Turbine vane assembly with cooling feature

Patent metadata

Field	Value
Publication number	US-10767495-B2
Application number	US-201916265575-A
Country	US
Kind code	B2
Filing date	Feb 1, 2019
Priority date	Feb 1, 2019
Publication date	Sep 8, 2020
Grant date	Sep 8, 2020

Abstract

A turbine vane assembly adapted for use in a gas turbine engine includes a support strut and a turbine vane arranged around the support strut. The support strut is made of metallic materials. The turbine vane is made of ceramic matrix composite materials to insulate the metallic materials of the support strut.

First claim

What is claimed is:

1. A turbine vane assembly for a gas turbine engine, the turbine vane assembly comprising
a ceramic matrix composite vane adapted to conduct hot gases flowing through a primary gas path of the gas turbine engine around the turbine vane assembly during use of the turbine vane assembly, the ceramic matrix composite vane includes an outer wall that defines an outer boundary of the primary gas path, an inner wall spaced apart radially from the outer wall relative to an axis to define an inner boundary of the primary gas path, and an aerofoil that extends between radially and interconnects the outer wall and the inner wall, and the aerofoil is formed to define an interior cavity that extends radially through the aerofoil,
a metallic support strut located in the interior cavity formed in the aerofoil and configured to receive force loads applied to the ceramic matrix composite vane by the hot gases during use of the turbine vane assembly, the metallic support strut having an outermost surface that faces the aerofoil, and the outermost surface of the metallic support strut being spaced apart from the aerofoil at all locations radially between the outer boundary and the inner boundary of the primary gas path to define a cooling channel between the metallic support strut and the aerofoil, and
a thermal barrier shield coupled to the outermost surface of the metallic support strut and spaced apart from the aerofoil at all locations radially between the outer boundary and the inner boundary of the primary gas path to reduce an amount of heat transfer to the metallic support strut from radiant, conductive, and convective heating caused by a temperature difference between the ceramic matrix composite vane and the metallic support strut during use of the turbine vane assembly, wherein the thermal barrier shield is a thermal barrier coating, and wherein the turbine vane assembly further includes a seal that engages the aerofoil and the metallic support strut to block fluid from flowing in the cooling channel.

2. The turbine vane assembly of claim 1 , wherein the aerofoil includes an outer surface that defines a leading edge, a trailing edge, a pressure side, and a suction side of the aerofoil.

3. The turbine vane assembly of claim 1 , wherein the thermal barrier coating comprises gadolinium oxide.

4. The turbine vane assembly of claim 1 , wherein the thermal barrier coating is a low emissivity coating that has a reflectivity between 0.5 and 1.

5. The turbine vane assembly of claim 1 , wherein the thermal barrier coating is a low emissivity coating that has a reflectivity between 0.6 and 0.95.

6. The turbine vane assembly of claim 1 , wherein the outer wall, the inner wall, and the aerofoil of the ceramic matrix composite vane are integrally formed from ceramic matrix composite materials such that the outer wall, the inner wall, and the aerofoil cooperate to form an integral, one-piece component.

7. A turbine vane assembly for a gas turbine engine, the turbine vane assembly comprising
a vane that includes an outer wall having a radial inner surface, an inner wall having a radial outer surface, and an aerofoil that extends between radially and interconnects the outer wall and the inner wall, and the aerofoil is formed to define an interior cavity that extends radially through the aerofoil,
a support strut located in the interior cavity formed in the aerofoil, the support strut spaced apart from the aerofoil at all locations radially between the radial inner surface and the radial outer surface to define a cooling channel between the support strut and the aerofoil, and
a thermal barrier shield located in the cooling channel and spaced apart from the aerofoil at all locations radially between the radial inner surface and the radial outer surface,
wherein the thermal barrier shield extends radially entirely between the radial inner surface and the radial outer surface, and
wherein the support strut includes a spar that extends radially into the interior cavity and a load transfer feature that extends circumferentially away from the spar and engages the aerofoil at a location radially outward of the radial inner surface.

8. The turbine vane assembly of claim 7 , wherein the thermal barrier shield extends radially outward beyond the radial inner surface and radially inward beyond the radial outer surface.

9. The turbine vane assembly of claim 7 , wherein the thermal barrier shield is coupled to the spar of the support strut.

10. The turbine vane assembly of claim 7 , wherein the support strut is completely solid and has a continuous outermost surface that it is formed without holes.

11. The turbine vane assembly of claim 7 , wherein the thermal barrier shield is continuous and formed without holes that extend either axially or circumferentially through the thermal barrier shield.

12. A turbine vane assembly for a gas turbine engine, the turbine vane assembly comprising
a vane that includes an outer wall having a radial inner surface, an inner wall having a radial outer surface, and an aerofoil that extends between radially and interconnects the outer wall and the inner wall, and the aerofoil is formed to define an interior cavity that extends radially into the aerofoil,
a support strut located in the interior cavity formed in the aerofoil, the support strut spaced apart from the aerofoil at all locations radially between the radial inner surface and the radial outer surface to define a cooling channel between the support strut and the aerofoil,
a thermal barrier shield located in the cooling channel and spaced apart from the aerofoil at all locations radially between the radial inner surface and the radial outer surface, wherein the thermal barrier shield extends radially entirely between the radial inner surface and the radial outer surface, and
a seal that engages the aerofoil and the support strut to block fluid from flowing into the cooling channel.

13. A method of making a turbine vane assembly, the method comprising
providing a metallic support strut, a ceramic matrix composite aerofoil formed to define an interior cavity therein, an outer wall, and an inner wall,
coating an outermost surface of the metallic support strut with a thermal barrier coating to define an insulated region of the metallic support strut,
locating the ceramic matrix composite aerofoil radially between the outer wall and the inner wall relative to an axis, and
arranging the metallic support strut in the interior cavity of the ceramic matrix composite aerofoil such that the insulated region of the metallic support strut extends radially at least between the outer wall and the inner wall and the thermal barrier coating is spaced apart from the ceramic matrix composite aerofoil at all locations in the insulated region, and
arranging a seal between the metallic support strut and the ceramic matrix composite aerofoil to block fluid from flowing into the cooling channel.

14. The method of claim 13 , further comprising doping the thermal barrier coating with gadolinium to form gadolinium oxide.

Assignees

• Rolls Royce Plc

Inventors

• Whittle Michael J
• Morrell Paul
• Razzell Anthony
• Adams Matthew

Classifications

• Y02T50/60

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

• F05D2260/231

  Preventing heat transfer · CPC title

• F05D2230/60

  Assembly methods · CPC title

• F01D5/284

  Selection of ceramic materials · CPC title

• F01D5/189 Primary

  the insert having a tubular cross-section, e.g. airfoil shape · CPC title