Snecma

Safran Aircraft Engines

Snecma Moteurs

Herakles

Snecma Propulsion Solide

Alain Dominique Gendraud

Romain Nicolas Lunel

Vincent François Georges Millier

Sebastien Jean Laurent Prestel

Mathieu Dakowski

Fabrice Marcel Noel Garin

Thierry Fachat

Anne-Marie Arraitz

Claire Dorin

Delphine Roussin

Simon Nicolas Morliere

Jean-Baptiste Arilla

Arnaud Lasantha Genilier

Nicolas Jean-Marc Marcel Beauquin

Vincent Philippot

Denis Amiot

Delphine Roussin-Moynier

Thierry Lequitte

Jean-Claude Christian Taillant

Olivier Renon

Georges Habarou

Stéphane Pierre Guillaume Blanchard

Nicolas Pommier

Nicolas Daniel Delaporte

Renaud James Martet

Patrick Joseph Marie Girard

Damien Bonneau

David Audeon

Sébastien Jean Laurent Prestel

Philippe Jean-Pierre Pabion

Fabien Stéphane Garnier

Jerome M Friedel

Stephane Blanchard

François Pierre Michel Comte

Hubert Illand

Jean-Baptiste Vincent Desforges

Delphine Leroux

Pascal Lefebvre

Christophe Lima

Jean-Luc Le Strat

Delphine Roussin-Leroux

Boucif Bensalah

Fabien Stephane Garnier

Josselin Sicard

Emmanuel Bernard Marie Charrier

Alberto Martin Matos

Fabrice Marcel Noël Garin

Sebastien Jean Prestel

Emmanuel Da Costa

Franck Roger Denis Denece

Bruno Andre Jean Varin

Florent Pierre Antoine Luneau

Reda Oudyi

Valentin Olivier Jean-Jacques Quesnel

Fabrice Michel Olivier Cot

Pascal Tatiossian

Gilles Jeannin

Alberto Martin-Matos

Vincent Millier

Cyril Loiseau

Nicolas Auguste Marcel Pommier

Claire Dorine Dorin-Coppee

Bruno André Jean Varin

Renaud Martet

Fabrice Michel Cot

Fabrice Marcel Garin

The invention proposes an aeronautical turbine engine assembly comprising an upstream casing () to which guide blading () is fastened, and a downstream casing () to which a sealing element () provided with an abradable material for rotor blading is fastened. This assembly further comprises a shroud ring () placed between the upstream casing and the downstream casing and fastening means () for detachably fastening the shroud ring. In order to be fastened to the upstream casing, the guide blading () of the turbine engine is mounted on a downstream hook () of the upstream casing, without being hooked onto the shroud ring (), and the downstream casing () has an upstream hook with which the sealing element () is engaged in order to be fastened to the downstream casing, or the shroud ring has an upstream hook on which the sealing element () is mounted so as to be fastened to the downstream casing.

Mounting of a sealing ring on an aeronautical turbine engine

The invention relates to an assembly () for a turbomachine turbine comprising a casing () including an annular wall () extending about an axis, and means of cooling the casing (), the said means of cooling comprising a collector housing () and at least one tube () extending circumferentially around the annular wall () of the casing () and connected to the housing (), with the tube () comprising cooling-air ejection orifices.

Assembly for a turbomachine turbine

A system for the axial retention of a holding ring for a bearing for guiding in rotation a rotary shaft of a turbomachine is disclosed. The system includes an annular bearing support and a bearing holding ring that is borne by the annular bearing support. The bearing holder ring includes an upstream ring configured to be brought into contact with the bearing support and a downstream ring that is elastically deformable. The bearing support has a first axial retention element, and the bearing holding ring has a second axial retention element. The first and second axial retention elements are configured to cooperate with one another to axially retain the bearing holding ring in the event of damage to the downstream ring.

System for the axial retention of a rolling element bearing

The present disclosure relates to a holding system for the dismantling of a blade wheel, the blade wheel comprising: a disc, a plurality of blades configured to be mounted circumferentially around the disc, the plurality of blades defining a plurality of inter-blade spaces, each of the inter-blade spaces being defined circumferentially between two adjacent blades of the plurality of blades, the holding system comprising a plurality of inserts, each of the inserts being configured to be inserted into each of the inter-blade spaces in a holding position so as to hold the relative position of the blades when the plurality of blades is dismantled from the disc.

Holding system for the dismantling of a blade wheel

A movable ring assembly () for a turbine engine turbine () that is mounted between two successive rotor discs () and () of said turbine (), said rotor discs () and () being fixed to each other by bolting, characterised in that it comprises:&#x2014;a fixing ferrule () fixed between the upstream () and downstream () discs by the bolting thereof; a part forming a movable ring (), said part supporting radial sealing members () and being fixed between the upstream disc () and the ferrule ();&#x2014;and a seal () interposed between said part forming the movable ring () and the ferrule ().

