Abstract: A device for centering and guiding a shaft of an aircraft turbine engine is provided. The device includes an outer ring of a rolling bearing extending about an axis and having orifices, an annular bearing support extending about the axis and at least partially about the ring, the support having orifices and a series of studs for connecting the ring to the support. The studs can be distributed about the axis and extend parallel to the axis. The ends of first studs can be engaged without clearance in the orifices, and ends of second studs can be engaged with clearances in the orifices of the ring and/or of the support, the clearances being configured so that the device has different stiffnesses in at least two directions perpendicular to the axis.

Abstract: A blade for a turbomachine fan, the blade being made from a composite material containing a plurality of longitudinal conductive fibres embedded in an electrically non-conductive matrix, the blade having at least one fastening region to which an identification medium for identifying the blade is fastened, the fastening region containing longitudinal fibres oriented along a preferred fibre axis, the identification medium contains an item of identification data for identifying the blade and at least one radio antenna having at least one communication lobe configured so as to receive a read request and transmit the item of identification data in return, the communication lobe being oriented along a radio axis that is angularly offset from the preferred fibre axis by less than 45°, preferably by less than 22.5°.

Abstract: A rear portion of a convergent-divergent nozzle turbojet engine includes a device for controlling the position of a divergent flap relative to that of a convergent flap, including a first connecting rod having a first end portion hinged to the convergent flap and an opposite second end portion, a second connecting rod having a first end portion hinged to the divergent flap and an opposite second end portion, and a third connecting rod having a first end portion hinged to a synchronization ring and an opposite second end portion. The second end portion of the first connecting rod is hinged to the second end portion of the second connecting rod and/or to the second end portion of the third connecting rod, and the second end portion of the third connecting rod is hinged to the second end portion of the second connecting rod.

Abstract: The present invention relates to a blade comprising: a structure of aerodynamic profile comprising two mutually opposite skins; and a spar comprising a fibrous reinforcement obtained by three-dimensional weaving and densified by the matrix, the spar comprising a blade root portion extending outside the structure of aerodynamic profile and an airfoil portion arranged inside the structure of aerodynamic profile between the two skins. Moreover, within the blade root portion, the fibrous reinforcement of the spar comprises a non-debound region and at least two debound regions extending radially from the non-debound region so as to form at least four separate branches.

Abstract: An arrangement for an aircraft turbine engine, including a shaft and a follower element rotatably coupled to the shaft by a spline connection, the arrangement including upstream and downstream connections for radially centring the follower element relative to the shaft; means for spraying a lubricant into a collection cavity; a passage for receiving lubricant, which passage opens into the collection cavity and into a cavity for lubricating the splines which is partially defined by the upstream and downstream radial centring connections; and a passage for discharging lubricant, which passage opens into the cavity for lubricating the splines and outside the arrangement.

Abstract: The invention relates to an assembly between an aircraft pylon (30) and a turbomachine (20) of longitudinal axis (X), the pylon (30) and the turbomachine (20) each comprising front and rear longitudinal regions, the assembly comprising a rear support (40) configured to connect the rear region of the turbomachine (20) to the rear region of the pylon (30), the assembly being characterized in that the rear support (40) comprises a sliding pivot connection arranged between the rear region of the turbomachine (20) and the rear region of the pylon (30), so as to allow the turbomachine (20) only translational and rotational movements along and about the longitudinal axis (X) and along and about a vertical axis (Z) and a rotational movement about a transverse axis (Y).

Abstract: A metal core for hot-forming a titanium-based alloy metal component is disclosed. The metal core has on an outer surface, intended to come into contact with the metal component, a layer of metal carbonitride-enriched material. The metal core comprises a nickel- or cobalt-based alloy. The metal core comprising a steel coating having an outer surface intended to come into contact with the metal component, the steel coating having a layer of metal carbonitride-enriched material. Processes for manufacturing and regenerating the metal core and a process for hot-forming a metal component using the metal core are also disclosed.

Abstract: A turbomachine has the upstream portion of the annular combustion chamber is fixed to the outer housing via individual pins with a swivel connection. The swivel connection is established between the pin and a part of the outer housing. The pin is slidable radially, or with a radial component, by a slide established between the pin and a tubular part of the outer annular shroud and/or the annular wall of the chamber bottom.

Abstract: An accessory gearbox for an aircraft turbine engine having a radial shaft connecting a high-pressure body of the turbine engine to a gear train having at least one mixed shaft coupled to a high-speed device. The accessory gearbox has a clutch system having a low-speed motor mounted thereon and configured to mate with the mixed shaft such that: in an engaged position of the clutch, the low-speed motor is coupled to the mixed shaft so as to drive the high-pressure body at low speed when the turbine engine is stationary and, in a disengaged position of the clutch, the low-speed motor is decoupled from the mixed shaft so as to be protected when the turbine engine is in operation.

Abstract: The turbomachine subassembly comprises:—at least one tow-pressure nozzle ring, the nozzle ring comprising a root (38) integral with the nozzle ring;—a clamp (40) configured so as to fasten the root to a casing of the turbomachine; and—at least two pads (150) interposed between the root and the clamp, each pad being assembled in a form-fitting manner with one of either the root or the clamp.

