Abstract: The present invention relates to a method for repairing a damaged thermal protection covering a nacelle structure for an aircraft turbojet engine by adding a repair thermal insulation blanket as well as a repair sheet onto the damaged part.

Abstract: A turbomachine part or assembly of parts includes a first blade and a second blade, and platform from which the first blade and the second blade extend, The platform has, between a pressure side of the first blade and a suction side of the second blade a non-axisymmetric surface defining at least one fin, and a transverse section of the fin is asymmetric. The transverse section includes a first and a second oblique faces joining at a dorsal edge, the first oblique face is oriented toward the pressure side of the first blade, the second oblique face is oriented toward the suction side of the second blade, and the first oblique face has a steeper slope than the second oblique face.

Abstract: An aircraft landing gear assembly comprising: a structural pin defining a first bore having a longitudinal insertion axis; a cap for covering the bore, the cap comprising: a base portion; a head portion; a body portion positioned between the base portion and the head portion; and a bolt which extends between the head portion and the base portion and comprises a threaded portion so that engagement of either the head portion or the base portion with the threaded portion moves either the head portion or the base portion relative to the other, the body being formed of an elastically deformable material such that upon tightening of the bolt, the base portion moves towards the head portion and the elastically deformable material moves radially outwards such that a contact surface of the elastically deformable material engages with an inner surface of the first bore in order to retain the cap.

Abstract: A method of monitoring the condition of an aircraft landing gear shock absorber, the shock absorber including at least one spring chamber containing a gas, the method comprising taking a plurality of measurements of the gas pressure and temperature, each gas pressure and temperature measurement pair being taken of at the same instant relative to an operating cycle of the shock absorber; calculating, based on each pair of gas pressure and temperature measurements, a first value; storing the first value in a log; determining a value trend based on the log; and in response to determining that the value trend is outside a first range of values, generating a first notification signal.

Abstract: Devices for distributing oil to a rolling bearing for an aircraft turbine engine include a rolling bearing including two rings, respectively an inner ring and an outer ring, an oil distribution ring configured to be mounted on a turbine engine shaft, said distribution ring including a first outer cylindrical surface for mounting the inner ring of the bearing, an oil recovery scoop supplying a lubricating circuit of the bearing, and an annular track of a dynamic seal. The distribution ring and the track are formed by a single-piece body, and the lubricating circuit is formed in said body and extends into the distribution ring and the track.

Abstract: Described are temporary privacy modules for non-passenger seating areas of aircraft cabin arrangements. The temporary privacy modules selectively enclose a portion of the non-passenger seating areas. The temporary privacy modules include a support and a privacy device. The privacy device is movable relative to the support between a stowed position and a deployed position. In the deployed position, the privacy device encloses the portion of the non-passenger seating area and defines a temporary privacy room within the non-passenger seating area.

Abstract: A blade for an aircraft gas turbine engine includes, in a longitudinal direction, a blade root, a shank and an aerofoil body, the aerofoil body extending in the longitudinal direction between the shank and a blade tip and in a transverse direction between a leading edge made of metal material and a trailing edge. The blade includes a blade core made of composite material having a three-dimensional woven fibrous reinforcement forming the blade root, the shank and a part of the aerofoil body. The blade also includes a skin made of composite material having a two-dimensional woven fibrous reinforcement surrounding the aerofoil body part of the blade core, the skin being interposed between the leading edge made of metal material and a front edge of the aerofoil body part of the blade core to define a thinned leading edge portion, the skin including one or more two-dimensional woven plies.

Abstract: A rotary assembly for a turbomachine including a rotor including at least two consecutive rotor stages provided with a plurality of blades and an annular rotor shroud connecting the two consecutive rotor stages, a stator including: at least one stator stage provided between the two consecutive rotor stages including a plurality of vanes, a turbomachine stator vane root, an annular clamping part and an annular support of abradable material, the root extending radially and being clamped axially between the annular support of abradable material and the annular clamping part, A space separates a radially internal end of the root and the annular support of abradable material. A turbojet engine including a rotary assembly as previously.

Abstract: The present invention relates to an aircraft seat comprising: a guide rail (26); a rack (34); an actuator (30) provided with a gear (31); at least one connection portion (35) providing a mechanical connection between the guide rail (26) and the rack (34), characterized in that the connection portion (35) included a zone (36) of mechanical weakness and in that a first bearing element (37) and a second bearing element (38) are disposed on either side of the rack (34), such that, if the deformation of the seat generates a movement of the gear (31), said gear (31) or the corresponding shaft (32) comes into abutment against one of the bearing elements (37, 38) and in this way deforms the connection portion (35) along the zone (36) of mechanical weakness.

