A hybrid three-layer system is presented. The hybrid three-layer system includes a two-layer composite system and an additively manufactured third layer comprising a lattice structure. The composite layer system includes a metallic substrate, a structured surface, and a thermal protection system. The structured surface may be additively manufactured onto the metallic substrate and includes structured surface features formed to project above the metallic substrate. Each of the structured surface features are separated from adjacent structured surface features by grooves. The thermal protection coating may be thermally sprayed onto the structured surface and is bonded to each of the structured surface features. The lattice structure is in contact with a surface of the metallic substrate of the composite layer system.

The invention relates to a seal carrier for a turbomachine, in particular a gas turbine, having a carrier base and at least one seal body, wherein the at least one seal body is connected to the carrier base, and wherein the at least one seal body is formed by a plurality of cavities arranged next to one another, in particular uniformly, in the peripheral direction and in the axial direction, wherein the cavities extend out from the carrier base in the radial direction and are delimited by a cavity wall. According to the invention, the seal body has a plurality of damping portions which are designed to locally damp or disrupt the flow of force in the seal body, wherein the carrier base is continuous in the region of the damping portions.

The invention is related to the gas turbine stators of the gas turbine engines applied in aviation. The gas turbine stator, in the outer housing of which sectors of the split honeycomb ring (made as double-layer one) are installed with support elements on the front and rear axial ends of the sector. In this invention, the layer of the sector facing the outer housing is made U-shaped in the plane, the support elements are made as separate rotary bodies distributed uniformly along the circumference and the front support elements (on the gas flow direction) are larger than rear ones in terms of geometrical dimensions by factor 1.1 . . . 1.5. Therefore, the implementation of the invention proposed with the characteristic features above, in conjunction with the known features of the invention claimed enables reduction of the gas turbine stator weight and improvement its reliability without compromising the turbine efficiency.

A structural and/or acoustic panel including an inner skin, an outer skin, a honeycomb structure and at least one sealing flange. The sealing flange has a U-shaped section directed toward the honeycomb structure and is clamped with the honeycomb structure between the inner and the outer skins. A method for manufacturing such a panel by soldering the component elements thereof is also disclosed.

Methods, apparatus, systems and articles of manufacture are disclosed. An inner shroud damper for a gas turbine engine includes: at least one carrier including a joint to couple to an inner shroud, the at least one carrier having a first side and a second side, and at least one mass damper coupled to the at least one carrier.

An abradable insert for a gas turbine engine, the abradable insert including: a base layer; a lattice layer connected to the base layer, wherein the lattice layer comprises a series of walls that define a plurality of cells; and a sheet layer connected to the lattice layer on an opposite side on the lattice layer from the base layer, wherein the sheet layer is curved and includes a direction of concavity that points away from the base layer, wherein the lattice layer and the sheet layer are integrally formed together and are a monolithic piece of material.

Fire retardant engine casing apparatus are disclosed. An example engine casing includes an inner shell circumferentially surrounding blades of a fan, a compressor, or a turbine, an outer shell positioned around the inner shell, and a fire retardant material between the inner shell and the outer shell.

A double walled stator housing includes a first stator housing wall, a second stator housing wall located radially outward from the first stator housing wall, and an air gap located between the first and the second stator housing walls. The housing also includes at least one support structure attached to the first stator housing wall and the second stator housing wall, spanning the air gap and configured to minimize heat transfer between the first wall and the second wall.

Disclosed is a movable vane (1) for a wheel (2) of an aircraft turbine engine, the vane (1) comprising a blade (4) delimited by an outer heel (8) comprising a first seal (14), the vane (1) comprising an internal circuit (16) suitable for receiving a first minor gas flow (f1), this circuit (16) comprising a supply cavity (17) opening at the root (9) via at least one inlet opening (18), characterised in that the circuit (16) comprises at least two channels (19) connected with the supply cavity (17) and each opening on an outer surface of the first seal (14) via a discharge opening such that a gas jet (J) of the first minor gas flow (f1) is capable of being discharged from each discharge opening, each channel (19) being oriented such that the corresponding gas jet (J) is capable of being projected towards a second minor gas flow (f2) escaping between the heel (8) and a directly adjacent member (22).

Turbofan blade blocking apparatus are disclosed. An example turbofan engine includes a containment casing, a fan, at least one of a trench filler or an abradable layer at least partially aligned with the fan in an axial direction defined by the turbofan engine, at least one of the trench filler or the abradable layer fixed to an inner radial surface of the containment casing, and a metal blocker coupled to the containment casing forward of the fan, an aft end of the metal blocker interfacing with at least one of the trench filler or the abradable layer.