A seal assembly for a gas turbine engine according to an example of the present disclosure includes a seal that has a main body extending circumferentially between opposed mate faces. The main body has a sealing portion and an engagement portion extending outwardly from sealing portion along at least one of the mate faces. The main body has a core that has one or more core plies having a first fiber construction and arranged to establish an internal cavity. An overwrap has one or more overwrap plies having a second fiber construction and arranged to follow a perimeter of the one or more core plies to establish the engagement portion and the sealing portion, and the second fiber construction differs from the first fiber construction. The first fiber construction establishes a first target fiber volume fraction, the second fiber construction establishes a second target fiber volume fraction. A method of fabricating a seal for a gas turbine engine is also disclosed.

A void filler material includes a ceramic rod and a fibrous overwrap. The void filler material may be used in a ceramic matrix composite. The method of making the ceramic matrix composite includes inserting the void filler material in voids of a preform and depositing a ceramic matrix on the preform and the void filler material using chemical vapor infiltration.

A seal includes a ceramic matrix composite ply having woven ceramic-based fibers in a ceramic-based matrix. The ceramic matrix composite ply has at least one bend formed about a bend axis and defines at least one rounded portion. A sealed assembly and a method of making a seal are also disclosed.

Methods for preparing ceramic matrix composites using melt infiltration are provided as well as the resulting ceramic matrix composites. The methods and products include the incorporation of a non-wetting coating to one or more sacrificial fibers. The one or more sacrificial fibers are removed, such as decomposed during pyrolysis, resulting in the formation of a plurality of functional features, in the form of regular and elongate channels along the ceramic matrix composite. During the removing of the one or more sacrificial fibers, the non-wetting coating remains on an interior surface of the plurality of functional features to block infiltration of an infiltrant to the plurality of functional features and deposition thereon. Alternatively, the sacrificial fibers may be removed subsequent to melt infiltration.

A method is provided in which a first tape is applied to an outer surface of a ceramic matrix composite (CMC). A second tape is applied to the first tape. The second tape is heated to at least a melting temperature of the second tape. During heating, the first tape is infiltrated with a molten material from the second tape, which forms a surface layer on the CMC

Method for manufacturing a turbomachine blade made of ceramic-matrix composite material comprising an airfoil and a platform, producing a first fibrous preform, the first fibrous preform forming an airfoil preform, producing a second fibrous preform comprising a housing, the second fibrous preform forming a platform preform, producing a strand of ceramic fibers, consolidating the fibrous preforms and the strand of ceramic fibers to form consolidated preforms and a consolidated strand, assembling the consolidated preforms by engagement and cooperation of the first consolidated preform in the housing, the consolidated strand being interposed between the first consolidated preform and the second consolidated perform, co-densifying the assembly to form the turbomachine blade.

A composite material part includes a fiber preform forming fiber reinforcement including a stack of at least two fiber plies, each of the fiber plies being made of an interlock weave three-dimensional fabric and each of the fiber plies having a number of warp yarn layers or a number of weft yarn layers that is greater than or equal to three; and a matrix present in the pores of the fiber preform.

A spar cap for a rotor blade of a wind power installation, having a longitudinal extent from a first end to a second end, a transverse extent orthogonal to the longitudinal extent, and a thickness orthogonal to the longitudinal extent and to the transverse extent. A method for producing a spar cap as mentioned at the outset. The spar cap has a longitudinal extent from a first end to a second end, a transverse extent orthogonal to the longitudinal extent, and a thickness orthogonal to the longitudinal extent and to the transverse extent, at least two tiers of a first fiber composite material, and at least one tier of a second fiber composite material, wherein the first fiber composite material has a matrix material and/or fibers which is/are different from that/those of the second fiber composite material, the second fiber composite material is disposed in a portion adjacent to the second end, in the direction of the thickness between the at least two tiers of the first fiber composite material, and the at least one tier of the second fiber composite material terminates ahead of the second end.

A braze alloy for joining or repairing ceramic matrix composite (CMC) components comprises a braze composition including silicon at a concentration from about 48 at. % to about 66 at. %, titanium at a concentration from about 1 at. % to about 35 at. %, and an additional element selected from aluminum, cobalt, vanadium, nickel, and chromium. The braze composition comprises a melting temperature of less than 1300 C.

Rotor assemblies and methods for manufacturing airfoils for rotor assemblies are provided. For example, a rotor assembly comprises a rotary structure extending circumferentially about an axial centerline of a gas turbine engine and an airfoil having a root and a tip. The root is coupled to the rotary structure and has a bulbous shape. The airfoil is formed from a plurality of composite plies, a portion of which defines at the root first and second end surfaces, which are in contact with the rotary structure and together define a chisel-shaped end of the root. In another embodiment, a rotor assembly comprises an outer rotor extending circumferentially about an axial centerline of a gas turbine engine and a composite outer rotor blade having a root and a tip. The blade root is coupled to the outer rotor and extends inward along a radial direction toward the axial centerline.