Safran Aircraft Engines

Snecma

Herakles

Antoine Jean-Philippe Beaujard

Elric Georges André Fremont

Tewfik Boudebiza

Julien Mateo

Didier Fabre

Dominique Coupe

Jacques Rene Bart

Clement Roussille

Clément Roussille

Catherine Pikovsky

Jean Pierre Mareix

Serge Benyamin

Guillaume Renon

Olivier Belmonte

Jonathan Goering

Michel Gilbert Roland Brault

Gilles Alain Charier

Nathalie Nowakowski

Nicolas Xavier Trappier

Nicolas Eberling-Fux

Nils Edouard Romain Bordoni

Philippe Pierre Vincent Bouiller

Nicolas Jerome Jean Tantot

Didier Jean-Louis Yvon

Nicolas Jérôme Jean Tantot

Augustin Marc Michel Curlier

Pierre-Alain Jean Philippe Reigner

Bastien Pierre Verdier

Fabien Roger Gaston Caty

Stephane Blanchard

Paco Maurer

Jean-Pierre Mareix

Arnaud Baudran

Romain Nunez

Jacques René Bart

Maxime Aurélien Rotenberg

Maryline Emmanuelle Besnard

David René Pierre Le Cair

Frédéric François Jean-Yves Patard

Frédéric Patard

Victor Henri Marius Varenne

Pascal Coat

Christophe Caplain

Nicolas Trappier

Elric Georges Andre Fremont

A turbomachine—for aircraft has a primary annular flow path and a cold flow channel extending around the primary annular flow path. The turbomachine includes an impeller, a compressor, a combustion chamber, a first turbine having a first turbine rotor, a second turbine having a second turbine rotor, a first connecting shaft and a second connecting shaft, an inter-turbine stator arranged between the first turbine and the second turbine, and a first cooling circuit having successively: a first cooling inlet located between the impeller and the compressor, a first passage extending into the inter-turbine stator, and a first cooling outlet extending into the cold flow channel.

Turbine engine for an aircraft

A turbine arrangement comprising a journal having a flange carrying a disc that is attached to the flange with bolts, wherein it includes a stationary circular trough surrounding the flange for collecting oil capable of passing radially between the disc and the flange carrying this disc.

Turbine arrangement incorporating an oil recovery circumferential trough

Double flow turbojet including an intermediate flow path dedicated to supplying with air via radial arms an exhaust casing of this turbojet

A ventilated chamber (), subject to an irruption of air at a pressure higher than that in a surrounding chamber () through labyrinth seals separating them and evacuation of excess air through an evacuation duct (), comprises a motor-driven compressor () in the evacuation duct to facilitate suction and evacuation, even when the engine speed does not allow a large overpressure in the chamber ().

Device for evacuation of a ventilated lubricated chamber of a turbomachine

A mechanical decoupler () at the inlet to a turbomachine is positioned on the outside of an intake casing, where radiating arms () meet an external casing () so as to partially unload a low-pressure shaft when a significant out-of-balance appears. Because it is positioned a long way from the bearing, the decoupler () can be designed with a greater degree of freedom at a location where there is more space available and where layout constraints are less of an issue. More specifically, it is housed in a cavity () of the external casing () which opens onto the flow path ().

Aircraft engine intake arrangement comprising a mechanical decoupler

The invention relates to a turbojet bearing housing including a fixed envelope traversed by a rotor, this envelope including a cylindrical end surrounding a seal ensuring the leak tightness of the cylindrical end of the envelope with the rotor, this envelope being equipped with a shell screwed onto the cylindrical end thereof, this shell including radial channels emerging opposite the seal while being arranged so that an air stream traversing the seal towards the housing mainly comes from these radial channels.

Shell for reducing overpressure in the vicinity of the upstream seal of a turbojet bearing housing

A turbojet includes an inlet casing, a low pressure compressor, a high pressure spool, a low pressure turbine, and an exhaust casing. The inlet casing and the low pressure compressor define an air inlet channel that divides downstream into a primary flow channel and a secondary flow channel. A low pressure shaft is supported by a first roller bearing, borne by the inlet casing, a second bearing borne by the intermediate casing, a third bearing borne by the exhaust casing, and an additional ball bearing borne by the intermediate casing and arranged between the first bearing and the second bearing. Such a configuration of bearings makes it possible to withstand an increased level of load and to be compatible with an increase in the rate of dilution.

Turbojet with bearing architecture optimised for the support of a low pressure shaft

An aircraft propulsion assembly, including a turbine engine comprising at least one gas generator configured to generate a main flow, which is supplied by a central jet to at least one power turbine, the central jet being surrounded by an outer fairing, and the power turbine driving, on the periphery thereof, at least one fan rotor. The aircraft propulsion assembly comprises first movable means which are arranged so as to divert at least some of the main flow from the central jet to the outside of the outer fairing and preferably upstream of the turbine engine so as to generate thrust reversal. An aircraft which uses the propulsion assembly, particularly on the rear tip of the fuselage of the aircraft.

