# US 10,371,383 B2 · Unitary flow path structure > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,371,383 B2 - Document ID: US10371383B2 - Status: Granted - Kind: B2 - Application number: 15/417,602 - Filed: 2017-01-27 - Published: 2018-08-02 - Granted: 2019-08-06 - Earliest claimed: 2017-01-27 - Assignees: General Electric Company - Applicants: General Electric Company - Inventors: Brandon ALlanson Reynolds; Jonathan David Baldiga; Andrew Scott Bilse; Michael Todd Radwanski; Ernesto Andres Vallejo Ruiz; Aaron Michael Dziech; Mark Eugene Noe - Web page: https://patentgrep.com/patent/US10371383B2 ## CPC - F23R 3/002: {Wall structures (F23R3/02 and F23R3/007 take precedence)} - F01D 5/06: Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections {(F01D5/022, F01D5/023 take precedence)} - F01D 5/225: {by shrouding} - F01D 9/023: {Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings} - F01D 9/04: forming ring or sector - F02C 3/14: characterised by the arrangement of the combustion chamber in the plant (combustion chambers per se F23R; F02C3/205 takes precedence) - F23R 3/10: for primary air (F23R3/06, F23R3/045 take precedence) - 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 2300/6033: Ceramic matrix composites [CMC] - F23R 2900/00005: Preventing fatigue failures or reducing mechanical stress in gas turbine components - 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 - 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) - F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D) - F02C: GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS (construction of turbines F01D; jet-propulsion plants F02K; construction of compressors or fans F04; generating combustion products of high pressure or high velocity F23R; using gas turbines in compression refrigeration plants F25B11/00) - 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 and gas turbine engines are provided. A flow path assembly may comprise a combustor dome positioned at a forward end of a combustor of a combustion section of a gas turbine engine, and a unitary outer wall including a combustor portion extending through the combustion section and a turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine. The combustor portion and the turbine portion are integrally formed as a single unitary structure. The flow path assembly also comprises an inner wall extending from the forward end of the combustor through at least the combustion section. The combustor dome extends radially from the unitary outer wall to the inner wall and is configured to move axially with respect to the inner wall and the unitary outer wall. Other flow path assemblies and gas turbine engine configurations are provided. ## Claims ### Claim 1 A flow path assembly of a gas turbine engine, the flow path assembly comprising: a combustor dome positioned at a forward end of a combustor of a combustion section of the gas turbine engine; an annular unitary outer wall including a combustor portion extending through the combustion section 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 shrould of a blade portion of the second turbine stage, and the combustor portion and the turbine portion being integrally formed as a single annular unitary structure; and an annular inner wall extending from the forward end of the combustor through at least the combustion section, wherein the combustor dome extends radially from the annular unitary outer wall to the annular inner wall, and wherein the combustor dome is configured to move axially with respect to the annular inner wall and the annular unitary outer wall. ### Claim 2 (depends on 1) The flow path assembly of claim 1, wherein an axial slip joint is formed between the combustor dome and each of the annular unitary outer wall and the annular inner wall. ### Claim 3 (depends on 1) The flow path assembly of claim 1, wherein the combustor dome comprises an outer wing and an inner wing, the outer wing extending along the annular unitary outer wall and the inner wing extending along the annular inner wall. ### Claim 4 (depends on 1) The flow path assembly of claim 1, wherein the annular unitary outer wall is configured to move radially with respect to the combustor dome. ### Claim 5 (depends on 4) The flow path assembly of claim 4, wherein the annular inner wall is configured to move radially with respect to the combustor dome. ### Claim 6 (depends on 1) The flow path assembly of claim 1, wherein the combustor portion of the annular unitary outer wall comprises an outer liner of the combustor of the combustion section. ### Claim 7 (depends on 1) The flow path assembly of claim 1, wherein the annular inner wall comprises an inner liner of the combustor and an inner band of the nozzle portion of the first turbine stage, and wherein the inner liner and the inner band are integrally formed as a single unitary structure. ### Claim 8 (depends on 1) The flow path assembly of claim 1, wherein the combustor portion and the turbine portion are integrally formed from a ceramic matrix composite material such that the annular unitary outer