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公开(公告)号：US20240413704A1
公开(公告)日：2024-12-12
申请号：US18207522
申请日：2023-06-08
Applicant: Raytheon Technologies Corporation
Inventor： Paul R. Hanrahan , Daniel B. Kupratis
Abstract: An aircraft system is provided that includes a first turbine engine, a second turbine engine and an electrical system. The electrical system includes a clutch, a first electric machine and a second electric machine electrically coupled to the first electric machine. The first electric machine is mechanically coupled to a first rotating assembly of the first turbine engine. The second electric machine is mechanically coupled to a second rotating assembly of the second turbine engine through the clutch.
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公开(公告)号：US12129802B2
公开(公告)日：2024-10-29
申请号：US17903681
申请日：2022-09-06
Applicant: Raytheon Technologies Corporation
Inventor： Paul R. Hanrahan
IPC: F02C7/36
CPC classification number: F02C7/36 , F05D2220/323 , F05D2260/40311 , F05D2260/902
Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a compressor section, a combustor section, a turbine section and a flowpath extending sequentially through the compressor section, the combustor section and the turbine section. The assembly also includes a turbine rotor, a propulsor rotor and an auxiliary turbine. The turbine rotor is within the turbine section. The turbine rotor is configured to rotatably drive the propulsor rotor. The auxiliary turbine includes an auxiliary turbine rotor. The auxiliary turbine rotor is configured to rotatably drive the propulsor rotor with the turbine rotor. The auxiliary turbine is configured to receive bleed gas from the flowpath.
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公开(公告)号：US11976598B2
公开(公告)日：2024-05-07
申请号：US17662204
申请日：2022-05-05
Applicant: Raytheon Technologies Corporation
Inventor： Paul R. Hanrahan , Daniel B. Kupratis
Abstract: A gas turbine engine includes a first spool associated with a primary combustor, a second spool associated with a secondary combustor, and a third spool, each spool including a compressor and a turbine mounted to a shaft. A transmission and accessory gearing are enclosed within an housing of an accessory gearbox. The transmission rotationally couples the third spool to the second spool and accessory gearing. A method of operating the gas turbine engine includes supplying a first fuel flow rate to the primary combustor and supplying a second fuel flow rate to the secondary combustor within an intermediate speed range of the gas turbine engine.
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公开(公告)号：US20230383709A1
公开(公告)日：2023-11-30
申请号：US18202709
申请日：2023-05-26
Applicant: Raytheon Technologies Corporation
Inventor： Paul R. Hanrahan , Roxanne M. Bochar
CPC classification number: F02K1/002 , F02K3/04 , F05D2220/323 , F05D2240/128
Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a bladed rotor and a thrust vectoring exhaust nozzle. The bladed rotor is rotatable about an axis. The thrust vectoring exhaust nozzle is configured to direct gas propelled by the bladed rotor out of the aircraft propulsion system along a first direction during a first mode and along a second direction during a second mode. The first direction is parallel with the axis or angularly offset from the axis by no more than five degrees. The second direction is angularly offset from the axis by at least seventy-five degrees. The thrust vectoring exhaust nozzle has a first exit area during the first mode and a second exit area during the second mode that is greater than the first exit area.
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公开(公告)号：US20230382549A1
公开(公告)日：2023-11-30
申请号：US18202734
申请日：2023-05-26
Applicant: Raytheon Technologies Corporation
Inventor： Paul R. Hanrahan
CPC classification number: B64D35/04 , F02C7/36 , B64D35/02 , B64D2027/026
Abstract: An assembly is provided for an aircraft. This aircraft assembly includes a carrier, a first gear system, a second gear system and an idler. The first gear system includes a first sun gear, a first ring gear and a plurality of first intermediate gears. Each of the first intermediate gears is between and is meshed with the first sun gear and the first ring gear. Each of the first intermediate gears is rotatably mounted to the carrier. The second gear system includes a second sun gear, a second ring gear and a plurality of second intermediate gears. Each of the second intermediate gears is between and is meshed with the second sun gear and the second ring gear. Each of the second intermediate gears is rotatably mounted to the carrier. The idler gear couples the first ring gear to the second ring gear.