Aircraft propulsion assembly comprising a thrust reverser

The present invention relates to an aircraft comprising a fuselage () and a propulsion unit at the rear of the fuselage, the propulsion unit comprising at least one fan rotor (), a nacelle () fairing the fan and at least one connection means () connecting the nacelle to the fuselage, the fan being rotated by the energy supplied by at least one gas-turbine gas generator () housed in the fuselage, said gas generator comprising auxiliary equipment cooled by a cooling circuit. The aircraft is characterised in that said cooling circuit comprises at least one heat exchanger exchanging heat with the ambient air housed in one of said connection means () and/or in said nacelle (). The cooling circuit optionally comprises also a heat exchanger exchanging heat with the ambient air, housed in the tail unit.

Aircraft with a propulsion unit comprising a fan at the rear of the fuselage

A turbine engine impeller comprises a disc bearing vanes and an annular plate. The roots of the vanes are fitted in grooves and the root of each vane is connected to a blade by an integrated platform. The platforms of said vanes cooperate to form a radially inner wall of a duct for a flow passing through the blades. The annular plate comprises a transverse annular portion forming an axial stop of the roots of the vanes. The integrated platform of each vane has no spoiler and stops axially to the right of transverse faces of the root. The transverse annular portion of said annular plate bears, between the integrated platforms and the disc, an annular spoiler configured to extend in an axial space between the impeller and an adjacent module of the turbine engine so that said annular plate bears annular knife-edge seals of a labyrinth sealing joint.

Impeller having a radial seal for a turbine engine turbine

The present invention relates to an aircraft comprising a fuselage (), a propulsion unit () at the rear tip of the fuselage and at least two engines () each driving an engine shaft (), the propulsion unit including at least one propeller () driven mechanically by the engine shafts (), characterized in that the propulsion unit () includes two propellers (), each of the propellers being driven by a propeller shaft (), the two propeller shafts being driven by the engine shafts () via a first differential epicyclic gear train (). Specifically, the input of the first differential gear train () is connected to the output of a second differential epicyclic gear train (), the second differential gear train () including two inputs (), each of said two inputs being connected mechanically to one of said engine shafts ().

Aircraft comprising a propulsion assembly including a pair of propellers at the rear of the fuselage

A rotor for a turbine engine includes a disk having cavities called primary cavities at its periphery; a plurality of blades each having a root of which the lower part is composed of a bulb locked axially in the primary cavities; a plurality of added platforms, each being arranged between two consecutive blades, wherein the platforms have: a substantially straight plate and a bulb extending radially under the plate, the bulb being locked axially in the secondary cavities arranged at the periphery of the disk, the secondary cavities being positioned between two consecutive primary cavities; a spoiler extending in the axial direction, the spoiler forming an annular sector facing at least two consecutive blades.

Rotor for turbine engine comprising blades with added platforms

The invention relates to a turbomachine rotor including a disk and at least one blade wherein the root comprises a bulb accommodated into a groove associated with the disk, wherein the bulb includes a slot that opens up radially inwards at its lower face, and in that the rotor includes a retaining part extending partly into the groove projecting radially outwards from the bottom face of the groove, of which a radially outer end of the retaining part fits into the slot of the bulb and bears radially inwards on the inner wall of the slot of the bulb.

Turbomachine rotor with optimised bearing surfaces

A gas turbine engine for propulsion of an aircraft, the engine extending axially from upstream to downstream and including a composite part, a metal part, and a device for flexible fixing of the parts, the fixing device including: a flexible connecting element including a first fixing portion connected to the metal part by an axial connection and a second fixing portion connected to the composite part by an axial connection, the connecting element including an axially extending free end; and a rigid locking element including a fixing portion connected to the metal part by an axial connection and an axially extending free end, the free end of the locking element being aligned radially with the free end of the connecting element to limit deformation of the connecting element while the engine is in operation.

Gas turbine engine comprising a composite component and a metal component which are connected by a flexible fixing device

An adapting part is designed to be held on a casing of a turbo-engine, partially covering the casing. The adapting part includes, in the extension of one of its axial ends, a first connector to engage with a first complementary connector associated with the casing to form a sliding connection. The adapting part also includes, at its other axial end, a second connector to be secured to a second complementary connector associated with the casing to form a rigid connection.

Method for holding an adapter piece on a tubular housing of a turbo engine, and corresponding adapter piece and holding system

Growing community of inventors

Moissy-Cramayel, France

Antoine Jean-Philippe Beaujard