Safran Aircraft Engines

Safran Ceramics

Snecma

Safran

Matthieu Arnaud Gimat

Yann Marchal

Sébastien Serge Francis Congratel

Lucien Henri Jacques Quennehen

Dominique Coupe

Nicolas Paul Tableau

Alexandre Bernard Marie Boisson

Thomas Alain De Gaillard

Clément Jarrossay

Jérémy Guivarc'h

Thibault Ruf

Caroline Jacqueline Denise Berdou

Gilles Gérard Claude Lepretre

Dominique Marie Christian Coupe

Aurélien Gaillard

Bruno Dambrine

Clément Marie Benoît Roussille

Audrey Myriam Laguerre

Patrick Joseph Marie Girard

Eric Bouillon

Eric Philippe

Adrien Jacques Philippe Fabre

Thierry Georges Paul Papin

Antoine Claude Michel Etienne Danis

Gilles Pierre-Marie Notarianni

Maxime François Roger Carlin

Julien André Roger Mateo

David René Pierre Le Cair

Arthur De Castro

Richard Mathon

François Charleux

Nicolas Xavier Trappier

Dominique Magnaudeix

Julien Mateo

Thierry Guy Xavier Tesson

Yann Didier Simon Marchal

Didier René André Escure

Baptiste Dorian Lawniczek

Frédéric Jean-Bernard Pouzadoux

Adrien Laurenceau

Hubert Jean-Yves Illand

Patrice Jean-Marc Rosset

Alain Jacques Michel Bassot

Rémy Dupont

Sébastien Marin

Jean-Noël Mahieu

Benjamin Lacombe

Noemie Souhaite

Jean-Nöel Mahieu

Thierry Papin

Clement Marie Benoit Roussille

Gwenael Marion

Jeremy Engel

Dominique Marie Coupe

Thierry Georges Papin

Alexandre Bernard Boisson

Caroline Jacqueline Berdou

A bladed assembly () includes two adjacent blades (). Each blade () a platform () located at a free radial end. The platform () of the first blade () includes a lateral face () which is located circumferentially opposite an associated lateral face () of the platform () of the second blade (). The platform () of the second blade () comes into contact with the platform () of the second blade () only via an excrescence () which projects circumferentially with respect to the lateral face () of the platform () of the first blade () in the direction of the platform () of the second blade (), and the excrescence () is received in a groove () formed in the platform () of the second blade.

Turbomachine blading assembly comprising means for limiting vibration between platforms

Blade of a turbomachine turbine including a airfoil extending along a first direction between an internal end and an external end, a base secured to said internal end, and a root secured to said external end and extending, along a second direction perpendicular to said first direction, between an upstream edge and a downstream edge and, along a third direction orthogonal to the first and second directions, between first and second lateral edges, the first side edge having, in a plane including the second and third directions, a shape complementary to that of the second lateral edge. The first lateral edge includes a first portion extending from the upstream edge, and a second portion externally protruding along the third direction at an acute angle with the first portion.

Turbine blade of a turbine engine with self-generated interlock contact force in operation

A turbomachine part of an annular assembly includes a structural body and a connecting edge integral with the structural body. Each connecting edge includes at least one groove intended to receive a sealing tab. Each connecting edge is made of composite material including a fibrous reinforcement consisting of randomly-oriented short fibres, the reinforcement being densified by a ceramic matrix.

Turbomachine part with connecting edge made of composite material with ceramic matrix and short fibres and method for the manufacture of same

A turbine engine nozzle extending about an axis and including a vane assembly mounted between an outer shroud and an inner shroud; and a position retaining device configured to be placed in an elastically prestressed state between the inner shroud and an inner platform of the vane assembly to apply a stress on the vane assembly towards the outer shroud. The contact between the inner shroud and the inner platform is obtained by one same surface of the position retaining device.

Guide vanes assembly with position-maintaining device

A turbomachine turbine nozzle extending around a central axis, including at least one radially outer shroud, at least one radially inner shroud, and at least one blade made of ceramic matrix composite material, distinct from the radially inner shroud and from the radially outer shroud, and extending radially between the radially inner shroud and the radially outer shroud, the blade being hollow and including a cavity opening at a radially inner end and at a radially outer end of the blade, the nozzle including at least one tubular mast arranged in the cavity of the blade and allowing routing the ventilation air passing through the cavity of the blade, the mast including a radially outer end attached to the radially outer shroud, and a radially inner end cooperating with a radial flange for positioning the radially inner shroud.

