# US 12,110,804 B2 · Super-cooled ice impact protection for a 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 12,110,804 B2 - Document ID: US12110804B2 - Status: Granted - Kind: B2 - Application number: 18/524,490 - Filed: 2023-11-30 - Published: 2024-03-21 - Granted: 2024-10-08 - Earliest claimed: 2018-12-14 - Assignees: ROLLS-ROYCE plc - Applicants: ROLLS-ROYCE plc - Inventors: Ian J. Bousfield; Duncan A. Macdougall - Web page: https://patentgrep.com/patent/US12110804B2 ## CPC - F01D 5/141: {Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147)} - F01D 5/142: {of the blades of successive rotor or stator blade-rows} - F01D 25/02: De-icing means for engines having icing phenomena - F02C 7/05: having provisions for obviating the penetration of damaging objects or particles - 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) - F04D 29/324: {Blades} - F01D 21/10: responsive to unwanted deposits on blades, in working-fluid conduits or the like - F05D 2220/32: in gas turbines - F05D 2220/3217: for the first stage of a compressor or a low pressure compressor - F05D 2240/24: for turbines - F05D 2240/303: related to the leading edge of a rotor blade - F05D 2240/307: related to the tip of a rotor blade - F05D 2260/94: Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF] - 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) - 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 - Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION ## Abstract A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8. ## Claims ### Claim 1 A gas turbine engine comprising: a fan mounted to rotate about a main longitudinal axis, an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft, a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft, the reduction gearbox having a gear ratio of at least 3, wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, wherein a ratio of a tip maximum thickness of the first blades to a tip maximum thickness of the second blades is less than 5; and wherein the ratio of the tip maximum thickness of the first blades to the tip maximum thickness of the second blades is greater than 2.5. ### Claim 2 (depends on 1) The gas turbine engine of claim 1, wherein the ratio of the tip maximum thickness of the first blades to the tip maximum thickness of the second blades is less than 4. ### Claim 3 (depends on 1) The gas turbine engine of claim 1, wherein a ratio of the tip maximum thickness of the second blades to the tip maximum thickness of the first blades is in the range from 0.30 to 0.45. ### Claim 4 (depends on 1) The gas turbine engine of claim 1, wherein the gear ratio is in the range from 3.1 to 4.2. ### Claim 5 (depends on 1) The gas turbine engine of claim 1, wherein the tip maximum thickness of the first blades is greater than 2.7 mm. ### Claim 6 (depends on 1) The gas turbine engine of claim 1, wherein the tip maximum thickness of the first blades is greater than 3.0 mm. ### Claim 7 (depends on 1) The gas turbine engine of claim 1, wherein the tip maximum thickness of the first blades is greater than 3.5 mm. ### Claim 8 (depends on 1) The gas turbine engine of claim 1, wherein the tip maximum thickness of the first blades is less than 5 mm. ### Claim 9 A gas turbine engine comprising: a fan mounted to rotate about a main longitudinal axis, an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft, a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft, wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, wherein a ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 3, the maximum leading edge radius of the first blades being the maximum radius that is defined by the leading edge of the first blades in circumferential cross-section and the maximum leading edge radius of the second blades being the maximum radius that is defined by the leading edge of the second blades in circumferential cross-section. ### Claim 10 (depends on 9) The gas turbine engine of claim 9, wherein the ratio of the maximum leading edge radius of the first blades to the maximum leading edge radius of the second blades is greater than 4. ### Claim 11 (depends on 9) The gas turbine engine of claim 9, wherein the maximum leading edge radius of the first blades is greater than 0.4 mm. ### Claim 12 (depends on 9) The gas turbine engine of claim 9, wherein the maximum leading edge radius of the first blades is less than 0.9 mm. ### Claim 13 (depends on 9) The gas turbine engine of claim 9, wherein the fan has a rotational speed at cruise conditions less than 2500 rpm. ### Claim 14 (depends on 13) The gas turbine engine of claim 13, wherein the fan has a fan diameter in the range from 220 cm and 300 cm. ### Claim 15 A gas turbine engine comprising: a fan mounted to rotate about a main longitudinal axis, an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft, a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft, wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, wherein a ratio of a tip maximum thickness of the first blades to a maximum leading edge radius of the first blades is less than 6.5, the maximum leading edge radius being the maximum radius that is defined by the leading edge of the first blades in circumferential cross-section, and wherein an overall pressure ratio at cruise conditions is greater than 40. ### Claim 16 (depends on 15) The gas turbine engine of claim 15, wherein the maximum leading edge radius of the first blades is greater than 0.40 mm. ### Claim 17 (depends on 15) The gas turbine engine of claim 15, wherein a minimum leading edge radius of the first blades is greater than 0.15 mm, the minimum leading edge radius of the first blades being the minimum radius that is defined by the leading edge of the first blades in circumferential cross-section. ### Claim 18 (depends on 15) The gas turbine engine of claim 15, wherein a tip maximum thickness of the first blades is greater than 3.0 mm. ### Claim 19 (depends on 15) The gas turbine engine of claim 15, wherein a bypass ratio at cruise conditions is less than 12.