Safran Aircraft Engines

Snecma

Christophe Paul Jacquemard

Adrien Jacques Philippe Fabre

Didier Gabriel Bertrand Desombre

Sebastien Tajan

Vivien Mickaël Courtier

Christophe Marcel Lucien Perdrigeon

Adrien Laurenceau

Cédric Zaccardi

Sébastien Emile Philippe Tajan

Romain Boudier

François Marie Paul Marlin

Sarah Chartier

Sébastien Tajan

Olivier Belmonte

Laurent Jablonski

Regis Eugene Henri Servant

Thierry Georges Paul Papin

Régis Eugène Henri Servant

Vincent Joudon

Emmanuel Pierre Dimitri Patsouris

Serge René Morreale

Anthony Binder

Clément Cottet

Dominik Igel

Paul Antoine Foresto

Clementine Charlotte Marie Mouton

Adrien Philippe Fabre

Sebastien Emile Philippe Tajan

Clémentine Charlotte Marie Mouton

Charles-Henri Michel Marie Derrez

Xavier Jean Yves Alain Agneray

Laurent Paul François Perrot

Fabrice Michel François René Cretin

Leny Toribio

An auxiliary tank for an aircraft turbine engine is provided, and in conjunction with a pump and associated auxiliary lubrication circuit, supplies oil to a reducer when, for example, a phase of free rotation of the fan is detected. Thus, the reducer is always lubricated, even during the phases of free rotation of the fan, thus ensuring a longer life of the reducer gears. The auxiliary tank is arranged with respect to the reducer so that oil is recovered from the reducer at least in part by the auxiliary tank.

Auxiliary oil tank for an aircraft turbine engine

Disclosed is a main shaft for an aircraft turbine engine drive line extending longitudinally from front to rear along an axis and having a rear portion configured for receiving a front portion of a rear shaft by interlocking, the main shaft having: —a front retaining member and a rear retaining member together defining a prison and each having a central opening, —a connecting ring, located in the prison, having a rear end configured to be screwed to a front end of the rear shaft extending into the central opening of the rear retaining member, the connecting ring having a gripping member formed on the inner surface and accessible from a front portion of the main shaft by a tool extending via the central opening of the front retaining member.

Aircraft turbine engine drive line shaft comprising a connecting ring for disassembling the drive line

The invention relates to an assembly for a turbine engine blade, the assembly comprising a fastener defining a well for receiving a blade root, and a clamping device for clamping the blade root against the fastener when the blade root is received in the well, wherein the clamping device comprises: at least one clamping member suitable for bearing simultaneously on the blade root and on the fastener in the well so as to exert a clamping force on the blade root, the clamping member defining a threaded hole, and a resilient layer arranged to bear in the well on the blade root, a rod comprising a thread engaging with the threaded hole so that a rotation of the rod with respect to the clamping member causes a variation in the clamping force exerted by the clamping member on the root.

Fastening assembly for a turbomachine blade

An aircraft propulsion assembly including a fan shaft which is connected to a turbine shaft by means of a reduction gear. The guiding of the fan shaft is carried out by a first bearing including cylindrical rollers and a second beating including two rows of balls with oblique contact.

Turbomachine fan assembly comprising a roller bearing and a double-row ball bearing with oblique contact

The invention concerns an aircraft turbine engine (), comprising a gas generator comprising at least one annular gas flow duct (), the duct () being defined by two annular housings, respectively an external housing () and an internal housing (), extending one around the other and connected together by at least one tubular arm () for the passage of a lubricating oil line (). According to the invention, the line () comprises a first fixed section () secured to the external housing (), a second fixed section () secured to a piece of equipment () of the turbomachine capable of moving or vibrating during operation relative to the housings (), and an intermediate section () connecting the first and second sections (), this intermediate section () having a generally elongate shape and comprising longitudinal ends engaged and capable of swiveling and/or sliding in ends of the first and second sections ().

Aircraft turbine engine

A planetary reduction gear () for an aircraft turbine engine, said reduction gear comprising a rotatable sun gear, a rotatable ring gear () and satellite gears meshed with the sun gear and the ring gear and carried by a planet carrier intended to be fixed to a stator, the reduction gear further comprising an annular trough for recovering and channelling lubrication oil that is mounted around the ring gear, characterised in that the reduction gear further comprises an annular deflector () fixed to the ring gear and configured to route the oil exiting radially towards the outside of the ring gear up to the trough by virtue of centrifugal forces, the trough comprising an annular chamber () which is axially remote relative to a median plane (P) substantially passing through the centre of the ring gear, and an annular duct () located on a side of said chamber and emerging in said chamber.

