Abstract: A coupling for a turbomachine, including a first and a second disc attached along an axis, and extending in majority respectively from a first and from a second side of a median plane normal to the axis, the first and the second discs being provided respectively with first and second teeth mutually engaging through the median plane to form a curvic connection. Several first teeth each include a first protuberance jutting out towards the axis while extending from the second side; several second teeth each include a second protuberance jutting out towards the axis of rotation while extending from the first side; the first protuberances are spaced apart from the second protuberances along the axis; at least one blocking member is brought to be supported against the first and second protuberances to block axially the first element relatively to the second element.

Abstract: A labyrinth sealing joint for a turbomachine, for example of an aircraft, includes a rotor element extending about an axis (A), and a stator element extending around the rotor element, the rotor element having two annular lips extending radially outwards and surrounded by at least one abradable element carried by the stator element. A plurality of gas passage cavities are arranged circumferentially next to one another between the two lips which are interconnected by connecting partitions, wherein at least one of the lips has at least one axial gas passage opening that leads into at least some of the cavities, wherein the partitions extend substantially axially between the lips and define sectors of gas passage spaces between them, the sectors of spaces being divided by separation walls for forming the cavities.

Abstract: 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 (E5) 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 (Z51) of the abradable element, said densified annular zo

Abstract: A bladed assembly for a turbine of a turbomachine comprises an outer platform comprising, at an axial end, a transverse end surface and a sealing device protruding from the transverse end surface and comprising at least two sealing ribs (102) extending transversely to an axis of the bladed assembly and each having two respective sides which, in an axial cross-section, are radially and axially inclined in the same direction with respect to the axis. The radial direction of the inclination of the sides of the sealing ribs, in an axial cross-section, is the same for at least two of the sealing ribs. This particularity confers flexibility on each of the sealing ribs, which allows to favour a surface contact between the sealing ribs and a radial stator wall facing which the transverse end surface, and in particular the sealing device, are intended to be arranged.

Abstract: An assembly comprising a vent tube intended to be mounted inside a turbine shaft of a turbomachine. The vent tube has a radially inner annular part from which at least one annular centering part extends radially outwards. The centering part has a groove and a sealing part of a sealing member is mounted in the groove. The sealing part is designed to come into sealing abutment against an inner surface of the turbine shaft. The sealing member has a protective part covering at least one area of the centering part situated upstream and/or downstream of the groove.

Abstract: The invention relates to a turbine assembly (1) comprising an annular structure extending circumferentially about an axial direction (DA) and comprising ring segments (10) arranged circumferentially end to end and comprising adjacent connection faces (13a), linked by linked by sealing tabs (21, 22) in the wall (11) and in the flange (12). The invention is characterised in that the grooves (31, 32) and tabs (21, 22) are curved, the tabs (21, 22) having a bending degree of freedom starting from their mounting position in the presence of an air pressure exerted from upstream to downstream between the adjacent connection faces (13a, 13b) of the at least two adjacent ring sectors (10) during operation of the turbine, the tab (22) having a second point (220) which is in contact with a point (213) of the tab (21).

Abstract: A labyrinth sealing joint for a turbomachine, for example of an aircraft, includes a rotor element extending about an axis (A), and a stator element extending around the rotor element, the rotor element having two annular lips extending radially outwards and surrounded by at least one abradable element carried by the stator element. A plurality of gas passage cavities are arranged circumferentially next to one another between the two lips which are interconnected by connecting partitions, wherein at least one of the lips has at least one axial gas passage opening that leads into at least some of the cavities, wherein the partitions extend substantially axially between the lips and define sectors of gas passage spaces between them, the sectors of spaces being divided by separation walls for forming the cavities.

Abstract: Labyrinth seal for a turbine engine, in particular of an aircraft, including a rotor element rotating about an axis of rotation, and a stator element extending around the rotor element The rotor element includes a series of annular lips extending radially outwards and surrounded by at least one abradable element carried by the stator element. Each lip includes an inner peripheral body portion, an outer peripheral body portion and an upstream annular face for impact of an air flow during operation. At least one lip has, looking from the upstream annular face, a first annular cavity with a concave rounded cross-section on its inner peripheral body portion and a second annular cavity with a concave rounded cross-section on its outer peripheral body portion.

