A casing assembly for a gas turbine engine includes a plurality of vane casing segments. Each vane casing segment includes an arcuate member and a pair of split-line flanges. Each split-line flange includes a first axial flange end and a second axial flange end. Each split-line flange is fixedly coupled to an adjacent split-line flange of an adjacent vane casing segment. Each vane casing segment includes at least one row of stator vanes welded to the arcuate member. Each split-line flange is at least partially and circumferentially inclined relative to a rotational axis of the gas turbine engine, such that the first axial flange end is circumferentially offset from the second axial-flange end. Further, each split-line flange includes an intersecting portion, such that at least the intersecting portion of each split-line flange is circumferentially inclined relative to the rotational axis.

Disclosed herein is a ceramic matrix composite airfoil comprising a triple bifurcated ply that defines a suction side, an outer platform, a pressure side and an inner platform of the airfoil, wherein the triple bifurcated ply comprises at least one ply that comprises a consolidated region, wherein the consolidated region is split into two bifurcated regions in three locations in three different directions.

The disclosure describes example techniques and assemblies for joining a first component and a second component. The techniques may include positioning the first and second component adjacent to each other to define a joint region between adjacent portions of the first component and the second component, the joint region being coated with an adhesion resistant coating. The techniques may also include positioning a braze material in the joint region, heating the braze material to form an at least softened material, and cooling the at least softened material to form a mechanical interlock including the braze material in the joint region joining the first and second components. The braze material does not metallurgically bond to the joint surface.

A method of manufacturing a compressor stator having: a first stator blade with a first leading edge and a first trailing edge; a second stator blade disposed a circumferential distance from the first stator blade, the second stator blade having a second leading edge disposed an axial distance from the first leading edge and a second trailing edge disposed an axial distance from the first trailing edge; the method comprising: using additive manufacturing to deposit and fuse together progressive layers of metal material commencing at a substrate to form the first stator blade, the second stator blade, at least one intermediate support structure disposed between the first stator blade and the second stator blade, and at least one primary support structure disposed between the substrate and at least one of: the first stator blade; and the second stator blade; and removing the primary support structure and the intermediate support structure.

An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a stationary engine structure. The stationary engine structure includes a diffuser, a combustor, an engine case and a plenum. The combustor is disposed within the plenum. The engine case forms a peripheral boundary of the plenum. A gas path extends sequentially through the diffuser, the plenum and the combustor. A first section of the stationary engine structure is formed as a first monolithic body. The first section includes the diffuser and the combustor. A second section of the stationary structure is formed as a second monolithic body. The second section is configured as or otherwise includes the engine case.

An end or intermediate casing for a combustion turbine engines includes prefabricated vanes of a first metal. Ends of the prefabricated vanes are then embedded within cast-in place inner and outer, annular-shaped ring castings, formed from a second metal having a lower melting point than the first metal. The respective ends of the prefabricated vanes include first and second shanks, with respective first and second surface features that are oriented transverse to the central axis of the vane are encapsulated in the molten second metal during the inner and outer ring casting. Once the castings harden, the first and second surface features, such as for example circumferential fillets projecting outwardly from the airfoil portion of the vane, inhibit separation of the vanes from the respective inner and outer rings.

A vane assembly for use with a gas turbine engine includes an outer wall, an inner wall, and a plurality of airfoils. The outer wall and the inner wall extend at least partway about an axis. At least one of the airfoils is coupled with the outer end wall and the inner end wall to transmit force loads through the vane assembly.

A method is for fabricating a part by additive manufacturing while sparing certain particularly sensitive surfaces of the part, and in particular surfaces that have an influence on the aerodynamics of the final part. The method includes the following steps: providing a digital model of a part that is to be fabricated, the part that is to be fabricated including at least one surface that is to be spared, and orienting the digital model relative to a construction direction wherein the part is to be constructed in such a manner that the surface that is to be spared presents a construction angle greater than 30°, preferably greater than 50°.

Vane assemblies are described. The vane assemblies include a platform, an airfoil extending from the platform, a forward rail extending from the platform and arranged along a forward side of the platform, and an aft rail extending from the platform and arranged along an aft side of the platform. At least one support beam is provided extending in a forward-aft direction between the forward rail and the aft rail and separated from the platform by a first distance. The at least one support beam has a thickness in a radial direction of 40% or less of a total radial extent from the platform to an outer diameter edge of at least one of the forward rail and the aft rail and the at least one support beam has a thickness in a circumferential direction of 30% or less of a total circumferential extent of vane assembly.

A vane assembly includes an aerofoil having a leading edge, a trailing edge, and a pressure surface and a suction surface defined between the leading edge and the trailing edge. The aerofoil includes a blade member forming the trailing edge, at least a portion of the pressure surface and at least a portion of the suction surface. The blade member is formed of a first material. The aerofoil further includes a spar at least partly enclosed by the blade member and forming at least a portion of the leading edge. The spar further forms at least one cooling channel and supports at least a portion of an interior surface of the blade member. The spar is formed of a second material different from the first material. The second material has a greater impact resistance than the first material.