Safran Aircraft Engines

Snecma

Snecma Moteurs

Laurent Jablonski

Philippe Gérard Edmond Joly

Jean-Marc Claude Perrollaz

Romain Nicolas Lagarde

Eric Jacques Boston

Charles Jean-Pierre Douguet

François Jean Comin

Michel Andre Bouru

Philippe Gerard Joly

Christophe Perdrigeon

Damien Merlot

Michel André Bouru

Régis Eugène Henri Servant

Vincent Joudon

Francois Jean Comin

Edouard Antoine Dominique Marie De Jaeghere

Adrien Jacques Philippe Fabre

Vivien Mickaël Courtier

Antoine Robert Alain Brunet

Clément Cottet

Hanna Reiss

Hervé Pohier

Regis Eugene Henri Servant

Christophe Paul Jacquemard

Christophe Marcel Lucien Perdrigeon

Vincent Paul Gabriel Perrot

Guillaume Sevi

Cédric Zaccardi

Kamel Benderradji

Jean-Noel Mahieu

Lilian Yann Dumas

Olivier Stephane Domercq

Jerome Talbotec

Clement Cottet

Vincent Perrot

Armel Touyeras

Anthony Lafitte

Thomas Nolwenn Emmanuel Delahaye

Rémi Roland Robert Mercier

Jean-Claude Christian Taillant

Vivien Mickael Courtier

Laurent Palmisano

Arnaud Langlois

Blaise Bergon

Jean-Noël Mahieu

Sebastien Juste

Delphine Hermance Maxime Parent

Marie-Oceane Parent

Benoit Fayard

Julien Sedlak

Philippe Joly

Pierre-Guillaume Bardin

Pauline Anne Freche

Noël Joseph Camille Robin

Jean Marc Claude Perrollaz

Joel Robert Mathieu

Thomas Etienne Camille Marie Van De Kerckhove

Sandrine Quevreux

Adrien Biscay

Adrien Jacques Fabre

Antoine Robert Brunet

Olivier Domercq

Benoît Fayard

Jérôme Talbotec

Disclosed is a variable pitch propeller blade for an aircraft turbine engine. The blade includes an airfoil connected to a root, the root having a body housed in an annular barrel that extends around a pitch axis of the blade, the body including a free end opposite the airfoil; a shoulder towards the airfoil; and a bulb located between the free end and the shoulder. The bulb can have a cross-section, referred to as the middle section, that has a maximum value Sm, and the barrel can be affixed to the body and covers and matches at least a portion of the bulb and the shoulder.

Aircraft turbine engine comprising variable-pitch propeller vanes

A system for controlling the pitch setting of a propeller blade for an aircraft turbine engine, includes a blade having a vane connected to a root. A hub accommodates the root of the blade, and a ring is mounted around the root and in the hub. Stops are also mounted around the root and in the hub. A nut is threadedly engaged with a thread of the ring and is configured to bear axially against the hub to secure the assembly.

System for controlling the pitch setting of a propeller vane for an aircraft turbine engine

A turbine engine assembly including a first rotor element and a second rotor element extending around a longitudinal axis X and coupled to one another by a gear coupling, the first rotor element including teeth and the second rotor element including complementary teeth which extend along the longitudinal axis and form the gear coupling. The first and second rotor elements each include a first and a second radial flange which are annular and arranged facing one another, and the assembly includes a sealing device configured to ensure that the gear coupling is sealed at the first and second flanges.

Turbine engine rotor element assembly equipped with a sealing device

An assembly comprising a propeller blade and its system for angularly adjusting the pitch of the blade including a bowl which is radially delimited by an annular wall extending about an axis for adjusting the pitch of the blade, the bowl having a bottom wall, a free lower end of the root being fitted axially into a complementary housing of the bottom wall in order to rotatably connect the bowl and the blade about the pitch-adjustment axis. The root of the blade comprises a first limiting face that engages with a first abutment face of the bowl to limit the rotation of the blade in the event of breakage of the lower end of the root.

Aircraft turbine engine comprising variable-pitch propeller blades

An assembly for a turbine engine extending along an axis, includes a stator, a low-pressure compressor shaft, a low-pressure compressor comprising a rotor comprising a drum rotatably coupled to the shaft of the low-pressure compressor, and a fan comprising a disk rotatably coupled to the shaft of the low-pressure compressor. The drum has a radially internal part rotatably coupled to the shaft of the low-pressure compressor and a radially external part fixed to the radially internal part by means of detachable fixing means. A first axial retention means is configured to axially and detachably retain the disk of the fan relative to the shaft of the low-pressure compressor. A second axial retention means is configured to axially and detachably retain the radially internal part of the drum relative to the shaft of the low-pressure compressor.

Assembly for a turbine engine

An assembly including a propeller blade and a pitch adjustment system for setting the pitch of the blade. A root of the blade has a bulb and the pitch adjustment system for setting the pitch of the blade includes a cup having an upper opening; —a retaining ring, which extends around the bulb, is restricted at least in terms of axial movement towards the opening relative to the cup and has a bearing face to inhibit axial movement of the root towards the opening The pitch adjustment system being further includes a lower seat for the root to engage the cup, the seat and/or the retaining ring being mounted so as to be capable of axial translation relative to the cup by at least one clamping mechanism.

