Abstract: A turbomachine blade including a blade body made of fiber-reinforced organic matrix composite material, the blade body having a pressure side and a head provided with an apex; an erosion protection film which is made of a thermoplastic polymer and adhered to the blade body; a reinforcement layer which is made of a thermosetting polymer and positioned on the erosion protection film on the head of the blade.

Abstract: A method for fabricating a part by selective melting or sintering of powder beds by high energy beam, the method including a) providing a material in the form of powder particles; b) depositing a first powder layer on a support; c) scanning a region of the first layer with the beam to heat the powder locally to a temperature higher than the powder sintering temperature, such that the powder particles as melted or sintered form a first single-piece element; d) depositing a second powder layer on the first powder layer; e) scanning a region of the second layer with the beam to heat the powder to a temperature higher than the powder sintering temperature, so that the particles of powder as sintered or melted form a second single-piece element; and f) repeating d) and e) for each new powder layer laid over a preceding layer until the part is formed.

Abstract: The invention concerns a process for the assembly of an epicyclic or planetary gear train of a planet gear carrier comprising one or more bores, each of the bores being intended to receive a pivot (18), the process comprising the following steps: a) manufacturing of at least one pivot (18) comprising a real axis (46) distinct from a theoretical axis; b) measuring the position of a real axis of each of said one or more bores of the carrier; c) for each bore of the carrier, mounting a pivot (18) in said bore and orienting it angularly so that the eccentricity of the pivot (18) at least partially makes up for the eccentricity of said bore.

Abstract: A turbine ring assembly including a ring support structure and a plurality of ring sectors made of ceramic matrix composite material forming a turbine ring, each sector presenting in a first section plane defined by an axial direction and a radial direction of the ring a portion forming an annular base having, in the radial direction, an inner face and an outer face from which there extend two attachment tabs defining between them a circumferentially-open annular cavity, the structure including two radial tabs between which the attachment tabs are held, and at least one injection orifice for injecting a stream of cooling air into the annular cavity. In a second section plane containing the axial direction and a direction orthogonal to the first section plane, the orifice forms a first feed angle relative to the direction orthogonal to the first section plane, the first feed angle lying in the range ?80° to +80°.

Abstract: A guide vane in an outer section of a turbofan engine contains oil cooling cavities which are located on one side of a main bending axis of the vane so that the differential thermal expansions compensate for the bending deformations caused by the aerodynamic forces exerted on the outer vane surface.

Abstract: A turbine engine assembly including an air flow guide assembly, including at least one guide vane and at least one structural arm, the vane and arm extending radially about an axis. The arm includes an upstream end portion having a guide vane profile and including a leading edge aligned with that of the vane; a downstream portion; and an intermediate portion including an upper surface extending between an upstream end point and a downstream end point. The upstream end point is separated from the leading edge of the arm by an axial distance of between 0.2c and 0.5c, c being the length of the axial chord of the vane, and the angle of the tangent to the upper surface at the upstream end point is equal to that at the downstream end point ±1 degree.

Abstract: A planet carrier for an epicyclic speed reduction gear, comprising a torque transmission member of longitudinal axis A and an annular cage extending around the axis A and connected to a longitudinal end of the member, said cage comprising two flanks extending essentially radially with respect to the axis A and connected by bridges, seats extending axially between the flanks and being intended for supporting planet gears mounted rotatably about the seats, characterized in that said bridges comprise bars which cross one another essentially in the shape of an X, and of which the ends located on the side of said member are connected to said member by a ring extending around said axis A and comprising a continuous internal annular channel.

Abstract: The invention relates to a turbine assembly (1) comprising an annular structure extending circumferentially about an axial direction (DA) and comprising ring segments (10) arranged circumferentially end to end and comprising adjacent connection faces (13a), linked by linked by sealing tabs (21, 22) in the wall (11) and in the flange (12). The invention is characterised in that the grooves (31, 32) and tabs (21, 22) are curved, the tabs (21, 22) having a bending degree of freedom starting from their mounting position in the presence of an air pressure exerted from upstream to downstream between the adjacent connection faces (13a, 13b) of the at least two adjacent ring sectors (10) during operation of the turbine, the tab (22) having a second point (220) which is in contact with a point (213) of the tab (21).

Abstract: The invention concerns an aircraft propulsion system comprising at least one member (1, 3, 4, 8) in contact with a turbulent flow of a stream (F), characterised in that said member is covered, at least partially, by a piezoelectric structure (S) such as a piezoelectric film, said structure (S) comprising a grooved structure (5, 6, 7, 9) comprising a series of grooves, in contact with the flow of the stream, the grooves extending in the direction of flow of the stream, the grooved structure comprising at least one geometric parameter (h, s, w) configured to adapt depending on at least one parameter of the flow of the stream and/or an operating point of the propulsion system and/or an engine speed of the propulsion system.

Abstract: An outlet guide vane for a fan of a turbomachine is provided. The outlet guide vane includes a blade made of composite material and having a radial inner end, a first metal fastener portion fastened to the radial inner end, a first platform extending transversely with respect to the blade in the vicinity of the radial inner end. The first metal fastener portion is to be mounted on and fastened to an upstream flange and to a downstream flange of a hub. The first platform is monolithic with the first metal fastener portion or with the blade. A rectifier wheel and a turbomachine including the outlet guide vane are also provided.

