Abstract: A unit for an aircraft thruster is provided that includes a first ferrule configured to be attached to a second ferrule and a first connection flange configured to be secured to a second corresponding connection flange secured to the second ferrule. The connection flange includes a portion forming a plane extending along a radial direction with respect to the first ferrule. A link between the first ferrule and the corresponding connection flange is offset to a distance from said plane corresponding to a length for damping a shear-generating flexural moment, generated by fluxes of forces passing through the first ferrule.

Abstract: The invention relates to a honeycomb core structure (202) for use in a structural panel for a jet engine nacelle, comprising at least one block with a honeycomb core (A, B) having a central part (5a, 5b) with core honeycomb cells (7a, 7b) and at least two lateral parts (9a, 9b) each comprising a number of honeycomb joining cells (11a, 11 b) of which at least one part of the honeycomb joining cells (11a, 11 b) has at least one additional wall (201a, 201 b) to form a join, the honeycomb core block(s) (A, B) being connected by at least one joint zone (213) formed by punching two additional walls (201a, 201b) superimposed one or the other and provided by honeycomb joining cells (11a, 11 b) belonging to the distinct lateral parts (9a, 9b). The invention further relates to a structural panel and a nacelle comprising such a structure (202) and a method for producing such a structure (202).

Abstract: The present disclosure concerns a component of an aircraft nacelle formed from at least one composite structure and one heating element. The nacelle includes frost protector, and the composite structure has a matrix reinforced by at least a material of which the heat conductivity at ambient temperature that is greater than or equal to 800 W·m?1·K?1, so as to provide transverse heat conductivity within the nacelle.

Abstract: The invention relates to a method for producing a part comprising a hollow body made from a composite material, using preforms (11, 12), said method including the following steps: a step in which the preforms are disposed around a suitable supporting element (1) such that the coupling edges of each preform (11, 12) overlap, and come into contact, with a coupling edge of another preform (12, 11); a step in which the coupling edges of each preform are joined along a seam; and a step in which the supporting element is removed and a mold core is inserted in its place before the assembly is placed inside a mold and resin is injected using an RTM/LRI method, prior to the polymerization of said resin.

Abstract: The present disclosure relates to a method for manufacturing composite parts and to manufacturing equipment. The present disclosure is useful in the field of aircraft construction. During a first phase, the part is designed, the geometry is computed, and the fabric architecture of the fabric preform and a mandrel for forming the first state of the preform are calculated. During a second phase, the preform fabric and the mandrel are produced. During a third phase, the mandrel is wrapped with the fabric, the preform is extracted from the mandrel in order to collapse a portion of the preform, and the preform is impregnated in order to obtain the desired part.

Abstract: A method of braiding reinforcing fibers on a mandrel (8) with a machine having a ring (9) carrying at least two series of reels of fibers, by moving the mandrel at a predetermined forward speed while moving the two series of reels along the ring (9) so that they cross while rotating in opposite directions and at a predetermined speed of rotation about an axis (AX) of the ring. The braid is formed on the mandrel (8) in the vicinity of a region of convergence (R) of the fibers that together define a conical shape (C). The method comprises: a step of reconfiguring the machine in which the angle (a2) at the apex of the cone (C) defined by the fibers takes on a new value (a2); and a step of restarting braiding in which the movement of the reels along the ring (9) and the forward movement of the mandrel (8) are re-established with a new speed of rotation and a new speed of advance.

Abstract: A thrust reverser for the nacelle of a turboreactor includes at least one fixed front frame that can be mounted downstream of a fan casing of the turboreactor and directly or indirectly supports at least one flow deviation vanes, in which at least part of the flow deviation vanes can be detached from the front frame and moved in translation independently therefrom when a maintenance operation is being carried out on the reverser.

Abstract: A nacelle for a turbojet engine is provided that includes an air intake upstream from the turbojet engine, a middle section intended to surround a fan of the turbojet engine and its case, and a downstream section. The middle section is made from at least one aerostructure capable of ensuring external aerodynamic continuity of the nacelle, and the aerostructure is rigidly bound, so as to be able to be taken apart, to at least one surrounding fixed structure.

Abstract: A flow straightening vane made of composite materials includes a core made up of three-dimensional woven textile rods which core is surrounded by a textile covering, this core and this covering being consolidated using a thermoset resin.

Abstract: A method of fabricating a connecting rod (6) out of composite material using three-dimensional fabric of woven reinforcing fibers, the connecting rod (6) extending in a main direction. The method comprises the steps of: cutting out one or more base pieces (7, 8) from a three-dimensional fabric; cutting out one or more reinforcing pieces (9, 10) from a three-dimensional fabric; installing the various pieces (7, 8, 9, 10) on a support in order to shape them; securing the various pieces (7, 8, 9, 10) to one another by stitching (12, 13, 14, 15) using reinforcing fibers in order to constitute a reinforcing fiber preform; installing the preform on a mandrel; and injecting resin into the preform and polymerizing the resin.

