A seal device according to the embodiment includes a seal fin provided between the rotation body and the stationary body and extending in a circumferential direction of the rotation body, and a swirl brake fin provided on the stationary body upstream of the seal fin. The swirl brake fin reduces a circumferential velocity component of the working fluid. The swirl brake fin has a negative pressure surface provided on a side of a rotation direction of the rotation body and a positive pressure surface provided on an opposite side to the negative pressure surface. The positive pressure surface extends in a direction opposite to the rotation direction of the rotation body from radially outward toward radially inward.

Thrust reverser composite cascade (1), comprising at least one longitudinal wall (15) and transverse walls (14) connecting to this longitudinal wall, characterized in that the longitudinal wall comprises at least one continuous longitudinal fibre bundle (19) and the transverse walls each comprise at least one continuous transverse fibre bundle (23) crossing the longitudinal bundle, so that the intersections (16) of the transverse and longitudinal walls are structurally bridged in both directions by the reinforcing continuous longitudinal and transverse fibre bundles.

A compressor of a gas turbine engine includes an impeller having a plurality of impeller blades. The compressor includes a diffuser downstream from the impeller that has a plurality of diffuser blades. Each diffuser blade extends from a hub to a shroud in a spanwise direction, and a leading edge of each diffuser blade is spaced apart from an impeller trailing edge of each of the plurality of impeller blades by a vaneless gap. Each diffuser blade includes a cutback region that extends from proximate the leading edge toward a trailing edge. The cutback region reduces a thickness of each of the diffuser blades such that a throat area defined between adjacent diffuser blades increases in the spanwise direction from the hub to the shroud and the vaneless gap increases in the spanwise direction from the hub to the shroud.

The present invention relates to a guide vane ring element for a turbomachine, such as a gas turbine, with at least one drilled hole for the at least partial uptake of a flange of a rotatable guide vane, wherein an edge of the drilled hole on the vane element side is designed convexly along the edge in a first circumferential portion and, in particular, has a radius.

A diffuser-deswirler for a gas turbine engine is provided. The diffuser-deswirler includes an inner shell and an outer shell spaced apart and configured for receiving a flow of compressed air from a compressor of a gas turbine engine. A plurality of vanes extend between the inner shell and the outer shell to define a plurality of fluid passageways and a splitter extends along the circumferential direction between adjacent vanes to split the flow of compressed air passing through the each fluid passageway. All of these components are manufactured as a single monolithic piece and are configured for diffusing and deswirling compressor air before passing it to a combustor for an improved combustion process.

A cascade array vane is provided that includes a vane, a first longitudinal portion, a second longitudinal portion, and at least one coupling feature. The vane includes an airfoil portion extending between a first end and a second end. The first longitudinal portion is disposed at the first end of the vane extending outwardly from the vane. The second longitudinal portion is disposed at the second end of the vane extending outwardly from the vane. The at least one coupling feature is incorporated with at least one of the first longitudinal portion or the second longitudinal portion.

A cascade box is provided that includes a plurality of lengthwise extending side turning vanes and a plurality of forward turning vanes. Each of the plurality of lengthwise extending side turning vanes is configured to have a side turning angle. Each of the plurality of forward turning vanes is attached to respective adjacent ones of the plurality of side turning vanes, and each forward turning vane is configured to have a forward turning angle, and at least one of the forward turning vanes is configured to have a negative forward turning angle.

A cascade thrust reverser assembly for a gas turbine engine includes a nacelle assembly defining a bypass passage. The cascade thrust reverser assembly includes a cascade assembly configured to be at least partially enclosed by the nacelle assembly, the cascade assembly comprising one or more cascade members, the one or more cascade members movable between a stowed configuration wherein the one or more cascade members define a first radial extent and a deployed configuration wherein the one or more cascade members define a second radial extent, wherein the one or more cascade members form a cascade segment in the deployed configuration, and wherein the second radial extent is greater than the first radial extent.

The disclosure relates to a flow deflector for a discharge valve system of a double flow turbomachine compressor. The flow deflector comprises a wall provided with a plurality of ejection channels positioned to discharge a discharge air flow. The ejection channels are arranged in rows along aligned directions, substantially parallel to an axial plane of the deflector. The ejection channels of each row are oriented at a defined angle between a normal line to the wall and axes of the ejection channels, and decreasing between an upstream edge and a downstream edge of the wall, as defined according to the direction of movement of the air flow.

An aircraft has a fuselage, left and right wings, and left and right engine assemblies connected to the wings. Each engine assembly has a nacelle, an engine housed in the nacelle, the engine having a plurality of rotors defining an uncontained rotor impact area, a pylon connecting the nacelle to its wing, at least one hydraulic actuator connected to at least one of the engine and the nacelle, at least one directional control valve hydraulically connected to the at least one hydraulic actuator, and at least one isolation valve hydraulically connected to the at least one directional control valve for selectively cutting off a supply of hydraulic fluid to the at least one directional control valve, the at least one isolation valve being disposed rearward of the uncontained rotor impact area and forward of a trailing edge of its corresponding wing.