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公开(公告)号：US11512599B1
公开(公告)日：2022-11-29
申请号：US17491828
申请日：2021-10-01
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Zachary Daniel Webster , Daniel Endecott Osgood , Kirk Douglas Gallier
Abstract: An apparatus and method for an engine component for a turbine engine having a working airflow separated into a cooling airflow and a combustion airflow. The engine component including a wall defining an interior and having an outer surface. A tip wall spanning first and second sides of the wall to close the interior. A tip rail extending from the tip wall and having an inner tip rail surface, which in combination with the tip wall, at least partially bounds a region defining a plenum. A rim formed in at least one of the outer surface and inner tip rail surface.
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公开(公告)号：US20220333490A1
公开(公告)日：2022-10-20
申请号：US17231263
申请日：2021-04-15
Applicant: General Electric Company
Inventor： Daniel Endecott Osgood , Kirk Douglas Gallier , Gregory Terrence Garay , Zachary Daniel Webster , Daniel Lee Durstock , Ricardo Caraballo
IPC: F01D5/18
Abstract: An airfoil for a turbine engine having a working airflow separated into a cooling airflow and a combustion airflow, the airfoil comprising a wall defining an interior and having an outer surface over which flows the combustion airflow, the outer surface defining a first side and a second side extending between a leading edge and a trailing edge to define a chord-wise direction; at least one cooling conduit located within the interior and fluidly coupled to the cooling airflow. A primary cooling passage having at least one inlet fluidly coupled to the at least one cooling conduit, a primary outlet on the outer surface. A passage connecting the at least one inlet to the primary outlet, the passage separated into a first portion and a second portion. The primary outlet spaced from the trailing edge a predetermined distance.
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公开(公告)号：US11268394B2
公开(公告)日：2022-03-08
申请号：US16817742
申请日：2020-03-13
Applicant: General Electric Company
Abstract: A nozzle assembly for a gas turbine engine and methods for assembling a nozzle assembly are provided. In one example aspect, the nozzle assembly includes an outer wall and an inner wall radially spaced from the outer wall. The outer wall defines a plurality of mounting openings spaced circumferentially from one another. The inner wall defines a plurality of mounting openings spaced circumferentially from one another. The mounting openings defined by the inner wall are positioned circumferentially between adjacent mounting openings defined by the outer wall. The nozzle assembly includes vanes that are inserted through the mounting openings of the outer wall in a radially inward direction and vanes that are inserted through the mounting openings of the inner wall in a radially outward direction in an alternating manner.
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公开(公告)号：US10934858B2
公开(公告)日：2021-03-02
申请号：US16365105
申请日：2019-03-26
Applicant: General Electric Company
Inventor： Eric Joseph Schroeder , Paul Hadley Vitt , Timothy John Swenson , Kirk Douglas Gallier , Aspi Rustom Wadia
Abstract: A turbine blade is described herein, the turbine blade including a blade root, a blade tip, and an airfoil extending between the blade root and the blade tip. The airfoil has opposite pressure and suction sides extending between a forward leading edge and an aft trailing edge of the airfoil, and a maximum thickness located between the leading edge and the trailing edge. The blade tip includes a winglet extending laterally outward from at least one of the pressure side and the suction side from a leading point between the leading edge and the maximum thickness aftward to a trailing point between the maximum thickness and the trailing edge.
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公开(公告)号：US20210017907A1
公开(公告)日：2021-01-21
申请号：US17060162
申请日：2020-10-01
Applicant: General Electric Company
Abstract: Features and methods for modulating a flow of cooling fluid to gas turbine engine components are provided. In one embodiment, an airfoil is provided having a flow modulation insert for modulating a flow of cooling fluid received in a cavity of a body of the airfoil. In another embodiment, a shroud is provided comprising a cooling channel for a flow of cooling fluid and an insert that varies in position to modulate the flow of cooling fluid through the cooling channel. In yet another embodiment, a method for operating a gas turbine engine having a cooling circuit for cooling one or more components of the gas turbine engine comprises increasing power provided to the engine and decreasing power provided to the engine to modulate a position of a flow modulation insert located in the cooling circuit and thereby modulate the flow of cooling fluid through the cooling circuit.
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Patent Agency Ranking
公开(公告)号：US10605095B2
公开(公告)日：2020-03-31
申请号：US15151860
申请日：2016-05-11
Applicant: General Electric Company
Inventor： Kirk Douglas Gallier , Darrell Glenn Senile , John Calhoun
Abstract: Airfoils for gas turbine engines are provided. In one embodiment, an airfoil formed from a ceramic matrix composite material includes opposite pressure and suction sides extending radially along a span and defining an outer surface of the airfoil. The airfoil also includes opposite leading and trailing edges extending radially along the span. The pressure and suction sides extend axially between the leading and trailing edges. The leading edge defines a forward end of the airfoil, and the trailing edge defining an aft end of the airfoil. Further, the airfoil includes a trailing edge portion defined adjacent the trailing edge at the aft end of the airfoil; a plenum defined within the airfoil forward of the trailing edge portion; and a cooling passage defined within the trailing edge portion proximate the suction side. Methods for forming airfoils for gas turbine engines also are provided.
