A turbine engine with a tip shroud and method for shaping the tip shroud where at least one pair of airfoils including a first and second airfoil each having an outer wall bounding an interior. Each airfoil extending between a pressure side and a suction side to define a circumferential direction and extending between a leading edge and a trailing edge to define an axial direction. Each airfoil extending between a root and a tip to define a radial direction. The first and second airfoils are circumferentially spaced to define an inlet between the leading edges and an outlet between the trailing edges where each airfoil is coupled to an inner platform at the root. A tip shroud circumscribing the airfoil operably coupled to the tip extending in the axial direction between axially spaced first and second planes where the tip shroud includes at least one scalloped portion.

The turbine exhaust case can have an outer shroud, an inner shroud internal to the outer shroud, an annular exhaust path between the outer shroud and the inner shroud, and a plurality of struts each having a length extending across the annular exhaust path from a radially inner end to a radially outer end, the struts circumferentially interspaced from one another, the struts each having a leading edge and a trailing edge, a stiff connection between the radially inner end and the inner shroud, and a point connection between the radially outer end and the outer shroud.

A component for use in a gas turbine engine is made from ceramic materials. The component is made from a core preform containing ceramic reinforcement fibers. A plurality of preform tows are applied on at least a portion of the core preform and extend along at least a portion of the core preform to provide an outermost surface. The core preform and the plurality of preform tows are infiltrated with ceramic matrix material to form a ceramic matrix composite component.

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.

A nozzle assembly is disclosed, including a CMC nozzle shell, a nozzle spar, and an endwall. The CMC nozzle shell includes a CMC composition and an interior cavity. The nozzle spar is partially disposed within the interior cavity and includes a metallic composition, a cross-sectional conformation, a plurality of spacers protruding from the cross-sectional conformation, the plurality of spacers contacting the CMC nozzle shell, and a spar cap. The endwall includes at least one surface in lateral contact with the spar cap and maintains a lateral orientation of the CMC nozzle shell and the nozzle spar relative to the endwall. The lateral orientation maintains a predetermined throat area of the nozzle assembly. A method for forming the nozzle assembly includes inserting the nozzle spar into the interior cavity, rotating the CMC nozzle shell and the nozzle spar laterally relative to the endwall, and maintaining the lateral orientation.

A method of fabricating an airfoil preform for a turbine engine by selective melting on a bed of powder, the preform including an airfoil and a removable support secured to the airfoil, the airfoil being fabricated layer by layer from a first edge to a second edge of the airfoil, the method including fabricating the removable support and the airfoil, the removable support being for securing to a fabrication platform and to a portion of a face of the airfoil situated near the first edge and facing the fabrication platform. The face of the airfoil facing the fabrication platform includes a flat extending away from the face, the flat being present over a portion of the face that is situated outside the first edge, the support being secured to the flat or both to the flat and to the portion of the face that is situated outside the first edge.

A vane for use with a stator assembly is disclosed herein. The vane having: an upper mounting portion; a lower mounting portion; an airfoil body extending between the upper mounting portion and the lower mounting portion; a first integrally formed metal fillet located between the upper mounting portion and the airfoil body, the first integrally formed metal fillet defining a first mounting surface located between the upper mounting portion and first integrally formed metal fillet; and a second integrally formed metal fillet located between the lower mounting portion and the airfoil body, the second integrally formed metal fillet defining a second mounting surface located between the lower mounting portion and second integrally formed metal fillet.

Shrouds and shroud segments for gas turbine engines are provided. In one exemplary aspect, a gas turbine engine includes a first rotor blade stage and a second rotor blade stage. The gas turbine engine also includes a shroud formed from a plurality of shroud segments that each have a body and a vane extending from the body. The vane is disposed within the flow path between the first rotor blade stage and the second rotor blade stage to form at least a portion of a nozzle stage. Further, the body of the shroud segment defines an outer wall of the flow path and spans from the first rotor blade stage to the second rotor blade stage. An impingement baffle segment couples the shroud segment with a casing of the gas turbine engine.

At least one turbine component for a gas turbine includes a base component formed by casting and an article. The base component includes a platform. The article on the upper surface of the platform is formed by additive manufacturing. The article has a proximal face sized and shaped to cover at least a portion of the upper surface of the platform of the turbine component and a contoured distal face opposite the proximal face. The contoured distal face has a contour surface serving as at least a portion of a hot gas path surface of the turbine component. The contour surface is arranged and disposed to provide a controlled flow pattern of a working fluid across the contour surface based on a clock mounting location of the turbine component in a turbine.

Methods of manufacturing a bladed stator element for a turbomachine include mounting a tool on a circumferential zone of an annular shell prior to welding vanes in the circumferential zone, welding radially outer ends of the vanes to the annular shell, dismantling the tool after welding the vanes in the circumferential zone, and repeatedly mounting and dismounting the tool on different circumferential zones of the annular shell so as to fix the vanes all around the annular shell.