Abstract: An assembly for the rear of a bypass turbomachine (10) comprises a primary nozzle (11) comprising a trailing edge defining a primary flow path portion and a secondary nozzle (110) defining a secondary flow path portion, defined about a longitudinal axis (X), said secondary nozzle being configured to eject a mixture of the flows coming from a secondary flow path (Vs) and from a primary flow path (Vp) of the turbomachine (10), the secondary nozzle being of convergent-divergent shape with a throat (112) corresponding to a minimum cross section of the nozzle (110), the secondary nozzle (110) comprising, at the throat (112), a periodic succession of lobes (116, 118) which are situated along the internal circumference of the secondary nozzle (110). The assembly also comprises a lobed mixer (130) at the downstream end of the primary nozzle (11), this having an alternation of hot lobes (134) extending inside the secondary flow path and of cold lobes (132) extending inside the primary flow path.

Abstract: A device for controlling the clearance between a rotor and a stator surrounding same, which is carried out by modifying the delivery flow carried in a rotor recess in order to provide a bleed flow that prevents the gases from the flow section from penetrating into the recess. A valve can partially close the delivery circuit, thus reducing the cooling of the rotor structure and allowing the expansion thereof and the reduction of the clearance with the stator, for certain operating speeds including cruise operating speeds. In addition, the delivery air is heated more at a lower flow rate, especially if it is made to follow a bypass provided with meanders in the hottest portions of the circuit.

Abstract: A ring for guiding variable-pitch blades and supporting an abradable coating for an aircraft turbomachine ring extends about an axis (X) and includes substantially radial orifices for mounting guiding sleeves intended each to receive a pivot of one of the blades. The ring may include a radially inner crown divided into sectors and supports the abradable coating. A radially outer annular ferrule may be divided into sectors and include the orifices for mounting the sleeves. In some embodiments, each ferrule sector is mounted about a crown sector by a circumferential sliding connection. The crown sector may include an immobilizer for rotationally immobilizing the sleeves of this ferrule sector about the axis.

Abstract: A method of fabricating an electronic power module by additive manufacturing, the electronic module including a substrate having an electrically insulating plate presenting opposite first and second faces, with a first metal layer arranged directly on the first face of the insulating plate, and a second metal layer arranged directly on the second face of the insulating plate. At least one of the metal layers is made by a step of depositing a thin layer of copper and a step of annealing the metal layer, and the method further includes a step of forming at least one thermomechanical transition layer on at least one of the first and second metal layers, the at least one thermomechanical transition layer including a material presenting a coefficient of thermal expansion that is less than that of the metal of the metal layer.

Abstract: The invention concerns a method for the surface treatment of a component, for example a turbomachine component, the component comprising a surface to be treated, the method comprising the following steps: loading a first dispenser with a chemical etching solution and a second dispenser with a rinsing solution, positioning the first dispenser and the second dispenser opposite the surface to be treated, moving the first dispenser and the second dispenser along the surface to be treated, such that the surface to be treated successively receives the chemical etching solution followed by the rinsing solution.

Abstract: A turbofan includes a fan, a casing positioned downstream of the fan and separating a primary flowpath from a secondary flowpath, a compressor, a combustion chamber and a turbine being arranged in the primary flowpath, the turbofan having a differential transmission coupled to the turbine, and a power supply device configured to provide additional power to the one provided by the turbine to drive the compressor.

Abstract: A fan includes: a fan disc; an inter-blade platform including a base and a radial tab, a second orifice being formed in the tab of the platform; and a lock having a downstream edge configured to bear against the tab of the platform. The one among the downstream edge and the yoke of the fan disc includes a pin, the other includes a first orifice, the pin being configured to enter the first orifice and the second orifice so as to block the platform relative to the fan disc.

Abstract: An overhead storage bin that includes a housing that defines a bin interior, a bucket or a door that is movable with respect to the housing between an open and a closed position, a projector positioned in or on the housing, and a display member positioned on the housing, wherein the projector is configured to project information onto the display member.

Abstract: The invention concerns a method for manufacturing a turbomachine flow straightener made from composite material having a blade extended by at least one attachment end, comprising the operations of: forming a fibrous blank from a three-dimensional fabric comprising warp threads oriented in a direction corresponding to the longitudinal direction of the blade to be manufactured; dividing at least one end of the fibrous blank by separating in order to constitute at least two sets of integrated layers; shifting at least one set of integrated layers in a planar configuration such that the warp threads are inclined in this set of integrated layers relative to the orientation of the warp threads in the fibrous blank portion of the blade.

Abstract: A tool for balancing a turbine engine module in a balancer, the module including a stator, a rotor and a shaft, the balancer including at least one power drive unit and a table equipped with two support of the shaft of the module. The tool includes at least: one stator support, fixed on the table of the balancer behind the supports, a tubular false bearing, configured to support a bearing for guiding the rotor in the stator and protruding outside of the stator, a guiding sleeve of the shaft, confining a support of the false bearing, and received on the supports of the balancer, a rear support plate of the stator, fixed to the support of the stator, and immobilising the rotor with respect to the stator, a front support plate of the stator, fixed to the stator support, and immobilising and centring the false bearing.