A rotor disc for a turbomachine, the disc extending circumferentially about an axis and including a plurality of cavities configured to receive blade roots, each cavity including a downstream radial wall configured to axially block the blade root in the cavity, each downstream radial wall including a channel of ventilation of the cavity, including an inlet orifice which opens into the cavity and an outlet orifice which opens onto a downstream surface of the disc. An assembly for a turbomachine including such a disc and an upstream retention ring and a turbomachine including such an assembly.

The invention relates to a vane (7) comprising: a structure made of composite material, a vane (7) root (16) attachment (9) comprising a base having a radial outer face (20) and in which a slot (24) configured to receive the vane (7) root (16) is formed and two platforms (26), extending on either side of the blade (18) opposite the radial outer face (20), and two stiffening parts (30), added and fixed on the radial outer face (20) on either side of the stilt (17) so as to tightly abut against the stilt (17).

A double ring axial seal for a disk and blade; the disk including an array of blade slots at a perimeter of the disk configured to receive a root portion of the blade, the blade includes a platform located between the root portion and an airfoil extending from the platform opposite the root portion, the platform including a receiver with an overhang proximate the root portion and the platform including a hook opposite the receiver proximate the root portion; a disk slot formed on a first side of the disk proximate the array of blade slots; a first axial seal ring coupled with the disk slot and the receiver; the first axial seal ring having a cutout proximate an outer perimeter of the first axial seal ring, the cutout sized to receive the overhang of the blade platform.

A rotor has formed therein a groove with which a moving blade including a dovetail portion and a platform portion connecting the dovetail portion and a blade portion engages. The groove opens in a surface intersecting the axis of rotation of the rotor, and includes a contacting portion extending obliquely with respect to the axis of rotation and coming into contact with the dovetail portion, a bottom portion of an end portion on the radially inner side of the rotor, a joining portion between the contacting portion and the bottom portion, a platform facing portion facing the platform portion, and a chamfered portion formed on an end surface in the groove-extending direction. The chamfered dimension of the joining portion is greater on the side with an acute angle between the groove and the end surface than on the side with an obtuse angle between the groove and the end surface.

The invention relates to an assembly for a turbine engine blade, the assembly comprising a fastener (9) defining a well (10) for receiving a blade root (7), and a clamping device (32) for clamping the blade root (7) against the fastener (9) when the blade root (7) is received in the well (10), wherein the clamping device (32) comprises: at least one clamping member (34; 36) suitable for bearing simultaneously on the blade root (7) and on the fastener (9) in the well (10) so as to exert a clamping force on the blade root (7), the clamping member (34; 36) defining a threaded hole, and a resilient layer (44) arranged to bear in the well on the blade root, a rod (38) comprising a thread engaging with the threaded hole so that a rotation of the rod (38) with respect to the clamping member causes a variation in the clamping force exerted by the clamping member on the root.

A method for manufacturing a turbo fan unit includes a step for preparing multiple fan blades and an other-side side plate, and a step for connecting each of the multiple fan blades to the other-side side plate by a welding process. In the preparing step, one of the fan blade and the side plate is prepared, in which a connecting-surface forming portion having a connecting surface and a welding projection protruded from the connecting surface is formed. The connecting surface connects one of the fan blade and the side plate to the other one of the fan blade and the side plate. In the connecting step, the welding projection is melted down and the connecting surface is connected to an opposing surface, which is the surface of one of the fan blade and the side plate and which is opposing to the connecting surface.

The invention relates to a retaining device for disassembling a bladed wheel of a turbine engine. The invention is characterized in that the retaining device (100) includes: an inter-blade ring sector (101), a plurality of retaining clips (105), each retaining clip (105) cooperating with one of the damping elements (40), the retaining clips (105) extending radially while being attached circumferentially to the ring sector (101).

A rotor blade is provided for a gas turbine engine. This rotor blade includes a rotor blade pair including a mount, a first airfoil and a second airfoil. The mount includes a forked body with a first leg and a second leg. The first airfoil is connected to the first leg. The second airfoil is connected to the second leg and arranged circumferentially next to the first airfoil.

A gas turbine engine includes a compressor section having a high pressure compressor, the high pressure compressor including an aft-most compressor stage. The gas turbine engine also includes a combustion section having a stage of discharge nozzles, the stage of discharge nozzles located downstream of the aft-most compressor stage and upstream of a diffuser cavity. The gas turbine engine also includes a high pressure spool drivingly coupled to the high pressure compressor, the high pressure spool forming in part a compressor discharge pressure seal and including a forward spool section. The forward spool section extends between the compressor discharge pressure seal and the aft-most compressor stage, the forward spool section defining an airflow cavity for providing a cooling airflow from the diffuser cavity to the aft-most compressor stage.

A flow inducer assembly and a method for cooling turbine blades of a gas turbine engine are presented. The gas turbine engine includes a rotor disk having circumferentially distributed disk grooves and turbine blades. Each turbine blade includes a blade root inserted into blade mounting section of the disk groove. Seal plates are attached to an aft side circumference of the rotor disk. The flow inducer assembly is integrated to each seal plate at a side facing away from the rotor disk. The flow inducer assembly is configured to function as a paddle due to rotation of the rotor disk and the seal plate therewith during operation of the gas turbine engine to drive ambient air as a cooling fluid into the disk cavity and enter inside of the turbine blade from the blade root for cooling the turbine blade.