Planetary reduction gear for an aircraft turbine engine

An auxiliary tank for an aircraft turbine engine has a curved general shape, and includes a cylindrical or frustoconical radially outer wall and three cylindrical or frustoconical radially inner walls arranged facing the radially outer wall, the three radially inner walls having a middle wall and two side walls arranged to either side of the middle wall. The inner volume of the tank forms a U-shape, the branches of which are formed by lateral volume portions between the side walls and the radially outer wall, and the base of which is formed by a middle volume portion between the middle wall and the radially outer wall.

The invention relates to an assembly for holding an epicyclic gear train in a turbomachine comprising an annular housing in which are engaged a first annular part and a second annular part which are locked against rotation in the housing by first annular interlocking means and second annular interlocking means, respectively, wherein the first annular interlocking means are configured to allow a first circumferential clearance between the first annular part and the housing, the second annular interlocking means are configured to allow a second circumferential clearance between the second annular part and the housing, and the first circumferential clearance is strictly less than the second circumferential clearance.

Assembly for retaining a gear train in a turbomachine

The invention relates to a vane () comprising:

Vane comprising a structure made of composite material and a metal stiffening part

The invention relates to an intermediate case () for a twin spool turbomachine for an aircraft, comprising a hub (), an outer shell () and outlet guide vanes () installed at their ends on the hub and on the outer shell, and each of at least some of the outlet guide vanes () performing a heat exchanger function and comprising a lubricant passage () designed to be cooled by the fan flow () following an outer surface of the outlet guide vane. According to the invention, the case also comprises at least one lubricant duct () passing along a circumferential direction of the hub () and at least part of which is made from a single casting with the hub, the lubricant duct () having at least one lateral opening communicating with the lubricant passage () of at least one of the vanes ().

Intermediate case for an aircraft turbomachine made from a single casting with a lubricant duct

The present disclosure describes a lubrication oil piping gutter of an aircraft turbomachine. The lubrication oil piping gutter can be used for a sun gear train reducer of the turbomachine. The gutter may include two lateral annular walls connected at their outer periphery by an annular bottom wall. In some embodiments, the annular bottom wall may include at least one oil evacuation opening in which is located a deflector configured to facilitate the flow of oil radially from the inside to the outside of the gutter.

Lubricating oil piping gutter of an aircraft turbomachine

A hub for a propeller having variable-pitch blades for a turbomachine with longitudinal rotation axis, including a plurality of blade platforms, each one being designed to receive a blade root, the platforms being distributed around the annular outer periphery of the hub concentrically with the longitudinal rotation axis, wherein each blade platform has the general geometric shape of a spherical cap, and wherein, between two adjacent blade platforms around the longitudinal rotation axis, the transverse dimension, in particular the diameter, of the hub can vary axially and radially.

Hub for propeller having variable-pitch blades, with radial and axial dimensioning variation

A pivot for an adjustable-pitch blade for a turbine engine fan hub, the pivot including a stud having coupling device for transmitting a twisting torque and an outside thread, a force-decoupling part positioned around the stud, provided with an outside thread, pressing transversely against the stud at its radially outer end, an outer rolling bearing having an inside ring mounted to press transversely against the force-decoupling part, an inner rolling bearing assembled around the force-decoupling part and held thereby, a pitch transmission ring provided with a coupling device co-operating, with the coupling device of the stud, a first clamping nut screwed onto the outside thread of the force-decoupling part in order to clamp the outer and inner rolling bearings around the fan hub, and a second clamping nut screwed onto the outside thread of the stud in order to provide clamping between the stud, the force-decoupling part, and the pitch transmission ring.

Variable pitch blade pivot for a turbine engine fan hub

A turbine engine is provided with a longitudinal rotation axis having at least one shaft with a radial axis, in particular a pitch change system for the blades of a propeller, said shaft traversing a radial passage of a substantially cylindrical case around the longitudinal axis. The turbine engine includes an annular oil guiding device around the radial shaft. The device has first and second annular parts nested in one another and secured to one another by hooping. The first part is secured by hooping to the radial shaft, and the second part is configured so as to be separated from the first part under the action of a force oriented along the radial axis and exerted on the second part by a member of the turbine engine forming a stop during a radial movement of the radial shaft.

Turbine engine with an oil guiding device and method for disassembling the turbine engine

The invention relates to an intermediate case for a twin spool turbomachine for an aircraft, comprising at least one lubricant passage and piece () associated with at least one of its vanes (), comprising a lubricant passage () extending along a passage axis () and opening up into the lubricant duct (). The passage axis () and the orifice axis () are also at a spacing from each other along a circumferential direction (), and a connection piece () through which a connection conduit () passes is firstly partially housed in the vane root () so as to create a communication between the connection duct () and the lubricant passage orifice (), and secondly is partially housed in an end-piece support device () or in the lubricant passage end-piece (), such that the connection conduit () communicates with the lubricant passage () in the end-piece.

Growing community of inventors

Hericy, France

Christophe Paul Jacquemard