Abstract: An apparatus includes a fitting connected to a body. The fitting has an engagement portion including a channel formed into a peripheral surface of it and an axis. The body includes an axial layer provided around at least some of the peripheral surface of the engagement portion of the end fitting and a hoop layer provided around the axial layer at a location corresponding to at least some of the channel so as to urge at least some of the axial layer into a portion of the channel. The channel is arranged such that at least some of it defines a pathway that is non-orthogonal with respect to the axis of the fitting, such that at least some of the surface area of the channel defines a load bearing surface for the transfer of torque between the fitting and body when a rotational force is applied to one of them.

Abstract: The invention relates to a method for manufacturing an arm (16) in composite material, this arm comprising a main body incorporating a bearing (22) designed to receive a shaft passing through it transversely, this method comprising the steps of: manufacturing a hub (11) by winding reinforcing fibers around a generally tubular support (12); manufacturing at least one sleeve (14); attaching the hub (11) to one end of the sleeve (14) and transversely to the sleeve (14) in order to form a mandrel; applying one or more layers of braided reinforcing fibers (21) around this mandrel with a machine for braiding reinforcing fibers (19); injecting and polymerizing resin into the layer or layers of braided fibers (21). The invention applies notably to an arm forming an undercarriage rocker arm of an aircraft.

Abstract: A method of manufacturing a mechanical component, a mechanical component thus obtained, and a winding device designed for implementing the method of manufacture. The mechanical component includes at least one insert made of metal matrix composite, within which matrix ceramic fibers extend, the composite insert being obtained from a plurality of coated filaments each including a ceramic fiber coated with a metal sheath, the method including manufacturing at least one insert by winding a bonded lap or bundle of coated filaments about a component of revolution. At least some of the winding is performed in a rectilinear direction.

Abstract: A method reinforcing an axisymmetric annular metal part by including a winding of composite material. A metal blank for the part is prepared, a cavity is formed therein that opens out into a coaxial inside face thereof, and that presents a right cross-section of axial extent that decreases from the inside towards the outside, a reinforcing yarn is wound in the cavity, the cavity is closed, the assembly is subjected to a hot isostatic compression process, and the blank is machined to obtain a final part.

Abstract: A method of fabricating a mechanical member for aircraft, including a plurality of operations of braiding and depositing layers of braided reinforcing fibers on a mandrel (11) by using braiding machine. Each operation comprises braiding a braided layer and depositing it by moving the mandrel (11) along a central axis of the braiding machine. Each of the various superposed braided layers comprises both longitudinal fibers (12, 12G) that are parallel to a main direction of the mandrel (11), and interlacing fibers that are inclined. At least one operation is configured to form and deposit a braided layer having, in at least one cross-section of the member, a density of longitudinal fibers that differs depending on whether consideration is given to one angular region (S1) or another angular region (S2) of the same extent around the center of gravity (G) of the mandrel (11) in the section under consideration.

Abstract: A method of fabricating an aircraft landing gear rocker beam comprising a body having a clevis (24) with two lugs (19, 21) parallel to a direction that is radial relative to the body, the method comprising the steps of: forming a first mandrel; applying one or more first braided fiber layers (11) on the first mandrel; fitting a second mandrel element on the assembly as constituted in this way, the second mandrel element extending between the two lugs (19, 21) of the clevis (24) that is to be formed; applying one or more second braided fiber layers (17) on the assembly as constituted in this way; injecting resin into the various layers (11, 17) and polymerizing the resin; and machining the resulting workpiece blank in order to cut away a region of the second braided layers (17) in order to separate the two lugs (19, 21) of the clevis (24) from each other.

Abstract: A process for manufacturing a structural part made of an organic matrix composite including: production of a fibrous structure, forming a preform by braiding rovings of a fibrous material on a mandrel that includes a reinforcement in its axial extension, impregnation of the preform with an organic resin; and curing of this resin, the reinforcement, forming the cover, having a bore with an axis perpendicular to the axis for housing a connecting member. Barbs, some of which are projecting, are incorporated into the reinforcement over at least part of the surface of which lying on either side of the bore, the braiding being carried out around the barbs so that the rovings at least partly criss-cross around the barbs.

Abstract: In the field of aviation, a method of fabricating a composite-material rod. The method comprises a first step of fabricating a mandrel comprising a sleeve of composite material and solid inserts of composite material secured to the ends of said sleeve to constitute a rigid whole. Then, one or more layers of braided fibers are applied around the mandrel by using a fiber-braiding machine. Next, a resin is injected into the layers of braided fibers so as to establish cohesion rigidly bonding the braided fiber layers and at least the ends of the mandrel in order to form a rod body out of composite material having reinforced ends. Finally, respective holes are drilled at each end of the rod body, each hole passing through the layers of braided fibers and also through the corresponding insert.

Abstract: During implementing a composite metal part by compaction of an insert having reinforcing fibers in a metal body or container, gas used for compaction may enter the cavity formed in the container for receiving the insert between a lid covering the insert and the container, which can prevent or degrade compaction and diffusion welding of fiber sheaths of the insert therebetween and/or with walls of the cavity. To solve this problem, the present method includes initiating isostatic compaction by a phase including raising and maintaining temperature, followed by a phase including hot-feeding pressurized gas, and machining an assembly to obtain the part. The temperature raising phase includes a diffusion pre-welding of material rigidly connecting the pressure-adjusted walls of the lid and the container. The method can be used for designing parts having a tensile and compression resistance, such as parts for aircraft landing gear.

