# US 11,112,119 B2 · Combustor assembly for a turbo machine > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 11,112,119 B2 - Document ID: US11112119B2 - Status: Granted - Kind: B2 - Application number: 16/170,305 - Filed: 2018-10-25 - Published: 2020-04-30 - Granted: 2021-09-07 - Earliest claimed: 2018-10-25 - Assignees: General Electric Company - Applicants: General Electric Company - Inventors: Mackenzie Christopher Hock - Web page: https://patentgrep.com/patent/US11112119B2 ## CPC - F23R 3/54: Reverse-flow combustion chambers - F01D 9/023: {Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings} - F23R 3/002: {Wall structures (F23R3/02 and F23R3/007 take precedence)} - F23R 3/007: {constructed mainly of ceramic components} - F05D 2240/35: Combustors or associated equipment - F05D 2300/6033: Ceramic matrix composites [CMC] - 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) - F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS ## Abstract A turbo machine including an annular liner assembly defining a reverse flow combustion chamber is generally provided. The liner assembly includes a first piece defining an inner diameter (ID) combustor inlet portion, an outer diameter (OD) combustor outlet portion, and an outer diameter turbine shroud portion, in which the first piece defines a substantially solid volume between the inner diameter combustor inlet portion and the outer diameter combustor outlet portion. ## Claims ### Claim 1 A turbo machine, the turbo machine defining a radial direction, the turbo machine comprising: an annular liner assembly defining a reverse flow combustion chamber therewithin, the liner assembly comprising a first piece and a second piece, the first piece defining an inner diameter (ID) combustor inlet portion, an outer diameter (OD) combustor outlet portion, and an outer diameter (OD) turbine shroud portion, wherein the first piece defines a substantially solid volume between the ID combustor inlet portion and the OD combustor outlet portion, the annular liner assembly further comprising a dome assembly and a deflector wall of the dome assembly, the deflector wall positioned at an inlet of the dome assembly of the annular liner assembly, the deflector wall extending along the radial direction between the first piece and the second piece, and the deflector wall aligned along the radial direction with at least a portion of a first turbine blade. ### Claim 2 (depends on 1) The turbo machine of claim 1, wherein the first piece is a single unitary piece defined from the ID combustor inlet portion to the OD combustor outlet portion and the OD turbine shroud portion. ### Claim 3 (depends on 2) The turbo machine of claim 2, wherein the OD turbine shroud portion is extended over at least the first turbine blade of a turbine section of the turbo machine. ### Claim 4 (depends on 3) The turbo machine of claim 3, wherein the OD turbine shroud portion of the first piece extended over at least the first turbine blade is defined directly inward in the radial direction of the ID combustor inlet portion. ### Claim 5 (depends on 3) The turbo machine of claim 3, wherein a primary combustion zone is defined at the combustion chamber directly outward in the radial direction of the ID combustor inlet portion of the first piece, and wherein the primary combustion zone is defined directly outward in the radial direction of the OD turbine shroud portion extended over at least the first turbine blade. ### Claim 6 (depends on 3) The turbo machine of claim 3, wherein a radial plane is defined from the deflector wall of the dome assembly, and wherein the OD turbine shroud portion of the first piece is extended at least to the radial plane over the first turbine blade. ### Claim 7 (depends on 1) The turbo machine of claim 1, wherein the annular liner assembly comprises a ceramic matrix composite material. ### Claim 8 (depends on 7) The turbo machine of claim 7, wherein the ceramic matrix composite material comprises silicon carbide (SiC), silicon, silica, or alumina matrix materials, or combinations thereof. ### Claim 9 (depends on 1) The turbo machine of claim 1, wherein the substantially solid volume is approximately 95% or greater solid between the ID combustor inlet portion and the OD combustor outlet portion. ### Claim 10 (depends on 1) The turbo machine of claim 1, wherein a radius is defined at the first piece between the ID combustor inlet portion and the OD combustor outlet portion. ### Claim 11 (depends on 10) The turbo machine of claim 10, wherein a volume of the first piece between the ID combustor inlet portion and the OD combustor outlet portion is equal to or less than the radius defined at the first piece. ### Claim 12 (depends on 1) The turbo machine of claim 1, further comprising: a nozzle assembly coupled to the annular liner assembly at the first piece OD combustor outlet portion. ### Claim 13 (depends on 12) The turbo machine of claim 12, wherein the nozzle assembly is defined as a single structure integral to the first piece of the annular liner assembly. ### Claim 14 (depends on 12) The turbo machine of claim 12, wherein the annular liner assembly further comprises: a vane assembly disposed downstream of the nozzle assembly, wherein the vane assembly is coupled to the first piece of the annular liner assembly. ### Claim 15 A gas generator for a gas turbine engine defining a radial direction, the gas generator comprising: an annular liner assembly defining a reverse flow combustion chamber therewithin, the annular liner assembly comprising a first piece and a second piece, the first piece defining an inner diameter (ID) combustor inlet portion, an outer diameter (OD) combustor outlet portion, and an outer diameter (OD) turbine shroud portion together formed integrally, wherein the first piece is approximately 95% or greater of a solid volume of a ceramic matrix composite material between the ID combustor inlet portion and the OD combustor outlet portion, the annular liner assembly further comprising a dome assembly and a deflector wall of the dome assembly, the deflector wall positioned at an inlet of the dome assembly of the annular liner assembly, the defector wall extending along the radial direction between the first piece and the second piece, and the deflector wall aligned along the radial direction with at least a portion of a first turbine blade. ### Claim 16 (depends on 15) The gas generator of claim 15, wherein the first piece is a single unitary piece forming the ID combustor inlet portion to the OD combustor outlet portion and the OD turbine shroud portion. ### Claim 17 (depends on 16) The gas generator of claim 16, wherein a radial plane is defined from the deflector wall, and wherein the ID combustor inlet portion of the first piece is defined from the radial plane to a radius defined at the first piece of the annular liner assembly. ### Claim 18 (depends on 17) The gas generator of claim 17, wherein the OD turbine shroud portion of the first piece is defined inward in the radial direction of the ID combustor inlet portion of the first piece. ### Claim 19 (depends on 18) The gas generator of claim 18, wherein a primary combustion zone is defined at the combustion chamber between the first piece and a second piece of the annular liner assembly, and wherein the primary combustion zone is defined, in the radial direction, directly between the first piece and the second piece of the annular liner assembly between the radius of the first piece and the deflector wall of the dome assembly. ### Claim 20 (depends on 18) The gas generator of claim 18, wherein the OD turbine shroud portion of the first piece is extended over the first turbine blade.