Safran

Nicolas Pierre Lanfant

Patrick Dunleavy

Dominique Marie Christian Coupe

Thierry Patrick Chauvin

Adrien Touze

Emilie Goncalves

Hugues Laurent Alglave

Robin Mandel

Pierre Faivre D'Arcier

Bastien Tranquart

Olivier Foussard

Romain Picon

Nicola Piccirelli

Sebastien Pautard

Didier Payen

Romain Venat

Cédric Pierre Michel Thomas

Sébastien Pautard

Loïc Sorgnard

Magali Mélanie Coulaud

Matthieu Stackler

Christophe Ravey

Maximilien Eddie Farrands

Florence Castagnet

Dominique Marie Coupe

Pierre Faivre D'arcier

Pierre François Robin Geraud-Grosheny

Pierre Faivre D'ARCIER

A method for manufacturing an acoustic attenuation structure including a complex acoustic multi-element panel extending along a horizontal and a vertical direction and an acoustic skin covering one of horizontal faces of the complex acoustic multi-element panel, the complex acoustic multi-element panel including complex acoustic elements each having a shape gradually narrowing between a base and a tip and partitions surrounding each complex acoustic element to form a plurality of acoustic cells, the partitions extending along the vertical direction from the base of the complex acoustic elements, the complex acoustic multi-element panel being produced by injection of a filled or unfilled thermoplastic or thermosetting material. The complex acoustic multi-element panel and the plurality of partitions are produced as a single part by injection of a filled or unfilled thermoplastic or thermosetting material.

Acoustic attenuation structure and method for manufacturing same

A method for manufacturing a turbomachine vane made of composite material and having a cavity, the method includes producing a core having the shape of the cavity of the vane to be manufactured, the core including a reinforcing structure occupying only a portion of the volume of the core, the core further including a sealed envelope defining the outer surface of said core; and forming a composite material skin around the core.

Method for manufacturing a vane comprising a reinforced cavity

A process for manufacturing a blade made of composite material for a turbomachine is provided. The blade includes an airfoil having a pressure side and a suction side which extend from a leading edge to a trailing edge of the airfoil. The blade further includes a metal sheath that extends along the leading edge of the airfoil. The process includes the steps of: a) placing a preform, made by three-dimensionally weaving fibers, in a mold, a polymerizable adhesive being inserted between the sheath and the edge of the preform; and b) injecting polymerizable resin into the mold to impregnate the preform so as to form the airfoil after solidifying, wherein the resin is injected within a time interval during which the adhesive reaches a freezing point.

Method for manufacturing a composite blade for an aircraft engine

A method for manufacturing a composite material part having a cellular structure that includes at least one cell delimited by walls. The method includes supplying at least a first core, providing a web of a first fibrous reinforcement comprising a plurality of long discontinuous fibers randomly distributed in a plane, and producing at least one strip of the first fibrous reinforcement. The method further includes producing a second fibrous reinforcement in the shape of a sock, inserting the first core into the sock, draping the strip around the sock containing the first core, placing the sock containing the first core and the wrapped strip in a mold, and thermocompression.

Method for manufacturing a composite material part having a cellular structure and corresponding part

A method for continuously manufacturing a complex acoustic multi-element panel for an acoustic attenuation structure, the method including supplying a thermoplastic resin film at the entry of a stamping system including at least one pair of complementary cylinders, a heating of the thermoplastic resin film upstream of the entry of the stamping system, and a passage of the thermoplastic resin film between the at least one pair of complementary cylinders, the pair of complementary cylinders including a male cylinder with teeth radially protruding from the cylinder and a female cylinder with cavities on its radial surface, the cavities of the female cylinder being of shapes complementary to the shapes of the teeth of the male cylinder and a separation of the stamped film at the exit of the stamping system using a separator disposed at the exit of the stamping system between the stamped film and the female cylinder.

