General Electric Patente

A turbine engine (10) including a combustor (34) with a combustor liner (38, 138) having dilution openings (62, 76, 78) and a geometry that changes along an axial direction. The combustor further having a baffle (70, 170, 270, 370) surrounding a combustor liner (38, 138) defining a combustion chamber (46) of the combustor. A method for controlling nitrogen oxides within the combustor, including injecting compressed air into the annular combustion chamber (46) through any of the dilution openings (62, 76, 78) described herein.

A turbine engine (10) can combust fuel to drive a turbine, which drives the engine. A fuel nozzle (100), (200), (300), (400), (500) can supply fuel for combustion or ignition of the fuel. The fuel nozzle (100), (200), (300), (400), (500) can include a central body (110) defining a first fuel passage (111). An annular first wall (121) circumferentially surrounds the central body (110) and is spaced therefrom to define a first airflow passage (131) therebetween. A set of first swirler vanes (104) is circumferentially spaced about the central body (110) and disposed within the first airflow passage (131). The fuel nozzle (100), (200), (300), (400), (500) can supply a mixture of fuel and air for combustion.

A gas turbine engine (10), comprising: a compressor section (12), a combustion section (14), and a turbine section (16) in a serial flow arrangement, with the combustion section comprising: a combustor liner (40) that at least partially defines a combustion chamber (50); and a fuel nozzle assembly (48), comprising: a liquid fuel supply (34) to supply a liquid fuel; a hydrogen fuel supply (36) to supply a gaseous hydrogen fuel; and a fuel nozzle body (49) fluidly coupled with the liquid fuel supply (34) and the hydrogen fuel supply (36) to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber (50).

An engine (10) can utilize a combustor (36) to combust fuel to drive the engine (10). A fuel nozzle assembly (130) can supply fuel to the combustor (36) for combustion or ignition of the fuel. The fuel nozzle assembly (130) can include a swirler (134) and a fuel nozzle (132) to supply a mixture of fuel and air for combustion. The fuel nozzle assembly (130) can be configured to increase lateral provision of fuels to reduce flame scrubbing on combustor liners (40, 42) for the combustor (36).

A gas turbine engine (10), comprising: a compressor section (12), combustion section (14), and turbine section (16) in serial flow arrangement, with the combustion section (14) comprising: a combustor liner (40) at least partially defining a combustion chamber (50); a wall (46) coupled to the combustor liner (40); a first fuel supply (34) to supply a first fuel; a gaseous fuel supply (36) to supply a gaseous hydrogen fuel; and a fuel nozzle (100) assembly coupled to the wall (46) and fluidly coupled to the first fuel supply (34) and the gaseous fuel supply (36), the fuel nozzle (100) assembly comprising: a main mixer (102); and a fuel nozzle (100) disposed such that the main mixer (102) is disposed at least partially around the fuel nozzle (100).

A gas turbine engine (10), comprising: a compressor section (12), a combustion section (14), and a turbine section (16) in a serial flow arrangement, with the combustion section (14) comprising: a combustor liner (40) that at least partially defines a combustion chamber (50); and a fuel nozzle assembly (248, 448, 648), comprising: a liquid fuel supply (34) to supply a liquid fuel; a hydrogen fuel supply (36) to supply a gaseous hydrogen fuel; and a fuel nozzle body (249, 449, 649) defining a fuel nozzle centerline, the fuel nozzle body (249, 449, 649) fluidly coupled with the liquid fuel supply (34) and the hydrogen fuel supply (36) to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber (50).

A turbine engine (10), comprising a compressor section (12), a combustion section (14), and a turbine section (16) in a serial flow arrangement, with the combustion section (14) further comprising a combustor liner (40) at least partially defining a combustion chamber (50), and a gaseous fuel nozzle assembly (48). The gaseous fuel nozzle assembly (48) comprises a set of mixing tubes (100), a set of fuel jets (114) and an air passage (120), the air passage (120) including a set of air passage outlets (126), with each air passage outlet of the set of air passage outlets (126) tangentially, fluidly coupled to a corresponding mixing tube air inlet (106).

A gas turbine engine (10), comprising a compressor section (12), combustion section (14), and turbine section (16) in serial flow arrangement, with the combustion section (14) comprising: a combustor liner (40) at least partially defining a combustion chamber (50); and a fuel injector assembly (48, 100, 200) comprising a mixing tube (38, 110, 210) having a mixing tube body (112, 212) defining a mixing channel (114, 214). The mixing tube body (112, 212) can have a set of fuel passages (132, 232) terminating in fuel orifices (136, 236) fluidly coupled to the mixing channel (114, 214). A set air flow passages (142, 242) terminating in air outlets (144, 244) can be fluidly coupled to the mixing channel (114, 214). At least some of the air outlets (144, 244) circumscribe a corresponding fuel orifice (136, 236).

