# US 10,767,485 B2 · Radial cooling system for gas turbine engine compressors > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,767,485 B2 - Document ID: US10767485B2 - Status: Granted - Kind: B2 - Application number: 15/864,925 - Filed: 2018-01-08 - Published: 2019-07-11 - Granted: 2020-09-08 - Earliest claimed: 2018-01-08 - Assignees: Raytheon Technologies Corporation - Applicants: United Technologies Corporation - Inventors: Matthew E Bintz; Enzo DiBenedetto; Paul E Coderre - Web page: https://patentgrep.com/patent/US10767485B2 ## CPC - F01D 5/087: {in the radial passages of the rotor disc} - 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)} - F02C 7/141: of working fluid - F02C 7/18: the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} - F01D 11/001: {for sealing space between stator blade and rotor} - F01D 11/02: by non-contact sealings, e.g. of labyrinth type (for sealing space between rotor blade tips and stator F01D11/08) - F04D 29/321: {for axial flow compressors} - F04D 29/582: {specially adapted for elastic fluid pumps} - F05D 2210/12: gaseous, i.e. compressible - F05D 2220/3218: for an intermediate stage of a compressor - F05D 2220/3219: for the last stage of a compressor or a high pressure compressor - F05D 2240/11: Shroud seal segments - F05D 2240/12: Fluid guiding means, e.g. vanes - F05D 2240/24: for turbines - F05D 2240/55: Seals - F05D 2260/20: Heat transfer, e.g. cooling - F05D 2260/209: using vortex tubes - F05D 2260/213: by the provision of a heat exchanger within the cooling circuit - 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) - F04D: NON-POSITIVE-DISPLACEMENT PUMPS (engine fuel-injection pumps F02M; ion pumps H01J41/12; electrodynamic pumps H02K44/02) - F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS ## Abstract A gas turbine engine is disclosed. The gas turbine engine includes a first rotor supporting a first plurality of circumferentially spaced rotor blades and a second rotor disposed axially downstream of the first rotor and supporting a second plurality of circumferentially spaced rotor blades, a first bore cavity between the first rotor and the second rotor, a first fluid passageway configured to provide cooled air to the first bore cavity and a first anti-vortex component positioned proximate the first bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the first bore cavity. ## Claims ### Claim 1 A gas turbine engine, comprising: a first rotor supporting a first plurality of circumferentially spaced rotor blades and a second rotor disposed axially downstream of the first rotor and supporting a second plurality of circumferentially spaced rotor blades; a first bore cavity between the first rotor and the second rotor; a second bore cavity between the second rotor and an aft hub; a fluid passageway configured to provide a cooled air to the first bore cavity and to the second bore cavity; a first anti-vortex component positioned within the first bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the first bore cavity; a second anti-vortex component positioned within the second bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the second bore cavity; and an orifice through the aft hub configured to introduce an aft cooled air flow path from the second bore cavity onto a downstream rim portion of the second rotor, wherein the downstream rim portion of the second rotor is disposed axially upstream and adjacent a plurality of exit guide vanes. ### Claim 2 (depends on 1) The gas turbine engine of claim 1, further comprising a vane inner shroud disposed radially outward of the first bore cavity and axially intermediate the first rotor and the second rotor, the vane inner shroud supporting a plurality of circumferentially spaced vanes on a radially outer surface and having a sealing surface on a radially inner surface, the first plurality of circumferentially spaced rotor blades, the second plurality of circumferentially spaced rotor blades and the plurality of circumferentially spaced vanes defining a portion of a core flow path. ### Claim 3 (depends on 2) The gas turbine engine of claim 2, further comprising a seal disposed between the radially inner surface of the vane inner shroud and the first bore cavity, the seal configured to obstruct passage of air from the core flow path from an axially downstream position of the plurality of circumferentially spaced vanes to an axially upstream position of the plurality of circumferentially spaced vanes via a seal flow space between the seal and the radially inner surface of the vane inner shroud. ### Claim 4 (depends on 3) The gas turbine engine of claim 3, further comprising a first gap disposed downstream of the plurality of circumferentially spaced vanes that provides an inlet passage for air from the core flow path to flow between the seal and the radially inner surface of the vane inner shroud. ### Claim 5 (depends on 4) The gas turbine engine of claim 4, further comprising a second gap disposed upstream of the plurality of circumferentially spaced vanes that provides an exit passage for air between the seal and the radially inner surface of the vane inner shroud to return to the core flow path. ### Claim 6 (depends on 5) The gas turbine engine of claim 5, further comprising a seal inlet configured to introduce a cooled air flow path from the first bore cavity into the seal flow space. ### Claim 7 (depends on 6) The gas turbine engine of