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公开(公告)号：US20240401483A1
公开(公告)日：2024-12-05
申请号：US18326684
申请日：2023-05-31
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Kirk Douglas Gallier , Zachary Daniel Webster , Brian Kenneth Corsetti
IPC: F01D5/18
Abstract: An airfoil assembly for a turbine engine which generates a hot fluid flow and provides a cooling fluid flow. The airfoil assembly includes a platform with radially spaced upper and lower walls, the platform extending between a platform leading edge and a platform trailing edge to define an axial direction and extending between a pair of slash faces, an airfoil having an airfoil wall extending radially between a root and a tip to define a span length, the platform wall extending from the airfoil wall proximate the root. A fillet extends between the heated surface and the airfoil wall and defines at least a portion of the root. The airfoil assembly further includes a set of film holes.
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公开(公告)号：US11988109B2
公开(公告)日：2024-05-21
申请号：US17973976
申请日：2022-10-26
Applicant: GENERAL ELECTRIC COMPANY
Inventor： Zachary Daniel Webster , Daniel Endecott Osgood , Kirk Douglas Gallier
CPC classification number: F01D5/187 , F01D5/087 , F05D2240/307 , F05D2260/20 , F05D2260/201 , F05D2260/232
Abstract: An apparatus and method for an engine component for a turbine engine. The engine component having an outer wall defining an interior and extending between a root and a tip to define a radial direction, a tip wall spanning the first side and second sides to close the interior at the tip. A tip rail extending from the tip wall and having an inner tip rail surface, an outer tip rail surface extending from at least one of the first or the second side, and radially terminating in an upper tip rail surface connecting the inner tip rail surface and the outer tip rail surface. A tip rim formed in at least one of the outer surface or the inner tip rail surface and spaced from the upper tip rail surface in the radial direction, and multiple cooling passages formed in the outer wall and fluidly coupling the at least one cooling conduit to the tip rim at corresponding passage outlets.
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公开(公告)号：US11846207B2
公开(公告)日：2023-12-19
申请号：US17685523
申请日：2022-03-03
Applicant: General Electric Company
CPC classification number: F01D9/042 , F01D5/18 , F01D25/246 , F02C7/28 , F01D5/282 , F01D5/284 , F05D2230/60 , F05D2300/6033
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：US20230399953A1
公开(公告)日：2023-12-14
申请号：US17836023
申请日：2022-06-09
Applicant: General Electric Company
Inventor： Jonathan Michael Rausch , Zachary Daniel Webster , Kevin Robert Feldmann , Andrew David Perry , Kirk Douglas Gallier , Daniel Endecott Osgood
CPC classification number: F01D5/186 , F01D25/12 , F05D2220/32 , F05D2240/30 , F05D2260/202
Abstract: A blade for a turbine engine with a wall separating a cooling fluid flow and a hot gas fluid flow and having a heated surface along which the hot gas fluid flow flows and a cooled surface facing the cooling fluid flow. A plurality of cooling holes each having a passage extending between an inlet at the cooled surface and an outlet at the heated surface. The outlet extending between an upstream end and a downstream end with respect to the hot gas fluid flow to define a distance, the passage defining a centerline forming a first angle (θ) with the heated surface.
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公开(公告)号：US11352887B2
公开(公告)日：2022-06-07
申请号：US16527390
申请日：2019-07-31
Applicant: General Electric Company
Inventor： Kirk Douglas Gallier
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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Patent Agency Ranking
6.发明申请公开(公告)号：US20220090504A1
公开(公告)日：2022-03-24
申请号：US17030595
申请日：2020-09-24
Applicant: General Electric Company
Inventor： Matthew Mark Weaver , Kirk Douglas Gallier
Abstract: A rotor blade for a gas turbine engine includes a structural member formed from a metallic material. The structural member, in turn, includes a base portion, a spar, and a tip cap, with the base portion at least partially forming a root of the rotor blade and a shank of the rotor blade. Furthermore, the structural member includes a fairing formed from a composite material. The fairing is, in turn, coupled to the structural member such that the fairing forms at least a portion of an airfoil of the rotor blade and a platform of the rotor blade. Additionally, the fairing including a first fairing panel and a second fairing panel in contact with the first fairing panel at a first split line and a second split line.
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公开(公告)号：US11221140B2
公开(公告)日：2022-01-11
申请号：US16907013
申请日：2020-06-19
Applicant: General Electric Company
Inventor： Jinjie Shi , Stephen Gerard Schadewald , Jason Paul Hoppa , Kirk Douglas Gallier , Christopher Edward Wolfe , Robert Proctor
Abstract: A seal assembly to seal a gas turbine hot gas path flow at an interface of a combustor liner and a downstream component, such as a stage one turbine nozzle, in a gas turbine. The seal assembly including a piston ring seal housing, defining a cavity, and a piston ring disposed within the cavity. The piston ring disposed circumferentially about the combustor liner. The piston ring is responsive to a regulated pressure to secure sealing engagement of the piston ring and outer surface of the combustor liner. The seal assembly includes at least one of one or more sectional through-slots, bumps or channel features to provide for a flow therethrough of a high-pressure (Phigh) bypass airflow exiting a compressor to the cavity. The high-pressure (Phigh) bypass airflow exerting a radial force on the piston ring.
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公开(公告)号：US10577949B2
公开(公告)日：2020-03-03
申请号：US15182651
申请日：2016-06-15
Applicant: General Electric Company
Inventor： Shawn Michael Pearson , Kirk Douglas Gallier , Ronald Scott Bunker
IPC: F01D5/28
Abstract: A component for a gas turbine engine includes a first region formed substantially of a first CMC material, wherein first region defines a first thermal conductivity. The component further includes a second region formed substantially of a second CMC material, wherein the second region defines a second thermal conductivity. Further, the component defines a thickness and the first region is positioned adjacent to the second region along the thickness, wherein the first thermal conductivity is different than the second thermal conductivity to alert a thermal profile of the component.
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公开(公告)号：US10450897B2
公开(公告)日：2019-10-22
申请号：US15212337
申请日：2016-07-18
Applicant: General Electric Company
Inventor： Kirk Douglas Gallier , Charles William Craig, III
Abstract: Shrouds and shroud segments for gas turbine engines are provided. In one embodiment, a shroud segment for a gas turbine engine having a rotor blade stage and a nozzle stage is provided. The shroud segment comprises a forward end defining an outer wall of the rotor blade stage and an aft end defining an outer wall of the nozzle stage. The aft end defines at least a portion of an opening therethrough for receipt of a nozzle, and the forward end and the aft end form a single, continuous component. In another embodiment, a gas turbine engine is provided, having a shroud with a forward end positioned near a leading edge of a plurality of rotor blades of a rotor blade stage and an aft end positioned near a trailing edge of a plurality of nozzles of a nozzle stage. Methods of assembling a gas turbine engine also are provided.
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公开(公告)号：US20170328217A1
公开(公告)日：2017-11-16
申请号：US15151860
申请日：2016-05-11
Applicant: General Electric Company
Inventor： Kirk Douglas Gallier , Darrell Glenn Senile , John Calhoun
CPC classification number: F01D5/187 , F01D5/147 , F01D5/18 , F01D5/284 , F01D5/288 , F01D9/041 , F01D25/005 , F01D25/12 , F05D2220/32 , F05D2230/50 , F05D2230/90 , F05D2260/202 , F05D2260/204 , F05D2300/6033 , F05D2300/611 , Y02T50/672 , Y02T50/673 , Y02T50/676
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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