# US 10,371,056 B2 · Multi-source turbine cooling air > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 10,371,056 B2 - Document ID: US10371056B2 - Status: Granted - Kind: B2 - Application number: 14/964,984 - Filed: 2015-12-10 - Published: 2017-06-15 - Granted: 2019-08-06 - Earliest claimed: 2015-12-10 - Assignees: United Technologies Corporation - Applicants: United Technologies Corporation - Inventors: Brian D. Merry; Gabriel L. Suciu; William K. Ackermann - Web page: https://patentgrep.com/patent/US10371056B2 ## CPC - F02C 7/185: {Cooling means for reducing the temperature of the cooling air or gas} - F01D 5/022: {with concentric rows of axial blades} - F01D 5/081: {Cooling fluid being directed on the side of the rotor disc or at the roots of the blades (F01D5/087 takes precedence)} - F01D 25/08: Cooling (of machines or engines in general F01P); Heating; Heat-insulation (of blade-carrying members, of blades F01D5/00) - F02C 6/08: the gas being bled from the gas-turbine compressor - F02C 7/12: Cooling of plants (of component parts, see the relevant subclasses, e.g. F01D; cooling of engines in general F01P) - 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) - F02K 3/06: with front fan - F05D 2220/32: in gas turbines - F05D 2260/213: by the provision of a heat exchanger within the cooling circuit - 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 gas turbine engine comprises a compressor section and a turbine section, with the turbine section having a first stage blade row and a downstream blade row. A higher pressure tap is tapped from a higher pressure first location in the compressor. A lower pressure tap is tapped from a lower pressure location in the compressor which is at a lower pressure than the first location. The higher pressure tap passes through a heat exchanger, and then is delivered to cool the first stage blade row in the turbine section. The lower pressure tap is delivered to at least partially cool the downstream blade row. ## Claims ### Claim 1 A gas turbine engine comprising: a compressor section and a turbine section, with said turbine section having a first stage blade row and a downstream blade row; a higher pressure tap tapped from a higher pressure first location in said compressor; a lower pressure tap tapped from a lower pressure location in said compressor which is at a lower pressure than said first location, said higher pressure tap passing through a heat exchanger, and then being delivered to cool said first stage blade row in said turbine section, and said lower pressure tap being delivered to at least partially cool said downstream blade row; and said higher pressure tap passing from said heat exchanger toward said turbine section, and split into a first path heading radially outwardly to cool an upstream end of said first stage blade row, and a second path moving radially inwardly of a hub mounting said first stage blade row and then moving radially outwardly to cool a downstream end of said first stage blade row. ### Claim 2 (depends on 1) The gas turbine engine as set forth in claim 1, wherein radially outwardly extending air from said higher pressure tap also cooling a vane mounted intermediate said first stage blade row and said downstream blade row. ### Claim 3 (depends on 2) The gas turbine engine as set forth in claim 2, wherein said radially outwardly extending air from said higher pressure tap also cooling an upstream end of said downstream blade row. ### Claim 4 (depends on 3) The gas turbine engine as set forth in claim 3, wherein said lower pressure tap passing radially inwardly of said first stage blade row, and axially beyond said downstream blade row and then radially outwardly to cool a downstream end of said downstream stage blade row. ### Claim 5 (depends on 4) The gas turbine engine as set forth in claim 4, wherein said downstream stage blade row is a second stage, and said first stage blade row and said second stage rotating together as a single rotor. ### Claim 6 (depends on 5) The gas turbine engine as set forth in claim 5, wherein a fan is positioned upstream of said compressor section and said fan delivering air into a bypass duct as propulsion air, and into said compressor section with a bypass ratio defined as the volume ratio of air delivered into said bypass duct compared to the volume of air delivered into said compressor, with said bypass ratio being greater than or equal to about 6.0. ### Claim 7 (depends on 6) The gas turbine engine as set forth in claim 6, wherein said bypass ratio is greater than or equal to about 10.0. ### Claim 8 (depends on 7) The gas turbine engine as set forth in claim 7, wherein a fan drive turbine rotor is positioned downstream of a turbine rotor including said first stage blade row and said downstream blade row, with said fan drive turbine driving said fan through a gear reduction. ### Claim 9 (depends on 8) The gas turbine engine as set forth in claim 8, wherein a gear ratio of said gear reduction is greater than or equal to about 2.3:1. ### Claim 10 (depends on 1) The gas turbine engine as set forth in claim 1, wherein radially outwardly extending air from said higher pressure tap also cooling a vane mounted intermediate said first stage blade row and said downstream blade row. ### Claim 11 (depends on 10) The gas turbine engine