Abstract: The present invention relates to a method of fabricating a metal link rod (6) that is longitudinally reinforced in its thickness by fiber reinforcement (2), the method comprising the following steps: making a metal blank including at least one longitudinal housing (11, 12) leading into an edge face of the blank; inserting fiber reinforcement (2) in the housing; connecting the fiber reinforcement to the blank by metal diffusion around the fibers of the fiber reinforcement (2); and finishing the blank in order to obtain a link rod (6). The invention also provides a link rod having a metal matrix in which reinforcing fibers are embedded that extend on either side of a longitudinal axis (X) of the link rod, the fibers leading into an edge face of the link rod.

Abstract: A process for manufacturing a metal part reinforced with ceramic fibers including machining at least one housing for an insert in a metal body having an upper face. At least one insert formed from ceramic fibers in a metal matrix is placed in the housing. The insert is covered with a cover. A vacuum is created in the interstitial space around the insert and the interstitial space is hermetically sealed under vacuum. The assembly, namely the metal body with the cover, is treated by hot isostatic pressure. The treated assembly is machined in order to obtain the part. The cover includes an element covering the insert in the slot and projecting from the upper face, and a sheet covering the upper face with said element. In particular, the insert is straight and the housing for the insert in the metal body forms a straight slot.

Abstract: A method and system improving energy efficiency of a turbojet engine by optimizing rotating speed of a fan and operability of an engine, by freeing the fan from exclusive control of a low-pressure (LP) shaft by providing combined control with a high-pressure (HP) shaft when cruising power has been reached. The turbojet engine include at least one LP turbine and one HP turbine coupled to coaxial LP shafts and HP shafts, respectively, which can drive LP and HP compressors, respectively. The LP compressors include a fan that forms a first primary air-intake compression stage. The LP and HP shafts are mounted on one of two driving mechanisms, an inner ring gear, and a planet carrier for a planetary gear train for driving the fan, the HP shaft being mounted on a disengagement mechanism and the fan being coupled to the planetary gear train via an outer driven ring gear.

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: The invention relates to a method of fabricating a structural part comprising a tubular portion from which there projects a structural element made of composite material, the method comprising the steps of: shaping a plane preform comprising a stack of a plurality of superposed fiber plies such that over at least a fraction of the preform that is to form a structural portion, the plies are secured to one another, and over at least another fraction of the preform that is to form a portion for connecting to the tube, the plies are separable so as to form two facing sheets; inserting a tube between the sheets of the connection portion so that the sheets surround the tube at least in part; and bonding the tube to the preform such that the tube forms the tubular portion and the preform forms the structural element of the structural part.

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 mechanical member, the method including a plurality of operations of braiding and depositing layers of braided reinforcing fibers on a mandrel (11) by using braiding machine. Each operation comprises braiding a braided layer and depositing it by moving the mandrel (11) along a central axis of the braiding machine. Each of the various superposed braided layers comprises both longitudinal fibers (12, 12G) that are parallel to a main direction of the mandrel (11), and interlacing fibers that are inclined. At least one operation is configured to form and deposit a braided layer having, in at least one cross-section of the member, a density of longitudinal fibers that differs depending on whether consideration is given to one angular region (S1) or another angular region (S2) of the same extent around the center of gravity (G) of the mandrel (11) in the section under consideration. The invention applies to fabricating structural elements in the field of aviation.

Abstract: The invention relates to a method for manufacturing a mechanical member in composite material, such as a connecting rod (19), comprising successive operations of applying layers of reinforcing fibres braided all about a mandrel, comprising, between two operations of applying layers of braided reinforcing fibres, a step: of placing at least one element formed of pultruded reinforcing fibres (14?), oriented longitudinally (AX); and of at least temporary attachment of each pultruded element (14?) to the last layer of braided reinforcing fibres, and in which a resin is then injected into the various braided layers before this resin is polymerized in order to form a rough connecting rod.

Abstract: The invention relates to a method for manufacturing an arm (16) in composite material, this arm comprising a main body incorporating a bearing (22) designed to receive a shaft passing through it transversely, this method comprising the steps of: manufacturing a hub (11) by winding reinforcing fibres around a generally tubular support (12); manufacturing at least one sleeve (14); attaching the hub (11) to one end of the sleeve (14) and transversely to the sleeve (14) in order to form a mandrel; applying one or more layers of braided reinforcing fibres (21) around this mandrel with a machine for braiding reinforcing fibres (19); injecting and polymerizing resin into the layer or layers of braided fibres (21). The invention applies notably to an arm forming an undercarriage rocker arm of an aircraft.

Abstract: A method for producing a metallic part including inner reinforcements of ceramic fibers, according to which: at least one recess for an insert is machined in a metallic body having an upper surface; at least one insert of ceramic fibers in a metallic matrix is arranged in the recess; the insert is covered with a cover; the gap around the insert is placed under a vacuum and hermetically sealed; the entire metallic body with the cover is treated by hot isostatic compaction; and the treated assembly is machined to produce the part. The insert is rectilinear, and the recess for the insert in the metallic body forms a rectilinear groove, the cover being dimensioned so as to be able to placed on the insert in the recess after having been shrunk by cooling and to establish a tight fit in the groove by dilation such as to close the space.