Various methods and assemblies are provided for producing composite components having formed in features. For example, a method may comprise depositing a composite material on a base tool; bringing a feature forming tool into contact with the composite material; and processing the composite material with the feature forming tool in contact with the composite material. The processed composite material forms a green state composite component. The feature forming tool comprises a sheet having one or more elements for interacting with one or more elements of the base tool to form one or more features of the composite component. In some embodiments, the method also may comprise sealing a bag around the feature forming tool and the composite material after bringing the tool into contact with the composite material and removing the bag and the feature forming tool from the green state composite component after processing.

Various methods and assemblies are provided for producing composite components having formed in features. For example, a method may comprise depositing a composite material on a base tool; bringing a feature forming tool into contact with the composite material; and processing the composite material with the feature forming tool in contact with the composite material. The processed composite material forms a green state composite component. The feature forming tool comprises a sheet having one or more elements for interacting with one or more elements of the base tool to form one or more features of the composite component. In some embodiments, the method also may comprise sealing a bag around the feature forming tool and the composite material after bringing the tool into contact with the composite material and removing the bag and the feature forming tool from the green state composite component after processing.

A method of forming a ceramic-metal composite part is described herein. The method includes maintaining molten metal in an interior of a housing in a liquefied state, the interior including a first chamber, a second chamber, and a port defined therebetween. The method further includes sealing the port such that the molten metal in the first chamber is maintained at a first liquid level, suspending a part at a height within the first chamber above the first liquid level, forming a pressure differential between the first chamber and the second chamber, unsealing the port such that molten metal from the second chamber flows into the first chamber, and resealing the port when the molten metal in the first chamber reaches a second liquid level such that the ceramic part is submerged in the molten metal.

A method of forming a ceramic-metal composite part is described herein. The method includes maintaining molten metal in an interior of a housing in a liquefied state, the interior including a first chamber, a second chamber, and a port defined therebetween. The method further includes sealing the port such that the molten metal in the first chamber is maintained at a first liquid level, suspending a part at a height within the first chamber above the first liquid level, forming a pressure differential between the first chamber and the second chamber, unsealing the port such that molten metal from the second chamber flows into the first chamber, and resealing the port when the molten metal in the first chamber reaches a second liquid level such that the ceramic part is submerged in the molten metal.

A combustion assembly for a gas turbine engine may be provided. The combustion assembly may include a ceramic matrix composite combustor shell, which may include a chamber defined by a wall of the ceramic matrix composite combustor shell, and the ceramic matrix composite combustor shell may include a ceramic matrix composite chute integral with the ceramic matrix composite combustor shell. The ceramic matrix composite chute may extend towards a midline of the chamber. A method for fabricating a ceramic matrix composite chute may be provided. At least one chute may be woven in three dimensions into a ceramic preform. A layup tool may be inserted into the chute. The chute may be enlarged with the layup tool. The ceramic preform may be formed into a ceramic matrix composite body, which includes a combustor shell and the chute.

A combustion assembly for a gas turbine engine may be provided. The combustion assembly may include a ceramic matrix composite combustor shell, which may include a chamber defined by a wall of the ceramic matrix composite combustor shell, and the ceramic matrix composite combustor shell may include a ceramic matrix composite chute integral with the ceramic matrix composite combustor shell. The ceramic matrix composite chute may extend towards a midline of the chamber. A method for fabricating a ceramic matrix composite chute may be provided. At least one chute may be woven in three dimensions into a ceramic preform. A layup tool may be inserted into the chute. The chute may be enlarged with the layup tool. The ceramic preform may be formed into a ceramic matrix composite body, which includes a combustor shell and the chute.

The invention relates to a method of fabricating a turbine engine blade out of composite material comprising fiber reinforcement densified by a matrix, the blade comprising an airfoil, a platform situated at a longitudinal end of the airfoil, and at least one functional element projecting from the outside face of the platform. The method comprises: making a single-piece fiber blank by multilayer weaving; shaping the fiber blank to obtain a single-piece fiber preform having a first portion (302) forming a preform for the blade airfoil (320) and a second portion (314) forming a preform for the platform (340) and at least one preform for a functional element (352; 354); and densifying the fiber preform with a matrix. The second preform portion comprises a set of yarn layers interlinked by weaving with at least one zone of non-interlinking being provided to make it possible to deploy the functional element preform relative to the first platform preform.

The invention relates to a method of fabricating a turbine engine blade out of composite material comprising fiber reinforcement densified by a matrix, the blade comprising an airfoil, a platform situated at a longitudinal end of the airfoil, and at least one functional element projecting from the outside face of the platform. The method comprises: making a single-piece fiber blank by multilayer weaving; shaping the fiber blank to obtain a single-piece fiber preform having a first portion (302) forming a preform for the blade airfoil (320) and a second portion (314) forming a preform for the platform (340) and at least one preform for a functional element (352; 354); and densifying the fiber preform with a matrix. The second preform portion comprises a set of yarn layers interlinked by weaving with at least one zone of non-interlinking being provided to make it possible to deploy the functional element preform relative to the first platform preform.

Disclosed herein are methods of forming substantially crystalline, dense silicon carbide fibers from infusible polysilazane fibers by utilizing a single stage pyrolysis. The pyrolysis is performed using a continuous process in a single furnace with a constant atmospheric condition. Also disclosed are substantially crystalline, dense silicon carbide fibers formed by these methods.

A method of production of edge protection strips (10) of ceramic material, includes the steps of: arranging a sheet-like article (20) of ceramic material elongated along a longitudinal axis (A) and provided with an exposed surface (S); incising the exposed surface (S) along two distinct incision planes, parallel to each other, orthogonal to the exposed surface (S) and each intersecting the exposed surface (S) along an incision line (I) parallel to the longitudinal axis (A); cutting the article (20) along two cutting planes each intersecting an incision plane along a cutting line (C) parallel to the incision lines (I) and mutually incident in an intersection line (X) included in the thickness of the article (20); and separating a substantially prismatic edge protection strip (10) from the article (20) with a first portion (S1) of exposed surface (S) internal to the incision lines (I).