Snecma

Safran Aircraft Engines

Francois Marie Paul Marlin

Philippe Verseux

Regis Eugene Henri Servant

Thierry Georges Paul Papin

Jacky Novi Mardjono

Dominique Coupe

Bruno Jacques Gerard Dambrine

Christophe Marcel Lucien Perdrigeon

Georges Jean Xavier Riou

Alain Marc Lucien Bromann

Antoine Phelippeau

Cedric Zaccardi

Romain Plante

Kaelig Merwen Orieux

Olivier Michael Molinari

Mathieu Gaillot

Didier Queant

Noel Robin

Florian Gaudry

Yacine Cheraga

Fabrice Cretin

Damien Lourit

Frederic Imbourg

Frederic Jacques Eugene Goupil

Kevin Michael Robert Contal

Steve Marcel Nicaise

Yvan Guezel

Christian Largillier

Brice Bruno Albert

Sebastien Nicolas Juigne

An annular intermediate casing for a turbomachine through which an aerodynamic airstream circulates, including a radially internal shroud, a radially external shroud, at least one stator blade including a vane with a leading edge, and at least one aero-dynamic member which is formed of a structural arm extending radially at least in part between the radially internal shroud and the radially external shroud. The aerodynamic member includes a leading edge which is aligned with the leading edge of the stator blade in one plane and which is borne by a profiled portion positioned upstream of the structural arm in the direction of circulation of the aerodynamic airstream, the profiled portion, the structural arm, the radially internal shroud and the radially external shroud being monobloc.

Intermediate flow-straightening casing with monobloc structural arm

To facilitate the removal of a bearing housed in an arm of a turbomachine casing, which bearing guides a transmission shaft of the turbomachine, a casing arm includes a body with a vertical portion and a base defining a portion of a shaft housing; an additional part removably attached to the body and defining a bearing housing opening in the shaft housing; and a removable shim inserted between the base and the additional part. A turbomachine including such a casing arm is also described, as well as a method for removing the bearing from such a turbomachine, including removing the shim and moving the additional part towards the base, thereby allowing access to the bearing.

Arm for turbine-engine casing comprising a removable additional piece

A ring of vanes for an aircraft turbine engine, the ring presenting an axial direction and a radial direction and including a first annular sheath presenting an inside surface and a second annular sheath presenting an inside surface facing the first inside surface of the first annular sheath, the first and second sheaths being coaxial and defining between them a flow passage for a gas stream. The ring further includes both partially annular acoustic treatment panels including resonant cavities and also vanes extending in the radial direction between the first and second sheaths. Each vane includes a radial airfoil with at least two attachment tabs at each radial end of the airfoil fastened to the first or second sheath, and the inside surface of at least one of the first and second annular sheaths and the corresponding attachment tabs are covered by acoustic treatment panels arranged between the airfoils.

Ring of turbojet vanes including an acoustic treatment structure

A coupling made of composite material including a polymer matrix reinforced by a fiber structure is disclosed. The coupling includes a structural portion reinforced by a main fiber structure, and a first machining portion reinforced by a first fiber structure that is distinct from the main fiber structure. The matrices of the structural portion and of the first machining portion are identical. The first machining portion is situated on at least a fraction of the main face of the structural portion and is machined in a first machining surface. There is no intersection between the first machining surface and the fibers of the main fiber structure.

Composite coupling with a machining portion

This invention relates to an assembly () for an aircraft turbofan comprising a flowpath splitter () downstream from which there is firstly a fan flowpath () inner delimitation ring () and secondly a compressor case (), the assembly comprising a plurality of outlet guide vanes () provided with an inner platform () located in the downstream continuity of the delimitation ring. According to the invention, the assembly comprises at least one part () to support the delimitation ring (), comprising a radially outer end () fixed to a downstream part of this ring () and a radially inner end () fixed to the compressor case ().

