# US 11,898,450 B2 · Flowpath assembly for gas turbine engine > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 11,898,450 B2 - Document ID: US11898450B2 - Status: Granted - Kind: B2 - Application number: 17/323,436 - Filed: 2021-05-18 - Published: 2022-11-24 - Granted: 2024-02-13 - Earliest claimed: 2021-05-18 - Assignees: Rtx Corporation - Applicants: Raytheon Technologies Corporation - Inventors: Raymond Surace; San Quach; Bryan P. Dube - Web page: https://patentgrep.com/patent/US11898450B2 ## CPC - F01D 11/12: using a rubstrip, e.g. erodible. deformable or resiliently-biased part - F01D 11/24: by selectively cooling-heating stator or rotor components - F01D 5/282: {Selecting composite materials, e.g. blades with reinforcing filaments} - F01D 11/20: Actively adjusting tip-clearance - F01D 25/005: {Selecting particular materials} - F01D 25/243: {Flange connections; Bolting arrangements (F01D25/265 takes precedence)} - F01D 5/147: {Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18)} - F01D 25/12: Cooling - F05D 2220/323: for aircraft propulsion, e.g. jet engines - F05D 2240/11: Shroud seal segments - F05D 2240/125: related to the tip of a stator vane - F05D 2300/20: Oxide or non-oxide ceramics - F05D 2300/6033: Ceramic matrix composites [CMC] - 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 An assembly for a gas turbine engine includes, among other things, an end wall including a main body extending between a first end portion and a second end portion to establish a seal face. The end wall includes a first attachment portion dimensioned to fixedly attach the main body to a static structure at a first attachment point. An airfoil extends radially inwardly from the end wall relative to an assembly axis. The airfoil includes an inner cavity extending between a first end portion and a second end portion, the first end portion adjacent the end wall of the airfoil. A spar member includes a spar body extending between a first end portion and a second end portion. The spar body extends at least partially through the inner cavity. The first end portion of the end wall is cantilevered from the first attachment point. ## Claims ### Claim 1 An assembly for a gas turbine engine comprising: an end wall including a main body extending between a first end portion and a second end portion to establish a seal face that bounds a gas path, the seal face dimensioned to establish a clearance gap with a rotatable blade along the gas path, and the end wall including a first attachment portion dimensioned to fixedly attach the main body to a static structure at a first attachment point; an airfoil extending radially inwardly from the end wall relative to an assembly axis, the airfoil including an inner cavity extending between a first end portion and a second end portion, the first end portion adjacent the end wall of the airfoil; and a spar member including a spar body extending between a first end portion and a second end portion, the spar body extending at least partially through the inner cavity, the first end portion of the spar member attachable to the static structure at a second attachment point axially between the first end portion of the end wall and the first attachment portion such that the first end portion of the end wall is cantilevered from the first attachment point. ### Claim 2 (depends on 1) The assembly as recited in claim 1, wherein the airfoil includes a fairing integrally formed with the end wall, and both the fairing and the end wall comprise a ceramic matrix composite (CMC) material. ### Claim 3 (depends on 1) The assembly as recited in claim 1, further comprising a spring member dimensioned to bias the first end portion of the end wall radially outwardly relative to the assembly axis. ### Claim 4 (depends on 1) The assembly as recited in claim 1, wherein the static structure includes a case dimensioned to extend circumferentially about the end wall relative to the assembly axis, the end wall is attachable to the case at the first attachment point, the first end portion of the spar member is attachable to the case at the second attachment point, and the second attachment point is axially spaced apart from the first attachment point relative to the assembly axis. ### Claim 5 (depends on 4) The assembly as recited in claim 4, wherein the end wall includes a flange that extends radially outwardly from the first end portion of the end wall relative to the assembly axis, and the flange is dimensioned to establish a spline interface with the first end portion of the spar member. ### Claim 6 (depends on 1) The assembly as recited in claim 1, wherein the airfoil includes a fairing, the fairing includes an airfoil section extending from a platform section, and the platform section radially opposes the end wall with respect to assembly axis. ### Claim 7 (depends on 6) The assembly as recited in claim 6, wherein the end wall includes a circumferential face dimensioned to at least partially surround an outer periphery of the airfoil section. ### Claim 8 (depends on 6) The assembly as recited in claim 6, wherein the fairing is integrally formed with the end wall. ### Claim 9 (depends on 6) The assembly as recited in claim 6, wherein the spar member includes an inner cavity dimensioned to convey cooling flow from a coolant source to a plenum radially inwardly of the platform section relative to the assembly axis. ### Claim 10 (depends on 6) The assembly as recited in claim 6, wherein the airfoil section is moveable relative to the spar member. ### Claim 11 (depends on 10) The assembly as recited in claim 10, further comprising a spring member dimensioned to bias the first end