# US 10,502,163 B2 · Geared turbofan arrangement with core split power ratio > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,502,163 B2 - Document ID: US10502163B2 - Status: Granted - Kind: B2 - Application number: 14/964,607 - Filed: 2015-12-10 - Published: 2016-05-12 - Granted: 2019-12-10 - Earliest claimed: 2013-11-01 - Assignees: United Technologies Corporation - Applicants: United Technologies Corporation - Inventors: Daniel Bernard Kupratis; Frederick M. Schwarz - Web page: https://patentgrep.com/patent/US10502163B2 ## CPC - F02K 3/06: with front fan - F02C 3/107: with two or more rotors connected by power transmission - 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) - F05D 2260/20: Heat transfer, e.g. cooling - F05D 2260/40311: of the epicyclical, planetary or differential type - F05D 2300/40: Organic materials - F05D 2300/603: Composites; e.g. fibre-reinforced - F05D 2300/606: Directionally-solidified crystalline structures - Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft - 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) - 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 - Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION ## Abstract A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including at least a first compressor section and a second compressor section. A power ratio is provided by the combination of the first compressor section and the second compressor section. A method of design a gas turbine engine is also disclosed. ## Claims ### Claim 1 A gas turbine engine comprising: a fan section including a plurality of fan blades; a compressor section, including at least a first compressor section and a second compressor section; a turbine section including at least one turbine to drive said second compressor section and a fan drive turbine to drive at least an epicyclic gear arrangement to drive said fan section and said first compressor section, said turbine section including at least two turbine stages upstream of said fan drive turbine; and wherein a power ratio is provided by the combination of said first compressor section and said second compressor section, with said power ratio being provided by a second power input to said second compressor section divided by a first power input to said first compressor section, said power ratio being equal to, or greater than, 1.0 and less than, or equal to, 1.4. ### Claim 2 (depends on 1) The gas turbine engine as set forth in claim 1, wherein said fan drive turbine includes six or fewer stages. ### Claim 3 (depends on 2) The gas turbine engine as set forth in claim 2, wherein said power ratio is less than 1.27. ### Claim 4 (depends on 2) The gas turbine engine as set forth in claim 2, wherein said fan section is configured to deliver a portion of air into said compressor section, and a portion of air into a bypass duct, and wherein a bypass ratio, which is defined as a volume of air passing to said bypass duct compared to a volume of air passing into said compressor section, is equal to or greater than 10. ### Claim 5 (depends on 2) The gas turbine engine as set forth in claim 2, wherein a gear ratio of said gear arrangement is greater than 2.6. ### Claim 6 (depends on 2) The gas turbine engine as set forth in claim 2, wherein an overall pressure ratio being provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to or greater than 36. ### Claim 7 (depends on 1) The gas turbine engine as set forth in claim 1, wherein said fan section defines a fan pressure ratio less than 1.50 across the fan blades alone, said first turbine section is configured to rotate at least 2.6 times faster than said fan section, and an overall pressure ratio being provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to or greater than 36. ### Claim 8 (depends on 1) The gas turbine engine as set forth in claim 1, wherein said first compressor section includes 3 or more stages and said second compressor section includes between 8 and 15 stages. ### Claim 9 (depends on 8) The gas turbine engine as recited in claim 8, wherein said fan drive turbine includes three 3 to six 6 stages. ### Claim 10 (depends on 9) The gas turbine engine as recited in claim 9, wherein said fan drive turbine defines a fan drive turbine pressure ratio that is greater than five. ### Claim 11 (depends on 7) The gas turbine engine as set forth in claim 7, wherein said power ratio is less than 1.27, said overall pressure ratio is equal to or greater than 50, and said gear arrangement is a planetary gear system. ### Claim 12 A gas turbine engine comprising: a fan section including a plurality of fan blades; a compressor section, including at least a first compressor section and a second compressor section, said first compressor section including 3 or more stages and said second compressor section including 6 or more stages; a turbine section including at least one turbine to drive said second compressor section and a fan drive turbine to drive at least a gear arrangement to drive said fan section, said turbine section including at least 2 turbine stages upstream of said fan drive turbine; and wherein a power ratio is provided by the combination of said first compressor section and said second compressor section, with said power ratio being provided by a second power input to said second compressor section divided by a first power input to said first compressor section, said power ratio being equal to or greater than 1.0; and