# US 10,371,047 B2 · Geared turbofan engine with targeted modular efficiency > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,371,047 B2 - Document ID: US10371047B2 - Status: Granted - Kind: B2 - Application number: 14/651,923 - Filed: 2014-09-24 - Published: 2016-09-15 - Granted: 2019-08-06 - Earliest claimed: 2013-10-16 - Assignees: United Technologies Corporation - Applicants: United Technologies Corporation - Inventors: Frederick M. Schwarz - Web page: https://patentgrep.com/patent/US10371047B2 ## CPC - F02C 3/107: with two or more rotors connected by power transmission - F01D 5/142: {of the blades of successive rotor or stator blade-rows} - F02K 3/06: with front fan - F05D 2220/327: to drive shrouded, high solidity propeller - F05D 2220/36: specially adapted for the fan of turbofan engines - F05D 2260/40311: of the epicyclical, planetary or differential type - 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) - F02K: JET-PROPULSION PLANTS (features of jet-propulsion plants common to gas-turbine plants, air intakes or fuel supply control of air-breathing jet-propulsion plants F02C7/00, F02C9/00) - 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 turbofan engine includes a fan section including a fan blade having a leading edge and hub to tip ratio of less than about 0.34 and greater than about 0.020 measured at the leading edge and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section. ## Claims ### Claim 1 A turbofan engine comprising; a fan section including a fan blade having a leading edge and hub to tip ratio of less than 0.34 and greater than 0.020 measured at the leading edge; a speed change mechanism with gear ratio greater than 2.6:1; a first compression section including a last blade trailing edge radial tip length that is greater than 67% of the radial tip length of a leading edge of a first stage of the first compression section; a second compression section including a last blade trailing edge radial tip length that is greater than 57% of a radial tip length of a leading edge of a first stage of the first compression section; and a turbine section including a fan drive turbine section and a second turbine section forward of the fan drive turbine section, wherein a mid-turbine frame is arranged between the fan drive turbine section and the second turbine section, the mid-turbine frame supporting a bearing system that supports rotation of at least one of the fan drive turbine and the second turbine section and includes a vane guiding core flow into the fan drive turbine section. ### Claim 2 (depends on 1) The turbofan engine as recited in claim 1, wherein the fan section provides a low fan pressure ratio less than 1.6 across a fan blade alone. ### Claim 3 (depends on 2) The turbofan engine as recited in claim 2, wherein the fan section provides a low fan pressure ratio between 1.45 and 1.20 across a fan blade alone. ### Claim 4 (depends on 3) The turbofan engine as recited in claim 3, wherein the fan section provides a bypass ratio greater than 8. ### Claim 5 (depends on 3) The turbofan engine as recited in claim 3, wherein the fan section provides a bypass ratio greater than 12. ### Claim 6 (depends on 1) The turbofan engine as recited in claim 1, wherein the fan section provides a bypass ratio greater than 8. ### Claim 7 (depends on 1) The turbofan engine as recited in claim 1, wherein the second turbine section includes at least two turbine stages forward of a first turbine blade of the fan drive turbine. ### Claim 8 (depends on 7) The turbofan engine as recited in claim 7, wherein the fan drive turbine includes at least three stages. ### Claim 9 (depends on 8) The turbofan engine as recited in claim 8, wherein the fan drive turbine is coupled to the drive the first compression section. ### Claim 10 (depends on 9) The turbofan engine as recited in claim 9, wherein at least one of the at least two turbine stages is coupled to drive the second compression section. ### Claim 11 (depends on 1) The turbofan engine as recited in claim 1, wherein the first compression section includes three stages and the second compression section includes eight stages. ### Claim 12 (depends on 1) The turbofan engine as recited in claim 1, wherein the fan includes a plurality of fan blades and the fan drive turbine includes a plurality of rotor stages and a ratio between the number of the plurality of fan blades and the number of the plurality of rotor stages is between 3.3 and 8.6. ### Claim 13 (depends on 12) The turbofan engine as recited in claim 12, wherein the plurality of fan blades is less than 20 fan blades. ### Claim 14 A turbofan engine comprising: a fan section providing a bypass ratio greater than 12; a speed change mechanism with gear ratio greater than 2.6 to 1; a first compression section including a last blade trailing edge tip length that is greater than 67% of the radial length of a first stage leading edge of the first compression section; and a second compression section including a last blade with a trailing edge tip that includes a radial length that is greater than 57% of a radial length of the first stage leading edge of the first compression section; and a turbine section including a fan drive turbine section and a second turbine section forward of the fan drive turbine section, wherein a mid-turbine frame is arranged between the fan drive turbine section and the second turbine section, the mid-turbine frame supporting a bearing system that supports rotation of at least one of the fan drive turbine and the second turbine section and includes a vane guiding core flow into the fan drive turbine section. ### Claim 15 (depends on 14) The turbofan engine as recited in claim 14, wherein the fan section includes a plurality of fan blades supported on a hub, wherein a leading edge of at least one of the fan blades includes a leading edge and a hub to tip ratio is less than 0.34 and greater than 0.020 measured at the leading edge. ### Claim 16 (depends on 15) The turbofan engine as recited in claim 15, wherein the fan section provides a fan pressure ratio between 1.45 and 1.20 across the plurality of fan blades alone. ### Claim 17 (depends on 14) The turbofan engine as recited in claim 14, wherein the second turbine section includes at least two turbine stages forward of a first turbine blade of the fan drive turbine. ### Claim 18 (depends on 17) The turbofan engine as recited in claim 17, wherein at least one of the at least two turbine stages is coupled to drive the second compression section. ### Claim 19 (depends on 14) The turbofan engine as recited in claim 14, wherein the first compression section includes three stages and the second compression section includes eight stages. ### Claim 20 (depends on 14) The turbofan engine as recited in claim 14, wherein the fan includes a plurality of fan blades and the fan drive turbine includes a plurality of rotor stages and a ratio between the number of the plurality of fan blades and the number of the plurality of rotor stages is between 3.3 and 8.6.