Abstract: The invention relates to a tab (5) for a sealing gasket device of a part of an aircraft gas turbine engine. The tab is arranged in respective slots of adjacent sectorized parts of the turbine engine nozzle. This tab has two bulging ends (51a, 51b) joined by a thinner intermediate part (51c) each having a connection with the other (53a, 53b), which can deform into a flat position by bracing the tab, when excess pressure is applied to its intermediate part from one side of the tab.

Abstract: An assembly comprising a vent tube intended to be mounted inside a turbine shaft of a turbomachine. The vent tube has a radially inner annular part from which at least one annular centering part extends radially outwards. The centering part has a groove and a sealing part of a sealing member is mounted in the groove. The sealing part is designed to come into sealing abutment against an inner surface of the turbine shaft. The sealing member has a protective part covering at least one area of the centering part situated upstream and/or downstream of the groove.

Abstract: The invention relates to a turbine assembly (1) comprising an annular structure extending circumferentially about an axial direction (DA) and comprising ring segments (10) arranged circumferentially end to end and comprising adjacent connection faces (13a), linked by linked by sealing tabs (21, 22) in the wall (11) and in the flange (12). The invention is characterised in that the grooves (31, 32) and tabs (21, 22) are curved, the tabs (21, 22) having a bending degree of freedom starting from their mounting position in the presence of an air pressure exerted from upstream to downstream between the adjacent connection faces (13a, 13b) of the at least two adjacent ring sectors (10) during operation of the turbine, the tab (22) having a second point (220) which is in contact with a point (213) of the tab (21).

Abstract: The invention relates to a coupling for a turbomachine, comprising a first and a second disc (17, 18) attached along an axis (AX), and extending in majority respectively from a first and from a second side (C1, C2) of a median plane (PM) normal to the axis (AX), the first and the second discs (17, 18) being provided respectively with first and second teeth (19, 21) mutually engaging through the median plane (PM) to form a curvic connection.

Abstract: A bladed assembly for a turbine of a turbomachine comprises an outer platform comprising, at an axial end, a transverse end surface and a sealing device protruding from the transverse end surface and comprising at least two sealing ribs (102) extending transversely to an axis of the bladed assembly and each having two respective sides which, in an axial cross-section, are radially and axially inclined in the same direction with respect to the axis. The radial direction of the inclination of the sides of the sealing ribs, in an axial cross-section, is the same for at least two of the sealing ribs. This particularity confers flexibility on each of the sealing ribs, which allows to favour a surface contact between the sealing ribs and a radial stator wall facing which the transverse end surface, and in particular the sealing device, are intended to be arranged.

Abstract: Labyrinth seal for a turbine engine, in particular of an aircraft, including a rotor element rotating about an axis of rotation (A), and a stator element extending around the rotor element (14), the rotor element including a series of annular lip(s) extending radially outwards and surrounded by at least one abradable element carried by the stator element, each lip comprising an inner peripheral body portion, an outer peripheral body portion, an upstream annular face of impact of an air flow during operation and a downstream annular face, wherein at least one lip includes, on the inner peripheral body portion thereof and/or the outer peripheral body portion thereof, through-orifices for the passage of air extending between the upstream and downstream annular faces.

Abstract: Labyrinth seal for a turbine engine, in particular of an aircraft, including a rotor element rotating about an axis of rotation (A), and a stator element extending around the rotor element, the rotor element including a series of annular lip(s) extending radially outwards and surrounded by at least one abradable element carried by the stator element, each lip including an inner peripheral body portion, an outer peripheral body portion (12b) and an upstream annular face for impact of an air flow during operation, wherein at least one lip has, looking from the upstream annular face, a first annular cavity with a concave rounded cross-section on its inner peripheral body portion and a second annular cavity with a concave rounded cross-section on its outer peripheral body portion.

Abstract: The invention relates to a tab (5) for a sealing gasket device of a part of an aircraft gas turbine engine. The tab is arranged in respective slots of adjacent sectorized parts of the turbine engine nozzle. This tab has two bulging ends (51a, 51b) joined by a thinner intermediate part (51c) each having a connection with the other (53a, 53b), which can deform into a flat position by bracing the tab, when excess pressure is applied to its intermediate part from one side of the tab.