Method and line for manufacturing a panel with multiple acoustic elements for an acoustic attenuation structure

A process for manufacturing a blade made of composite material for a turbomachine. The blade includes an airfoil having a pressure side and a suction side which extend from a leading edge to a trailing edge of the airfoil. The blade further includes a metal sheath that extends along the leading edge of the airfoil. The process includes the steps of: placing a preform, produced by three-dimensionally weaving fibers, in a mold, the sheath being positioned on an edge of the preform intended to form the leading edge of the airfoil; and injecting polymerizable resin into the mold to impregnate the preform so as to form the airfoil after solidifying. At least one double-sided adhesive film may be inserted between the sheath and the edge of the preform prior to injection of the resin.

Composite blade for an aircraft engine and methods for manufacturing and repairing same

A method for manufacturing an acoustic panel by welding of at least two acoustic components made of thermoplastic material, a thermal gradient being applied on the acoustic components during the welding operation so as to apply a temperature higher than the glass transition or melting temperature of the material on the area to be welded and so as to apply a temperature lower than or equal to the geometric stabilization temperature of the material on the parts furthest from the area to be welded.

Method for manufacturing an acoustic panel by welding

A process for shaping, by compaction, a fibrous structure obtained by three-dimensional or multilayer weaving between a plurality of layers of warp yarns and a plurality of layers of weft yarns. The yarns of each plurality of layers of warp yarns and layers of weft yarns are coated with a sizing composition. The shaping includes a pre-compaction stage followed by a stage of final compaction of the fibrous structure. During pre-compaction and final compaction, the fibrous structure is at a temperature greater than or equal to the softening temperature of the sizing composition. The process further includes the cooling of the fibrous preform obtained after final compaction of the fibrous structure.

Method for shaping a fibrous preform by compacting in order to produce a composite material part

A mould for obtaining an element is provided. The mould includes an outer face and a plurality of parallel ribs and/or grooves formed on the outer face. One of the walls of the mould includes ribs and/or grooves of which at least a portion of the ribs and/or grooves can be retracted, at least temporarily, such that at least a portion of the surface of the wall is smooth.

Method and mould for producing an aerodynamic element comprising riblets

The invention concerns an aircraft propulsion system comprising at least one member () 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 () 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.

Aircraft propulsion system comprising a member covered with a grooved structure

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:

Method for manufacturing a composite material part comprising a body extended by a reinforced attachment end

A method for making an aerodynamic element, particularly for an aircraft, including an external face and a plurality of parallel ribs and/or grooves formed on the external face, the method including making the element and its ribs and/or its grooves simultaneously with a mold, including a step to supply a film made of a deformable material that includes ribs and/or grooves complementary to the ribs and/or grooves of the element, a step to position the film on a wall of the mold, a step in which the element is molded, a step in which the element is separated from the mold simultaneously with the film, and a step in which the film is separated from the element.

Method and mould for creating an aerodynamic element comprising riblets

A blade of a turbomachine includes a blade body of composite material having a fiber reinforcement having a three-dimensional weave and densified by a matrix, the reinforcement having a first part extended by a second, end, part including two segments separated from each other; and an insert having a pi-shaped section, the insert having a platform part and two longitudinal flanges separated from each other, the platform part including a housing delimited by a bottom wall and a rim, the bottom wall including an opening communicating with the space between the two flanges, the first part of the fiber reinforcement being clamped between the two flanges of the insert, the segments of the second part of the fiber reinforcement being folded against the bottom wall of the housing.

Turbomachine blade and method for the manufacture of same

A method of fabricating a composite material part having a body secured to at least one platform present at one end of the body, the method including using the fiber blank to form a preform for the part that is to be made by folding out segments of the second and third portions that are not interlinked with the first portion on opposite sides of the first portion, by shaping the folded-out segments of the second and third portions to form preform portions for a platform of the part to be fabricated, and by folding down the segment of the first portion in order to fill in the groove present in the gap between the second and third portions; and densifying the preform with a matrix in order to obtain a part made of composite material having at least one integrated platform.

Growing community of inventors

Moissy-Cramayel, France

Nicolas Pierre Lanfant