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Patent Agency Ranking
公开(公告)号：US20230382522A1
公开(公告)日：2023-11-30
申请号：US18202733
申请日：2023-05-26
Applicant: Raytheon Technologies Corporation
Inventor： Paul R. Hanrahan
IPC: B64C29/00
CPC classification number: B64C29/0025
Abstract: A propulsion system is provided for an aircraft. This aircraft propulsion system includes a gas turbine engine core, a first propulsor rotor and a second propulsor rotor. The gas turbine engine core includes a compressor section, a combustor section, a turbine section and a rotating structure. The rotating structure includes a turbine rotor within the turbine section. The first propulsor rotor is rotatably driven by the rotating structure during a first mode and a second mode. The first propulsor rotor includes a plurality of variable pitch blades. The variable pitch blades include a first blade configured to pivot between a thrust position and an idle position. The first blade is in the thrust position during the first mode. The first blade is in the idle position during the second mode. The second propulsor rotor is rotatably driven by the rotating structure during the second mode.
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公开(公告)号：US11753990B2
公开(公告)日：2023-09-12
申请号：US17864982
申请日：2022-07-14
Applicant: RAYTHEON TECHNOLOGIES CORPORATION
Inventor： Daniel Bernard Kupratis , Paul R. Hanrahan , Arthur M. Salve, Jr.
Abstract: A method of distributing power within a gas turbine engine is disclosed. In various embodiments, the method includes driving a high pressure turbine having a first stage and a second stage with an exhaust stream from a combustor, the first stage connected to a high pressure turbine first stage spool and the second stage connected to a high pressure turbine second stage spool; driving a high pressure compressor connected to a high pressure compressor spool via a differential system, the differential system having a first stage input gear connected to the high pressure turbine first stage spool, a second stage input gear connected to the high pressure turbine second stage spool and an output gear assembly connected to the high pressure compressor spool; and selectively applying an auxiliary input power into at least one of the high pressure compressor spool and the high pressure turbine.
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公开(公告)号：US20230159175A1
公开(公告)日：2023-05-25
申请号：US17535200
申请日：2021-11-24
Applicant: Raytheon Technologies Corporation
Inventor： Jon E. Sobanski , Paul R. Hanrahan , Joseph B. Staubach , Jill Klinowski
Abstract: A propulsor includes a propulsor body and a prop assembly in rotational communication with the propulsor body. The prop assembly includes a plurality of prop blades configured for rotation about an axial centerline of the propulsor. The plurality of prop blades are rotatable between a deployed position and a stowed position. The propulsor further includes at least one linkage having a first linkage end and a second linkage end. The first linkage end of the at least one linkage is rotatably mounted to the propulsor body and the second linkage end is configured to be rotatably mounted to an aircraft body. The propulsor further includes a first motor coupled to the at least one linkage and configured to rotate the at least one linkage relative to the propulsor body between a first rotational position and a second rotational position.
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公开(公告)号：US11639690B1
公开(公告)日：2023-05-02
申请号：US17662202
申请日：2022-05-05
Applicant: Raytheon Technologies Corporation
Inventor： Daniel B. Kupratis , Paul R. Hanrahan
IPC: F02C7/36 , F02C6/02 , F02C6/08 , F02C9/20 , F02C9/42 , F02K3/075 , F02K3/12 , F02C9/18 , F02C7/32 , F02K3/077
Abstract: A gas turbine engine includes a first spool of a primary flow path and a second spool of a secondary flow path. The second spool is nonconcentric with the first spool. The second spool includes a boost compressor and a load compressor in fluid communication with an inlet plenum. An inlet duct assembly and an outlet duct assembly place the secondary flow path in communication with the primary flow path. The gas turbine includes a controller operable to vary open areas of variable inlet guide vanes to control a flow division between the boost compressor and the load compressor.
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公开(公告)号：US20220341327A1
公开(公告)日：2022-10-27
申请号：US17238793
申请日：2021-04-23
Applicant: Raytheon Technologies Corporation
Inventor： Paul R. Hanrahan
IPC: F01D5/08
Abstract: Gas turbine engines and rotor arms thereof are described. The gas turbine engines include a first disk, a second disk, and a rotor arm arranged between and connecting the first disk to the second disk, wherein a cavity is defined at least between the rotor arm and the first disk. The rotor arm includes a radial portion having an inner diameter end and an outer diameter end, an axial portion having a first end and a second end, wherein the first end of the axial portion is connected to the outer diameter end of the radial portion, at least one entrance flow path defined within the radial portion extending from the inner diameter end to the outer diameter end, and at least one exit aperture arranged proximate the second end of the axial portion.
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