wall is a ceramic matrix composite component. ### Claim 9 (depends on 8) The flow path assembly of claim 8, wherein the annular inner wall and the combustor dome comprise a ceramic matrix composite material. ### Claim 10 A flow path assembly of a gas turbine engine, the flow path assembly extending from a combustor through a turbine section of the gas turbine engine, the turbine section including a first turbine stage immediately downstream of the combustor and a second turbine stage immediately downstream of the first turbine stage, the flow path assembly comprising: a combustor dome positioned at a forward end of the combustor; an inner wall defining an inner boundary of a flow path, the inner wall comprising an inner liner of the combustor; and a unitary outer wall defining an outer boundary of the flow path, the unitary outer wall having a combustor portion extending from the forward end through the combustor and having a turbine portion extending through at least the first turbine stage and the second turbine stage, 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 annular unitary structure and wherein the combustor dome is a separate component from the inner wall and the unitary outer wall. ### Claim 11 (depends on 10) The flow path assembly of claim 10, wherein an axial slip joint is formed between the combustor dome and each of the unitary outer wall and the inner wall such that the combustor dome is configured to move axially with respect to the inner wall and the unitary outer wall. ### Claim 12 (depends on 11) The flow path assembly of claim 11, wherein the unitary outer wall is configured to move radially with respect to the combustor dome, and wherein the inner wall is configured to move radially with respect to the combustor dome. ### Claim 13 (depends on 11) The flow path assembly of claim 11, wherein the combustor dome is attached to a fuel nozzle assembly. ### Claim 14 (depends on 10) The flow path assembly of claim 10, wherein the combustor dome comprises an outer wing and an inner wing, the outer wing extending along the unitary outer wall and the inner wing extending along the inner wall. ### Claim 15 (depends on 14) The flow path assembly of claim 14, wherein the outer wing extends forward from a body of the combustor dome, and wherein the inner wing extends forward from the body of the combustor dome. ### Claim 16 (depends on 10) The flow path assembly of claim 10, wherein the combustor dome comprises an outer end and an inner wing extending forward from a body, wherein the inner wing is pinned with the inner wall at the forward end of the combustor, and wherein a spring element is pinned with the unitary outer wall at the forward end of the combustor, the spring element pressing against the body of the combustor dome to preload the outer end against the unitary outer wall. ### Claim 17 A gas turbine engine, comprising: a combustion section including a combustor; and a turbine section including a first turbine stage positioned immediately downstream of the combustion section and a second turbine stage positioned immediately downstream of the first turbine stage, wherein the combustion section and the turbine section define a flow path, wherein the combustion section includes an inner liner defining an inner boundary of the flow path through the combustion section, an outer liner defining an outer boundary of the flow path through the combustion section, and a combustor dome positioned at a forward end of the combustor and extending between the inner liner and the outer liner, wherein each of the first turbine stage and the second turbine stage of the turbine section includes a nozzle portion and a blade portion, each nozzle portion comprising an inner band defining an inner boundary of the flow path through the nozzle portion and comprising an outer band defining an outer boundary of the flow path through the nozzle portion, each blade portion comprising a shroud defining an outer boundary of the flow path through the blade portion, wherein the outer liner, the outer bands, and the shrouds are integrally formed such that the outer liner, the outer bands, and the shrouds are a single unitary outer wall, and wherein the combustor dome is a separate component from the inner liner and the single unitary outer wall. ### Claim 18 (depends on 17) The gas turbine engine of claim 17, wherein an axial slip joint is formed between the combustor dome and each of the single unitary outer wall and the inner liner. ### Claim 19 (depends on 17) The gas turbine engine of claim 17, wherein the inner liner, the single unitary outer wall, and the combustor dome are each formed from a plurality of ceramic matrix composite tapes each comprising a reinforcement material of unidirectional arrays of continuous fiber tows. ### Claim 20 (depends on 17) The gas turbine engine of claim 17, wherein each of the combustor dome, the inner liner, and the single unitary outer wall comprises a ceramic matrix composite material, the ceramic matrix composite material comprising ceramic fibers pre-impregnated with matrix material.