Abstract: A manufacturing method of a turbomachine airfoil, such as an outlet guide vane airfoil, comprising positioning a first fibrous wall preform on a first mold portion, placing at least one core on the first wall preform, positioning a second fibrous wall preform on the core, assembling a second mold portion to the first mold portion so as to form a mold around the first and second wall preforms, applying a hardening treatment to the first and second wall preforms, removing the core, and positioning a reinforcing structure between the first wall preform and the second wall preform.

Abstract: Device for perforating an acoustic annular panel for a turbine engine, wherein it includes a perforation needle support ring, the ring having an axis A of revolution and being intended to be inserted substantially coaxially inside the panel, the ring being sectored and including ring sectors regularly distributed around the axis and each one radially mobile opposite the axis from a radially inner rest position up to a perforation position of the radially outer panel, and at least one rotating actuator of rotation axis A and mechanically connected to the ring sectors such that a rotation of the actuator leads to a radial movement of the ring sectors.

Abstract: An acoustic treatment panel intended to be disposed on at least one wall of a turbojet engine in contact with a fluid flow, the panel including a first acoustically reflective plate, a second plate and a plurality of cavities mounted between the first plate and the second plate and including a plurality of cells. The second plate is a one-piece plate through which a plurality of channels pass, each opening out on the one hand onto a first orifice formed on a first face of the second plate and on the other hand onto a second orifice formed on a second face of the second plate, the length of each channel extending between its first orifice and its second orifice being greater than the thickness of the second plate.

Abstract: An assembly for a rear of a dual-flow turbomachine having a longitudinal axis includes a secondary nozzle and a heating system. The secondary nozzle is defined about the longitudinal axis and ejects a mixture of the flows coming from a secondary vein and a primary vein of the turbomachine. The secondary nozzle is of a convergent-divergent form with a neck corresponding to a minimal cross-cross-section of the secondary nozzle. The heating system is located on a portion of the internal circumference of the secondary nozzle longitudinally in the region of the neck or upstream from the neck.

Abstract: A nacelle for a turbomachine is provided and includes an inner wall that is annular about a longitudinal axis of the nacelle, the annular inner wall being designed to surround part of the turbomachine. The nacelle further includes an annular outer wall surrounding the annular inner wall. The annular outer wall includes a first acoustically porous part, and the annular inner wall including a second acoustically porous part. The first and second porous parts are arranged facing one another so as to allow soundwaves, emitted by the turbomachine housed in the nacelle, to pass through the annular inner wall and then the annular outer wall so as to escape from the nacelle.

Abstract: An airplane propelled by a turbine engine having at least one fan, the turbine engine being integrated in the rear of a fuselage of the airplane, extending it rearwards, the airplane further including at least one acoustic baffle forming panel connected to the fuselage of the airplane and arranged below the turbine engine.

Abstract: An aircraft including a wing and an aeroengine having a longitudinal main axis and including a ducted nacelle with at least one fan is disclosed. The wing includes an acoustic treatment surface on a bottom portion of its outer shell on either side of the longitudinal main axis and over a width perpendicular to the longitudinal main axis that is not greater than three times the diameter of the fan. The acoustic treatment surface includes a layer of porous material for attenuating acoustic waves coming from the fan.

Abstract: A method of a coating by additive manufacturing on a turbomachine casing, including depositing on an internal surface of the turbomachine casing a filament of an abradable material to create a three-dimensional scaffold of filaments forming an ordered array of channels.

Abstract: The invention proposes an assembly for a rear of a dual-flow turbomachine (10) having a longitudinal axis (X), comprising: a secondary nozzle (110) defined about the longitudinal axis (X), said secondary nozzle being configured to eject a mixture of the flows coming from a secondary vein (Vs) and a primary vein (Vp) of the turbomachine (10), the secondary nozzle being of convergent-divergent form with a neck (112) corresponding to a minimal cross-cross-section of the secondary nozzle (110), a heating system located on at least one portion of the internal circumference of the secondary nozzle longitudinally in the region of the neck and/or upstream from the neck (112).

Abstract: A flow-straightener vane of a bypass turbomachine includes a plurality of vane sections stacked radially with respect to a longitudinal axis (X) along a stacking line (L) between a root end and a tip end. Each vane section has a pressure-face surface and a suction-face surface extending axially between an upstream leading edge and a downstream trailing edge. Between the leading and trailing edges of each vane section there is formed a profile chord (CA) the length of which is substantially constant between the tip end and the root end, and the stacking line (L) exhibits a curvature in a plane passing more or less through the axis (X) and through the stacking line (L), situated in the vicinity of the tip end and oriented from downstream towards upstream.