Abstract: A method for manufacturing a nacelle air intake lip including a plurality of components includes the following steps. The first step is positioning, inside a mold, two components within a set of fibrous layers encompassing all or part of the components so that a preform is formed. Next step is performing impregnation of the obtained preform to obtain molding of the nacelle air intake lip.

Abstract: A method for manufacturing an air-intake structure for a turbojet-engine nacelle includes two component parts, i.e. one air-intake lip structure and one outer cowl for the air-intake structure. In particular, the method includes the following steps: separately manufacturing the component parts such that each integrates one overlapping portion with a part to which is to be attached. The overlapping portion is configured such that an outer surface of the outer overlapping portion provides continuity with an outer surface of the outer cowl associated with the inner overlapping portion. To secure the overlapping portions, glue or a fastener may be used.

Abstract: The invention relates to a method of braiding reinforcing fibers on a mandrel (8) with a machine having a ring (9) carrying at least two series of reels of fibers, by moving the mandrel at a predetermined forward speed while moving the two series of reels along the ring (9) so that they cross while rotating in opposite directions and at a predetermined speed of rotation about an axis (AX) of the ring, the braid being formed on the mandrel (8) in the vicinity of a region of convergence (R) of the fibers that together define a conical shape (C), the method being characterized in that it comprises: a step of reconfiguring the machine in which the angle (a2) at the apex of the cone (C) defined by the fibers takes on a new value (a2); and a step of restarting braiding in which the movement of the reels along the ring (9) and the forward movement of the mandrel (8) are re-established with a new speed of rotation and a new speed of advance.

Abstract: An acoustic panel is provided that includes at least one cellular core arranged between at least an internal skin and an external skin, wherein the external skin incorporates at least one fastener able to collaborate in a disconnectable fashion with a complementary fastener associated with a structure to which it is to be attached.

Abstract: A turbojet casing adapted to receive a plurality of vanes, the casing including attachment means (21, 30) for attaching one end of each vane to the casing, the casing being characterized in that the attachment means extend on a face of the casing facing away from the vanes, the casing including orifices (26) for passing the ends of the vanes so that they can co-operate with the attachment means of the casing. A turbojet including such a casing.

Abstract: The invention relates to a method of making a composite material part, the method comprising the steps of: producing a mandrel; braiding a fiber covering (11) around the mandrel fitting as closely as possible to the shapes of the mandrel; locally applying at least one fiber patch (12a, 12b, 12c) on the braided covering; repeating the two steps of braiding and patch application as often as needed to obtain a preform; and delivering resin into the preform and polymerizing it.

Abstract: A turbojet vane, in particular a compressor guide vane, characterized in that the vane comprises the following elements: an elongate one-piece front portion (2) cut out from a pultruded section member comprising fibers bonded together by resin and forming a leading edge; an elongate one-piece rear portion (3) cut out from a pultruded section member comprising fibers bonded together by resin and forming a trailing edge; and a skin (5) based on resin-impregnated fibers extending between the leading edge and the trailing edge in order to form side faces of the vane and covering at least zones (6, 7) of the leading edge (2) and of the trailing edge (3); these elements being assembled together by hot compaction.

Abstract: The invention relates to a method of fabricating a composite material part in which one or more layers of braided, or woven, or indeed draped reinforcing fibers are applied onto a mandrel having the general shape of a body of revolution. After one or more layers have been applied onto the mandrel, an operation is performed of winding at least one tie helically around and along the assembly constituted by the mandrel and each layer of reinforcing fibers that it carries, so as to press each layer of reinforcing fibers against the mandrel. The invention relates to fabricating composite material parts such as beams, connecting rods, or arms made out of carbon fibers.

Abstract: The present invention relates to a cellular panel deployable from a compacted form to a deployed form, including, in the compacted form thereof, a plurality of consecutive sheets substantially parallel to one another and substantially perpendicular to an expansion direction, each sheet being discretely joined at a plurality of junction points with the following and/or preceding sheet, the junction points being substantially regularly spaced apart along the lines substantially parallel to the sheets and alternating with the preceding and/or following junction points, characterized in that, on at least one line parallel to the expansion direction, one or more junction points are excised or are not made and the portions of the sheets associated with said excised or unmade junction points are excised.

Abstract: The invention relates to a honeycomb core structure (202) for use in a structural panel for a jet engine nacelle, comprising at least one block with a honeycomb core (A, B) having a central part (5a, 5b) with core honeycomb cells (7a, 7b) and at least two lateral parts (9a, 9b) each comprising a number of honeycomb joining cells (11a, 11b) of which at least one part of the honeycomb joining cells (11a, 11b) has at least one additional wall (201a, 201b) to form a join, the honeycomb core block(s) (A, B) being connected by at least one joint zone (213) formed by punching two additional walls (201a, 201b) superimposed one o the other and provided by honeycomb joining cells (11a, 11b) belonging to the distinct lateral parts (9a, 9b). The invention further relates to a structural panel and a nacelle comprising such a structure (202) and a method for producing such a structure (202).