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公开(公告)号：US10830149B2
公开(公告)日：2020-11-10
申请号：US16059415
申请日：2018-08-09
Applicant: United Technologies Corporation
Inventor： Nathan Snape , Gabriel L. Suciu , Brian D. Merry
Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor is connected to be driven with at least one rotor in the main compressor section. A source of pressurized air is selectively sent to the cooling compressor to drive a rotor of the cooling compressor to rotate, and to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine. An intercooling system is also disclosed.
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公开(公告)号：US10808608B2
公开(公告)日：2020-10-20
申请号：US16027496
申请日：2018-07-05
Applicant: United Technologies Corporation
Inventor： Gabriel L. Suciu , Nathan Snape , Christopher M. Dye
Abstract: A gas turbine engine includes an engine static structure housing that includes a compressor section and a turbine section. A combustor section is arranged axially between the compressor section and the turbine section. A core nacelle encloses the engine static structure to provide a core compartment. An oil tank is arranged in the core compartment and is axially aligned with the compressor section. A heat exchanger is secured to the oil tank and arranged in the core compartment.
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公开(公告)号：US10738703B2
公开(公告)日：2020-08-11
申请号：US15928506
申请日：2018-03-22
Applicant: United Technologies Corporation
Inventor： Frederick M. Schwarz , Nathan Snape
Abstract: A gas turbine engine includes a plurality of rotating components housed within a main compressor section and a turbine section. A first tap is connected to the main compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A cooling air valve is configured to selectively block flow of cooling air across the heat exchanger. A cooling compressor is connected downstream of the heat exchanger. A shut off valve stops flow between the heat exchanger and the cooling compressor. A second tap is configured to deliver air at a second pressure which is higher than the first pressure. A mixing chamber is connected downstream of the cooling compressor and the second tap. The mixing chamber is configured to deliver air to at least one of the plurality of rotating components. A system stops flow between the cooling compressor and the plurality of rotating components. A controller is configured to modulate flow between the heat exchanger and the plurality of rotating components under certain power conditions of the gas turbine engine. The controller is programmed to control the cooling air valve, the shut off valve and the system such that flow is stopped between the heat exchanger and the cooling compressor only after the cooling compressor has been stopped.
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公开(公告)号：US10731560B2
公开(公告)日：2020-08-04
申请号：US15907942
申请日：2018-02-28
Applicant: United Technologies Corporation
Inventor： Nathan Snape , Gabriel L. Suciu , Brian Merry , Jesse M. Chandler , Frederick M. Schwarz
Abstract: A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap and configured to deliver air to an aircraft fuselage. A cooling compressor is connected downstream of the heat exchanger. A high pressure feed is configured to deliver air at a second pressure which is higher than the first pressure. The cooling compressor is configured to deliver air to at least one of the plurality of rotating components. A valve assembly that can select whether air from the first tap or air from the high pressure feed is delivered to the aircraft pneumatic system.
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公开(公告)号：US10458332B2
公开(公告)日：2019-10-29
申请号：US15407758
申请日：2017-01-17
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Frederick M. Schwarz , Nathan Snape
Abstract: A high pressure compressor has a downstream most end. A housing surrounds the compressor section and a turbine section. A low pressure turbine has a downstream most end. A first tap selectively taps high pressure cooling air from a location downstream of the downstream most end in the high pressure compressor and passes the high pressure cooling air through a heat exchanger. A second tap taps compressed air from a location upstream of the downstream most end in the high pressure compressor, and passes air over the heat exchanger, cooling the high pressure cooling air. A chamber is defined between the core engine housing and a nacelle airflow wall, and the second tap air flows through the chamber. The second tap air moves from the chamber into a core engine flow at a location downstream of the downstream most end of the low pressure turbine.
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Patent Agency Ranking
公开(公告)号：US10371055B2
公开(公告)日：2019-08-06
申请号：US14837009
申请日：2015-08-27
Applicant: United Technologies Corporation
Inventor： Nathan Snape , Gabriel L. Suciu , Brian D. Merry
Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor is connected to be driven with at least one rotor in the main compressor section. A source of pressurized air is selectively sent to the cooling compressor to drive a rotor of the cooling compressor to rotate, and to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine. An intercooling system is also disclosed.
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公开(公告)号：US20190063325A1
公开(公告)日：2019-02-28
申请号：US16059415
申请日：2018-08-09
Applicant: United Technologies Corporation
Inventor： Nathan Snape , Gabriel L. Suciu , Brian D. Merry
CPC classification number: F02C7/185 , F02C3/04 , F02C6/08 , F02C7/277 , F02C7/32 , F02K3/025 , F02K3/065 , F02K3/075 , F05D2260/213 , Y02T50/671 , Y02T50/676
Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor is connected to be driven with at least one rotor in the main compressor section. A source of pressurized air is selectively sent to the cooling compressor to drive a rotor of the cooling compressor to rotate, and to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine. An intercooling system is also disclosed.
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公开(公告)号：US20190003391A1
公开(公告)日：2019-01-03
申请号：US16111870
申请日：2018-08-24
Applicant: United Technologies Corporation
Inventor： Nathan Snape , James D. Hill , Gabriel L. Suciu , Brian Merry
IPC: F02C7/14
Abstract: An aircraft thermal management system includes a first fluid system containing a first fluid, a fluid loop containing a thermally neutral heat transfer fluid, a second fluid system containing a second fluid, a first heat exchanger configured to transfer heat from the first fluid to the thermally neutral heat transfer fluid, and a second heat exchanger configured to transfer heat from the thermally neutral heat transfer fluid to the second fluid. The fluid loop is configured to provide the thermally neutral heat transfer fluid to the first heat exchanger at a pressure that matches the pressure of the first fluid.
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公开(公告)号：US10072569B2
公开(公告)日：2018-09-11
申请号：US14806809
申请日：2015-07-23
Applicant: United Technologies Corporation
Inventor： Gabriel L. Suciu , Nathan Snape , Christopher M. Dye
CPC classification number: F02C3/04 , F01D25/12 , F01D25/18 , F02C7/06 , F02C7/12 , F02C7/14 , F02C7/32
Abstract: A gas turbine engine includes an engine static structure housing that includes a compressor section and a turbine section. A combustor section is arranged axially between the compressor section and the turbine section. A core nacelle encloses the engine static structure to provide a core compartment. An oil tank is arranged in the core compartment and is axially aligned with the compressor section. A heat exchanger is secured to the oil tank and arranged in the core compartment.
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公开(公告)号：US20180202363A1
公开(公告)日：2018-07-19
申请号：US15407758
申请日：2017-01-17
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Frederick M. Schwarz , Nathan Snape
Abstract: A high pressure compressor has a downstream most end. A housing surrounds the compressor section and a turbine section. A low pressure turbine has a downstream most end. A first tap selectively taps high pressure cooling air from a location downstream of the downstream most end in the high pressure compressor and passes the high pressure cooling air through a heat exchanger. A second tap taps compressed air from a location upstream of the downstream most end in the high pressure compressor, and passes air over the heat exchanger, cooling the high pressure cooling air. A chamber is defined between the core engine housing and a nacelle airflow wall, and the second tap air flows through the chamber. The second tap air moves from the chamber into a core engine flow at a location downstream of the downstream most end of the low pressure turbine.
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