CMC guide vane for a turbomachine turbine

A nozzle for a turbomachine turbine. The nozzle includes blades made of ceramic matrix composite material, and at least one metal liner passing through a respective blade. The liner is connected in a sealed manner to a metal internal shroud so as to guide a ventilation fluid through the blade. The nozzle can be configured to maintain effective sealing in spite of the different thermal expansions of the blade and of the internal shroud.

Turbine nozzle having blading made of ceramic matrix composite through which a metal ventilation circuit passes

A woven fibrous preform of a hollow aerodynamic profile of a turbomachine vane or blade, has a non-interlinked trailing edge and a draping of two 3D woven fibrous textures.

Fibrous preform for the aerodynamic profile of a turbomachine blade

A turbine comprises a casing, an outer metal shroud, an inner metal shroud and an annular distributor having a plurality of CMC ring sectors, each sector comprising a mast, an inner platform, an outer platform and at least one blade having a hollow profile that defines an inner housing, the inner and outer platforms each having an opening communicating with said inner housing, and the mast passing through said openings and the inner housing and being secured to said casing and connected to said annular sector. Each blade comprises at least one first radial shoulder projecting axially towards the inside of the blade, and each mast comprises at least one second shoulder projecting axially towards the outside of the mast configured to radially cooperate with a first shoulder and radially press the blade against the mast.

Turbomachine turbine having a CMC nozzle with load spreading

A method for manufacturing a turbine nozzle vane made of ceramic matrix composite material, wherein the vane is manufactured using a first fibrous preform including a hollow central section intended to form a fibrous reinforcement of an airfoil of the vane to be obtained, and a pair of second fibrous preforms each having an opening with a shape of the airfoil of the vane to be obtained.

Method for manufacturing a guide vane from a ceramic matrix composite material

A turbine stator blade made of ceramic matrix composite material includes a hollow blade profile and has a trailing edge and a leading edge, the blade including a first portion including an extrados face and a second portion distinct from the first portion including an intrados face, the first and second portions being connected to one another by a connection interface present at least on the trailing edge or leading edge, the connecting interface including a region of overlap between the first and second portions present on at least one longitudinal end of the blade profile and intended to be present outside a flow path of a gas stream of the turbine, the blade also including a platform present at a longitudinal end of the blade profile and that includes a first portion integral with the extrados face and a second part integral with the intrados face.

Turbine stator blade made of ceramic matrix composite material

A method for producing a hollow part made of a ceramic matrix composite material. The method includes shaping a hollow fibrous preform. A core of oxidizable material is housed or inserted into the preform. The method also includes consolidating the preform and extracting the core by oxidising the core.

Method for producing a hollow part made of a ceramic matrix composite material

The invention relates to an assembly comprising a turbomachine casing () and a nozzle made of ceramic matrix composite material having a blade (), the frontside and backside surfaces of which delimit an internal cavity, and which is connected at its radially external end by a connecting part () of the nozzle to the casing (), said connecting part () extending substantially radially outwards, the blade () and said connecting part () being formed in one piece, said assembly being characterized in that the connecting part () is fastened to a first radial wall () integral with the casing () by fastening means ().

Attachment of a nozzle to a casing of a turbomachine

The invention provides a blade () comprising a blade body () made of organic matrix composite material reinforced by fibers and a leading edge shield () made of a material that withstands point impacts better than the composite material of the blade body. The leading edge shield () is assembled on the blade body () and the blade () includes bonding reinforcement () adhesively bonded on the leading edge shield () and on the blade body (). The invention also provides a turbojet including a fan including such blades, and also a method of fabricating such a blade.

Blade comprising a leading edge shield and method for producing the blade

The invention relates to a compaction assembly comprising a shaping mold () delimiting an upwardly open housing, capable of receiving a woven preform cut out beforehand (), and a vertically moveable compaction tool () and forming, with the shaping mold, (), a compaction assembly for the preform placed beforehand in the housing. The compaction tool () includes at least one foot portion (A). The compaction tool () comprises at least three separate compaction blocks (-). Application to the manufacturing of composite fan blades for a turbomachine.

Method and compaction assembly for manufacturing a composite turbomachine blade

An assembly comprising ring sectors made of a first material, the sectors being placed circumferentially end to end and suspended from an outer casing, and radial positioning means comprising at least one annular tongue made of a different material, for fastening the ring sectors to the outer casing, by an annular support. The radial positioning means further comprise an eccentric spreader engaged on one side with said at least one annular tongue and, on the other side, with the sectorized annular element, and acting to enable the distance between the annular tongue and the sectorized annular element to be adjusted radially.

Assembly for a spreader connection between a turbine casing and a turbine engine ring element

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Moissy-Cramayel, France

Matthieu Arnaud Gimat