An apparatus and method relating to a film-hole of a component of a turbine engine comprising including forming the hole in the component and applying a coating to the component such that the coating fills in portions of the film-hole.

A tubular ventilation sleeve for a turbomachine distributor, in particular for an aircraft, the sleeve having a generally elongate shape along an axis (A-A) and including a perforated tubular wall around said axis, one of the axial ends of the sleeve being open and the other being closed by a bottom wall, wherein it further includes support beams when the sleeve is made by additive manufacturing, the beams extending inside the sleeve between the tubular wall and the bottom wall and having a longitudinal cross-section with a generally triangular shape, two sides of which are respectively connected to the tubular wall and the bottom wall and the last side of which is free and extends inside the sleeve, perforations in the tubular wall being provided between the support beams.

An assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a gas turbine engine component that has a first interface portion, and a support that has a mounting portion and a second interface portion, the mounting portion attachable to an engine static structure, a first retention feature that releasably secures the first interface portion to the support in a first installed position of the gas turbine engine component, and a second retention feature dimensioned to secure the first interface portion to the second interface portion in a second installed position of the gas turbine engine component. The first installed position differs from the second installed position, and one of first and second retention features is dimensioned to carry the gas turbine engine component in response to release of another one of the first and second retention features. A method of sealing for a gas turbine engine is also disclosed.

An aerofoil component defines an in use leading edge and a trailing edge. The leading edge has at least one serration defining an apex and a nadir. The leading edge has a generally chordwise extending slot located at the nadir of each serration.

Damping systems are provided for a rotor blade platform. The damping system may include a blade platform defining a damper pocket and a CMC wedge damper positioned within the damper pocket. The CMC wedge damper has at least one damper angled surface parallel to a longitudinal axis. The damper pocket comprises a pocket angled surface positioned about the at least one damper angled surface.

A damping device for being situated between a housing wall of a housing of a thermal gas turbine and a casing ring is provided. The casing ring includes an area radially internal with regard to a rotation axis of a rotor of the thermal gas turbine and facing rotating moving blades of the gas turbine. The damping device includes at least sectionally a porous damping structure. A method for manufacturing this type of damping device as well as to a thermal gas turbine, in particular an aircraft engine, in which this type of damping device is situated in a housing of the gas turbine between a housing wall and a casing ring are also provided.

A fan containment system arranged to be fitted around an array of radially extending fan blades mounted on a hub in an axial gas turbine engine. Each fan blade has a respective tip. The system includes: a cylindrical fan case including a hook projecting in a radially inward direction and positioned axially forward of the radial array of fan blades; a fan track liner disposed on the radially inner surface of the fan case; and a damaging tool which projects radially inwards from the fan case towards the tips. The damaging tool has a tip radially outward of the fan blade tips. The damaging tool is configured that in the event that one of the fan blades is released from the hub, the tip of the damaging tool damages the fan blade tip of the released fan blade to promote penetration of the fan blade into the fan track liner.

An aircraft including a pylon attached to a gas turbine engine and a mounting system attaching the engine to the pylon. The mounting system including a first and a second frame each of three elongate members arranged in a triangle, each frame respectively arranged such that a core of the engine is positioned extending through an area defined between the three elongate members of each frame. Each frame forming at least part of a load bearing connection between the pylon and the engine. Each frame consisting of two portions, each portion corresponding to each side of the engine as attached to the pylon. The triangle formed by each frame being symmetrical about a plane separating the two portions. The engine is attached to the mounting system such that both frames are positioned axially forward of a radially extending projection of a first turbine stage in the core.

A method for coating a turbomachine guide vane including a root and a tip, an extrados face and an intrados face connected to one another by a leading edge and a trailing edge, the method including completely covering one of the faces of the vane with a polymer coating of thickness (e.sub.1) provided with grooves, removing the grooves from a part of the polymer coating in such a way that the polymer coating includes a grooved zone and a non-grooved zone, coating the non-grooved zone with a coat of paint of thickness (e.sub.3) such that the thickness of the coat of paint superimposed on the non-grooved zone is substantially equal to the thickness (e.sub.1) of the grooved zone.

A lubricant nozzle for a planetary gear set speed reducer of a turbomachine, the nozzle having a generally elongate shape and including a body with a longitudinal axis B, the body having a longitudinal inner cavity in fluid communication with a lubricant inlet located at a longitudinal end of the body and with lubricant outlet apertures that are provided in an annular wall of the body and that extend substantially radially relative to axis B, wherein the apertures are formed in at least one boss of the body, which boss projects radially outwards on the wall and extends, about the axis, at an angular extent that does not exceed 180°.