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: Assembly (2) comprising an annular turbine distributor (3) for turbomachinery (1), an annular structural element (4) of the turbomachinery (1), coaxial with the distributor (3) and at least one device (5) for attaching the annular distributor (3) with the annular structural element (4), characterised in that: —the annular distributor (3) is crossed by at least two axial passages, and —the or each attachment device (5) comprises: at least two pins (15), rigidly connected to the annular structural element (4) and passing through one of said passages of the distributor (3), a first transverse stop (16) for stopping the distributor (3) axially and passing through two consecutive pins (15), and two second stops (17) rigidly connected to the structural element (4) to retain, in the pins (15), the first transverse stop (16) for stopping the distributor (3) radially.

Abstract: A support case for a turbomachine ring sector includes an annular body around a longitudinal axis of the case, and a downstream edge for attachment of the ring sector. The downstream edge extends from the body radially towards the inside of the case. The downstream edge extends circumferentially along a junction region between the body and a fixed vanes stage. The junction region includes a central portion and two lateral portions located on each side of the central portion, along a circumferential direction of the case with respect to its longitudinal axis. The central portion projects radially outwards from the case with respect to the lateral portions.

Abstract: A mobile vane for a turbine engine, including a root designed to be inserted into a receiving element of a rotor disk for a turbine engine, a platform carried by the root, and a blade extending from the platform. The platform includes an upstream edge. The upstream edge includes a lug for engaging in a locking notch of the disk in such a way as to hold the vane axially in relation to the disk, according to the longitudinal direction of the receiving element.

Abstract: A turbine blade for an aircraft turbomachine including a root, an airfoil and a platform inserted between the airfoil and the root and delimiting a gas circulation flowpath, the platform having two axial ends each forming an angel wing of which at least one has an internal cavity that will be supplied with air from the root of the blade. At least one of the two angel wings is drilled with at least one bleed hole for passage of a bleed flow from the internal cavity that will limit/prevent gas recirculation outside the flowpath.

Abstract: The present invention relates to an assembly (10) comprising: an annular nozzle (4) of a turbine of a turbine engine, a structural annular element (6) of the turbine engine, the nozzle (4) and the structural annular element (6) each comprising a radial flange (16, 18), the flanges being applied axially onto one another, and at least one member (12) for axially retaining the flange (16) of the nozzle (4) with the flange (18) of the structural annular element (6), with the member being applied to one of the flanges and being able to axially retain the two flanges one on top of the other.

Abstract: Assembly (2) comprising an annular turbine distributor (3) for turbomachinery (1), an annular structural element (4) of the turbomachinery (1), coaxial with the distributor (3) and at least one device (5) for attaching the annular distributor (3) with the annular structural element (4), characterised in that: —the annular distributor (3) is crossed by at least two axial passages, and —the or each attachment device (5) comprises: at least two pins (15), rigidly connected to the annular structural element (4) and passing through one of said passages of the distributor (3), a first transverse stop (16) for stopping the distributor (3) axially and passing through two consecutive pins (15), and two second stops (17) rigidly connected to the structural element (4) to retain, in the pins (15), the first transverse stop (16) for stopping the distributor (3) radially.

Abstract: The present invention relates to an assembly (10) comprising: an annular nozzle (4) of a turbine of a turbine engine, a structural annular element (6) of the turbine engine, the nozzle (4) and the structural annular element (6) each comprising a radial flange (16, 18), the flanges being applied axially onto one another, and at least one member (12) for axially retaining the flange (16) of the nozzle (4) with the flange (18) of the structural annular element (6), with the member being applied to one of the flanges and being able to axially retain the two flanges one on top of the other.

Abstract: The present invention relates to a turbine assembly (10) of a turbine engine (1), comprising at least: a first bladed rotor (12), a bladed stator (13) and a second bladed rotor (14) arranged in series, the rotors (12, 14) being mounted on a shaft (2); a sealing plate (20) extending between the stator (13) and the shaft (2) and separating a first recess (C1) arranged between the first rotor (12) and the stator (13), from a second recess (C2) arranged between the stator (13) and the second rotor (14); and pressure-reducing means (300, 31) positioned inside the first recess (C1), the assembly being characterized in that said pressure-reducing means (300, 31) comprise a plurality of substantially radial recompression fins (300) extending into the first recess (C1).

Abstract: A support case for a turbomachine ring sector includes an annular body around a longitudinal axis of the case, and a downstream edge for attachment of the ring sector. The downstream edge extends from the body radially towards the inside of the case. The downstream edge extends circumferentially along a junction region between the body and a fixed vanes stage. The junction region includes a central portion and two lateral portions located on each side of the central portion, along a circumferential direction of the case with respect to its longitudinal axis. The central portion projects radially outwards from the case with respect to the lateral portions.

