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公开(公告)号：US11073088B2
公开(公告)日：2021-07-27
申请号：US16280548
申请日：2019-02-20
Applicant: General Electric Company
Inventor： Alan Roy Stuart , Thomas Ory Moniz
IPC: F02C7/36
Abstract: A gas turbine engine includes a frame member extending generally along an axial direction and defining a plurality of slots spaced along a circumferential direction; and a gearbox including a ring gear, a plurality of planet gears, and a sun gear, the plurality of planet gears each positioned at least partially in a respective slot of the plurality of slots defined by the frame member such that the frame member extends through the gearbox.
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公开(公告)号：US10723471B2
公开(公告)日：2020-07-28
申请号：US15622290
申请日：2017-06-14
Applicant: General Electric Company
Inventor： Alan Roy Stuart , Thomas Ory Moniz , Jeffrey Donald Clements , Joseph George Rose
Abstract: A system for mounting an engine to an aircraft includes an engine forward mount angled toward the forward end of the engine at a first angle. At least two thrust links extend between an engine aft mount to a link support connection at a second angle. The engine aft mount is angled toward the forward end of the engine at a third angle. A projection of a load vector of the engine forward mount onto a vertical plane extending through the axis of rotation of the engine and a projection of a load vector of each of the at least two thrust links onto the vertical plane intersect the axis of rotation of the engine within a first vertical plane segment extending between a forward end of a nose of a fan assembly and forward of a forward mount interface.
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公开(公告)号：US10655537B2
公开(公告)日：2020-05-19
申请号：US15412197
申请日：2017-01-23
Applicant: General Electric Company
Inventor： Alan Roy Stuart , Jeffrey Donald Clements , Richard Schmidt , Thomas Ory Moniz
IPC: F01D1/24 , F02C3/04 , F02C7/36 , F02K3/06 , F02C3/067 , F02K3/072 , F01D5/03 , F01D5/06 , F01D5/14
Abstract: The present disclosure is directed to a method of operating a gas turbine engine with an interdigitated turbine section. The engine includes a fan rotor, an intermediate pressure compressor, a high pressure compressor, a combustion section, and a turbine section in serial flow arrangement. The turbine section includes, in serial flow arrangement, a first stage of a low speed turbine rotor, a high speed turbine rotor, a second stage of the low speed turbine rotor, an intermediate speed turbine rotor, and one or more additional stages of the low speed turbine rotor. The low speed turbine rotor is coupled to the fan rotor via a low pressure shaft. The intermediate speed turbine rotor is coupled to the intermediate pressure compressor via an intermediate pressure shaft. The high speed turbine rotor is coupled to the high pressure compressor via a high pressure shaft. The method includes rotating the low speed turbine rotor in a first direction along the circumferential direction; rotating the high speed turbine rotor in a second direction opposite of the first direction along the circumferential direction; and rotating the intermediate speed turbine rotor in the second direction.
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公开(公告)号：US10519860B2
公开(公告)日：2019-12-31
申请号：US15451498
申请日：2017-03-07
Applicant: General Electric Company
Inventor： Thomas Ory Moniz , Alan Roy Stuart , Jeffrey Donald Clements , Brandon Wayne Miller , Darek Tomasz Zatorski
IPC: F01D25/16 , F02C7/06 , F02C3/067 , F01D5/06 , F01D1/24 , F01D11/00 , F02C7/28 , F02K3/072 , F01D25/18
Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction. The gas turbine engine defines an upstream end and a downstream end along the longitudinal direction and includes a turbine frame defined around the axial centerline. The turbine frame includes a first bearing surface, a second bearing surface, and a third bearing surface. The first bearing surface corresponds to a first turbine rotor, the second bearing surface corresponds to a second turbine rotor, and the third bearing surface corresponds to a third turbine rotor, and each turbine rotor is independently rotatable.
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公开(公告)号：US20190218913A1
公开(公告)日：2019-07-18
申请号：US15870313
申请日：2018-01-12
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Arnab Sen , Jeffrey Douglas Rambo , Rajesh Kumar , Bhaskar Nanda Mondal , Alan Roy Stuart , Robert Proctor , Christopher Charles Glynn
Abstract: An apparatus and method for cooling a portion of a turbine engine comprising an outer casing defining an axial centerline, a turbine section through which a flow of combustion gasses flows in a forward to aft direction, an outer drum located between the outer casing and the turbine section defining an annular cavity therebetween. A set of seals extends between the outer casing and outer drum to define at least one cooled cavity.
