# US 11,994,067 B2 · Gas turbine engine > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 11,994,067 B2 - Document ID: US11994067B2 - Status: Granted - Kind: B2 - Application number: 17/940,055 - Filed: 2022-09-08 - Published: 2023-06-01 - Granted: 2024-05-28 - Earliest claimed: 2022-09-08 - Assignees: ROLLS-ROYCE plc - Applicants: ROLLS-ROYCE plc - Inventors: Natalie C Wong; Jonathan A Cherry; Paul R Davies; David A Jones; Andrew J Newman; Benjamin J Sellers; Stephen J Bradbrook - Web page: https://patentgrep.com/patent/US11994067B2 ## CPC - F02C 7/12: Cooling of plants (of component parts, see the relevant subclasses, e.g. F01D; cooling of engines in general F01P) - F02C 6/20: Adaptations of gas-turbine plants for driving vehicles - F02C 7/32: Arrangement, mounting, or driving, of auxiliaries - F05D 2220/323: for aircraft propulsion, e.g. jet engines - F05D 2220/76: an electrical generator - F05D 2260/213: by the provision of a heat exchanger within the cooling circuit - 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 ## Abstract An aircraft gas turbine engine includes a heat exchanger module, and a core engine including an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine. The high-pressure compressor is connected to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor is connected to the low-pressure turbine by a second shaft. The heat exchanger module includes a central hub and heat transfer elements extending radially from the central hub and spaced in a circumferential array, for transferring heat energy from a fluid within the heat transfer elements to an inlet airflow passing over the heat transfer elements prior to entry of the airflow into an inlet to the core engine. The gas turbine engine further includes a first electric machine connected to the first shaft and positioned downstream of the heat exchanger module, and a second electric machines connected to the second shaft. ## Claims ### Claim 1 A gas turbine engine for an aircraft, the gas turbine engine comprising, in axial flow sequence, a heat exchanger module, and a core engine, the core engine comprising, in axial flow sequence, an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine, the high-pressure compressor being rotationally connected to the high-pressure turbine by a first shaft, the intermediate-pressure compressor being rotationally connected to the low-pressure turbine by a second shaft, the heat exchanger module being in fluid communication with the core engine by an inlet duct, the heat exchanger module comprising a central hub and a plurality of heat transfer elements extending radially outwardly from the central hub and spaced in a circumferential array, for transfer of heat energy from a first fluid contained within the heat transfer elements to an inlet airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the core engine, and wherein the gas turbine engine further comprises a first electric machine and a second electric machine, the first electric machine is rotationally connected to the first shaft, the first electric machine is positioned downstream of the heat exchanger module relative to a direction of the airflow, and the second electrical machine is rotationally connected to the second shaft, the first electric machine is coaxial with the first shaft, and the entire first electric machine is positioned upstream of the intermediate-pressure compressor. ### Claim 2 (depends on 1) The gas turbine engine as claimed in claim 1, wherein the second electric machine is accommodated within the central hub. ### Claim 3 (depends on 1) The gas turbine engine as claimed in in claim 1, wherein the second electric machine is positioned axially downstream of the low-pressure turbine. ### Claim 4 (depends on 1) The gas turbine engine as claimed in claim 1, wherein the first electric machine is positioned axially upstream of the core engine. ### Claim 5 (depends on 1) The gas turbine engine as claimed in claim 1, wherein at least one of the first electric machine and the second electric machine, comprises an axial length LEM and a diameter DEM, and wherein a ratio of the axial length to the diameter (LEM/DEM) for the respective electric machine is in a range between 0.5 to 2.0. ### Claim 6 (depends on 1) The gas turbine engine as claimed in claim 1, further comprising an electrical energy storage unit, and wherein the electrical energy storage unit is configured to store electrical energy that may be generated by at least one of the first electrical machine and the second electrical machine. ### Claim 7 (depends on 6) The gas turbine engine as claimed in claim 6, wherein the electrical energy storage unit is a battery. ### Claim 8 (depends on 6) The gas turbine engine as claimed in claim 6, wherein the electrical energy storage unit is a capacitor. ### Claim 9 (depends on 1) The gas turbine engine as claimed in claim 1, wherein the gas turbine engine further comprises a fan assembly, the fan assembly being rotationally connected to the second shaft. ### Claim 10 (depends on 9) The gas turbine engine as claimed in claim 9, wherein the second electrical machine is rotationally connected to the fan assembly. ### Claim 11 (depends on 9) The gas turbine engine as claimed in claim 9, the fan assembly comprising a plurality of fan blades extending radially from a hub, the plurality of fan blades defining a fan diameter (DFAN), and wherein the fan diameter DFAN is within the range of 0.3 m to 1.4 m. ### Claim 12 (depends on 9) The gas turbine engine as claimed in claim 9, the gas turbine engine further comprising an outer casing, the outer casing enclosing the sequential arrangement of fan assembly, compressor module, combustor module, and turbine module, an annular bypass duct being defined between the outer casing and the sequential arrangement of compressor module, combustor module, and turbine module, a bypass ratio being defined as a ratio of a mass air flow rate through the bypass duct to a mass air flow rate through the sequential arrangement of modules, and wherein the bypass ratio is less than 4.0. ### Claim 13 (depends on 9) The gas turbine engine as claimed in claim 9, wherein the fan assembly has two or more fan stages, at least one of the fan stages comprising a plurality of fan blades defining the fan diameter DFAN. ### Claim 14 (depends on 1) An aircraft comprising a gas turbine engine as claimed in claim 1.