Movable ring assembly for a turbine engine turbine

An assembly includes a securing stud held by a locking key that rotates with the stud and is provided with bearing surfaces for sitting on an edge of one of the parts to be assembled.

Assembly comprising a locked securing stud

A turbomachine module, comprising a mobile wheel rotatably mounted inside a housing of the module and surrounded by a sectorised sealing ring () that comprises an annular row of sectors, each ring sector comprising at least one circumferential hook that is configured to engage with an annular hooking rail of the housing, the module further comprising a sectorised protective annular shim () that is interposed between the hooks of the ring sectors and the rail of the housing and that comprises an annular row of sectors of shim, characterised in that the edges () of the circumferential ends of the sectors of shim are not aligned with the edges () of the circumferential ends of the ring sectors along the longitudinal axis of the module.

Turbomachine module

A turbine stage for a turbine engine, the stage including a nozzle and a wheel mounted inside a sectorized ring carried by a casing. The nozzle is attached to the casing and is retained axially downstream by bearing against an annular split ring. Each ring sector includes at its upstream end a member of C-shaped section that is engaged on the casing rail and that holds the split ring radially. A radially inner wall of the C-shaped member of each ring sector extends inside the split ring over an entire axial dimension thereof and its upstream end portion is engaged in at least one recess of the nozzle.

Turbine stage for a turbine engine

A rotor wheel for a turbine engine, such as an airplane turboprop or turbojet, the rotor wheel including a rotor disk including teeth at its outer periphery defining slots for axially mounting and radially retaining blade roots. An annular lip includes an annular rim extending axially downstream from and radially towards an inside of a radial retaining mechanism formed to project axially from an upstream face of the disk, and a sealing mechanism is arranged radially inside the annular lip and upstream ends of platforms of the blades.

Rotor wheel for a turbine engine

A turbine for a gas turbine engine comprising at least one turbine stator and one turbine rotor, said turbine stator comprising at least one inner annular platform, said inner platform comprising a radial wall delimiting at least one cavity between the turbine stator and the turbine rotor, wherein the turbine rotor comprises at least one elongate element positioned substantially axially and defining a sealing baffle in said cavity, and in that the turbine stator comprises at least one element made of abradable material designed to engage with said elongate element of the turbine rotor so as to form a labyrinth seal.

Labyrinth seal for gas turbine engine turbine

A turbomachine assembly in which a foil is configured to cover mainly one of bulbs of a disc and to be held, radially with respect to the disc, by the bulb of the disc and a pocket for a blade that can collaborate therewith, when these two are effectively collaborating, and the bulb of the disc includes at least one longitudinal cavity configured to form, with the foil, when the foil is covering the bulb of the disc, a secondary passage through which a secondary cooling air flow can pass.

Cooling for the retaining dovetail of a turbomachine blade

An interface element to be mounted between a blade root and a blade root housing provided in a turbine disc of a gas turbine engine to limit wear between the root and the housing. The interface element includes: a base, configured to be aligned with a lower part of the root; and first and second upper side walls connected to the base and configured to surround the blade root up to an upper portion thereof, the first upper side wall including at least one ventilation opening to allow a flow of cooling air flowing through the rotor disc housing to flow over the upper portion of the root via the ventilation opening.

Interface element between a blade root and a blade root housing of a turbine disc, and turbine rotor comprising an interface element

A turbine ring assembly includes a ring support structure and a plurality of ring sectors, each including a single piece of ceramic matrix composite material. Each ring sector includes a first portion forming an annular base with an inside face defining an inside face of the turbine ring and an outside face from which there extends two tab-forming portions including ends that are engaged in housings in the ring support structure. The ring sectors present a section that is substantially &#x3c0;-shaped and the ends of the tabs are held without radial clearance by the ring support structure. The tabs can have a free length in meridian section that is not less than three times their mean width.

Turbine ring assembly

The invention relates to a low pressure turbine () of a turbomachine comprising a nozzle () belonging to an upstream stage (B) and a nozzle () belonging to a downstream stage (C), the nozzle () forming segments () having an external structure () comprising:

Low pressure turbine for an aircraft turbomachine, comprising a segmented nozzle with an improved design

A turbine stage of a turbine engine, the stage including a rotor wheel mounted inside a sectorized ring carried by an outer casing, the outer casing including at least a circumferential rim housed in an annular cavity to attach a downstream end of the ring sector. A bottom wall of the annular cavity of the ring sector remains radially spaced apart from the circumferential rim of the outer casing to provide a thermally insulating space between them and includes a radial positioning mechanism acting on the circumferential rim.

Insulating a circumferential rim of an outer casing of a turbine engine from a corresponding ring sector

Growing community of inventors

Vernou la Celle sur Seine, France

Alain Dominique Gendraud