A method for producing a hollow part made of a ceramic matrix composite material. The method includes shaping a hollow fibrous preform. A core of oxidizable material is housed or inserted into the preform. The method also includes consolidating the preform and extracting the core by oxidising the core.

A component for a turbine engine is provided. The component includes a main body portion having a flow path surface; and a protective layer formed of a chopped fiber material, the protective layer cohered to at least a portion of the flow path surface of the main body portion.

A turbine for a gas turbine engine includes, among other things, a shaft rotatable about a longitudinal axis, a turbine rotor including one or more rows of turbine blades and a disk assembly coupled to the shaft. The disk assembly includes one or more disks each having an attachment region extending radially between an inner boundary and an outer boundary, the outer boundary is established by an outer periphery of the respective disk, the attachment region defines an array of slots distributed about the outer periphery, each of the slots extends radially inwardly from the outer boundary to the inner boundary, and each of the slots is dimensioned to receive a root section of a respective one of the turbine blades to mount the turbine blades to the disk assembly.

A method of forming a ceramic matrix composite component having an internal cooling circuit includes wrapping at least a first sheet around a first mandrel, wrapping at least a second sheet around a second mandrel, creating a first plurality of holes in the first sheet corresponding to a plurality of openings in the first mandrel, creating a second plurality of holes in the second sheet corresponding to a plurality of openings in the second mandrel, aligning the first mandrel and the second mandrel such that the first plurality of holes face and are aligned with the second plurality of holes, wrapping at least a third sheet around both the first mandrel and second mandrel to form a preform, the preform comprising each of the first sheet, the second sheet, and the third sheet, and densifying the preform. The first sheet, second sheet, and third sheet are formed from a ceramic fiber material.

A propeller includes a rotatable hub and at least two propeller blades coupled to the rotatable hub. Each of the propeller blades is formed from a combination of a first material having a negative Poisson's ratio (NPR) and a second material having a positive Poisson's ratio (PPR). The first material and the second material can be layered or can be formed as a matrix with one of the first or second material embedded in the other. In a layered configuration, a layer of the first material is positioned between adjacent layers of the second material, and the layers can be connected by tabs of NPR material. The combination of the NPR and PPR materials improve the strength and impact resilience of the propeller blades compared to conventional materials.

The invention relates to a composite blade (5) for a turbine engine rotor, for example, an unducted propeller, comprising a skin (6) made of woven fibres forming the outer profile of the blade and an attachment (11) with cylindrical geometry in the direction of the span emerging from the blade root (9), which attachment is intended to retain the blade (5) on a hub of the rotor, characterised in that it further comprises a spar (16) having a hollow tube structure made of braided carbon fibres, fixed to the attachment (11) and extending inside the skin (6) over at least part of the span of the blade (5). The invention also relates to a propeller comprising said blade and to a method for manufacturing said blade.

A turbine of a turbo engine includes a casing and a nozzle stage including an outer metal shroud secured to the casing, an inner metal shroud, and a plurality of ring sectors made of ceramic-matrix composite material each having an internal platform, an external platform and at least one blade extending between the external platform and the internal platform and secured thereto. The internal platform of each sector includes a first and a second tab protruding radially inwardly from respectively an upstream end and a downstream end of the internal platform, and the inner shroud includes a shoulder protruding radially outwardly and being in abutment upstream against the first tab, and a clamp downstream of the shoulder including two jaws extending radially outwardly and held under stress towards each other against the second tab.

A method may include: in a used metal turbomachine blade including a root including a shank, a platform coupled to the shank and an airfoil coupled to the platform, removing the airfoil, leaving a remaining base including the platform, the shank and the root. The method may also form a radially extending opening through the platform into the shank, and insert a ceramic shank nub extending from a ceramic airfoil into the radially extending opening of the remaining base. The ceramic airfoil is fixedly attached to the remaining base. The method allows reuse of the metal shank while providing the lower cooling requirements of a ceramic airfoil.

An assembly for forming a gas turbine engine according to an example of the present disclosure includes, among other things, a layup tool including a main body extending along a longitudinal axis and a flange extending radially from the main body, the flange defining an edge face slopes towards the main body to an axial face. At least one compression tool has a tool body having a first tool section and a second tool section extending transversely from the first tool section. The first tool section is translatable along a retention member in a first direction substantially perpendicular to the edge face such that relative movement causes the second tool section to apply a first compressive force on a composite article trapped between the axial face of the flange and the second tool section. A method of forming a gas turbine engine component is also disclosed.

Fibrous preform for a composite blade and also a composite blade formed by means of such a preform, a rotor and a rotating machine comprising such a blade, the preform comprising a first longitudinal section, configured to form a blade root, and a second longitudinal section, extending from the first longitudinal section, configured to form a portion of an airfoil, wherein the first longitudinal section has a first thickness at its upper end and wherein the second longitudinal section comprises at least one set-back zone having a thickness at least three times less than the first thickness, said set-back zone occupying at least 50% of the second longitudinal section.