Abstract: The turbine (1) has a rotor (2) with blades (4) arranged on a periphery of disks (5) in radial planes (3), and a tie rod (6) extending along slots (7) in the disks. An annular spacer (15, 15?) fixes a position of the tie rod relative to a center line (M) of the disks, and is arranged in a channel (10, 11). The spacer has through-openings, which are arranged radially relative to the tie rod or to its center line, and extend coaxially. The channel carries a cooling medium and is limited by a separating pipe (13, 14) radially outwards. The through-openings serve for flow of the medium.

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 trunnion (23) for a high-pressure turbine (11), to be arranged between a shaft of a low-pressure turbine (9) and an inner surface (34) of a seal mounting (26) of the low-pressure turbine (8), the trunnion (23) being characterized in that it includes a drop-launching extension (32) arranged such as to extend opposite a flared portion (33) of the inner surface (34) of the seal mounting (26), such that when the trunnion (23) is rotated about the shaft of the low-pressure turbine (9), oil (H2), which tends to penetrate between the trunnion (23) and the seal mounting (26), is thrown by centrifugal effect from the drop-launching extension (32) toward the flared portion (33) of the inner surface (34) of the seal mounting (26).

Abstract: A turbine wheel for a turbine engine, comprising a disk carrying blades each having a platform, carrying an impeller connected by a stilt to a root. Each blade comprises an upstream radial wall and a downstream radial wall extending inwards from the platform of the blade. The roots of the blades are engaged into slots on the periphery of the disk, so that the radial walls of the blades and the platforms of the blades are circumferentially arranged end-to-end, and define inter-blade cavities radially inside the platforms, each accommodating an inter-blade sealing and vibration damping member. At least one member comprises an elastically deformable element configured so as to be elastically prestressed against the walls of a cavity in the member mounting position.

Abstract: A rotating assembly for a turbine engine, comprising a disc having an outer periphery having alternating slots and teeth, blades radially extending from the disc and roots of which are axially engaged in the slots, with spaces called slot cavities being provided between the roots of the blades and the slots, platforms laterally extending from the blades and circumferentially arranged end-to-end, so as to form spaces called inter-blade cavities, and a downstream annular shroud, comprising an outer annular sealing lip opposite the downstream ends of the platforms. The downstream shroud further comprises an intermediate annular sealing lip opposite the downstream faces of the teeth of the disc, radially between the slot cavities and the inter-blade cavities.

Abstract: The present invention relates to a turbine assembly (10) of a turbine engine (1), comprising at least: a first bladed rotor (12), a bladed stator (13) and a second bladed rotor (14) arranged in series, the rotors (12, 14) being mounted on a shaft (2); a sealing plate (20) extending between the stator (13) and the shaft (2) and separating a first recess (C1) arranged between the first rotor (12) and the stator (13), from a second recess (C2) arranged between the stator (13) and the second rotor (14); and pressure-reducing means (300, 31) positioned inside the first recess (C1), the assembly being characterised in that said pressure-reducing means (300, 31) comprise a plurality of substantially radial recompression fins (300) extending into the first recess (C1).

Abstract: A mobile vane for a turbine engine, including a root designed to be inserted into a receiving element of a rotor disk for a turbine engine, a platform carried by the root, and a blade extending from the platform. The platform includes an upstream edge. The upstream edge includes a lug for engaging in a locking notch of the disk in such a way as to hold the vane axially in relation to the disk, according to the longitudinal direction of the receiving element.

Abstract: The invention relates to a turbomachine (30) comprising: a compressor stage and a turbine stage, each stage comprising at least one disk (42); and a tubular shaft (31) sleeve (33) extending along the axis (32) of the turbomachine, wherein the sleeve (33) comprises at least one tab (40) extending from an outer radial surface (41) of the sleeve and facing the disk (42), the tab (40) being designed to come into contact with the disk (42) when the sleeve (33) is in rotation about the axis (32) of the turbomachine.

Abstract: A turbine blade for an aircraft turbomachine including a root, an airfoil and a platform inserted between the airfoil and the root and delimiting a gas circulation flowpath, the platform having two axial ends each forming an angel wing of which at least one has an internal cavity that will be supplied with air from the root of the blade. At least one of the two angel wings is drilled with at least one bleed hole for passage of a bleed flow from the internal cavity that will limit/prevent gas recirculation outside the flowpath.