# US 2019/0316477 A1 · Gas turbine engine component for acoustic attenuation > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 2019/0316477 A1 - Document ID: US20190316477A1 - Status: Published application - Kind: A1 - Application number: 15/951,289 - Filed: 2018-04-12 - Published: 2019-10-17 - Earliest claimed: 2018-04-12 - Assignees: United Technologies Corporation - Applicants: United Technologies Corporation - Inventors: Gary D. Roberge - Web page: https://patentgrep.com/patent/US20190316477A1 ## CPC - F01D 9/02: Nozzles; Nozzle boxes; Stator blades; Guide conduits {, e.g. individual nozzles (nozzle boxes F01D9/047)} - F01D 25/04: Antivibration arrangements - F01D 5/08: Heating, heat-insulating or cooling means {(specially adapted for radial flow machines or engines F01D5/04)} - F01D 11/001: {for sealing space between stator blade and rotor} - 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}) - F05D 2220/323: for aircraft propulsion, e.g. jet engines - F05D 2240/55: Seals - F05D 2250/283: honeycomb - F05D 2260/20: Heat transfer, e.g. cooling - F05D 2260/96: Preventing, counteracting or reducing vibration or noise - F05D 2260/963: by Helmholtz resonators - 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 section for a gas turbine engine according to an example of the present disclosure includes, among other things, a rotor having a hub carrying a plurality of blades, the hub rotatable about a longitudinal axis, and a seal extending outwardly from the hub to establish a sealing relationship with a plurality of vanes distributed about the longitudinal axis. A flow guide assembly is secured to an engine static structure such that a flow path is defined between the hub and the flow guide assembly. The flow path has an inlet portion defined along the seal and an outlet portion, and the flow guide assembly includes an acoustic liner that extends along the flow path. ## Claims ### Claim 1 A section for a gas turbine engine comprising: a rotor including a hub carrying a plurality of blades, the hub rotatable about a longitudinal axis, and a seal that extends outwardly from the hub to establish a sealing relationship with a plurality of vanes distributed about the longitudinal axis; and a flow guide assembly secured to an engine static structure such that a flow path is defined between the hub and the flow guide assembly, the flow path including an inlet portion defined along the seal and an outlet portion, and the flow guide assembly including an acoustic liner that extends along the flow path. ### Claim 2 (depends on 1) The section as recited in claim 1, wherein the seal is a knife edge seal that includes one or more knife edge portions supported by a neck portion, the neck portion extending radially outward from the hub with respect to the longitudinal axis. ### Claim 3 (depends on 2) The section as recited in claim 2, wherein the acoustic liner is radially inward of the neck portion with respect to the longitudinal axis, and the hub and the flow guide assembly are dimensioned such that the flow path slopes radially inward from the neck portion to the outlet portion with respect to the longitudinal axis. ### Claim 4 (depends on 2) The section as recited in claim 2, wherein the knife edge seal establishes the sealing relationship with an abradable honeycomb structure mounted to the plurality of vanes in response to rotation of the rotor about the longitudinal axis. ### Claim 5 (depends on 1) The section as recited in claim 1, wherein the acoustic liner defines at least one resonant chamber dimensioned with respect to an acoustic frequency range relating to the flow path. ### Claim 6 (depends on 5) The section as recited in claim 5, wherein the acoustic liner includes first and second face sheets that establish the at least one resonant chamber, and the acoustic liner includes a honeycomb core disposed in the at least one resonant chamber, the honeycomb core including a plurality of honeycomb cells. ### Claim 7 (depends on 6) The section as recited in claim 6, wherein surfaces of the first face sheet bounding the flow path define a plurality of perforations that interconnect the flow path and the plurality of honeycomb cells. ### Claim 8 (depends on 7) The section as recited in claim 7, wherein the at least one resonant chamber includes a first resonant chamber adjacent the inlet portion and a second resonant chamber adjacent the outlet portion, the acoustic frequency range relating to the first resonant chamber differing from the acoustic frequency range relating to the second resonant chamber. ### Claim 9 (depends on 7) The section as recited in claim 7, wherein the seal is a knife edge seal that includes one or more knife edge portions supported by a neck portion, the neck portion