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Patent Agency Ranking
公开(公告)号：US20180340469A1
公开(公告)日：2018-11-29
申请号：US15605164
申请日：2017-05-25
Applicant: General Electric Company
Inventor： Alan Roy Stuart , Jeffrey Donald Clements , Richard Schmidt , Thomas Ory Moniz
Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor comprising a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor comprising a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section; a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor. The first turbine bearing defines an outer air bearing along an outer diameter of the first turbine bearing and adjacent to the hub of the low speed turbine rotor. The first turbine bearing defines an inner air bearing along an inner diameter of the first turbine bearing and adjacent to the HP shaft.
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公开(公告)号：US20180320632A1
公开(公告)日：2018-11-08
申请号：US15427340
申请日：2017-02-08
Applicant: General Electric Company
Inventor： Jeffrey Donald Clements , Darek Tomasz Zatorski , Alan Roy Stuart
Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section and a gear assembly within or downstream of the turbine section. The turbine section includes a first rotating component and a second rotating component along the longitudinal direction. The first rotating component includes one or more connecting airfoils coupled to a radially extended rotor, and the second rotating component includes an inner shroud defining a plurality of inner shroud airfoils extended outward of the inner shroud along the radial direction. The second rotating component is coupled to a second shaft connected to an input accessory of the gear assembly, and the first rotating component is coupled to an output accessory of the gear assembly. The output accessory rotates the first rotating component about the axial centerline at a first speed and wherein the second rotating component rotates about the axial centerline at a second speed.
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公开(公告)号：US20180298784A1
公开(公告)日：2018-10-18
申请号：US15485515
申请日：2017-04-12
Applicant: General Electric Company
Inventor： Thomas Ory Moniz , Alan Roy Stuart , Jeffrey Donald Clements , Brandon Wayne Miller , Darek Tomasz Zatorski , Gert Johannes van der Merwe , Joel Francis Kirk , Richard Wesling
CPC classification number: F01D25/183 , F01D25/162
Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and wherein the gas turbine engine defines an upstream end and a downstream end along the longitudinal direction, and wherein the gas turbine engine defines a core flowpath extended generally along the longitudinal direction. The gas turbine engine includes a turbine frame defined around the axial centerline, the turbine frame comprising a first bearing surface disposed inward along the radial direction. The gas turbine engine further includes a turbine rotor assembly including a bearing assembly coupled to the first bearing surface of the turbine frame and the turbine rotor assembly. The turbine rotor assembly further includes a first turbine rotor disposed upstream of the turbine frame and a second turbine rotor disposed downstream of the turbine frame. The first turbine rotor and the second turbine rotor are rotatable together about the axial centerline.
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公开(公告)号：US20180209335A1
公开(公告)日：2018-07-26
申请号：US15412197
申请日：2017-01-23
Applicant: General Electric Company
Inventor： Alan Roy Stuart , Jeffrey Donald Clements , Richard Schmidt , Thomas Ory Moniz
CPC classification number: F02C3/04 , F02C3/067 , F02C7/36 , F02K3/06 , F02K3/072 , F05D2220/32 , F05D2260/40311
Abstract: The present disclosure is directed to a method of operating a gas turbine engine with an interdigitated turbine section. The engine includes a fan rotor, an intermediate pressure compressor, a high pressure compressor, a combustion section, and a turbine section in serial flow arrangement. The turbine section includes, in serial flow arrangement, a first stage of a low speed turbine rotor, a high speed turbine rotor, a second stage of the low speed turbine rotor, an intermediate speed turbine rotor, and one or more additional stages of the low speed turbine rotor. The low speed turbine rotor is coupled to the fan rotor via a low pressure shaft. The intermediate speed turbine rotor is coupled to the intermediate pressure compressor via an intermediate pressure shaft. The high speed turbine rotor is coupled to the high pressure compressor via a high pressure shaft. The method includes rotating the low speed turbine rotor in a first direction along the circumferential direction; rotating the high speed turbine rotor in a second direction opposite of the first direction along the circumferential direction; and rotating the intermediate speed turbine rotor in the second direction.
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公开(公告)号：US20180128206A1
公开(公告)日：2018-05-10
申请号：US15347301
申请日：2016-11-09
Applicant: General Electric Company
Inventor： Richard David Cedar , David Baker Riddle , Alan Roy Stuart , Jeffrey Donald Clements
IPC: F02K1/72
CPC classification number: F02K1/72 , F02K3/06 , F05D2220/327 , F05D2260/904 , Y02T50/671
Abstract: A turbofan engine is provided having a fan and a nacelle assembly enclosing the fan. The nacelle assembly includes a thrust reverser system having one or more cascade segments configured to translate at least partially along an axial direction of the turbofan engine. The turbofan engine further includes a core operable with the fan and at least partially enclosed by the nacelle. The core includes a turbine section having a low pressure turbine defining an exit diameter. A ratio of the exit diameter of the low pressure turbine to a fan diameter of the fan is less than 0.5, providing for a more compact turbofan engine.
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