An airfoil assembly (100) comprising: an airfoil portion (102) having an outer wall (108) extending between a root (114) and a tip (116) in a spanwise direction (Sd), and between a leading edge (110) and a trailing edge (112), the outer wall (108) defining an interior (118) of the airfoil portion (102); and a spar (106) coupled to the airfoil portion (102) and extending into the interior (118), the spar (106) comprising a spar centerline (122), a spar body (124) defining a perimeter surface (126), and a plurality of pockets (128) provided within the spar body (124), each pocket of the plurality of pockets (128) forming a respective cavity (130) within the spar body (124).

A gas turbine engine including a fan section, a compressor section, combustion section. An airfoil is provided in one of the fan section, the compressor section, or the turbine section, and includes an outer wall defining an interior. A first portion is positioned within the interior and at least one disengagement feature is positioned within the first portion to disengage or separate within the first portion in response to a force exceeding a predetermined threshold.

A fuel injector (100) for a gas turbine engine (10) includes a compression section (12), combustion section (14), and turbine section (16) in serial flow arrangement. The fuel injector (100) arranges in the combustion section (14) and includes an inner nozzle (104) defining a fuel injector axis (114) and an outer nozzle (106) in annular arrangement about the inner nozzle (104). The outer nozzle (106) at least partially defines a mixing region (156). An inner air passage (150) and an outer air passage (152) exhaust to the mixing region (156). A fuel body (180) positions between the inner air passage (150) and the outer air passage (152) and has an outer fuel passage (154) exhausting to the mixing region (156).

A ceramic matrix composite (CMC) component (10; 120; 200) comprises a ceramic core (12; 122; 202) having one or more interlocking elements (140; 140A; 140B; 170; 220) such as protrusions (142), locking elements (150; 150A; 150B; 170) or apertures (222) formed therein. A CMC facesheet (14; 16; 124; 126; 204; 206) is positioned against a surface (18; 20; 128; 130; 208; 210) of the ceramic core and is in mechanical engagement with the interlocking elements to secure the facesheet to the core, with at least a portion of the facesheet disposed between at least one interlocking element and the surface. The CMC facesheet may comprise a fibre preform (32; 166; 180; 230) with an interface coating (42) and a ceramic matrix (52; 168; 182; 232), and may be bonded to the ceramic core via a bonding layer (62) formed by reaction between the interface coating and a reaction material at the interface, such as silicon from the ceramic core or a reaction layer.

A method for forming a ceramic matrix composite ("CMC") component includes applying a CMC fiber preform to a ceramic core where the CMC fiber preform is formed from a plurality of CMC fiber tows. An interface coating is formed on at least a portion of the plurality of CMC fiber tows before or after the CMC fiber preform is formed from the plurality of CMC fiber tows. A ceramic matrix is formed on the CMC fiber preform with the CMC fiber preform applied to the ceramic core. The CMC fiber preform and the ceramic matrix form a CMC facesheet. A reaction material is located at or adjacent an interface of the CMC facesheet and the ceramic core. The CMC facesheet and the ceramic core are thermally processed to react the interface coating with the reaction material to form a bonding layer at the interface.

A propulsion system (100, 300, 500, 700, 900, 1100, 1300) and a method of operating the propulsion system (100, 300, 500, 700, 900, 1100, 1300). The propulsion system (100, 300, 500, 700, 900, 1100, 1300) includes a turbine engine (10, 910, 1310) including a turbo-engine (16, 716), a first propulsor (38a, 738a) including a first propulsor shaft (45a, 745a), and a first gearbox assembly (46a, 746a) having a gear assembly (50). The propulsion system (100, 300, 500, 700, 900, 1100, 1300) includes a second propulsor (38b, 738b) that is remote from the turbine engine (10, 910, 1310) and includes a second gearbox assembly (46b, 746b) including a gear assembly (50). The propulsion system (100, 300, 500, 700, 900, 1100, 1300) further includes a lubrication system (600, 800, 1000, 1400) including a first gearbox lubrication system (602a, 802a), an engine lubrication system (620), and a second gearbox lubrication system (602b, 802b). The first gearbox lubrication system (602a, 802a) is disposed within the first gearbox assembly (46a, 746a) and supplies lubricant to the gear assembly (50). The engine lubrication system (620, 1620) supplies lubricant to the engine bearings (37). The first gearbox lubrication system (602a, 802a) is fluidly separate from the engine lubrication system (620, 1620). The second gearbox lubrication system (602b, 802b) is disposed within the second gearbox assembly (46b, 746b) and supplies lubricant to the gear assembly (50).

