Abstract: A bearing support journal (100) for an internal shaft of a gas turbine engine comprising a means of retaining a sealing sleeve surrounding said shaft, the retaining means comprising: a full radial annular flange (132) formed upstream on an axial cylindrical part (130) of the journal (100), said full flange (132) being intended to butt axially against a radial annular flange of the sealing sleeve, and tangential means (135) of immobilizing the sealing sleeve with respect to the journal (100), these being created downstream on the axial cylindrical part (130) of the journal (100). The invention also relates to assembly of a journal (100) and of a sealing sleeve.

Abstract: The invention relates to a turbojet compressor comprising a shroud and a plurality of blades, each fixed via its root to the shroud. Level with each blade root, the shroud presents a suction-side ramp and a pressure-side ramp extending from the trailing edge of the blade to beyond its leading edge. The pressure-side and suction-side ramps join together upstream from the leading edge of the blade to form a projection presenting a profile that is twisted about the axis of the shroud so as to force the gas stream flowing in each flow passage to go round the blade root essentially on the suction-side thereof, and the suction-side ramp presents a profile that is inclined relative to the outside surface of the shroud so as to deflect the gas stream flowing in the corresponding flow passage from the suction-side surface of the blade towards the pressure-side surface of the adjacent blade.

Abstract: The present invention relates to an aircraft turbomachine fan (1) comprising a turbomachine inlet cone (20) that has an exterior surface (19) intended to be closely followed by a flow section (22) of the turbomachine, an annular balancing flange (34) that rotates as one with a fan disk (4) about an axis of rotation (2) and is equipped with a plurality of weight-fixing holes (24), the fan also comprising one or more balancing weights (26) mounted fixedly on said annular balancing flange. Furthermore, the annular balancing flange (34) is positioned inside the inlet cone (20) in such a way that its weight-fixing holes (24) are isolated from the flow section (22). According to the invention, the flange (34) is produced as a single piece with a fan blade retaining ring (16).

Abstract: The present invention relates to the field of blades, particularly fan blades, intended for turbojets, particularly of the aeronautical type. Its objective is to provide a fan blade the features of which will enable the number of fan blades to be reduced while at the same time providing satisfactory performance. According to the invention, the fan blade comprises a plurality of aerodynamic parts which are superposed in a radial direction Z and the number of aerodynamic profiles varies from one aerodynamic part to the other.

Abstract: Wedging blades attached to the disk of a rotor having slots. According to the invention, each wedging spacer is made of an elastically deformable material and at least one longitudinal segment of the spacer presents a transverse profile of arcuate shape to define two lateral contact zones.

Abstract: A turbomachine fan (10) comprising blades (12) and inter-blade platforms (18) having radial flanges (20) fastened to flanges (22) of a rotor disk (16) by means of threaded rods inserted into bushes (26) mounted in orifices in the flanges of the platforms (18), wherein each bush (26) comprises at least one deformable part (30) interposed transversely between the threaded rod (24) and an edge of the orifice in which the bush (26) is mounted, this deformable part (30) being intended to allow the platform (18) to pivot in the event of an impact of a blade (12) against the platform (18) and to absorb some of the energy of the impact.

Abstract: A turbomachine rotor disk (32) comprising inter-blade platforms (44) fixed to ribs (16) delimited by cavities (14) in which blade roots (20) are retained, and protective liners (30) mounted between the flanks of the cavities (14) of the disk (32) and the blade roots (20), wherein the protective liners (30) have a C-shaped cross section so that they can be fitted translationally and retained radially on the ribs (16) of the disk (32) and constitute means of locking the platforms (44) on the ribs (16) of the disk (32).

Abstract: The present invention relates to a device for attaching a turbomachine stator vane in an annular casing (20) of the turbomachine, the vane comprising a tip formed of a platform (14) with two edges (141, 142), one of them forming an upstream edge (141) and the other forming a downstream edge (142), an airfoil (11) and a root (13), the annular casing (20) comprising a circumferential groove (23) formed in its interior wall with an upstream flank (21) and a downstream flank (22), the vane being held by engagement of the platform (14) in the groove (23), the upstream edge against the upstream flank (21) and the downstream edge (142) against the downstream flank (22) of the groove (23), the downstream flank (22) of the groove (23) being in the form of a slot (22) opening axially in the upstream direction, the downstream edge of the platform (23) being engaged in said downstream flank (22) and a fastening means (40) being attached in order to hold the platform (14) in the groove on the upstream side, wherein the va