Cooling air for gas turbine engine with thermally isolated cooling air delivery

Patent metadata

Field	Value
Publication number	US-11428121-B2
Application number	US-201916375892-A
Country	US
Kind code	B2
Filing date	Apr 5, 2019
Priority date	Apr 6, 2018
Publication date	Aug 30, 2022
Grant date	Aug 30, 2022

Abstract

A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A flowpath is defined between a rotating surface and a non-rotating surface. The flowpath is connected downstream of the heat exchanger and is configured to deliver air to at least one of the plurality of rotating components. At least a portion of the non-rotating surface and the rotating surface includes a base metal. An insulation material is disposed on a surface along the flowpath.

First claim

The invention claimed is:

1. A gas turbine engine comprising:
a plurality of rotating components housed within a compressor section and a turbine section;
a first tap connected to said compressor section and configured to deliver air at a first pressure;
a heat exchanger connected downstream of said first tap;
a rotating surface radially outward of a non-rotating surface defining a flowpath therebetween and wherein the non-rotating surface faces the rotating surface, wherein the flowpath is connected downstream of said heat exchanger and is configured to deliver air to at least one of said plurality of rotating components, wherein at least a portion of said non-rotating surface and said rotating surface comprising a base metal;
an insulation material disposed on a surface along the flowpath;
wherein said insulation material being provided outwardly of said base metal on at least a portion of both said rotating surface and said non-rotating surface;
wherein there is a combustor radially outward of said non-rotating surface, and a chamber intermediate said combustor and said non-rotating surface connected to receive compressed air downstream of a downstream most location in said compressor section;
wherein said insulation material on said rotating surface is a coating;
wherein said rotating surface is an outer surface of a shaft connecting a high pressure turbine rotor in said turbine section to a high pressure compressor rotor in said compressor section; and
wherein said insulation material on said non-rotating surface includes a ceramic fiber blanket formed of bulk fibers, and wherein said base metal is radially inward of said ceramic fiber blanket and an outer wall of said non-rotating structure is attached on an opposed radial side of said ceramic fiber blanket relative to said base metal.

2. The gas turbine engine as set forth in claim 1 , wherein there is a downstream most location in a high pressure compressor within said compressor section and the first tap is at an upstream location relative to the downstream most location.

3. The gas turbine engine as set forth in claim 1 , wherein there is a high pressure compressor in the compressor section with a downstream most location and said first tap is at a location where air will have passed downstream of the downstream most location.

4. The gas turbine engine as set forth in claim 1 , wherein said at least one rotating component includes at least a downstream most portion of a high pressure compressor within the compressor section.

5. The gas turbine engine as set forth in claim 1 , wherein said at least one rotating component includes at least an upstream most blade and vane in a high pressure turbine which is part of said turbine section.

6. The gas turbine engine as set forth in claim 5 , wherein said coating including an outer ceramic topcoat facing the insulation material on said non-rotating surface.

7. The gas turbine engine as set forth in claim 6 , wherein there is a metallic bond coat intermediate said ceramic topcoat and the underlying base metal in said rotating surface.

8. The gas turbine engine as set forth in claim 7 , wherein there is a thermally-grown oxide coating intermediate said metallic bond coat and said ceramic topcoat.

9. The gas turbine engine as set forth in claim 5 , wherein fluid conduits are connected to a location downstream of said heat exchanger, to communicate air downstream of the heat exchanger into said flow path, and at least some of said fluid conduits being provided with insulation.

10. The gas turbine engine as set forth in claim 5 , wherein said ceramic fiber blanket is formed with a alumina-silica fibers.

11. The gas turbine engine as set forth in claim 10 , wherein fluid conduits are connected to a location downstream of said heat exchanger, to communicate air downstream of the heat exchanger into said flow path, and at least some of said fluid conduits being provided with insulation.

12. The gas turbine engine as set forth in claim 5 , wherein fluid conduits are connected to a location downstream of said heat exchanger, to communicate air downstream of the heat exchanger into said flow path, and at least some of said fluid conduits being provided with insulation.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Schwarz Frederick M
• Snape Nathan

Classifications

• Y02T50/60

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

• F02C6/08

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

• F01D5/081 Primary

  Cooling fluid being directed on the side of the rotor disc or at the roots of the blades (F01D5/087 takes precedence) · CPC title

• F02C7/24

  Heat or noise insulation (air intakes having provisions for noise suppression F02C7/045; turbine exhaust heads, chambers, or the like F01D25/30; silencing nozzles of jet-propulsion plants F02K1/00) · CPC title

• F05D2300/21

  Oxide ceramics · CPC title