Safran Aircraft Engines

Safran Transmission Systems

Snecma

Julien Fabien Patrick Becoulet

Michel Gilbert Roland Brault

Alexandre Jean-Marie Tan-Kim

Arnaud Nicolas Negri

Paul Ghislain Albert Levisse

Emmanuel Fabrice Marie Baret

Romain Guillaume Cuvillier

Emmanuel Pierre Dimitri Patsouris

Didier Jean-Louis Yvon

Franck Emmanuel Bosco

Olivier Belmonte

Guillaume Patrice Kubiak

Olivier Formica

Paco Maurer

Nicolas Stoliaroff-Pepin

Gilles Alain Marie Charier

Vincent François Georges Millier

Serge Rene Morreale

Olivier Renon

Guillaume Pierre Mouly

Tewfik Boudebiza

Serge René Morreale

Patrice Jocelyn Francis Gedin

Laurent Soulat

Jean Charles Olivier Roda

Amélie Argie Antoinette Chassagne

Yanis Benslama

Pierre-Alain Jean Philippe Reigner

Simon Loïc Clément Lefebvre

Patrice Jean-Marc Rosset

Romain Truco

Maeva Daphné Gros-Borot

Jean-Claude Christian Taillant

Alexandre Tan-Kim

Fabrice Joel Luc Chevillot

Xavier Jean Yves Alain Agneray

Maxime Paul Numa Givert

Jérémy Dievart

Sylvain Bories

Nicolas Auguste Marcel Pommier

The invention relates to an aircraft turbomachine subassembly comprising an annular duct () extending concentrically about a longitudinal axis and in which an air stream flows from upstream to downstream, the annular duct comprising a duct portion () defining a swan neck having an upstream inlet section and a downstream outlet section. A plurality of radial arms (B, B) extend circumferentially in the duct portion (). One or several intermediate walls (P, P) extend between two circumferentially adjacent arms (B, B) which form a pair of circumferentially adjacent arms among the plurality of circumferentially adjacent arms so as to locally separate, between the two arms of the pair of arms, the annular duct portion () into a radially inner portion (Ci) and a radially outer portion (Ce) further away from the longitudinal axis than the radially inner portion.

Turbomachine subassembly comprising a gooseneck of improved configuration, and turbomachine comprising such a subassembly

Disclosed is an assembly for a turbomachine that includes connection elements having respective radial flexibilities, wherein a ratio of a first of the respective radial flexibilities to a third of the respective radial flexibilities is strictly lower than 10% and/or a ratio of the first of the respective radial flexibilities to a second of the respective radial flexibilities is strictly lower than 4%.

Flexibilities in a gas turbine engine with speed reducer

The invention relates to a turbomachine for an aircraft, comprising:

Aircraft turbomachine comprising a device for lubricating a bearing

A turbomachine includes a low-pressure shaft having an internal passage for a first air stream and a first segment. A second segment includes a first frustoconical portion, and an intermediate lubrication chamber wherein at least one bearing is arranged. The intermediate lubrication chamber is arranged in an intermediate casing. An annular air chamber is configured to be supplied with air from a low-pressure compressor and is at least partly axially delimited by the first frustoconical portion. A radial separating wall is mounted around the first segment, and the first frustoconical portion further includes an opening to enable the first air stream to pass from the air chamber to the internal passage.

Turbomachine for an aircraft

A turbomachine for an aircraft comprising a stator, a turbine shaft and an electric machine arranged to the rear of the turbine shaft, and comprising an electric machine stator and an electric machine rotor, a rotationally fixed coupling carriage capable of moving in axial translation relative to the electric machine rotor between coupled and uncoupled positions, coupling members which are designed to engage with one another after the coupling carriage has moved from the uncoupled position to the coupled position and to transmit a torque from the turbine shaft to the electric machine rotor when they are engaged, and an elastic return member designed to elastically return the coupling carriage in axial translation to the coupled position.

Aircraft turbomachine comprising an electric machine

A device for centering and guiding a shaft of an aircraft turbine engine is provided. The device includes an outer ring of a rolling bearing extending about an axis and having orifices, an annular bearing support extending about the axis and at least partially about the ring, the support having orifices and a series of studs for connecting the ring to the support. The studs can be distributed about the axis and extend parallel to the axis. The ends of first studs can be engaged without clearance in the orifices, and ends of second studs can be engaged with clearances in the orifices of the ring and/or of the support, the clearances being configured so that the device has different stiffnesses in at least two directions perpendicular to the axis.

Device for centering and guiding a shaft of an aircraft turbine engine

A fan module for an aircraft turbine engine includes a fan having variable pitch blades and an oil transfer device configured to provide a transfer of oil between a stator and a rotor. The oil transfer device includes a stator ring having internal oil ducts, a shaft inserted into the ring and having internal oil ducts, and an annular support of the ring configured to deform elastically to allow movements of the ring in the radial direction. A plain bearing is located between the ring and the shaft, and roller bearings are mounted between the ring and the shaft on either side of the plain bearing.

