Safran Aircraft Engines

Safran Ceramics

Nicolas Xavier Trappier

Stéphane Pierre Guillaume Blanchard

Joao Antonio Amorim

Caroline Marie Frantz

Baptiste Dorian Lawniczek

Simon Jean-Marie Bernard Cousseau

Patrice Jean-Marc Rosset

Antoine Hervé Dos Santos

Maxime Aurelien Rotenberg

Patrick Joseph Marie Girard

Quentin De Richaud

Michel Gilbert Roland Brault

Sébastien Serge Francis Congratel

Matthieu Arnaud Gimat

Antoine Jean-Philippe Beaujard

Arnaud Nicolas Negri

Romain Guillaume Cuvillier

Guillaume Patrice Kubiak

Nils Edouard Romain Bordoni

Julian Nicolas Girardeau

Romain Jean-Claude Ferrier

Nora El Ghannam

Maxime Aurélien Rotenberg

Matthieu Etienne Attali

Helene Myriam Condat

Laurent Paul François Perrot

Pierrot Guillou

Benoit Jean Henri Gomanne

Thomas Gerard Daniel Riviere

Pascal Grégory Casaliggi

Arthur De Castro

Didier René André Escure

Nadège Hugon

A bladed assembly () includes two adjacent blades (). Each blade () a platform () located at a free radial end. The platform () of the first blade () includes a lateral face () which is located circumferentially opposite an associated lateral face () of the platform () of the second blade (). The platform () of the second blade () comes into contact with the platform () of the second blade () only via an excrescence () which projects circumferentially with respect to the lateral face () of the platform () of the first blade () in the direction of the platform () of the second blade (), and the excrescence () is received in a groove () formed in the platform () of the second blade.

Turbomachine blading assembly comprising means for limiting vibration between platforms

A turbine rotor of a turbine engine of an aircraft, the rotor comprising a rotor disc provided at its periphery with a plurality of cavities for receiving the root of a blade. Each blade includes a circumferential coupling rib, a plurality of platforms, a main wall that is extended at its two ends by an upstream radial wall and by a downstream radial wall, the two longitudinal edges of each platform comprising a groove for receiving part of said coupling rib of one of the two blades between which this platform is assembled, and in that the height of each upstream radial wall is such that said wall extends radially inwards beyond the bottom of the cavity, until it is in front of the upstream face of the rotor disc.

Turbine rotor and platform for such a rotor

A rotary assembly for a turbomachine including a rotor including at least two consecutive rotor stages provided with a plurality of blades and an annular rotor shroud connecting the two consecutive rotor stages, a stator including: at least one stator stage provided between the two consecutive rotor stages including a plurality of vanes, a turbomachine stator vane root, an annular clamping part and an annular support of abradable material, the root extending radially and being clamped axially between the annular support of abradable material and the annular clamping part, A space separates a radially internal end of the root and the annular support of abradable material. A turbojet engine including a rotary assembly as previously.

Turbomachine rotary assembly comprising an annular clamping part

An assembly for a turbine engine turbine includes a turbine rotor disc centered on a longitudinal axis and a turbine shaft centered on the longitudinal axis and driven in rotation by the rotor disc. Torque from the rotor disc is transmitted to the shaft, wherein the rotor disc is locked in translation relative to the shaft in the direction of the longitudinal axis by a screwed member on the shaft. Torque from the rotor disc is transmitted from the rotor disc to the screwed member when the torque ceases being transmitted from the rotor disc to the shaft. The screwed member has an unscrewing direction identical to the direction of rotation of the rotor disc in operation.

Turbine overspeed disengagement device for a turbine engine

A turbine shaft of a turbomachine extending along a longitudinal axis (A) includes a main tubular portion with a length Land a maximum outer main diameter. The shaft further includes a secondary tubular segment with a length Land a maximum outer secondary diameter greater than the maximum outer main diameter. The secondary tubular segment extends radially from the shaft such that the main tubular portion extends on either side of the secondary tubular segment along the longitudinal axis (A). The secondary tubular segment includes a weak point that is configured to break the shaft into two distinct parts when the value of a tangential stress applied to the shaft exceeds a predetermined threshold value, and thus cause retraction of the turbine.

Turbine shaft of a turbomachine and method for protecting against overspeed of said shaft

Assembly including a rotor disc, an adjacent stator and a plurality of sealing elements secured to the rotor disc, the stator including an inner platform and a root bearing at least one abradable element configured to cooperate with the sealing elements, the sealing elements being placed in an enclosure formed by the abradable element, the enclosure being open to the inside and delimited axially by an upstream abradable edge and a downstream abradable edge, the enclosure being delimited radially by an outer abradable edge, at least one of the sealing elements including a first lip configured to cooperate with the upstream abradable edge or the downstream abradable edge, and a second, separate lip configured to cooperate with the outer abradable edge.