Aircraft turbine engine comprising variable pitch propeller blades

Disclosed is a compressor of a dual-flow aircraft turboshaft engine extending longitudinally along an axis X, the compressor comprising a low-pressure compressor, a high-pressure compressor and an intermediate casing axially connecting the low-pressure compressor and the high-pressure compressor, the high-pressure compressor comprising an inlet guide stator comprising a retaining ring, an outer high-pressure compressor casing and a plurality of vanes, the compressor comprising an immobilising device comprising at least one first member secured to the intermediate casing and at least one second member secured to the retaining ring and configured to cooperate with the first member in order to prevent tangential movement while allowing axial movement and radial movement of the retaining ring relative to the intermediate casing according to the axis X.

Aircraft turboshaft engine compressor comprising a device for immobilising a retaining annulus

A turbomachine assembly includes a casing, first and second rotors, and a damper. The first rotor includes a disk and blades and is movable in rotation relative to the casing. The second rotor is movable relative to the casing around a longitudinal axis. The damper damps a movement of the first rotor relative to the second rotor. The damper includes first to third parts. The first part bears on the first rotor in a first area extending over a first angular sector around the longitudinal axis and applies a first centrifugal force on the first rotor. The second part bears on the first rotor in a second area that is smaller than the first angular sector and extends over a second angular sector around the longitudinal axis. The third part bears on the second rotor and applies a second centrifugal force on the second rotor.

Assembly for turbomachine

The present invention relates to an assembly for a turbomachine, comprising: a first rotor; a second rotor, and a damper () configured to damp a displacement of the first rotor relative to the second rotor, the damper comprising: a first portion () bearing against the first rotor and having a first thickness, a second portion () bearing against the second rotor and having a second radial thickness, and a third portion () connecting the first portion () to the second portion () and having a third radial thickness, wherein the third radial thickness is greater than at least one of the first radial thickness and the second radial thickness.

Assembly for a turbomachine

The present invention relates to a turbomachine assembly, comprising: a casing (), a first rotor () which is movable in rotation with respect to the casing () about a longitudinal axis (X-X), and comprising: *a disk (), and *a plurality of blades () capable of flapping with respect to the disk () during a rotation of the first rotor () with respect to the casing (), a second rotor () which is movable in rotation with respect to the casing () about the longitudinal axis (X-X), and a damper () which is configured to damp a displacement of the first rotor () with respect to the second rotor () in a plane orthogonal to the longitudinal axis (X-X), the displacement being caused by a flapping of at least one blade () among the plurality of blades (), the damper () comprising: o a first bearing part (): *bearing against the first rotor (), and *being configured to apply a first centrifugal force (C) to the first rotor (), o a second bearing part (): *bearing against the second rotor (), and *being configured to apply a second centrifugal force (C) on the second rotor (), and o a linking part (): *connecting the first bearing part () to the second bearing part (), and being thinned relative to the first bearing part () and the second bearing part (), and o a flyweight () which is fixedly mounted on the damper ().

Turbomachine assembly having a damper

The invention relates to a bladed disk () of a fan, comprising: a hub () comprising an outer radial platform () designed so as to define an inner gas flow stream in the turbomachine, a plurality of blades () comprising a root () connected to the platform (), a leading edge () and a trailing edge (), a groove formed in the platform () around part of the root () of each blade () in an area adjacent to the leading edge () and/or the trailing edge (), and a joint () placed in the groove () in such a way that it extends in the extension of the radially outer face () of the platform () in order to ensure a continuity of the inner flow stream.

Bladed disk flexible in the lower part of the blades

The invention relates to an assembly () for a turbomachine comprising: characterized in that the damping device () is annular while extending circumferentially around the turbomachine longitudinal axis (X-X) and in that the damping device () comprises a first radial external surface () supported with friction against the first module () as well as a second radial external surface () supported with friction against the second module (), so as to couple the modules () in order to damp their respective vibrational movements during operation.

Damping device

A turbomachine assembly is provided. The assembly includes: a first rotor module including a first blade, a second rotor module, connected to the first rotor module, and including a second blade of smaller length than the first blade, and a damping device including a plurality of first radial external surfaces supported with friction against the first module. The damping device is stepped, and includes a second radial external surface supported with friction against the second module, so as to couple the modules for the purpose of damping their respective vibrational movements during operation.

An assembly for a turbomachine including a first rotor module () with a first blade (), and a second rotor module (), connected to the first rotor module (). The second rotor module includes a second blade with a smaller length than the first blade (). A damping device () is attached to the second rotor module () and includes a radial external surface () supported with friction against the first module (), so as to couple the modules () for the purpose of damping their respective vibrational movements during operation.

The invention relates to a turbomachine assembly () comprising:

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Melun, France

Laurent Jablonski