as set forth in claim 10, wherein said radially outwardly extending air from said higher pressure tap also cooling an upstream end of said downstream blade row. ### Claim 12 (depends on 11) The gas turbine engine as set forth in claim 11, wherein said lower pressure tap passing radially inwardly of said first stage blade row, and axially beyond said downstream blade row and then radially outwardly to cool a downstream end of said downstream stage blade row. ### Claim 13 (depends on 12) The gas turbine engine as set forth in claim 12, wherein said downstream stage blade row is a second stage, and said first stage blade row and said second stage rotating together as a single rotor. ### Claim 14 (depends on 1) The gas turbine engine as set forth in claim 1, wherein said lower pressure tap passing radially inwardly of said first stage blade row, and axially beyond said downstream blade row and then radially outwardly to cool a downstream end of said downstream stage blade row. ### Claim 15 (depends on 14) The gas turbine engine as set forth in claim 14, wherein said downstream stage blade row is a second stage, and said first stage blade row and said second stage rotating together as a single rotor. ### Claim 16 (depends on 15) The gas turbine engine as set forth in claim 15, wherein a fan is positioned upstream of said compressor section and said fan delivering air into a bypass duct as propulsion air, and into said compressor section with a bypass ratio defined as the volume ratio of air delivered into said bypass duct compared to the air delivered into said compressor, with said bypass ratio being greater than or equal to about 6.0. ### Claim 17 A gas turbine engine comprising: a compressor section and a turbine section, with said turbine section having a first stage blade row and a downstream blade row: a higher pressure tap tapped from a higher pressure first location in said compressor; a lower pressure tap tapped from a lower pressure location in said compressor which is at a lower pressure than said first location, said higher pressure tap passing through a heat exchanger, and then being delivered to cool said first stage blade row in said turbine section, and said lower pressure tap being delivered to at least partially cool said downstream blade row; and wherein a fan is positioned upstream of said compressor section and said fan delivering air into a bypass duct as propulsion air, and into said compressor section with a bypass ratio defined as the volume ratio of air delivered into said bypass duct compared to the air delivered into said compressor, with said bypass ratio being greater than or equal to about 6.0. ### Claim 18 (depends on 17) The gas turbine engine as set forth in claim 17, wherein said bypass ratio is greater than or equal to about 10.0. ### Claim 19 (depends on 17) The gas turbine engine as set forth in claim 17, wherein said higher pressure tap passing from said heat exchanger toward said turbine section, and split into a first path heading radially outwardly to cool an upstream end of said first stage blade row, and a second path moving radially inwardly of a hub mounting said first stage blade row and then moving radially outwardly to cool a downstream end of said first stage blade row. ### Claim 20 A gas turbine comprising: a compressor section and a turbine section, with said turbine section having a first stage blade row and a downstream blade row; a higher pressure tap tapped from a higher pressure first location in said compressor; a lower pressure tap tapped from a lower pressure location in said compressor which is at a lower pressure than said first location, said higher pressure tap passing through a heat exchanger, and then being delivered to cool said first stage blade row in said turbine section, and said lower pressure tap being delivered to at least partially cool said downstream blade row; said higher pressure tap passing from said heat exchanger toward said turbine section, and split into a first path heading radially outwardly to cool an upstream end of said first stage blade row, and a second path moving radially inwardly of a hub mounting said first stage blade row and then moving radially outwardly to cool a downstream end of said first stage blade row; wherein said lower pressure tap passing radially inwardly of said first stage blade row, and axially beyond said downstream blade row and then radially outward to cool a downstream end of said downstream stage blade row; wherein said downstream stage blade row is a second stage, and said first stage blade row and said second stage rotating together as a single rotor; and wherein a fan drive turbine rotor is positioned downstream of a turbine rotor including said first stage blade row and said downstream blade row, with said fan drive turbine driving said fan through a gear reduction. ### Claim 21 (depends on 20) The gas turbine engine as set forth in claim 20, wherein said higher pressure tap passing from said heat exchanger toward said turbine section, and split into a first path heading radially outwardly to cool an upstream end of said first stage blade row, and a second path moving radially inwardly of a hub mounting said first stage blade row and then moving radially outwardly to cool a downstream end of said first stage blade row.