wherein an overall pressure ratio is provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to, or greater than, 36. ### Claim 13 (depends on 12) The gas turbine engine as set forth in claim 12, wherein said fan section defines a fan pressure ratio less than 1.50 across the fan blades alone, a gear ratio of said gear arrangement is greater than 2.6, said turbine section includes at least 2 turbine stages upstream of said fan drive turbine, and said fan drive turbine includes 3 to 6 stages. ### Claim 14 (depends on 12) The gas turbine engine as set forth in claim 12, wherein said power ratio is between 1.0 and 1.4. ### Claim 15 (depends on 12) The gas turbine engine as set forth in claim 12, wherein said fan section defines a fan pressure ratio less than 1.45 across the fan blades alone. ### Claim 16 (depends on 15) The gas turbine engine as recited in claim 15, wherein said fan drive turbine defines a fan drive turbine pressure ratio that is greater than five (5). ### Claim 17 (depends on 16) The gas turbine engine as set forth in claim 16, wherein said power ratio is less than 1.27. ### Claim 18 (depends on 17) The gas turbine engine as recited in claim 17, wherein said fan section is configured to deliver a portion of air into said compressor section, and a portion of air into a bypass duct, and wherein a bypass ratio, which is defined as a volume of air passing to said bypass duct compared to a volume of air passing into said compressor section, is equal to or greater than 12. ### Claim 19 (depends on 13) The gas turbine engine as set forth in claim 13, wherein said gear arrangement is a planetary gear system, and said overall pressure ratio is equal to or greater than 60. ### Claim 20 A gas turbine engine comprising: a fan section; a compressor section, including at least a first compressor section and a second compressor section; a turbine section including at least one turbine to drive said second compressor section and a fan drive turbine to drive at least a gear arrangement to drive said fan section; wherein a power ratio is provided by the combination of said first compressor section and said second compressor section, with said power ratio being provided by a second power input to said second compressor section divided by a first power input to said first compressor section, said power ratio being equal to or less than 1.4; and wherein an overall pressure ratio is provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to or greater than 36. ### Claim 21 (depends on 20) The gas turbine engine as set forth in claim 20, wherein said first compressor section includes 3 or more stages, said second compressor section includes 6 or more stages, said turbine section includes at least 2 turbine stages upstream of said fan drive turbine, and said fan drive turbine includes three 3 to six 6 stages. ### Claim 22 (depends on 20) The gas turbine engine as set forth in claim 20, wherein said power ratio is greater than or equal to 1.0. ### Claim 23 (depends on 20) The gas turbine engine as set forth in claim 20, wherein said fan section is configured to deliver a portion of air into said compressor section, and a portion of air into a bypass duct, and wherein a bypass ratio, which is defined as a volume of air passing to said bypass duct compared to a volume of air passing into said compressor section, is equal to or greater than 10. ### Claim 24 (depends on 20) The gas turbine engine as set forth in claim 20, wherein said overall pressure ratio is equal to or greater than 60. ### Claim 25 A method of designing a gas turbine engine comprising: providing a fan section; providing a compressor section, including a first compressor section and a second compressor section; providing a turbine section including at least one turbine to drive said second compressor section and a fan drive turbine to drive said fan section via a gear arrangement; wherein a power ratio is provided by the combination of said first compressor section and said second compressor section, with said power ratio being provided by a second power input to said second compressor section divided by a first power input to said first compressor section, said power ratio is equal to or less than 1.4 at a predetermined design target; and wherein an overall pressure ratio is provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to or greater than 36 at the predetermined design target. ### Claim 26 (depends on 25) The method as set forth in claim 25, wherein said predetermined design target is defined at sea level and at a static, full-rated takeoff power condition. ### Claim 27 (depends on 25) The method as set forth in claim 25, wherein said predetermined design target is defined at a cruise condition. ### Claim 28 (depends on 25) The method as set forth in claim 25, wherein said power ratio is greater than or equal to 1.0. ### Claim 29 (depends on 25) The method as set forth in claim 25, wherein said first compressor section includes 3 or more stages, said second compressor section includes 6 or more stages, said second compressor section is positioned downstream of said first compressor section, said fan drive turbine includes three 3 to six 6 stages, and said at least one turbine includes fewer stages than said fan drive turbine. ### Claim 30 (depends on 29) The method as set forth in claim 29, wherein said power ratio is less than 1.27, and said overall pressure ratio is equal to or greater than 60.