An aircraft turbomachine rotor comprising a rotor disc extending transversely with respect to a longitudinal axis X, a plurality of blades and a damping device comprising: a support ring configured to extend transversely with respect to the longitudinal axis X and configured to be mounted on the outer periphery of the rotor disc, and a plurality of damping members that are secured to the support ring and extend projecting upstream from the support ring, each damping member being configured to extend at least under a platform of a blade so as to exert a radially outward force so as to damp the radial movement of said blade when the turbomachine is in operation.

A method for disassembling/assembling a gas turbine including a seal plate disposed on a first side of a rotor disc in an axial direction of the rotor disc, and a seal plate restraint part for restricting movement of the seal plate in a radial direction of the rotor disc. The method includes a seal-plate-restraint-state switching step of operating the seal plate restraint part from a second side in the axial direction via a clearance between a platform of a blade and a region of an outer peripheral surface of the rotor disc except a blade groove for receiving the blade to switch between a seal plate non-restraint state where the seal plate restraint part does not restrict movement of the seal plate in the radial direction and a seal plate restraint state where at least a part of the seal plate restraint part protrudes in the axial direction.

A rotor of a gas turbine, having two adjacent rotor disks having a plurality of blade-holding grooves for receiving rotor blades, distributed around the periphery thereof, and having an axially extending peripheral ring projection radially beneath the blade-holding grooves. A peripheral rotor component is fixed to the ring projections, between the rotor disks. In order to protect the periphery, the rotor disk or the rotor component includes at least two recesses arranged on the periphery in a distributed manner, in each of which engaging shoulders of the rotor component or the rotor disk engage.

A turbine blade including an aerofoil and a shroud. The shroud includes a first abutment surface configured to face a second abutment surface of a first circumferentially adjacent turbine blade. The shroud further includes a second abutment surface configured to face a first abutment surface of a second circumferentially adjacent turbine blade. The shroud further includes an inner platform surface extending at least circumferentially between the first abutment surface and the second abutment surface. The shroud further includes a first recessed surface extending at least radially and circumferentially from the first abutment surface to the inner platform surface. The first recessed surface defines a first recessed region configured to receive a flow of a cooling fluid from the first circumferentially adjacent turbine blade.

The present disclosure relates generally an additive manufactured rotor for a seal assembly. The rotor surrounds a central axis and extends along the central axis from a first axial end to an opposite second axial end. The rotor is rotatable about an axis of rotation that is co-axial with the central axis. The rotor can include an outer ring and an inner ring that is connected to the outer ring by a radial web. The outer ring has a rotor face for providing a sealing surface at the first axial end. The outer ring can define a first annular cavity and the inner ring can define a second annular cavity that is radially separated from the first annular cavity. The radial web can define a fluid channel for providing fluid communication between the first annular cavity of the outer ring and the second annular cavity of the inner ring. The first and second annular cavities can be concentric with the axis of rotation.

A method for manufacturing a stage of a steam turbine comprising the steps of milling a block of material to define a sector having a plurality of blades, each blade having an external surface; machining an opening in the external surface of at least one of the blades; machining a cavity in fluid communication with the opening; the step of machining the cavity being performed by wire electric discharge machining.

An inner band assembly for at turbine nozzle of a rotary machine that includes a centerline axis, the inner band assembly comprising a platform portion, a first flange portion coupled to the platform portion and obliquely oriented with respect to the centerline axis, and a second flange coupled to the first flange and obliquely oriented with respect to the first flange. Wherein the platform portion and the first flange intersect.

A rotor blade for a gas turbine, in particular an aircraft gas turbine, including a blade root, a blade neck that adjoins the blade root in the radial direction, an airfoil that adjoins the blade neck in the radial direction, a radially outer partition wall that forms a radially inner delimiting section of an annular space of a gas turbine, an axially front partition wall and an axially rear partition wall that are connected to the radially outer partition wall so that the partition walls surround the blade neck on three sides, the partition walls protruding beyond the blade neck in the circumferential direction. For placement in a blade root receptacle of a rotor disk, the rotor blade is provided with a blade root protective plate that is situated between the blade root and the rotor disk. The blade root protective plate includes at least one sealing section that extends in the axial direction from the front partition wall to the rear partition wall, and whose radial outer side is situated opposite from the radially outer partition wall when the blade root protective plate is situated at the blade root. One or multiple ribs are situated at the blade neck for supporting the sealing section and are integrally joined to the blade neck.

A vibration suppression device for a rotary machine according to at least one embodiment of the present disclosure is a vibration suppression device for a rotor of a rotary machine including a damper pin movably provided inside a gap of the rotor, the damper pin including a magnet, and a magnetic force generation portion provided in the rotor at a periphery of the gap. The magnetic force generation portion is configured to exert, against the magnet, a magnetic force in a direction pushing the damper pin away from a stick region of the damper pin located on a radially outward side of the rotor in the gap.

A turbine blade for use in a gas turbine engine includes localized thickening of a neck pocket of the turbine blade to meet strength requirements while minimizing the weight of the turbine blade. More specifically, a convex spline is positioned within a suction side neck pocket of the turbine blade adjacent a leading edge of the neck pocket to increase the strength of the blade.