Method of making a spar cap includes: providing a plurality of composite strips, each strip being of constant cross section defined by first and second sides and edges, the first and second sides comprising first and second abutment surfaces, the strip being of uniform thickness between the abutment surfaces, a first edge region of the strip comprising a first edge being of relatively reduced thickness, the first side of the strip comprising an edge surface, and the strip having a peel ply layer at least partially covering the first abutment surface and the edge surface; removing the peel ply layers; stacking the strips such that the first abutment surface abuts an abutment surface of an adjacent strip to define an interface region, such that a clearance region is defined; supplying resin to the respective clearance regions and causing the resin to infiltrate into the interface regions; and curing the resin.

A turbine blade includes a core element having a base portion, a tip portion, and an intermediate portion extending between the base portion and the tip portion. The intermediate portion includes a non-uniform cross-section and is a high-strength fiber material. The turbine blade further includes a shell disposed around the core element, and the volume between the core element and the shell forms a void.

A turbomachine rotor blade, said blade having, at the distal end thereof, a heel comprising: a platform (2) having a first edge (201) on the lower side and a second edge (202) on the upper side, at least one sealing member having a first end portion (301) on the lower side and a second end portion (302) on the upper side, said sealing member having a sealing top that extends radially outwards from said platform (2) between said first and second end portions (301, 302), characterized in that, for at least one sealing member, the heel (105) comprises, on at least at one of the edges (201, 202), a portion forming a bowl (5) extending along the end portion (301, 302) of the sealing member which corresponds to the edge (201, 202), the portion forming the bowl (5) being suitable for receiving a deposit of anti-wear material.

A method of manufacturing a hollow aerofoil component for a gas turbine engine includes using a capping panel to cover a pocket in a pocketed aerofoil body. During manufacture, the outer surface of the capping panel is located relative to the pocketed aerofoil body. This ensures that the outer surface of the capping panel is located as accurately as possible. This means that the capping panel can be made to be as thin as possible, which in turn reduces weight and material wastage. Once the capping panel has been located in position, it may be welded to the aerofoil body in order to produce the hollow aerofoil component.

A cooling jacket for a hollow airfoil of a turbine nozzle of a turbomachine, includes a main body including a central intake duct central defining a first ventilation air circulation area and connected to suction and pressure faces including at least two rows of drill holes by two separating walls defining second and third ventilation air circulation areas, an outer plate including first, second and third holes to allow the ventilation air respectively into the first, second and third ventilation air circulation areas, and an inner plate including a central opening to expel air from the first ventilation air circulation area, the outer and inner plates being secured by respectively soldering to the main body to form a one-piece unit with three ventilation air circulation areas, independent and airtight with respect to one another, before its installation in the hollow airfoil of the nozzle.

A hollow ceramic matrix composite (CMC) turbine bucket with an internal reinforcement lattice structure has improved vibration properties and stiffness. The lattice structure is formed of thin-walled plies made of CMC. The wall structures are arranged and located according to high stress areas within the hollow bucket. After the melt infiltration process, the mandrels melt away, leaving the wall structure to become the internal lattice reinforcement structure of the bucket.

An attachment assembly for a ceramic matrix composite (CMC) vane includes an inner diameter ring having a strut recess and a first flange on a downstream side configured for engaging an inner platform of the CMC vane, an outer diameter ring having a support structure disposed to form a gap relative to a portion of an outer platform of the CMC vane during a non-operating engine condition, and a strut cantilevered from an inner side of the outer diameter ring and disposed to pass through a hollow portion within an airfoil of the CMC vane to engage the strut recess in the inner diameter ring to support the inner diameter ring and the CMC vane. The inner and outer diameter rings and the strut may be metallic and carry the loads rather than the CMC carrying the loads.