Abstract: A cascade type thrust reverser device for a turbomachine of an aircraft, comprising a cascade including first partitions, second partitions intersecting the first partitions, and cavities, and a casing including a housing into which the cascade can be inserted in a first direction, the casing and the cascade being in relative translation with respect to one another in the first direction. The casing comprises a perforated wall intended to be in contact with an air flow and including orifices and wall strips with no orifices and intended to face the first walls of the cascade when the device is in a first position in which the cascade is entirely positioned in the housing.

Abstract: An acoustic attenuation liner includes a porous coating with a mesh structure, an acoustically absorbent material covered with the porous coating and having a first volume, and at least one cavity created between the acoustically absorbent material and the porous coating and which has a second volume, such as Vcav>3%, wherein the mesh structure has wires separating between them openings passing through the structure, some of which have, when flat, a minimum elementary surface area.

Abstract: A cascade type thrust reverser device for a turbomachine of an aircraft, comprising a thrust reverser cascade and a casing, the cascade including first cavities, and the casing comprising an opening defining a housing wherein said cascade can be inserted in a first direction, and the casing and said cascade being in relative translation with respect to one another in the first direction between a first position of the device in which the cascade is entirely positioned in the housing and a second position of the device in which said cascade is at least partially outside said housing. The casing comprises an acoustic treatment panel including second cavities extending in a second plane parallel to the first plane, each first cavity facing a second cavity when the device is in the first position to form an acoustic treatment cell.

Abstract: Internal structure for a turbomachine nacelle, the internal structure being intended to surround at least one part of a compartment capable of receiving a gas generator of the turbomachine, the internal structure comprising a ventilation cavity of said compartment, the ventilation cavity being provided with a main ventilation outlet and an auxiliary ventilation outlet separate to the main ventilation outlet, the internal structure comprising a shut-off member movable relative to the ventilation cavity between a flow position and a shut-off position in which the shut-off member shuts off the auxiliary ventilation outlet to a greater extent than in the flow position.

Abstract: A cascade for a thrust reversal device intended to be mounted on a turbomachine of an aircraft, the cascade including first partitions extending in a first direction, second fixed partitions extending in a second direction orthogonal to the first direction, and a frame inside which the first and second partitions extend, the frame including at least two fixed walls extending according to the first direction, and at least one part of each first partition extending between two second partitions. At least one first partition is mobile according to the second direction between a first position wherein the first partition is distant, in the second direction, from the fixed walls to form a plurality of resonating cavities with the first partitions and/or the fixed walls, and a second position wherein the one first partition is in contact with a fixed wall or another first partition.

Abstract: The invention relates to acoustic management, on an aircraft turbomachine (3,12) or on a nacelle (1,10), via a panel (30,32). On a support (38) is reserved a recess (34), recessed with respect to a surrounding general surface (36) for contact with moving air. The recess (34) is adapted to receive the panel, as another so-called surface for contact with moving air. The support (38) and/or the panel comprise removable connecting elements for, in the recess (34), mounting it removably with respect to the support, the panel being an acoustic panel or a non-acoustic panel.

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: A method of a coating by additive manufacturing on a turbomachine casing includes depositing on an internal surface of the turbomachine casing a filament of an abradable material to create a three-dimensional scaffold of filaments. A filamentary material deposition system is positioned from the internal surface of the casing; a first layer of the coating is deposited over 360?; a rotation of the filamentary material deposition system is carried out by a first predetermined angle and the filamentary material deposition system is positioned from the deposited layer; a second layer of coating is deposited on the first coating layer, on a sector of the casing; a displacement is carried out corresponding to the first sector already covered, then for the following sectors until 360° is covered; and after having carried out a rotation of the filamentary material deposition system.

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: A cascade for a thrust reversal device intended to be mounted on a turbomachine of an aircraft, the cascade including first partitions extending in a first direction, second fixed partitions extending in a second direction orthogonal to the first direction, and a frame inside which the first and second partitions extend, the frame including at least two fixed walls extending according to the first direction, and at least one part of each first partition extending between two second partitions. At least one first partition is mobile according to the second direction between a first position wherein the first partition is distant, in the second direction, from the fixed walls to form a plurality of resonating cavities with the first partitions and/or the fixed walls, and a second position wherein the one first partition is in contact with a fixed wall or another first partition.