Fan module provided with an oil transfer device

The invention relates to a planetary gear reducer () for an aircraft turbine engine (), the reducer () comprising a sun gear () and a ring gear () which are centred on a longitudinal axis X and a planet carrier () which carries at least one planet gear () rotatably mounted about a planet axis A parallel to the longitudinal axis X, the sun gear () being rotatable about the longitudinal axis, the planet gear () meshing both with the sun gear () and the ring gear (), the planet carrier () being movable about the longitudinal axis and the ring gear () being rotationally fixed. According to the invention, the reducer () comprises an electric machine () integrated therewith and which comprises a rotor () mounted on the planet carrier () in such a way as to be rotated about the longitudinal axis X and a stator () mounted on the ring gear ().

Speed reducer equipped with an electric machine

A turbomachine including a low-pressure shaft; a stator supporting the low-pressure shaft by bearings including a rear bearing; an electrical machine located at a rear portion of the turbomachine and including a rotor disconnectably driven by the low-pressure shaft and a stator surrounding the rotor and attached to the stator of the turbomachine; the turbomachine includes a journal for supporting the rotor of the electrical machine, the journal including a shaft segment releasably attached to a rear end of the low-pressure shaft by extending the low-pressure shaft, a frame to which the rotor of the electrical machine is attached and which surrounds the shaft segment, and at least one bearing mounted between the frame and the shaft segment; and a sleeve that is axially movable about the low-pressure shaft and/or the shaft segment that is rotationally connected to the low-pressure shaft and the shaft segment.

Turbomachine comprising a rear integrated electrical machine

A turbomachine fan module includes a fan shaft and a lubrication chamber traversed axially by at least part of the fan shaft. The module further includes at least one first bearing configured to support and guide the fan shaft in rotation. The fan shaft has a first tubular part closed by a cover, wherein the first tubular part has a distal end and a proximal end that are opposite each other along a longitudinal axis X and an annular squealer cavity, formed in an inner surface of the fan shaft, between the distal end and the proximal end. The cover has a body extending along the longitudinal axis X and mounted in the first tubular portion so as to cover and sealingly close the annular squealer cavity.

Fan module comprising improved sealing means for a lubricant chamber

A scoop for an aircraft turbine engine, this scoop including a body which is movable between at least two positions, including a first position in which it is configured to capture part of a first air stream flowing in a first direction, wherein the body is movable into a second position in which it is configured to capture part of a second air stream flowing in a second direction which is opposite to the first direction, and wherein the body is mounted to move freely between the at least two positions so as to automatically adopt the first position when the first air stream flows, and to automatically adopt the second position when the second air stream flows.

Scoop for an aircraft turbine engine

A device for centering and guiding a shaft of an aircraft turbine engine, the device including an outer ring of a rolling bearing, the ring extending about an axis; an annular bearing support extending about the axis and at least partially about the ring; at least one series of connecting elements for linking the ring to the support, the connecting elements being interposed between an inner cylindrical rim of the ring and an outer cylindrical rim of the support that extends about the inner rim, and each having a first radially outer end for linking to the outer cylindrical rim and a second radially inner end for linking to the inner cylindrical rim, the ring, the support and the connecting elements being integrally formed, the connecting elements being enclosed in an annular housing that is radially delimited by the rims and laterally enclosed by an annular web.

A device for centering and guiding a shaft of an aircraft turbine engines. The device includes an outer ring of a rolling bearing, the ring extending about a main axis and having orifices arranged around the axis axis; an annular bearing support extending about the main axis and at least partially about the ring, the support having orifices and openings arranged about the axis; and a series of studs for linking the ring to the support, the studs being distributed about the main axis and having elongation axes substantially parallel to the main axis, the body of each of the studs passing through one of the openings. Some of the studs, have bodies that pass with a first positive clearance and a second virtually zero clearance through the first openings, the first and second clearances being configured so that the device has different displacement amplitudes.

An enclosure for a turbomachine includes a turbomachine drive shaft rotating about a longitudinal axis (X) by means of two roller bearings, an upstream bearing and a downstream bearing, each having an inner ring carried by the drive shaft. The two bearings share a single integral outer ring that has an upstream end and a downstream end connected to one another by a section of studs. The single outer ring is carried by an upstream base plate and a downstream base plate of a bearing support configured to be attached to a stationary structure of the turbomachine.

Fused soft and hard bearing rings

An aircraft turbine engine includes a turbine shaft having a first axis of rotation, a propulsion propeller having a second axis of rotation parallel to and spaced from the first axis, and a mechanical reduction gear coupled to the turbine shaft and rotating the propeller. The reduction gear has a sun gear connected to the turbine shaft, a ring gear, and at least two planet gears, each including a first external toothing that is meshed with an external toothing of the sun gear. A second external toothing is located within the ring gear and is meshed with an internal toothing of the ring gear.

Growing community of inventors

Moissy-Cramayel, France

Julien Fabien Patrick Becoulet