A turbine engine (100) has a compression section, a combustion section (102), a turbine section in serial flow arrangement and defining a working airflow (Fw) path (182). The combustion section (102) has a circumferential casing (104), a combustor (106), and a fuel nozzle (130). The circumferential casing (104) defines an interior (108). The turbine engine (100) includes a deswirler assembly (134). The deswirler assembly (134) couples the working airflow (Fw) path (182) of the compression section to the interior (108). The turbine engine (100) includes a fuel supply system (140).

A mixer for blending fuel and air in a combustor of a turbine engine. The mixer includes a central body having a central passageway and a central axis. The mixer includes a plurality of tubes positioned radially around the central axis and circumferentially around a periphery of the mixer. Each of the tubes of the mixer includes opposed openings and a tangential opening. Each of the tubes of the mixer includes a cylindrical interior mixing passage configured to receive air flow from the opposed openings and the tangential opening and a fuel flow. The opposed openings are configured to spread the fuel flow laterally and the tangential opening is configured to spread the fuel flow tangentially.

A combustion section (102, 202, 302) for a turbine engine (100, 200, 300) comprises a circumferential casing (104) defining an interior (108), and a combustor (106, 206, 306) arranged within the interior (108). The combustor includes an annular liner (116, 216, 316) that defines a combustion chamber (124). Multiple fuel nozzles (130) open into the combustion chamber (124). A fuel supply system (142, 242, 342) is provided within the interior (108) and includes a fuel manifold (148, 248, 348) and multiple fuel line branches (146, 246, 346). Each fuel line branch (146, 246, 346) extends between a respective portion of the fuel manifold (148, 248, 348) and a corresponding portion of the combustor (106, 206, 306) or a respective one of the fuel nozzles (130), thereby supplying fuel to the combustion chamber (124).

A turbine engine (10) comprising a compression section (12), combustion section (14), and turbine section (16) is serial flow arrangement, with the combustion section (14) including an injector (100) for providing a mixture of fuel and air for combustion. The injector (100) includes a body (114), an inner nozzle (102) provided within the body (114) and defining an injector axis (112), and an outer nozzle (104) in annular arrangement about the inner nozzle. (102) A first fuel passage (120) fluidly couples to the inner nozzle (102) and a second fuel passage (164) fluidly couples to the outer nozzle (104). A first set of air conduits (174) are in annular arrangement about the body (114) interior of the outer nozzle (104) and a second set of air conduits (176) are in annular arrangement about the body (114) exterior of the outer nozzle (104).

An insertion tool (102) with an implement gripping mechanism is provided. The insertion tool (102) includes an elongated section (106) defining an implement path (108) extending from a proximal end (110) to a distal end (112), a handle (104) coupled to the proximal end (110) of the elongated section (106), a grip mechanism (132) coupled to the handle (104) and configured to selectively secure an implement (116) inserted into the implement path (108) through the handle (104), and a grip actuator (134) operable to cause the grip mechanism (132) to secure the implement (116) to the handle (104).

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A method (100) for protecting a wind turbine (10) from an extreme change in wind direction includes receiving a wind direction and/or a wind speed at the wind turbine (10). When a change in the wind direction or the wind speed exceeds a predetermined threshold, the method (100) includes determining a margin to stall and/or zero lift of the at least one rotor blade (22) of the wind turbine (10) as a function of an angle of attack or change in the angle of attack at a blade span location of at least one rotor blade (22) of the wind turbine (10). The method (100) also includes implementing a corrective action for the wind turbine (10) (without shutting down the wind turbine (10)) when the margin to stall and/or zero lift exceeds a predetermined value so as to avoid stall and/or negative lift on the at least one rotor blade during operation of the wind turbine (10).

An aircraft engine (10) assembly comprising: a gas turbine engine (10) having a high pressure compressor (22), a high pressure turbine (28), a high pressure shaft (34) coupling the high pressure compressor (22) with the high pressure turbine (28), a low pressure turbine, and a low pressure shaft (36) coupled to the low pressure turbine, the high pressure turbine (28) located forward of the high pressure compressor (22), and the low pressure turbine located on a forward end of the gas turbine engine (10); an intake channel (54) of the gas turbine engine (10) configured to receive an incoming flow of air and form an intake flow of air, the intake channel (54) configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine (10), the incoming flow direction reverse of the first axial direction; and an electric machine (74, 74a, 74b) coupled with the low pressure shaft (36) and located on a side of the high pressure compressor (22) opposite of the high pressure turbine (28) and proximate the intake channel (54), the electric machine (74, 74a, 74b) in heat exchange communication with the intake flow of air such that the electric machine (74, 74a, 74b) transfers heat to the incoming flow of air within the intake channel (54) when the electric machine (74, 74a, 74b) is operated.

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