Abstract: An aircraft propulsion unit including two distinct fan spools, and including a first duct and a second duct extending downstream of the two fan spools, the propulsion unit including a control ring disposed on one of the ducts, downstream of one of the fan spools, with at least one annular internal wall extending in an interior space of the duct, the control unit of the control ring being configured to modify the shape of the internal wall, and to cause an air passage cross section in the duct to vary at the ring, a device for acquiring acoustic signals generated by acoustic waves propagating downstream of the fan spools, and a device for processing the acoustic signals, configured to measure a dephasing between the acoustic waves and to control the control ring depending on the dephasing.

Abstract: A propulsion assembly for aircraft, the assembly including a turbojet having at least one unducted propulsion propeller; and an attachment pylon for attaching the turbojet to a structural element of the aircraft, the pylon being positioned on the turbojet upstream from the propeller and having an airfoil extending transversely between a leading edge and a trailing edge, the trailing edge of the airfoil of the pylon includes a cutout extending longitudinally over a fraction of the trailing edge facing at least a portion of the propeller, the cutout being configured to increase locally the distance between the trailing edge and the propeller, the cutout presenting an outline having a curved shape presenting at least two points of inflection.

Abstract: An acoustic panel for an aircraft propulsion unit, comprises a core and an entry layer in contact with the core, the entry layer comprising lower elements and upper elements, each comprising an inner face, an outer face and at least one through-hole, the inner faces of the lower elements and upper elements being in contact with each other and forming a channel defining a baffle and opening into a cell of the core.

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 turbomachine vane includes an electroacoustic source including two membranes fixed on a support arranged such that the membranes and the support delimit a first cavity. The membranes are arranged on two opposite sides of the first cavity between the two aerodynamic surfaces of the vane. A first of the aerodynamic surfaces includes a first region arranged facing a first of the membranes for the passage of acoustic waves. The membranes and membrane vibration device are configured such that the membranes vibrate in phase opposition and along a same emission direction by applying forces to the support such that the resultant force is approximately zero. The mechanical energy lost in the form of vane deformations due to membrane vibrations can thus be minimized.

Abstract: An acoustic absorber cell (1) for mounting on a turbojet wall in contact with a fluid flow, the cell (1) having a resonant cavity (2) presenting a wall (30) having at least one orifice (4) associated with a neck (5) having a tube (51) extending inside the resonant cavity (2) and projecting from the wall (30). The neck (5) also having a tubular ring (52) coaxial with the tube (51) and movable relative to the tube (51), and the cell (1) includes electromagnetic actuator (6) for moving the ring (52) along the tube (51), and controller (7) for controlling the electromagnetic actuator (6) and configured to control movement of the ring (52) along the tube (51) as a function of the operating speed of the turbojet between at least first and second positions defining respective first and second resonant frequencies of the resonant cavity (2).

Abstract: A turboprop including a propeller including a blade extending in a direction, which also includes a root, a leading edge, a trailing edge, and a wing tip, and an inner air stream channel, wherein the inner air stream channel includes an inlet located at the root of the blade and an outlet leading to the trailing edge of the blade transversely directed in relation to the main elongation direction, such that an inner stream of air flowing in the inner air stream channel by entering via the inlet adjacent to the root of the blade is discharged via the outlet adjacent to the trailing edge of the blade by forming a stream of blown air that moves away from the trailing edge in a direction which is transverse to the main elongation direction and which has a component in the direction of a skeleton line of the blade at the trailing edge.

Abstract: A blade for a turbine engine propeller, in particular a propfan engine, comprising a protruding part on the leading edge thereof, wherein said blade comprises means for controlling the position of the protruding part along the leading edge thereof.

Abstract: A stator designed to be placed radially in a flow which passes through one or more rotors which share the same axis of rotation, with a leading edge and a trailing edge. The leading edge and trailing edge are connected by a lower face and an upper face, wherein at least one of the faces of the stator has radial undulations which extend axially from the leading edge to the trailing edge. The radial undulations can have at least two bosses in the same azimuth direction, the amplitude of which is at least one centimeter on at least part of the axial length of the stator. A propulsion assembly formed by the rotor and the stator, and to a turbine engine comprising such assembly is also provided.

Abstract: The invention relates to an acoustic attenuation liner comprising: a porous coating (39) with a mesh structure, an acoustically absorbent material (41) covered with the porous coating and having a first volume (Vabs) and at least one cavity (45) created between the acoustically absorbent material (41) and the porous coating (39) and which has a second volume (Vcav), such as Vcav>3% (Vcav+Vabs), wherein the mesh structure has wires (390) separating between them openings passing through said structure, some of which have, when flat, a minimum elementary area So.