Abstract: Internal structure for a turbomachine nacelle, the internal structure being intended to surround at least one part of a compartment capable of receiving a gas generator of the turbomachine, the internal structure comprising a ventilation cavity of said compartment, the ventilation cavity being provided with a main ventilation outlet and an auxiliary ventilation outlet separate to the main ventilation outlet, the internal structure comprising a shut-off member movable relative to the ventilation cavity between a flow position and a shut-off position in which the shut-off member shuts off the auxiliary ventilation outlet to a greater extent than in the flow position.

Abstract: A cascade for a thrust reverser device intended to be mounted on a turbomachine of an aircraft, the cascade including first partitions extending in a first direction, seconds partitions extending in a second direction orthogonal to the first direction, the second partitions extending in a third direction secant to a plane comprising first and second directions between a first and a second end. Each second partition forms, with reference planes parallel to the second direction and to the third direction, a plurality of different angles at positions distinct from the height of the second partitions separating the first end from the second end of the second partitions, each angle being formed between a reference plane and the tangent to the second partition taken at the intersection between said corresponding reference plane and the second partition at the position of the height of the corresponding second partition.

Abstract: A cascade type thrust reverser device for a turbomachine of an aircraft, comprising a cascade including first partitions, second partitions intersecting the first partitions, and cavities, and a casing including a housing into which the cascade can be inserted in a first direction, the casing and the cascade being in relative translation with respect to one another in the first direction. The casing comprises a perforated wall intended to be in contact with an air flow and including orifices and wall strips with no orifices and intended to face the first walls of the cascade when the device is in a first position in which the cascade is entirely positioned in the housing.

Abstract: A cascade type thrust reverser device for a turbomachine of an aircraft, comprising a thrust reverser cascade and a casing, the cascade including first cavities, and the casing comprising an opening defining a housing wherein said cascade can be inserted in a first direction, and the casing and said cascade being in relative translation with respect to one another in the first direction between a first position of the device in which the cascade is entirely positioned in the housing and a second position of the device in which said cascade is at least partially outside said housing. The casing comprises an acoustic treatment panel including second cavities extending in a second plane parallel to the first plane, each first cavity facing a second cavity when the device is in the first position to form an acoustic treatment cell.

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 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: The invention relates to acoustic management, on an aircraft turbomachine or on a nacelle, via a panel (30,32). On a support (38) is reserved a recess (34), recessed with respect to a surrounding general surface (36) for contact with moving air. The recess (34) is adapted to receive the panel, as another so-called surface for contact with moving air. The support (38) and/or the panel comprise removable connecting elements for, in the recess, mounting the panel removably with respect to the support, the panel being an acoustic panel (30) or a non-acoustic panel (32).

Abstract: The invention relates to acoustic management, on an aircraft turbomachine (3,12) or on a nacelle (1,10), via a panel (30,32). On a support (38) is reserved a recess (34), recessed with respect to a surrounding general surface (36) for contact with moving air. The recess (34) is adapted to receive the panel, as another so-called surface for contact with moving air. The support (38) and/or the panel comprise removable connecting elements for, in the recess (34), mounting it removably with respect to the support, the panel being an acoustic panel or a non-acoustic panel.

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 supply duct of a compressor of a turbine engine, formed from internal and external walls of revolution around an axis and opposite one another to define a circulation stream of a fluid, is provided. The stream allows the fluid to be routed from the inlet of the duct to the inlet of the compressor. The radius of the external wall at the inlet of the duct is greater than the radius of the duct at the inlet of the compressor. The duct includes a portion for which the radius of the external wall along the portion is less than the radius of the external wall at the inlet of the compressor, and the radius of the internal wall along the portion of the duct is less than the radius of the internal wall at the inlet of the compressor.