# US 12,110,830 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 12,110,830 B2 - Document ID: US12110830B2 - Status: Granted - Kind: B2 - Application number: 18/137,123 - Filed: 2023-04-20 - Published: 2023-08-31 - Granted: 2024-10-08 - Earliest claimed: 2018-08-10 - Assignees: ROLLS-ROYCE plc - Applicants: ROLLS-ROYCE plc - Inventors: Pascal Dunning; Roderick M. Townes; Michael J. Whittle - Web page: https://patentgrep.com/patent/US12110830B2 ## CPC - F02C 7/36: Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) - F01D 5/02: Blade-carrying members, e.g. rotors (rotors of non-bladed type F01D1/34; stators F01D9/00 {; selecting particular materials F01D5/28}) - F01D 5/282: {Selecting composite materials, e.g. blades with reinforcing filaments} - F01D 5/284: {Selection of ceramic materials} - F01D 5/3084: {the blades being made of ceramics} - F01D 25/005: {Selecting particular materials} - F02C 3/073: the compressor and turbine stages being concentric - F02C 7/18: the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} - F02C 9/18: by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} - F05D 2220/3212: the first stage of a turbine - F05D 2240/11: Shroud seal segments - F05D 2240/12: Fluid guiding means, e.g. vanes - F05D 2240/30: Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor - F05D 2260/40311: of the epicyclical, planetary or differential type - F05D 2300/6033: Ceramic matrix composites [CMC] - Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft - 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) - 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 highly efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power. ## Claims ### Claim 1 A gas turbine engine for an aircraft comprising: an engine core comprising: a first turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft; the gas turbine engine further comprising: a fan comprising a plurality of fan blades; and a gearbox that receives an input from the first core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, wherein: the mass of the second turbine is no greater than 9% of the total dry mass of the gas turbine engine, and the maximum temperature at entry to a most axially upstream rotor stage of the second turbine at a maximum take-off condition is in an inclusive range of from 1800K to 2100K. ### Claim 2 (depends on 1) The gas turbine engine according to claim 1, wherein the mass of the second turbine is in a range of from 3% to 9% of the total dry mass of the gas turbine engine. ### Claim 3 (depends on 1) The gas turbine engine according to claim 1, wherein the total mass of the turbine is in a range of from 7% to 17% of the total dry mass of the gas turbine engine. ### Claim 4 (depends on 1) The gas turbine engine according to claim 1, wherein the mass of the second turbine is no greater than 8% of the total dry mass of the gas turbine engine. ### Claim 5 (depends on 1) The gas turbine engine according to claim 1, wherein the second turbine comprises at least one ceramic matrix composite component. ### Claim 6 (depends on 5) The gas turbine engine according to claim 5, wherein the mass of ceramic matrix composite in the second turbine is in a range of from 1% to 20% of the total mass of the second turbine. ### Claim 7 (depends on 1) The gas turbine engine according to claim 1, wherein an overall pressure ratio of the gas turbine engine at cruise conditions is greater than 40. ### Claim 8 (depends on 7) The gas turbine engine according to claim 7, wherein the overall pressure ratio of the gas turbine engine at cruise conditions is less than 75. ### Claim 9 (depends on 1) The gas turbine engine according to claim 1, wherein a bypass ratio of the gas turbine engine at cruise conditions is in an inclusive range of from 10 to 12.5. ### Claim 10 (depends on 1) The gas turbine engine according to claim 1, wherein a specific thrust of the gas turbine engine is in an inclusive range of from 90 Nkg−1s to 100 Nkg−1s. ### Claim 11 A gas turbine engine for an aircraft comprising: an engine core comprising: a first turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft; the gas turbine engine further comprising: a fan comprising a plurality of fan blades; and a gearbox that receives an input from the first core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, wherein: the mass of the second turbine is no greater than 9% of the total dry mass of the gas turbine engine, and a bypass ratio of the gas turbine engine at cruise conditions in an inclusive range of from 10 to 12.5. ### Claim 12 (depends on 11) The gas turbine engine according to claim 11, wherein the bypass ratio of the gas turbine engine at cruise conditions is in an inclusive range of from 10 to 11.5. ### Claim 13 (depends on 11) The gas turbine engine according to claim 11, wherein the mass of the second turbine is in a range of from 3% to 9% of the total dry mass of the gas turbine engine. ### Claim 14 (depends on 11) The gas turbine engine according to claim 11, wherein an overall pressure ratio of the gas turbine engine at cruise conditions is greater than 40. ### Claim 15 (depends on 11) The gas turbine engine according to claim 11, wherein the second turbine comprises two rotor blade rows. ### Claim 16 A gas turbine engine for an aircraft comprising: an engine core comprising: a first turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft; the gas turbine engine further comprising: a fan comprising a plurality of fan blades; and a gearbox that receives an input from the first core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, wherein: the mass of the second turbine is a range of from 3% to 9% of the total dry mass of the gas turbine engine, and an overall pressure ratio of the gas turbine engine at cruise conditions is greater than 40 and less than 60. ### Claim 17 (depends on 16) The gas turbine engine according to claim 16, wherein the first and second turbine comprise at least one ceramic matrix composite component and the total mass of ceramic matrix composite in both the first and second turbines is in a range of from 1% to 15% of the total mass of the first and second turbines. ### Claim 18 (depends on 16) The gas turbine engine according to claim 16, wherein a specific thrust of the gas turbine engine is in an inclusive range of from 90 Nkg−1s to 100 Nkg−1s. ### Claim 19 (depends on 16) The gas turbine engine according to claim 16, wherein the gearbox has a gear ratio in a range of from 3 to 4. ### Claim 20 (depends on 16) The gas turbine engine according to claim 16, wherein the second turbine comprises two or three rotor blade rows.