Aircraft turbofan comprising an intermediate ring with simplified downstream support

A rotor of a turbomachine, such as a multi-flow turbojet engine fan, has a disc including, on the rim thereof, substantially axial slots with a dovetailed cross-section, and blades mounted individually in the slots, an axial wedge positioned between the root of the blades and the bottom of the slots, and a transverse lock providing upstream axial immobilization of the blades in the slot thereof. The lock is guided in radial notches formed in the flanks of the slots and radially supported against the wedge. The axial wedge is immobilised in the upstream direction, butting up against a transverse annular component secured to the disc.

Turbomachine rotor with a means for axial retention of the blades

A one-piece soundproofing panel which is attached within an circumferential air blower casing of a turbomachine, with an axis, in which circulates an airflow from upstream to downstream, is disclosed. The panel includes a circumferential rigid seat with a longitudinal cylindrical part, arranged to axially extend with respect to the axis A of the turbomachine, having an outside surface, intended to come to extend facing with an internal surface of the air blower casing, and an internal surface to which is applied a soundproofing coating; and an annular fastening flange, formed at the upstream end of the longitudinal cylindrical part, radially outwardly extending, the fastening flange being arranged to cooperate with a fastening flange of the air blower casing.

Soundproofing panel for turbomachine and turbomachine comprising such a panel

A coupling made of composite material including a polymer matrix reinforced by a fiber structure is disclosed. The coupling includes a structural portion reinforced by a main fiber structure, and a first machining portion reinforced by a first fiber structure that is distinct from the main fiber structure. The matrices of the structural portion and of the first machining portion are identical. The first machining portion is situated on at least a fraction of the main face of the structural portion and is machined in a first machining surface, and there is no intersection between the first machining surface and the fibers of the main fiber structure.

A composite material casing of varying thickness, in particular for a gas turbine fan, is made by forming fiber reinforcement and densifying the reinforcement with a matrix. The fiber reinforcement is made by winding superposed layers of a fiber texture onto a mandrel, the texture being made by three-dimensional weaving with varying thickness.

Method of manufacturing a gas turbine casing out of composite material, and a casing as obtained thereby

A turbomachine comprising high- and low-pressure compressor shafts guided in bearings isolated from an internal enclosure by a sealing end plate, and radial pressurization tubes connecting the enclosure to an air passage passing through the intermediate casing, the ends of these tubes being engaged in sealed manner in radial ducts of the intermediate casing and in radial chimneys of the sealing end plate, the chimneys being of a length is sufficient to enable the ends of the tubes to be moved in translation therein between a service position and a mounting position for the tubes.

Mounting tubes for pressurizing an internal enclosure in a turbomachine

The relief device is designed to divert a portion of a main flow toward a by-pass flow in a turbojet that includes a main flow stream, a by-pass flow stream and an intermediate casing. The relief device includes inner openings of the inner shroud, outer openings of the outer shroud, a relief circuit connecting the inner openings to the outer openings, and a blanking device capable of opening and closing the relief circuit. The relief circuit includes upstream holes of the intermediate casing, downstream holes of the intermediate casing, and an annular channel defined between the intermediate casing, the outer shroud and a connecting partition. The connecting partition rests on the intermediate casing so as to encompass the downstream holes of the intermediate casing in the annular channel and resting on the outer shroud so as to encompass the outer openings of the outer shroud in the annular channel.

Relief device for a turbojet and a turbojet comprising same

A panel for supporting abradable material in a turbojet, the panel comprising a rigid support for fastening to an inside wall of a fan casing and having one face covered in a laminated structure carrying a layer of abradable material, the laminated structure comprising sheets of fibers embedded in a polymer, and comprising an upstream portion covered by the abradable material and a downstream portion that extends beyond the abradable material, the thickness of the downstream portion being greater than the thickness of the upstream portion and being capable of withstanding ice impacts.

Growing community of inventors

Villiers sous Grez, France

Francois Marie Paul Marlin