Turbine cooling system

Patent metadata

Field	Value
Publication number	US-10633996-B2
Application number	US-201615354430-A
Country	US
Kind code	B2
Filing date	Nov 17, 2016
Priority date	Nov 17, 2016
Publication date	Apr 28, 2020
Grant date	Apr 28, 2020

Abstract

A turbine assembly for a gas turbine engine is disclosed herein. The turbine assembly includes a turbine vane and an annular turbine shroud. The turbine vane has an airfoil extending through a primary gas path and is formed to include cooling passages sized to conduct cooling air from an inlet aperture to a discharge aperture. The turbine shroud is adapted to extend around a bladed turbine wheel to resist gasses from passing over turbine blades without interacting with the turbine blades. The turbine shroud includes a ceramic matrix composite blade track having a radially-inwardly facing inner surface arranged in confronting relation with the primary gas path.

First claim

What is claimed is:

1. A turbine section assembly, the assembly comprising
a turbine vane including an airfoil extending through a primary gas path, the turbine vane formed to include cooling passages sized to conduct cooling air from an inlet aperture to a discharge aperture, the inlet aperture radially inboard of the primary gas path, the discharge aperture radially outboard of the primary gas path, and the cooling passages extending from the inlet aperture to the discharge aperture,
an annular turbine shroud adapted to extend around a bladed turbine wheel to resist gasses from passing over turbine blades without interacting with the turbine blades, the turbine shroud including a ceramic matrix composite blade track having a radially-inwardly facing inner surface arranged in confronting relation with the primary gas path, and
a passageway extending from the discharge aperture of the cooling passages to the ceramic matrix composite blade track to conduct cooling air from the turbine vane to the ceramic matrix composite blade track radially outward of the primary gas path after the cooling air has been passed through and warmed in the turbine vane so that the ceramic matrix composite blade track is actively cooled at a location radially outward of the inner surface by warmed cooling air to manage the temperature and the thermal gradient across the ceramic matrix composite blade track during operation of the turbine section assembly;
wherein the turbine vane includes an inner wall and an outer wall defining an exterior of the turbine vane, the inner wall and the outer wall cooperate to define a gap therebetween configured to receive cooling air provided thereto, wherein the gap acts as one of the cooling passages, thereby conducting cooling air from the inlet aperture to the discharge aperture.

2. The turbine section assembly of claim 1 , wherein the ceramic matrix composite blade track includes a runner providing the inner surface and an attachment feature extending radially outward from a radially-outwardly facing outer surface of the runner and the warmed cooling air is applied to the outer surface of the runner during operation of the turbine section assembly.

3. The turbine section assembly of claim 2 , wherein the runner is formed to include internal air passages sized to conduct the warmed cooling air at least partway through the runner during operation of the turbine section assembly.

4. The turbine section assembly of claim 1 , wherein the turbine vane is formed to include central cavities located internal to the gap and the discharge aperture is fluidly coupled to the central cavities.

5. The turbine section assembly of claim 1 , wherein the turbine vane includes internal cavities located internal to an exterior of the turbine vane, the internal cavities are fluidly coupled to the discharge aperture, and at least one of the cooling passages extends through the internal cavities to conduct cooling air through the turbine vane to the discharge aperture along a serpentine path during operation of the turbine section assembly.

6. The turbine section assembly of claim 1 , wherein the passageway is arranged radially outward of the primary gas path.

7. The turbine section assembly of claim 1 , wherein a source of cooling air is configured to provide cooling air to the inlet aperture of the turbine vane such that warmed cooling air at a first pressure is provided to the ceramic matrix composite blade track by the turbine vane during operation of the turbine section assembly, gasses at a second pressure are passed along the primary gas path during operation of the turbine section assembly, and the first pressure is greater than the second pressure.

8. A turbine section assembly, the assembly comprising
a turbine vane formed to include cooling passages that extend from an inlet aperture to a discharge aperture, the inlet aperture radially inboard of a primary gas path and the discharge aperture radially outboard of the primary gas path,
a turbine shroud including a ceramic matrix composite blade track, and
a passageway extending from the discharge aperture of the cooling passages to the ceramic matrix composite blade track to provide cooling air passed through the turbine vane to the ceramic matrix composite blade track to actively cool the ceramic matrix composite blade track during operation of the turbine section assembly;
wherein the turbine vane includes an inner wall and an outer wall defining an exterior of the turbine vane, the inner wall and the outer wall cooperate to define a gap therebetween configured to receive cooling air provided thereto, wherein the gap acts as one of the cooling passages, thereby conducting cooling air from the inlet aperture to the discharge aperture.

9. The turbine section assembly of claim 8 , wherein the ceramic matrix composite blade track includes a runner having a radially-inwardly facing inner surface and a radially-outwardly facing outer surface and an attachment feature extending radially outward from the runner and the cooling air provided to the ceramic matrix composite blade track by the passageway is applied to the outer surface during operation of the turbine section assembly.

10. The turbine section assembly of claim 8 , wherein the turbine vane is formed to include central cavities located internal to the gap and the discharge aperture is fluidly coupled to the central cavities.

11. The turbine section assembly of claim 8 , wherein the turbine vane includes internal cavities located internal to an exterior of the turbine vane, the internal cavities are fluidly coupled to the discharge aperture, and at least one of the cooling passages extends through the internal cavities to conduct cooling air through the turbine vane to the discharge aperture along a serpentine path during operation of the turbine section assembly.

12. The turbine section assembly of claim 11 , wherein a source of cooling air is configured to provide cooling air to the inlet aperture of the turbine vane such that cooling air at a first pressure is provided to the ceramic matrix composite blade track by the turbine vane during operation of the turbine section assembly, gasses at a second pressure are passed along a primary gas path located radially inward of the passageway during operation of the turbine section assembly, and the first pressure is greater than the second pressure.

13. A gas turbine engine comprising
a compressor section configured to provide cooling air,
a combustor section configured to discharge pressurized gasses along a primary gas path, and
a turbine section configured to receive the cooling air provided by the compressor section and the pressurized gasses discharged by the combustor section along the primary gas path, the turbine section including an assembly having
a turbine vane formed to include cooling passages sized to conduct the cooling air from an inlet aperture to a discharge aperture, the inlet aperture radially inboard of the primary gas path and the discharge aperture radially outboard of the primary gas path,
a turbine shroud having a ceramic matrix composite blade track, and
a passageway extending from the discharge aperture of the cooling passages to the ceramic matrix composite blade track to provide cooling air passed through the turbine vane to the ceramic matrix composite blade track to actively cool the ceramic matrix composite blade track during operation of the gas turbine engine;

wherein the turbine vane includes an inner wall and an outer wall defining an exterior of the turbine vane, the inner wall and the outer wall cooperate to define a gap therebetween configured to receive cooling air prov

Assignees

• Rolls Royce Corp

Inventors

• Barker Brett J
• Varney Bruce E

Classifications

• F05D2300/6033

  Ceramic matrix composites [CMC] · CPC title

• F01D5/187

  Convection cooling · CPC title

• F01D11/08 Primary

  for sealing space between rotor blade tips and stator (specially-shaped blade tips therefor F01D5/20) · CPC title

• F01D5/188

  with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall · CPC title

• F01D11/24

  by selectively cooling-heating stator or rotor components · CPC title