portion of the end wall radially outwardly relative to the assembly axis. ### Claim 12 (depends on 11) The assembly as recited in claim 11, further comprising: a spring plate; and a seal member captured between the spring plate and the end wall to establish a sealing relationship with an outer periphery of the airfoil section. ### Claim 13 (depends on 12) The assembly as recited in claim 12, further comprising: a spring member that biases the spring plate and the seal member against the main body of the end wall such that the seal member is trapped between the spring plate and the main body of the end wall; wherein the spring plate and the seal member circumscribe a perimeter of the outer periphery of the airfoil section and a perimeter of the spar body of the spar member. ### Claim 14 (depends on 6) The assembly as recited in claim 6, wherein: the end wall includes an opening defined through a thickness of the main body; and the spar body and the airfoil section extend through the opening such that the first end portion of the spar member and the first end portion of the airfoil are radially outward of the opening relative to the assembly axis. ### Claim 15 (depends on 1) The assembly as recited in claim 1, wherein the airfoil is a turbine vane, and the rotatable blade is a turbine blade. ### Claim 16 A gas turbine engine comprising: a section including an array of blades rotatable in a gas path about a longitudinal axis; an end wall including an array of arc segments arranged circumferentially about the array of blades relative to the longitudinal axis, each of the arc segments including a main body extending in an axial direction between a first end portion and a second end portion relative to the longitudinal axis to establish a seal face that bounds the gas path, the seal face dimensioned to establish a clearance gap with the blades, and each of the arc segments including a first attachment portion fixedly attached to a static structure at a first attachment point such that the first end portion of the arc segment is cantilevered from the first attachment point; an array of vanes adjacent to the array of blades, wherein the vanes extend inwardly from the seal face of a respective one of the arc segments and at least partially across the gas path; and an array of spar members fixedly attached to the static structure, wherein each of the spar members extends at least partially through an inner cavity of a respective one of the vanes. ### Claim 17 (depends on 16) The gas turbine engine as recited in claim 16, wherein each of the arc segments comprises a ceramic material. ### Claim 18 (depends on 16) The gas turbine engine as recited in claim 16, wherein the arc segments are moveable in a radial direction relative to the spar members with respect to the longitudinal axis. ### Claim 19 (depends on 18) The gas turbine engine as recited in claim 18, wherein each of the spar members includes an inner cavity dimensioned to convey cooling flow from a coolant source to a plenum radially inward of the respective vane relative to the longitudinal axis. ### Claim 20 (depends on 16) The gas turbine engine as recited in claim 16, wherein each of the vanes includes an airfoil section and a platform section that bounds the gas path, the airfoil section extending radially between the platform section and the respective arc segment relative to the longitudinal axis. ### Claim 21 (depends on 20) The gas turbine engine as recited in claim 20, wherein the airfoil section is integrally formed with the main body of a respective one of the arc segments. ### Claim 22 (depends on 21) The gas turbine engine as recited in claim 21, further comprising: at least one spring member dimensioned to bias a first end portion of the respective one of the arc segments radially outwardly relative to the longitudinal axis; and wherein the at least one spring member is dimensioned to provide a radial reaction force that is between approximately 70 percent and approximately 95 percent of a peak total radial aerodynamic load acting on the respective one of the arc segments and a respective one of the vanes axially forward of the first attachment portion relative to the longitudinal axis in operation. ### Claim 23 (depends on 20) The gas turbine engine as recited in claim 20, further comprising: a plurality of seal members dimensioned to establish a sealing relationship with an outer periphery of a respective one of the airfoil sections; and wherein each of the arc segments extends in a circumferential direction between a first mate face and a second mate face with respect to the longitudinal axis, the first mate face dimensioned to establish an intersegment gap with the second mate face of an adjacent one of the arc segments, and the first and second mate faces including respective slots dimensioned to receive a respective one of the seal members such that the seal member spans across the intersegment gap. ### Claim 24 (depends on 20) The gas turbine engine as recited in claim 20, wherein each of the arc segments includes a plurality of openings, and each of the openings at least partially receives a spar body of a respective one of the spar members and the airfoil section of a respective one of the vanes such that the arc segment is radially aligned with at least two of the spar members and at least two of the vanes relative to the longitudinal axis. ### Claim 25 (depends on 24) The gas turbine engine as recited in claim 24, wherein each of the openings circumscribes a perimeter of the spar body of the respective one of the spar members and circumscribes a perimeter of the airfoil section of a respective one of the vanes.