Safran Aircraft Engines

Eddy Keomorakott Souryavongsa

Thomas Alain De Gaillard

Sébastien Emile Philippe Tajan

Teddy Fixy

Régis Eugène Henri Servant

Clément Pierre Postec

Adrien Jacques Philippe Fabre

Vivien Mickaël Courtier

Anthony Binder

Gilles Pierre-Marie Notarianni

Benjamin Bulot

Anthony Lafitte

Dominique Gerhardt Mayhew

Vivien Mickael Courtier

Stéphane Ibarzo

Regis Eugéne Henri Servant

Arnaud Martin

Xavier Bertrand Miqueu

The present invention relates to a blade comprising: —a structure made of composite material; —a blade root attachment part comprising a wall delimiting a cavity, a first opening formed in the wall and a second opening located under the blade root portion, the structure made of composite material extending through the first opening; —two locking parts configured to axially abut against a shoulder of the attachment part and to bear against the blade root portion; and —a cover for compressing the blade root portion against the locking parts.

Blade comprising a structure made of composite material and associated manufacturing method

A blade for an aircraft gas turbine engine includes, in a longitudinal direction, a blade root, a shank and an aerofoil body, the aerofoil body extending in the longitudinal direction between the shank and a blade tip and in a transverse direction between a leading edge made of metal material and a trailing edge. The blade includes a blade core made of composite material having a three-dimensional woven fibrous reinforcement forming the blade root, the shank and a part of the aerofoil body. The blade also includes a skin made of composite material having a two-dimensional woven fibrous reinforcement surrounding the aerofoil body part of the blade core, the skin being interposed between the leading edge made of metal material and a front edge of the aerofoil body part of the blade core to define a thinned leading edge portion, the skin including one or more two-dimensional woven plies.

Vane made of composite material having a three-dimensional woven fibrous reinforcement and two-dimensional woven skin and method for manufacturing same

Fan or propeller vane () for an aircraft turbomachine, the vane being made from a composite material and comprising a blade () and a base (), the base being formed by a longitudinal end () of a spar () which is formed by a fibrous reinforcement formed from threads woven in three dimensions and a portion () of which extends inside the blade (), the blade () having an aerodynamic profile which is defined by a skin () which is formed by woven threads and which surrounds the portion of the spar, the spar () and the skin () being embedded in a polymerised resin, characterised in that the portion () of the spar comprises projecting longitudinal stiffening members () which together delimit spaces () for receiving longitudinal inserts () which are formed from a honeycomb material.

Fan or propeller vane for an aircraft turbomachine and method for manufacturing same

The invention relates to a fan blade () comprising: &#x2014;an aerofoil () made of a composite material comprising a fibrous reinforcement densified by a polymer matrix, a leading edge () and a trailing edge (), and &#x2014;a structural shield () fitted and attached to the leading edge () or the training edge (), and &#x2014;at least one stiffener () formed integrally and in one piece with the structural shield (), said stiffener () extending in a cavity () formed between the shield () and the leading edge () or the trailing edge () so as to increase a stiffness of the blade ().

Fan blade comprising a thin shield and a stiffener

A sealing arrangement is provided between inner and outer coaxial ring-shaped walls of an aircraft engine. The outer wall has an axial end with a U-shaped cross-section parallel to axis, the opening therein being oriented axially in a first direction. The inner wall has an axial end with a U-shaped cross-section parallel to axis, the opening therein being oriented axially in the opposite direction. The U-shaped end of the outer wall includes a ring-shaped free edge axially engaged in the opening in the axial end of the inner wall. The U-shaped end of the inner wall includes a ring-shaped free edge axially engaged in the opening in the axial end of the outer wall. The walls define therebetween a ring-shaped gas channel having a substantially S-shaped axial cross-section.

Sealing arrangement between two ring-shaped walls of an aircraft engine

A system for changing the pitch of blades of at least one turbomachine rotor is provided. The system generally includes a control means acting on a connecting mechanism connected to the blades of the rotor and having a body mobile in translation along a longitudinal axis with respect to a fixed body, load-transfer bearing mounted on the mobile body cooperating with the connecting mechanism, and means for lubricating the bearing having a lubricant duct and extending radially above the fixed and mobile bodies. The duct generally includes first and second telescopic tubular parts that slide coaxially with respect to one another, the first part connected to the fixed body and the second part connected to the mobile body, and means for spraying lubricant into the bearing mounted on the mobile body and lubricant supply conduit mounted on the mobile body to connect the duct to the spraying means.

Pitch-changing system equipped with means for lubricating a load-transfer bearing

A system changes the pitch of blades of a turbine engine propeller having a plurality of blades. The system includes a plurality of links, each link being connected to one of the propeller blades at a first interface, and a control means acting on the link and having a body movable in translation along a longitudinal axis. A load transfer module is arranged between the links and the control means and is connected to the links at a plurality of second interfaces. A first adjustment element is configured to adjust an axial position of the second interfaces along the axis to provide adjustment of the pitch of all of the blades simultaneously. A plurality of second adjustment elements are configured to adjust an axial distance between each first interface and the corresponding second interface to allow the pitch of each blade to be adjusted individually.

Pitch-change system equipped with means for adjusting blade pitch and corresponding turbine engine

A controller for controlling a system for changing the pitch of blades of a turbine engine propeller, the controller includes a fixed member () and a member () which is movable in translation along a longitudinal axis (X) relative to the fixed member (), and an anti-rotation device () configured so as to prevent the rotation of the movable member () relative to the fixed member () about the axis (X). The anti-rotation device includes a longitudinal structural crossmember () which is mounted by a first and a second end () on the fixed member, and support and guide unit () integrally in translation the movable member along the axis (X), the crossmember extending radially outside the movable member and the fixed member relative to a radial axis (Y) which is perpendicular to the axis (X) and being guided through the support and guide unit.

Means for controlling a pitch change system comprising an anti-rotation device, a pitch change system equipped with said control means, and a corresponding turbine engine

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Moissy-Cramayel, France

Eddy Keomorakott Souryavongsa