claim 6, wherein the seal comprises a labyrinth seal having one or more knife edges. ### Claim 8 (depends on 6) The gas turbine engine of claim 6, wherein the first anti-vortex component comprises a first anti-vortex tube. ### Claim 9 (depends on 6) The gas turbine engine of claim 6, wherein the fluid passageway includes a removal orifice extending into the core flow path at a removal location. ### Claim 10 (depends on 9) The gas turbine engine of claim 9, wherein the removal location is positioned upstream of the first rotor. ### Claim 11 (depends on 10) The gas turbine engine of claim 10, further comprising a heat exchanger positioned within the fluid passageway intermediate the removal orifice and an inlet orifice positioned between the first rotor and the second rotor. ### Claim 12 (depends on 11) The gas turbine engine of claim 11, wherein the first anti-vortex component comprises a first anti-vortex tube and the second anti-vortex component comprises a second anti-vortex tube. ### Claim 13 (depends on 12) The gas turbine engine of claim 12, wherein the first anti-vortex tube has a first length and the second anti-vortex tube has a second length greater than the first length. ### Claim 14 A compressor for a gas turbine engine, comprising: a first rotor supporting a first plurality of circumferentially spaced rotor blades, a second rotor disposed axially downstream of the first rotor and supporting a second plurality of circumferentially spaced rotor blades and a third rotor disposed axially downstream of the second rotor and supporting a third plurality of circumferentially spaced rotor blades; a first bore cavity between the first rotor and the second rotor, a second bore cavity between the second rotor and the third rotor and a third bore cavity between the third rotor and an aft hub; a fluid passageway configured to provide cooled air to the first bore cavity, the second bore cavity and the third bore cavity; a first anti-vortex component positioned within the first bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the first bore cavity; a second anti-vortex component positioned within the second bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the second bore cavity; and a third anti-vortex component positioned within the third bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the third bore cavity; and an orifice through the aft hub configured to introduce an aft cooled air flow path from the third bore cavity onto a downstream rim portion of the third rotor, wherein the downstream rim portion of the third rotor is disposed axially upstream and adjacent a plurality of exit guide vanes. ### Claim 15 (depends on 14) The compressor of claim 14, further comprising: a first vane inner shroud disposed radially outward of the first bore cavity and axially intermediate the first rotor and the second rotor, the first vane inner shroud supporting a first plurality of circumferentially spaced vanes on a first radially outer surface and having a first sealing surface on a first radially inner surface, a second vane inner shroud disposed radially outward of the second bore cavity and axially intermediate the second rotor and the third rotor, the second vane inner shroud supporting a second plurality of circumferentially spaced vanes on a second radially outer surface and having a second sealing surface on a second radially inner surface, the first plurality of circumferentially spaced rotor blades, the second plurality of circumferentially spaced rotor blades and the third plurality of circumferentially spaced rotor blades, and the first plurality of circumferentially spaced vanes and the second plurality of circumferentially spaced vanes defining a portion of a core flow path. ### Claim 16 (depends on 15) The compressor of claim 15, further comprising: a first seal disposed between the first radially inner surface of the first vane inner shroud and the first bore cavity, the first seal configured to obstruct passage of air from the core flow path from an axially downstream position of the first plurality of circumferentially spaced vanes to an axially upstream position of the first plurality of circumferentially spaced vanes via a first seal flow space between the first seal and the first radially inner surface of the first vane inner shroud and a second seal disposed between the second radially inner surface of the second vane inner shroud and the second bore cavity, the second seal configured to obstruct passage of air from the core flow path from an axially downstream position of the second plurality of circumferentially spaced vanes to an axially upstream position of the second plurality of circumferentially spaced vanes via a second seal flow space between the second seal and the second radially inner surface of the second vane inner shroud; a first seal inlet configured to introduce a first cooled air flow path from the first bore cavity into the first seal flow space and a second seal inlet configured to introduce a second cooled air flow path from the second bore cavity into the second seal flow space. ### Claim 17 (depends on 16) The compressor of claim 16, wherein the first anti-vortex component comprises a first anti-vortex tube and the second anti-vortex component comprises a second anti-vortex tube. ### Claim 18 (depends on 16) The compressor of claim 16, further comprising a heat exchanger positioned within the fluid passageway.