extends radially outward from the hub with respect to the longitudinal axis, and the inlet portion extends between the neck portion and surfaces of the first face sheet. ### Claim 10 (depends on 1) The section as recited in claim 1, wherein the hub and the flow guide assembly are dimensioned such that the flow path slopes radially inward from the inlet portion to the outlet portion with respect to the longitudinal axis. ### Claim 11 (depends on 1) The section as recited in claim 1, wherein the section is a high pressure compressor section of the gas turbine engine. ### Claim 12 A gas turbine engine comprising: a fan section including a plurality of fan blades rotatable about an engine longitudinal axis; a compressor section that defines a core flow path, the compressor section including a first compressor and a second compressor downstream of the first compressor; a combustor section in fluid communication with the compressor section; a turbine section that drives the compressor section and the fan section; and wherein at least one of the compressor section and the turbine section includes a rotor assembly, the rotor assembly comprising: a rotor including a hub carrying a plurality of blades, the hub rotatable about the engine longitudinal axis, and a seal that extends outwardly from the hub to establish a sealing relationship with a plurality of vanes distributed about the engine longitudinal axis; and a flow guide assembly arranged about the longitudinal axis, the flow guide assembly secured to an engine static structure such that an annular flow path is defined between the hub and the flow guide assembly, the flow path including an inlet portion defined along the seal and an outlet portion, and the flow guide assembly including an acoustic liner that extends about the longitudinal axis to bound the flow path. ### Claim 13 (depends on 12) The gas turbine engine as recited in claim 12, wherein the second compressor section comprises the rotor assembly, and the outlet portion communicates with a cooling plenum that extends radially inward of a combustor of the combustor section with respect to the engine longitudinal axis. ### Claim 14 (depends on 12) The gas turbine engine as recited in claim 12, wherein the seal includes one or more knife edge portions supported by a neck portion, the neck portion extending radially outward from the hub with respect to the engine longitudinal axis. ### Claim 15 (depends on 12) The gas turbine engine as recited in claim 12, wherein walls of the hub that define the flow path slope radially inward from the neck portion toward the engine longitudinal axis such that inlet portion is radially outward of the outlet portion. ### Claim 16 A method of sealing of a gas turbine engine, comprising: rotating a knife edge seal about an engine longitudinal axis, the knife edge seal including one or more knife edge portions supported by a neck portion, and the neck portion extending radially outward from a hub with respect to the engine longitudinal axis; communicating flow between a core flow path and an annular flow path, the flow path defined between the hub and a flow guide assembly such that the flow path slopes towards the engine longitudinal axis; and wherein the flow guide assembly includes an acoustic liner that extends about the longitudinal axis to bound the flow path. ### Claim 17 (depends on 16) The method as recited in claim 16, wherein the hub is a compressor hub that carries a plurality of blades rotatable about the engine longitudinal axis. ### Claim 18 (depends on 17) The method as recited in claim 17, wherein the acoustic liner defines at least one resonant chamber dimensioned with respect to an acoustic frequency range such that the step of communicating the flow causes the at least one resonant chamber to at least partially attenuate acoustic energy in the acoustic frequency range relating to the annular flow path. ### Claim 19 (depends on 18) The method as recited in claim 18, wherein the acoustic liner includes a honeycomb core disposed in the at least one resonant chamber, surfaces of the acoustic liner define a plurality of perforations interconnecting the flow path and a plurality of honeycomb cells of the honeycomb core, and the plurality of perforations are defined with respect to the acoustic frequency range. ### Claim 20 (depends on 18) The method as recited in claim 18, wherein the flow path extends between an inlet portion and an outlet portion, the at least one resonant chamber includes a first resonant chamber adjacent the inlet portion and a second resonant chamber adjacent the outlet portion, and the acoustic frequency range relating to the first resonant chamber differs from the acoustic frequency range relating to the second resonant chamber.