A method of applying a braze component to a honeycomb structure may comprise: applying at least a partial vacuum within a chamber, the chamber defined at least partially by a vacuum device and a cover, the honeycomb structure disposed within the chamber, the braze component disposed between the honeycomb structure and the cover; pulling the cover towards the braze component in response to applying the partial vacuum; and pulling the braze component into a plurality of hexagonal cells defined by the honeycomb structure in response to pulling the cover towards the braze component.

The present invention relates to a labyrinth seal for a turbine engine, in particular of an aircraft, comprising a rotor element and a stator element extending around the rotor element, the rotor element being suitable for rotating relative to the stator element about an axis of rotation having an axial direction (DA), the rotor element comprising an annular lip having an outer radial end extending towards an abradable element (57) carried by the stator element, the outer radial end of the annular lip having a corrugation in the axial direction (DA) and a non-zero axial expanse (E.sub.5) associated with the corrugation, the abradable element (57) comprising a plurality of cells (50a, 50b) arranged adjacent to one another along the axial direction (DA) and an ortho-radial direction (O), the cells (50a, 50b) comprising walls which extend in an essentially radial direction, the cells being distributed with a first cell density in a first densified annular zone (Z.sub.51) of the abradable element, said densified annular zone (Z.sub.51) being located opposite the radial end of the lip, said densified annular zone having an axial expanse less than or equal to the axial expanse of the outer radial end of the lip, the cells being distributed according to a reference density of cells outside said first zone, the first density being greater than the reference density.

A rotor for an axial turbomachine includes a drum formed of several parts including: composite rings made of composite material and metallic rings interposed between the composite rings. The metallic rings carry the rotor blades. The metallic rings have an axial branch axially overlapping the composite rings and at least one radial branch in contact with the composite rings.

A seal for a rotor is provided. The rotor defines an axial direction, a circumferential direction, and a radial direction. The seal includes a first flexible element coupled to the rotor. The first flexible element extends in a first direction that is within forty five degrees of the axial direction.

A seal support structure is provided for a circumferential seal. In one embodiment, the seal support structure includes an engine support structure, a seal support, and a shoulder joining the engine support and seal support. The shoulder offsets the engine support from the seal support, and the shoulder and the seal support structure are configured to dampen vibration for the circumferential seal. The seal support structure may employ one or more dampening elements or materials to interoperate with a seal support structure to dampen vibration to a seal system.

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 turbomachine housing element, having a flow channel for accommodating a rotor blade assembly and a first cavity at least partially produced by primary shaping; the first cavity being adapted for passive thermal insulation and/or the assemblable turbomachine housing element not being adapted for the active circulation of fluid through the first cavity; and/or, between the first cavity and the flow channel, a separate seal being attached to turbomachine housing element; and/or the first cavity extending in the axial direction of the flow channel over at least 20% of a minimum axial length of the turbomachine housing element at the level of the cavity and/or over a minimum axial length of the separate seal and/or being filled with air or a thermally insulating fluid, whose specific thermal conductivity λ is at least 10% lower than the specific thermal conductivity λ of air.

A brush seal for sealing a leakage gap in an axial flow path between a relatively higher fluid pressure region and a relatively lower fluid pressure region, that includes an outer housing and an inner housing located at least partially within and configured for radial displacement relative to the outer housing, wherein the upstream facing inner surface of the outer housing is free of any protrusion toward the first downstream facing outer contact surface and second downstream facing outer contact surface.

A sealing ring element of a turbomachine includes: a sealing portion with a first area and a second area, with the inner surface of a first area being at the same radial distance from the axis of the turbomachine. The sealing portion includes an annular cavity which opens into an inner surface of the second area and extends into the first area, the annular cavity defining an upstream lateral wall and/or a downstream lateral wall forming an angle which is strictly between 0 and 90°.

A turbine arrangement includes a turbine rotor arrangement, a turbine seal arrangement and a turbine outlet stator vane arrangement. Turbine rotor arrangement includes a rotor and a plurality of turbine blades that extend radially. Each turbine blade has a turbine shroud. Turbine seal arrangement is spaced radially around the turbine shrouds. Turbine outlet stator vane arrangement includes radially inner and outer annular members arranged coaxially and a plurality of vanes extending radially between the radially inner and outer annular members. The vanes are arranged downstream of the turbine blades. Liner is spaced radially inwardly from a radially inner surface of the annular member to define a chamber. Turbine shrouds and the upstream end of the liner are arranged such that in operation any leakage flow of gas between the turbine shrouds and the turbine seal arrangement flows into the chamber to manage the temperature of the radially outer annular member.