# US 10,370,990 B2 · Flow path assembly with pin supported nozzle airfoils > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,370,990 B2 - Document ID: US10370990B2 - Status: Granted - Kind: B2 - Application number: 15/440,320 - Filed: 2017-02-23 - Published: 2018-08-23 - Granted: 2019-08-06 - Earliest claimed: 2017-02-23 - Assignees: General Electric Company - Applicants: General Electric Company - Inventors: Brandon ALlanson Reynolds; Jonathan David Baldiga - Web page: https://patentgrep.com/patent/US10370990B2 ## CPC - F01D 9/042: {fixing blades to stators (fixing stator-rings in the casing or to each other F01D25/246)} - F01D 9/023: {Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings} - F01D 25/005: {Selecting particular materials} - F01D 25/24: Casings (modified for heating or cooling F01D25/14); Casing parts, e.g. diaphragms, casing fastenings (casings for rotary machines or engines in general F16M {; special arrangements in stators dealing with breaking-off of part of rotor F01D21/045}) - F23R 3/002: {Wall structures (F23R3/02 and F23R3/007 take precedence)} - F23R 3/007: {constructed mainly of ceramic components} - F23R 3/10: for primary air (F23R3/06, F23R3/045 take precedence) - F23R 3/60: Support structures; Attaching or mounting means - F05D 2220/32: in gas turbines - F05D 2230/53: by integrally manufacturing a component, e.g. by milling from a billet or one piece construction - F05D 2240/35: Combustors or associated equipment - F05D 2260/30: Retaining components in desired mutual position - F05D 2300/6033: Ceramic matrix composites [CMC] - F23R 2900/00017: Assembling combustion chamber liners or subparts - F23R 2900/00018: Manufacturing combustion chamber liners or subparts - 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) - F23R: GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS (fluidised bed combustion apparatus specially adapted for operation at superatmospheric pressures F23C10/16) - 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 Flow path assemblies for gas turbine engines are provided. For example, a flow path assembly comprises an inner wall and an outer wall defining an opening therethrough, as well as a nozzle airfoil having an inner end radially opposite an outer end that is positioned within the opening defined in the outer wall such that the outer end projects radially outwardly from the outer wall. The flow path assembly also comprises a pin and a first support member. The pin extends from the first support member to the outer end of the nozzle airfoil to support the nozzle airfoil. In some embodiments, the outer end of the nozzle airfoil defines a flange configured to receive the pin and a radial stop configured to press against a radially outward external support, and the pin is configured to transfer a load on the nozzle airfoil to an external structure. ## Claims ### Claim 1 A flow path assembly of a gas turbine engine, the flow path assembly comprising: an inner wall; a unitary outer wall including a combustor portion extending through combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage and a second turbine stage of a turbine section of the gas engine, wherein the turbine portion comprises an outer band of a nozzle portion of the first turbine stage, a shroud of a blade portion of the first turbine stage, an outer band of a nozzle portion of the second turbine stage, and a shroud of a blade portion of the second turbine stage, the combustor portion and the turbine portion being integrally formed as a single unitary structure, and the unitary outer wall defining an opening therethrough; a nozzle airfoil having an inner end radially opposite an outer end, the outer end positioned within the opening defined in the unitary outer wall such that the outer end projects radially outwardly from the unitary outer wall; a pin; and a first support member, wherein the pin extends from the first support member to the outer end of the nozzle airfoil to support the nozzle airfoil. ### Claim 2 (depends on 1) The flow path assembly of claim 1, wherein the pin extends through the outer end of the nozzle airfoil to a second support member such that the pin extends from the first support member to the second support member through the nozzle airfoil. ### Claim 3 (depends on 1) The flow path assembly of claim 1, wherein the pin extends through a flange defined at the outer end of the nozzle airfoil. ### Claim 4 (depends on 1) The flow path assembly of claim 1, wherein the nozzle airfoil defines a radial stop on an aft side of the outer end, and wherein the radial stop is configured to press against a radially outward external support. ### Claim 5 (depends on 1) The flow path assembly of claim 1, wherein the nozzle airfoil defines an axial stop on the outer end, and wherein the axial stop is configured to press against a radially outward external support. ### Claim 6 (depends on 1) The flow path assembly of claim 1, wherein the pin is integrally formed with the first support member. ### Claim 7 (depends on 1) The flow path assembly of claim 1, the combustor portion and the turbine portion being integrally formed from a