General Electric Company

David Alan Frey

Kirk Douglas Gallier

Sean Robert Keith

Leslie Eugene Leeke, Jr

Alexander Martin Sener

Charles William Craig, Iii

Ronald Eugene Mc Rae, Jr

Jeffrey Douglas Rambo

Richard William Albrecht, Jr

Samir Armando Salamah

Brian Keith

Brandon ALlanson Reynolds

Daniel Patrick Kerns

Jamil Darnell Grant

Steven Robert Brassfield

Daniel Endecott Osgood

Herbert Chidsey Roberts

Dean Thomas Lenahan

Michael Anthony Ruthemeyer

Ronald Eugene McRae, Jr

Robert Ingram Ackerman

Lawrence Paul Timko

Daniel John Harris

Edward F Pietraszkiewicz

Mark E Ulliman

Carl D Wright

Brandon Allanson Reynolds

A variable guide vane assembly is provided for a turbine defining a core air flowpath. The variable guide vane assembly includes an airfoil band defining a flowpath surface and a cavity. The variable guide vane assembly further includes an airfoil including a first end extending at least partially into the cavity of the airfoil band and an opposite second end, the airfoil extending generally along an axis between the first end and the second end and being moveable generally about the axis relative to the airfoil band. The variable guide vane assembly further includes a sealing element operable to form a seal between the first end of the airfoil and the airfoil band.

Variable guide vane sealing

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.

Nozzle assembly with alternating inserted vanes for a turbine engine

Flow path assemblies and methods for forming such flow path assemblies for gas turbine engines are provided. For example, a flow path assembly for a gas turbine engine has a boundary structure, an airfoil, and a locking feature. The boundary structure and the airfoil are formed from a composite material. The boundary structure defines an opening and a cutout proximate the opening, and the airfoil is sized to fit within the opening of the boundary structure. The locking feature is received within the cutout defined by the boundary structure to interlock the airfoil with the boundary structure.

Flow path assemblies for gas turbine engines and assembly methods therefore

An apparatus and method relating an airfoil for a turbine engine with an outer wall bounding an interior and defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a span-wise direction. The airfoil includes a first cooling passage extending in the span-wise direction within the interior and a second cooling passage defining an impingement surface and located proximate the first cooling passage, an interior wall separating the first cooling passage from the second cooling passage, and at least one cooling hole passing through the interior wall.

Component for a turbine engine with a cooling hole

An ingestion restricting device is provided, which is particularly useful in a turbine engine. The ingestion restricting device can include a first component defining a first edge and including a plurality of ceramic matrix composite (CMC) plies forming outer, inner, and intermediate layers. The ingestion restricting device can also include a second component adjacent to the first component and defining a second edge. The first edge of the first component is opposite the second edge of the second component and defines a gap therebetween. Also, at least one ply of the plurality of CMC plies in the outer layer of the first component comprises an integral protrusion extending from the first edge a distance sufficient to traverse the gap and to overlap the second edge of the second component.

CMC ply overlap ingestion restrictor

An apparatus and method of sealing for an engine component between two surfaces arranged in spaced relationship to define a gap. A groove located is located in one of the two surface for receiving a seal having a body with a multi-faceted cross-sectional shape defining multiple seal facets. At least one of the multiple seal facets contacts one or both of the two surfaces.

Turbine engine with a seal

A component, such as for a turbine engine, can include an airfoil with an outer wall defining an exterior surface bounding an interior and defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a span-wise direction. The component can also include at least one cooling passage within the interior.

Turbine engine airfoil

CMC components having microchannels and methods for forming microchannels in CMC components are provided. For example, a method for forming microchannels in a CMC component comprises laying up a plurality of body plies for forming a body of the CMC component; laying up a microchannel ply on the plurality of body plies that has at least one void therein for forming at least one microchannel; laying up a cover ply on the microchannel ply to define an outer layer of the CMC component; and processing the laid up body plies, microchannel ply, and cover ply to form the CMC component. In another embodiment, the method comprises applying an additive matrix to the body plies to define at least one microchannel. In still other embodiments, the method comprises machining at least one microchannel in the plurality of body plies.

CMC components having microchannels and methods for forming microchannels in CMC components

Ceramic matrix composite (CMC) airfoils and methods for forming CMC airfoils are provided. In one embodiment, an airfoil is provided that includes opposite pressure and suction sides extending radially along a span and opposite leading and trailing edges extending radially along the span. The leading edge defines a forward end of the airfoil, and the trailing edge defines an aft end of the airfoil. A trailing edge portion is defined adjacent the trailing edge at the aft end, and a pocket is defined in and extends within the trailing edge portion. A heat pipe is received in the pocket. A method for forming an airfoil is provided that includes laying up a CMC material to form an airfoil preform assembly; processing the airfoil preform assembly; defining a pocket in a trailing edge portion of the airfoil; and inserting a heat pipe into the pocket.

Ceramic matrix composite component cooling

Airfoils, additively manufactured airfoils, and methods of manufacturing airfoils are provided. For example, an airfoil comprises opposite pressure and suction sides that extend axially from a leading edge to a trailing edge and radially spaced apart inner and outer ends. The airfoil also comprises an outer wall defining the pressure and suction sides and leading and trailing edges. A rib extends within the airfoil from the pressure side to the suction side of the outer wall and radially from the inner to the outer end. The airfoil further comprises a first pre-impingement chamber surrounded by a first post-impingement chamber and a first dividing wall segment separating the first pre-impingement and first post-impingement chambers and having a plurality of cooling holes defined therein. The outer wall, rib, and first dividing wall segment are integrally formed as a single monolithic component.

Airfoil with tunable cooling configuration

Flow path assemblies and methods for forming such flow path assemblies for gas turbine engines are provided. For example, a method for assembling an airfoil with a boundary structure to form a flow path assembly is provided. The method includes machining an opening into the boundary structure. The opening is sized to receive an airfoil or other component. The method also includes machining a cutout into the boundary structure proximate the opening. A locking feature is inserted into the cutout. When the airfoil is inserted into the opening, the locking feature interlocks the airfoil with the boundary structure. To seal the airfoil with the boundary structure, the airfoil is pressed against or into the boundary structure. When the airfoil is pressed, the locking feature is compressed such that a seal is formed between the airfoil and the boundary structure to seal the flow path assembly.

The present disclosure is directed to a shroud assembly for a gas turbine engine that includes a first shroud segment, a second shroud segment spaced from the first shroud segment, and a third shroud segment spaced from the first shroud segment. The first and the second shroud segments define a gap therebetween, and the first and the third shroud segments define a gap therebetween. A first seal portion is positioned at least partially on the first and the third shroud segments and at least partially covers the gap defined by the first and the third shroud segments. A second seal portion is positioned at least partially on the first and the second shroud segments and at least partially covers the gap defined by the first and the second shroud segments. A third seal portion extends between and couples the first and the second seal portions.

Growing community of inventors

West Chester, OH, United States of America

David Alan Frey