# US 10,767,495 B2 · Turbine vane assembly with cooling feature > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,767,495 B2 - Document ID: US10767495B2 - Status: Granted - Kind: B2 - Application number: 16/265,575 - Filed: 2019-02-01 - Published: 2020-08-06 - Granted: 2020-09-08 - Earliest claimed: 2019-02-01 - Assignees: ROLLS-ROYCE plc - Applicants: Rolls-Royce plc - Inventors: Michael J. Whittle; Paul Morrell; Anthony Razzell; Matthew Adams - Web page: https://patentgrep.com/patent/US10767495B2 ## CPC - F01D 5/189: {the insert having a tubular cross-section, e.g. airfoil shape} - F01D 5/284: {Selection of ceramic materials} - F01D 25/162: {Bearing supports} - F05D 2220/32: in gas turbines - F05D 2230/60: Assembly methods - F05D 2230/90: Coating; Surface treatment - F05D 2240/12: Fluid guiding means, e.g. vanes - F05D 2240/55: Seals - F05D 2260/231: Preventing heat transfer - F05D 2300/00: Materials; Properties thereof - F05D 2300/6033: Ceramic matrix composites [CMC] - F05D 2300/611: Coating - Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft - F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D) - F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS - Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION ## 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. ## Claims ### Claim 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. ### Claim 2 (depends on 1) 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. ### Claim 3 (depends on 1) The turbine vane assembly of claim 1, wherein the thermal barrier coating comprises gadolinium oxide. ### Claim 4 (depends on 1) 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. ### Claim 5 (depends on 1) 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. ### Claim 6 (depends on 1) 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. ### Claim 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. ### Claim 8 (depends on 7) 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. ### Claim 9 (depends on 7) The turbine vane assembly of claim 7, wherein the thermal barrier shield is coupled to the spar of the support strut. ### Claim 10 (depends on 7) 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. ### Claim 11 (depends on 7) 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. ### Claim 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. ### Claim 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. ### Claim 14 (depends on 13) The method of claim 13, further comprising doping the thermal barrier coating with gadolinium to form gadolinium oxide.