Abstract: Device for distributing oil from a rolling bearing (8) for an aircraft turbine engine, comprising: —a rolling bearing (8), —an oil distributor ring (5) and —an annular track (26) of a dynamic seal (22), characterised in that it further comprises a nut (16) screwed on to a thread (5d) of the distributor ring and bearing axially against an axial end of the inner ring so as to clamp it axially, and in that the annular track is configured to bear axially against the distributor ring and comprises rotating locking elements (27) engaging with the additional elements (28) of the nut.

Abstract: Device (30) for controlling an airflow guiding system (20), comprising at least one actuator (31) configured to translate a control rod (32) between a first and a second end position of a nominal operating range in which at least one vane (21a) of the airflow guiding system (20) can be moved between a first and a second angle, the control rod (32) being connected to the vane (21a) by a control lever (33) comprising a first control rod (36) and a second control rod (37) which are hinged together. The actuator (31) is configured to bring the control rod (32) into a safety position located beyond the second end position of the nominal operating range and to orient the vane (21a) at a safe pitch angle between the first angle and the second angle.

Abstract: A system for controlling the pitch setting of a propeller blade for an aircraft turbine engine is provided. The system generally including a blade having an airfoil connected to a root; a cup having an annular wall extending about a pitch axis of the blade and a lower axial end enclosed by a bottom wall; a locking ring that extends around the root and inside the cup; a safety element that ensures the retention of the root relative to the cup. The at least one safety element can have a generally elongate shape and can pass through aligned holes in the bottom wall of the cup and in the free end of the root.

Abstract: A non-return device for coaxial rotation transmission includes coaxial input and output shafts, a frame for guiding the rotation of the shafts. A locking element is urged in radial translation through a channel of the output shaft between a radial locking position, in which the locking element projects from the channel so as to prevent a rotation of the output shaft by abutting against the frame, and a radial unlocking position, in which the locking element is retracted so as to allow axial rotation of the output shaft. The input shaft includes means for radially switching the locking element between the locking and unlocking positions thereof.

Abstract: A component (1, 2) for supporting at least one bearing (3) for a turbine engine (10) comprising: two coaxial walls, internal (4) and external (5) walls respectively, defining a gas flow vein (6) between them and interconnected by a row of arms (7); an external ferrule (50) comprising an internal peripheral edge (51) connected to the external wall (5) and an external peripheral edge (52) connected to an external mounting flange (53); an internal ferrule (40) comprising an external peripheral edge (41) connected to the internal wall (4) and an internal peripheral edge (42) comprising an internal mounting flange (43); at least one of the ferrules (4, 5), which at the peripheral edge (41, 51) thereof is connected to the corresponding wall (4, 5), having a general shape which is corrugated about an axis (X-X) of the component (1, 2).

Abstract: A device for fabricating a three-dimensional part by selectively melting a powder bed, the device including a first tank for containing a first powder and provided with a first powder dispenser valve, a second tank for containing a second different powder and provided with a second powder dispenser valve, a first and a second monitoring device for monitoring the quantity of first powder delivered by the first valve and the quantity of second powder delivered by the second valve, a mixer chamber in communication with the first and second valves and including a third powder dispenser valve, and a mixer for mixing the powder particles in the chamber, a support for receiving the powder delivered by the third valve and on which the parts is to be fabricated, a powder spreader for spreading powder on the support, and a heater member for locally melting the powder spread on the support.

Abstract: A method for manufacturing a turbine engine vane a root connected to a blade extending in a longitudinal direction includes the steps of providing a root; and providing mold with a first cavity and a second cavity that together define a recess in which the vane is formed. The recess includes a first space in which the blade is formed and a second space in which the root is formed. The method further includes the steps of providing aluminum strips; positioning a fibrous reinforcement; arranging the vane root in the second space; and injecting a foam comprising aluminum or injecting an aluminum alloy into the first space of the recess of the mold such that the foam impregnates the fibrous reinforcement.

Abstract: This inertial navigation unit comprises a navigation data acquisition unit (1) and a first navigation stage (2) capable of calculating hybrid navigation values by combining navigation data.

Abstract: A surface heat exchanger for an aircraft turbomachine includes a support wall, a panel parallel to the support wall, partitions connecting the wall to the panel to define channels in which an air flow flows, and fins situated in the channels. The panel has a central part parallel to the wall and a downstream part that is inclined with respect to the wall. An upstream end is connected to the central part and a downstream end is situated at a distance from the wall and delimits with the latter a main outlet of the channels. The downstream part has fixed flaps disposed one after another so as to delimit between one another additional outlets of the channels.

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 method of determining a service interval for an actuator of a landing gear assembly is suitable for used with a landing gear assembly in which the actuator positions the landing gear assembly between a stowed position and a deployed position. The method includes the steps of determining an actuator travel value for a time interval and increasing a total actuator accumulated travel value by the actuator travel value for the time interval. The method further includes the step of comparing the total actuator accumulated travel value to a predetermined value.