ceramic matrix composite (CMC) material such that the unitary outer wall is a CMC component, and wherein the inner wall is formed from a CMC material. ### Claim 8 A flow path assembly of a gas turbine engine, the flow path assembly comprising: an inner wall; a unitary outer wall including a combustor portion extending through a combustion section of the gas turbine engine and turbine portion extending through at least a first turbine stage and a second turbine stage of a turbine section of the gas turbine engine, wherein the turbine portion comprises an outer band of a nozzle portion of the first turbine stage, a shroud of a blade portion of the first turbine stage, an outer band of a nozzle portion of the second turbine stage, and a shroud of a blade portion of the second turbine stage, the combustor portion and the turbine portion being integrally formed as a single unitary structure, and the unitary outer wall defining an opening therethrough; a nozzle airfoil having an inner end radially opposite an outer end, the outer end positioned within the opening defined in the unitary outer wall such that the outer end projects radially outwardly from the unitary outer wall; and a pin, wherein the outer end of the nozzle airfoil defines a flange and a radial stop, wherein the flange is configured to receive the pin and the radial stop is configured to press against a radially outward external support, and wherein the pin is configured to transfer a load on the nozzle airfoil to a structure external to the flow path assembly. ### Claim 9 (depends on 8) The flow path assembly of claim 8, wherein an aperture is defined through the flange for receipt of the pin. ### Claim 10 (depends on 9) The flow path assembly of claim 9, wherein the pin extends from a first support member to a second support member through the aperture in the flange. ### Claim 11 (depends on 8) The flow path assembly of claim 8, wherein the radial stop is defined at an aft side of the outer end of the nozzle airfoil. ### Claim 12 (depends on 11) The flow path assembly of claim 11, wherein the flange is defined at a forward side of the outer end of the nozzle airfoil. ### Claim 13 (depends on 8) The flow path assembly of claim 8, wherein the nozzle airfoil is formed from a ceramic matrix composite (CMC) material. ### Claim 14 (depends on 8) The flow path assembly of claim 8, the combustor portion and the turbine portion being integrally formed from a ceramic matrix composite (CMC) material such that the unitary outer wall is a CMC component, and wherein the inner wall is formed from a CMC material. ### Claim 15 A flow path assembly of a gas turbine engine, the flow path assembly comprising: an inner wall; a unitary outer wall defining a plurality of openings therethrough, the unitary outer wall including a combustor portion extending through a combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage and a second turbine stage of a turbine section of the gas turbine engine, wherein the turbine portion comprises an outer band of a nozzle portion of the first turbine stage, a shroud of a blade portion of the first turbine stage, an outer band of a nozzle portion of the second turbine stage, and a shroud of a blade portion of the second turbine stage, and the combustor portion and the turbine portion being integrally formed as a single unitary structure; a plurality of nozzle airfoils, each of the plurality of nozzle airfoils having an inner end radially opposite an outer end, the outer end of each of the plurality of nozzle airfoils positioned within one of the plurality of openings defined in the unitary outer wall such that the outer end projects radially outwardly from the unitary outer wall; and a plurality of pins, wherein the outer end of each of the plurality of nozzle airfoils defines a flange and a radial stop, wherein the flange of each of the plurality of nozzle airfoils is configured to receive one pin of the plurality of pins and the radial stop is configured to press against a radially outward external support, and wherein each pin of the plurality of pins is configured to transfer a load to a structure external to the flow path assembly. ### Claim 16 (depends on 15) The flow path assembly of claim 15, wherein an aperture is defined through the flange of each of the plurality of nozzle airfoils for receipt of the one pin of the plurality of pins. ### Claim 17 (depends on 16) The flow path assembly of claim 16, wherein each pin of the plurality of pins extends from a first support member to a second support member. ### Claim 18 (depends on 17) The flow path assembly of claim 17, wherein the plurality of pins are integrally formed with the first support member. ### Claim 19 (depends on 15) The flow path assembly of claim 15, wherein each of the plurality of nozzle airfoils is formed from a ceramic matrix composite (CMC) material. ### Claim 20 (depends on 19) The flow path assembly of claim 19, wherein the combustor portion and the turbine portion are integrally formed from a CMC material such that the unitary outer wall is a CMC component, and wherein the inner wall is formed from a CMC material.