Abstract: Assembly for a turbomachine blade, comprising: a fastener (9) defining a pocket (10) for receiving a blade root (11), and comprising two flanks defining between them a passage leading into the pocket (10) and forming stops preventing the root (11) from exiting from the pocket (10) via the passage, a shim (26a, 26b) adapted to be received in the pocket (10) with the root (11), the shim (26a, 26b) having a first support surface (28a, 28b) to bear on the root (11) and a second support surface (30a, 30b) opposite the first support surface (28a, 28b) to bear against the fastener (9), the support surfaces being fixed relative to each other and oriented such that movement of the shim (26a, 26b) in the pocket (10) relative to the fastener (9) varies the value of force exerted by the shim (26a, 26b) on the root (11).

Abstract: A method for generating an alarm is implemented by an alarm-generating device installed in an aircraft. It comprises: a step of obtaining information delivered by a sensor of said device and representative of an oil flow rate inside an oil filter of the aircraft; a step of obtaining a temperature delivered by a sensor of said device and representative of a temperature of the oil inside said filter; a step of obtaining a differential pressure at said filter; and a step of generating an alarm if the differential pressure exceeds an alarm threshold obtained on the basis of said information and of said temperature.

Abstract: A distributor for a turbine of a turbomachine. The distributor includes a radially inner platform, a blade and a root which is coated with a track made of an abradable material. According to the invention, the distributor includes a leakage gas reintroduction duct. The duct includes an inlet through the track made of an abradable material, an outlet opening through a downstream surface of the root, and a duct intermediate portion which extends from the inlet up to the outlet. The duct intermediate portion is orientated with a tangential component.

Abstract: A turbine engine nozzle extending about an axis and including a vane assembly mounted between an outer shroud and an inner shroud; and a position retaining device configured to be placed in an elastically prestressed state between the inner shroud and an inner platform of the vane assembly to apply a stress on the vane assembly towards the outer shroud. The contact between the inner shroud and the inner platform is obtained by one same surface of the position retaining device.

Abstract: A propulsion unit including: a gas generator including a compressor, a combustion chamber, a turbine, at least two compressed-air propulsion modules, each propulsion module including: a fan, a compressed-air turbine configured to drive the rotation of the fan, a manifold allowing the respective turbines of the compressed-air propulsion modules to be supplied with compressed air, wherein the manifold is configured to collect and mix: at least a portion of the flow that has passed through the combustion chamber of the gas generator, typically as it leaves the turbine, and at least one bypass flow, the bypass flow being a flow of air which is not passed through the combustion chamber of the gas generator.

Abstract: The turbine (12) for a turbine engine (100) extends around a major axis (X-X) and comprises: —a casing (4) comprising an annular hook (38), —a movably mounted impeller (16), —a ring (30) extending opposite the impeller in a direction radial to the major axis (XX), —a distributor (14) comprising a blade provided with a platform (11), the platform being extended radially outwards by a spoiler (22, 24), the spoiler (22, 24) being radially mounted on the hook (38), and —foils (40) each having a profiled trough shape in a direction circumferential to the axis, the foils extending in succession in the circumferential direction, each foil (40) extending between the spoiler (22, 24) and the hook (38). The turbine comprises stops (162) to prevent the foils (40) from moving in the circumferential direction. Each foil (40) comprises a protrusion (162) arranged so as to extend circumferentially opposite another protrusion (162) of another adjacent foil (40).

Abstract: A turbomachine turbine nozzle extending around a central axis, including at least one radially outer shroud, at least one radially inner shroud, and at least one blade made of ceramic matrix composite material, distinct from the radially inner shroud and from the radially outer shroud, and extending radially between the radially inner shroud and the radially outer shroud, the blade being hollow and including a cavity opening at a radially inner end and at a radially outer end of the blade, the nozzle including at least one tubular mast arranged in the cavity of the blade and allowing routing the ventilation air passing through the cavity of the blade, the mast including a radially outer end attached to the radially outer shroud, and a radially inner end cooperating with a radial flange for positioning the radially inner shroud.

Abstract: In order to fill a reservoir while avoiding laborious manual filling operations, there is added to the existing device, including a supply duct, a pump and a measuring probe, a stop valve designed to interrupt the filling as soon as the desired level has been reached, which is detected for example by an air intake which controls the closure of the valve. Two level indicators are added and make it possible to ascertain either that optimum filling has been achieved or that a malfunction is present in the system. Application to the systems for oil filling of aircraft engine reservoirs from another reservoir arranged within the aircraft, generally common to all the engines.

Abstract: A ventilation ring for a bearing support member of an aircraft turbine engine includes two or more tubular walls and one or more spacer walls. The two or more tubular walls extend opposite each other, and one or more of the tubular walls includes a main portion and a base, the base having a thickness greater than a thickness of the main portion. The one or more spacer walls connect the two or more tubular walls. The one or more spacer walls have one or more apertures extending into the base and a rim around the aperture. An entirety of the rim has a thickness greater than a thickness of the spacer wall at a distance from the rim.

Abstract: The present invention relates to a fan rotor with variable pitch blades, comprising a rotor disc, equipped at its periphery with a plurality of rotary fasteners (16), each fastener (16) comprising a cell (17) for receiving the root (150) of a blade (15).