A hollow vane assembly including an open body including an interior; at least one cover support structure formed in said open body proximate the interior; a cover brazed to the open body to form at least one flow passage; and at least one ply formed in the cover.

An improved stator assembly for use in a gas-turbine engine is disclosed. The stator assembly may comprise a vane, an inner diameter (ID) ring, an outer diameter (OD) ring, a vane disposed between the ID ring and the OD ring, a potting component coupling the vane to at least one of the OD ring or the ID ring, and a potting embedded component disposed within the potting component. The potting embedded component may prevent disbond of the potting component during operation of the gas-turbine engine.

A gasket comprising at least one spacer and configured to prevent contact between a mount link of a turbine mount assembly and a turbine frame mount of an aircraft engine, which is a turbine engine in some embodiments, is provided. The at least one spacer includes a pin cutout to receive a turbine frame pawl pin of the engine mount and a pin groove that is expandable so as to enable the turbine frame pawl pin to pass through and into the pin cutout. The gasket may include a grommet for receiving a connecting ring that is coupled to a streamer grommet of a streamer. The streamer may act as an identifier for the gasket.

A non-metallic engine case inlet compression seal for a gas turbine engine includes a non-metallic longitudinal leg section that extends from the non-metallic arcuate interface section and a non-metallic mount flange section that extends from the longitudinal leg section.

An assembly for a turbine engine includes an inner platform, and outer platform and a plurality of structural inlet guide vanes. The outer platform circumscribes the inner platform. The structural inlet guide vanes are arranged around an axis, and extend radially between and are connected to the inner platform and the outer platform. A first of the structural inlet guide vanes includes a structural vane body that is configured from or otherwise includes an organic matrix composite.

An inner shroud segment may include an inner housing and an outer housing. The inner housing may have a radial curve centered relative to an axis with a radial wall and a bottom wall that define a first channel. The outer housing may have a first axial wall, a first circumferential wall, and a second axial wall that define a second channel. The outer housing may also be disposed within the first channel with the radial wall of the inner housing contacting the first axial wall, the first circumferential wall, and/or the second axial wall. A compliant material may be disposed within the second channel and coupled to the radial wall and the first axial wall, the first circumferential wall, and/or the second axial wall.

The present invention provides an apparatus for providing a balancing weight in a groove on a rotor disk in a gas turbine engine. The apparatus comprises an elongate reservoir for housing a hardenable material and an inflatable balloon in fluid communication with the distal end of the elongate reservoir. An actuator is provided for forcing the hardenable material from the elongate reservoir to inflate the inflatable balloon with hardenable material within the groove on the rotor disk. A sealing device for sealing the inflated inflatable element to form the balancing weight is also provided.

A stator of a low-pressure compressor of an axial-flow turbomachine. The stator includes an annular row of stator blades including radial extremities which pass through the openings of an internal shroud, and which include radial retaining slots having tapers formed by hooks. The stator includes a ring for retaining the blades on the internal shroud. The ring is curved circumferentially in order to be inserted into a plurality of retaining slots and exhibits the form of a strip having an arched transversal profile which is in abutment against the tapers, in such a way as to maintain the ring in the interior of the slots. The shroud includes an annular layer of abradable material made from silicone, which encloses the ring in such a way as to block the curvature of the arched profile of the ring in order to prevent it from disengaging from the tapers of the slots.

A fan section for a gas turbine engine is provided. The fan section having a fan hub with a slot and a fan blade with an airfoil extending from a root to a tip, the airfoil having a leading edge and the root is received in the slot, wherein a first platform is secured to the fan hub and arranged between adjacent fan blades of the fan section, the first platform having a first platform seal including a platform seal leading edge and a base is secured to a side of the first platform and a first winglet extends from the platform seal leading edge and contacts the airfoil leading edge.

The invention relates to a fan blink (28) for an aircraft turbomachine, comprising a hub, an annular platform (42) and fan blades (44) arranged projecting from the annual platform. It also comprises a mechanical discharge slit (52) from a trailing edge (49) of the fan blade, associated with at least one of the fan blades (44), for the case of ingestion of a bird, the slit being made on the annular platform (42) going around the trailing edge (49).