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公开(公告)号：US20190218918A1
公开(公告)日：2019-07-18
申请号：US16365105
申请日：2019-03-26
Applicant: General Electric Company
Inventor： Eric Joseph Schroeder , Paul Hadley Vitt , Timothy John Swenson , Kirk Douglas Gallier , Aspi Rustom Wadia
CPC classification number: F01D5/20 , F01D5/141 , F01D5/145 , F01D5/147 , F01D5/187 , F01D9/04 , F02C3/06 , F02C7/18 , F05D2220/36 , F05D2230/10 , F05D2240/307 , F05D2240/35 , F05D2260/20 , Y02T50/673 , Y02T50/676
Abstract: A turbine blade is described herein, the turbine blade including a blade root, a blade tip, and an airfoil extending between the blade root and the blade tip. The airfoil has opposite pressure and suction sides extending between a forward leading edge and an aft trailing edge of the airfoil, and a maximum thickness located between the leading edge and the trailing edge. The blade tip includes a winglet extending laterally outward from at least one of the pressure side and the suction side from a leading point between the leading edge and the maximum thickness aftward to a trailing point between the maximum thickness and the trailing edge.
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公开(公告)号：US20170342842A1
公开(公告)日：2017-11-30
申请号：US15130013
申请日：2016-04-15
Applicant: General Electric Company
Inventor： Kirk Douglas Gallier
CPC classification number: F01D5/186 , B23P15/04 , B23P2700/06 , B28B1/30 , B28B11/12 , F01D5/187 , F01D5/282 , F01D5/284 , F01D9/041 , F05D2220/32 , F05D2230/50 , F05D2240/128 , F05D2240/30 , F05D2260/202 , F05D2260/203 , F05D2300/6033 , Y02T50/672 , Y02T50/673 , Y02T50/676
Abstract: An airfoil for a gas turbine engine is provided that includes a first portion formed from a first plurality of plies of a ceramic matrix composite material and defining an inner surface of the airfoil, as well as a second portion formed from a second plurality of plies of a ceramic matrix composite material and defining an outer surface of the airfoil. The first portion and the second portion define a non-line of sight cooling aperture extending from the inner surface to the outer surface of the airfoil. In one embodiment, a surface angle that is less than 45° is defined between a second aperture and the outer surface. A method for forming an airfoil for a gas turbine engine also is provided.
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公开(公告)号：US20170167275A1
公开(公告)日：2017-06-15
申请号：US14967069
申请日：2015-12-11
Applicant: General Electric Company
Inventor： Eric Joseph Schroeder , Paul Hadley Vitt , Timothy John Swenson , Kirk Douglas Gallier , Aspi Rustom Wadia
CPC classification number: F01D5/20 , F01D5/141 , F01D5/145 , F01D5/147 , F01D5/187 , F01D9/04 , F02C3/06 , F02C7/18 , F05D2220/36 , F05D2230/10 , F05D2240/307 , F05D2240/35 , F05D2260/20 , Y02T50/673 , Y02T50/676
Abstract: A turbine blade is described herein, the turbine blade including a blade root, a blade tip, and an airfoil extending between the blade root and the blade tip. The airfoil has opposite pressure and suction sides extending between a forward leading edge and an aft trailing edge of the airfoil, and a maximum thickness located between the leading edge and the trailing edge. The blade tip includes a winglet extending laterally outward from at least one of the pressure side and the suction side from a leading point between the leading edge and the maximum thickness aftward to a trailing point between the maximum thickness and the trailing edge.
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公开(公告)号：US11994039B2
公开(公告)日：2024-05-28
申请号：US17711405
申请日：2022-04-01
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Kirk Douglas Gallier , Charles William Craig, III
CPC classification number: F01D25/246 , F01D9/042 , F01D11/005 , F05D2300/6033 , Y02T50/60
Abstract: A gas turbine engine having a rotor blade stage with a plurality of circumferentially spaced rotors, a nozzle stage adjacent the rotor blade stage and including an outer nozzle end, at least one stop, and a shroud. The shroud having a forward end positioned radially outward from the circumferentially spaced rotor blades, and an aft end axially aft of the forward end. The at least one stop confronting at least a portion of the outer nozzle end.
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