high pressure compressor, a combustor, and a turbine section;
said high pressure compressor having a downstream most location and compressed air in a chamber radially outward of said combustor being air downstream of said downstream most location in said high pressure compressor, and said compressed air passing to said combustor;
a tap for tapping compressed cooling air and moving the compressed cooling air through a heat exchanger, then returning said compressed cooling air back into a core housing and passing the compressed cooling air through a conduit radially inward of said core housing, and radially outwardly of said combustor, and said compressed cooling air passing from said conduit radially inwardly to cool said turbine section;
wherein said conduit is defined between inner and outer core housing walls and extends in an upstream direction to be aligned with a compressor diffuser positioned downstream of said downstream most location in said high pressure compressor; and
said inner core housing wall is positioned outward of the chamber that supplies the compressed air to said combustor and defines said chamber with said combustor, such that the compressed air in said chamber contacts said inner core housing wall;
wherein said conduit extends generally circumferentially about a center axis of said gas turbine engine; and
wherein components extend through said inner and outer core housing to provide functions with regard to said combustor and a location through which said components extend is sealed to limit leakage of said compressed cooling air in said conduit radially outwardly.

18. The gas turbine engine as set forth in claim 17 , wherein said components include a fuel injector nozzle.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Roberge Gary D

Classifications

• F01D5/288 Primary

  Protective coatings for blades · CPC title

• F04D29/5826

  Cooling at least part of the working fluid in a heat exchanger · CPC title

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title

• F05D2260/213

  by the provision of a heat exchanger within the cooling circuit · CPC title

• F05D2260/606

  Bypassing the fluid · CPC title