# US 10,502,135 B2 · Buffer system for communicating one or more buffer supply airs throughout 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 10,502,135 B2 - Document ID: US10502135B2 - Status: Granted - Kind: B2 - Application number: 13/366,445 - Filed: 2012-02-06 - Published: 2013-08-01 - Granted: 2019-12-10 - Earliest claimed: 2012-01-31 - Assignees: United Technologies Corporation - Applicants: William K. Ackermann; Philip S. Stripinis - Inventors: William K. Ackermann; Philip S. Stripinis - Web page: https://patentgrep.com/patent/US10502135B2 ## CPC - F02C 7/047: Heating to prevent icing - B64D 15/04: Hot gas application - F01D 25/125: {of bearings} - F02C 7/06: Arrangements of bearings (bearings F16C); Lubricating ({of turbo machines F01D25/18; of machines or} engines in general F01M) - F02C 7/141: of working fluid - 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) - B64D: EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT - F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D) - Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION ## Abstract A gas turbine engine includes a buffer system. The buffer system can include a first bleed air supply and a conditioning device that conditions the first bleed air supply to render the buffer supply air at an acceptable temperature to anti-ice the hardware of said fan section. ## Claims ### Claim 1 A gas turbine engine, comprising: a buffer system, wherein said buffer system includes: a first bleed air supply; a conditioning device that conditions said first bleed air supply to render a buffer supply air at an acceptable temperature for pressurizing at least one bearing structure and at least a portion of anti-icing hardware of a fan section of the gas turbine engine, wherein separate buffer supply airs are communicated to said at least one bearing structure and said at least a portion of anti-icing hardware; a sensor configured to detect a flight condition of the gas turbine engine; and a controller configured to control operation of said conditioning device and command communication of said buffer supply air to said anti-icing hardware in response to detection of said flight condition, wherein a portion of said buffer supply air is communicated in a downstream direction to condition other portions of the gas turbine engine separate from said at least one bearing structure and said hardware. ### Claim 2 (depends on 1) The gas turbine engine as recited in claim 1, wherein said hardware includes a a spinner. ### Claim 3 (depends on 2) The gas turbine engine as recited in claim 2, wherein said buffer system communicates said buffer supply air to an interior portion of said spinner to anti-ice said spinner. ### Claim 4 (depends on 3) The gas turbine engine as recited in claim 3, comprising a passageway that communicates said buffer supply air from said buffer system to said interior portion. ### Claim 5 (depends on 1) The gas turbine engine as recited in claim 1, wherein said hardware includes a static nosecone. ### Claim 6 (depends on 1) The gas turbine engine as recited in claim 1, wherein said first bleed air supply is sourced from a stage of one of the fan section and a compressor section of the gas turbine engine. ### Claim 7 (depends on 1) The gas turbine engine as recited in claim 1, wherein said conditioning device is a heat exchanger. ### Claim 8 (depends on 1) The gas turbine engine as recited in claim 1, wherein the gas turbine engine is a geared turbofan engine. ### Claim 9 (depends on 1) The gas turbine engine as recited in claim 1, wherein said first bleed air supply is sourced from the fan section of the gas turbine engine. ### Claim 10 (depends on 1) The gas turbine engine as recited in claim 1, wherein said first bleed air supply is sourced from a high pressure compressor of the gas turbine engine. ### Claim 11 (depends on 10) The gas turbine engine as recited in claim 10, wherein said first bleed air supply is sourced from a middle stage of said high pressure compressor. ### Claim 12 A gas turbine engine, comprising: a fan section; a compressor section in fluid communication with said fan section; a combustor in fluid communication with said compressor section; a turbine section in fluid communication with said combustor; at least one shaft that interconnects at least a portion of said compressor section and said turbine section; a bearing structure that supports said at least one shaft; and a buffer system that communicates a first buffer supply air to at least a spinner of said fan section to anti-ice said spinner and a second buffer supply air separate from said first buffer supply air to at least a portion of said bearing structure, wherein said buffer system includes a first bleed air supply, a conditioning device that conditions said first bleed air supply to render said first buffer supply air and said second buffer supply air at an acceptable temperature, a sensor configured to detect a flight condition of the gas turbine engine, and a controller configured to control operation of said conditioning device to command said second buffer supply air at an adequate differential pressure for sealing said bearing structure and configured to command communication of said first buffer supply air in response to detection of said flight condition, wherein a portion of either of said first buffer supply air and said second buffer supply air is communicated in a downstream direction to condition other portions of the gas turbine engine separate from said bearing structure and said spinner. ### Claim 13 (depends on 12) The gas turbine engine as recited in claim 12, comprising a geared architecture that drives said fan section at a lower speed than said compressor section. ### Claim 14 (depends on 12) The gas turbine engine as recited in claim 12, wherein the gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6). ### Claim 15 (depends on 12) The gas turbine engine as recited in claim 12, wherein the gas turbine engine includes a low Fan Pressure Ratio of less than 1.45. ### Claim 16 (depends on 12) The gas turbine engine as recited in claim 12, wherein said compressor section includes a high pressure compressor that is a source of said first bleed air supply that is communicated to said buffer system.