Sealing between a rotor disc and a stator of a turbomachine

A turbomachine includes a low-pressure shaft that drives a fan shaft in rotation by a coupling assembly including a first shaft on which are formed plural first axial grooves, a second shaft on which are formed plural second axial grooves, and a coupling device including rolling elements and an annular cage positioned between the first shaft and the second shaft. The rolling elements are positioned between one of the first axial grooves and one of the second axial grooves to couple in rotation the first and the second shaft. Each first and second groove has a first and a second substantially planar surface inclined with respect to one another and extending along the axis.

Torque transmission device with reduced friction

The invention relates to an assembly () for a turbine of a turbomachine, comprising: a first rotor disk (), a second rotor disk (), a part forming a mobile ring (), comprising a system for preventing rotation of the mobile ring () relative to the rotor disks (), said system comprising: a rotor disk securing flange () having a plurality of teeth () that are distributed circumferentially about the turbomachine longitudinal axis (X-X), and a mobile ring securing flange () having a plurality of lugs () that are distributed circumferentially about the turbomachine longitudinal axis (X-X), the engagement of the rotor disk securing flange () with the mobile ring securing flange () ensuring, by means of the teeth () and the lugs (), that the mobile ring () does not rotate relative to the rotor disks ().

Assembly for a turbine of a turbomachine comprising a mobile sealing ring

The present invention relates to a rotor () of a turbomachine turbine, comprising a rotor disc () and a plurality of blades () distributed at its periphery, said rotor disc () comprising a plurality of mainly axial cells (), each blade () comprising a root (), retained in one of the cells () of said rotor disc, each root being dimensioned so as to form a space () between the bottom of the cell () and the radially inner face () of the root. This rotor is remarkable in that said root comprises a mainly axial blind hole () opening onto the upstream face of the root and a plurality of mainly radial air ejection orifices (), each air ejection orifice () opening into said blind hole () and onto the radially inner face () of the root located facing the bottom of the cell ().

Rotor, turbine equipped with such a rotor and turbomachine equipped with such a turbine

The invention relates to a system for rotating a fan () of a turbojet engine () about a first rotation axis &#x3a9;, including a reduction gear () made up of a planetary gearset placed at the centre of the fan (), which includes: a central sun gear (); an outer planet gear () attached to the fan (); at least one satellite gear () arranged between the central sun gear () and the outer planet gear () in order to transmit a rotary movement between the central sun gear () and the outer planet gear (); characterised in that the outer planet gear () includes an inner portion (a) which meshes with the satellite gear (), as well as an outer portion (b) to which blades () of the fan () are directly attached.

System for rotating a fan of a turbojet engine

The invention relates to an aircraft propulsion assembly comprising a cradle receiving a turbine engine comprising at least one propeller having a longitudinal axis of rotation, a gas turbine engine having a longitudinal axis of rotation offset from the axis, and a reduction gear by means of which said propeller receives drive power from said engine, wherein the propeller and the gas turbine engine are designed such that axes and are offset from one another within said cradle at least by a given value in a transverse direction, the axis of the gas turbine engine being transversely closer to a proximal lateral side of the cradle than to an opposite distal lateral side of the cradle in order to create a lateral space between said engine and said distal lateral side of the cradle, thereby forming at least one region for installing equipments, components or accessories of said turbine engine.

Turbine engine having horizontally offset axes

The rotor disk (B) for a compressor comprises, relative to the rotational axis of the disk: a radial web (), blades () at the outer periphery of the web, a bore () at the inner periphery of the web, and a cylindrical side wall () extending the web in the vicinity of the ter periphery of same and having an air supply port (), and&#x2014;a centripetal air collection device (). Advantageously, the device () comprises a cylindrical support () and at least one air supply tube (), the inlet of which is turned towards the port () and the outlet of which is turned towards the bore () in the web, the disk comprising an inner radial flange () extending from the cylindrical side wall (), the cylindrical support () of the device () being attached to said inner radial flange (), and a ring () extending from the web (), the cylindrical support () being centred on the ring ().

Rotor disk having a centripetal air collection device, compressor comprising said disc and turbomachine with such a compressor

A turbine rotor is fitted to a turbomachine and includes disks each containing a coupling portion with recesses each holding a root of a blade, and coupled to one another by a first annular ferrule attached to one of them close to the recesses, and second rotationally coupled ferrules, which are respectively coupled radially to the disks, which each consists of at least two semi-annular sectors, and which form with the associated disk a space which communicates with the recesses of the coupling portion of this latter disk. In addition, each first ferrule includes through-holes enabling air to enter this space, and then the recesses, intended to cool the coupling portion which couples its disk and the blade roots.

Turbine rotor with air separation ferrules for cooling of blade and disk coupling portions, for a turbomachine

A rotationally symmetrical part, such as a disk, for a turbine engine rotor. The part has a rotational axis (A) and includes, an annular row of teeth defining grooves therebetween for therein holding fir tree-shaped rotor blade roots. Each tooth includes a first side flank, having at least two projecting portions for holding a blade root and separated from each other by a hollow portion, and a second side flank including at least two projecting portions for holding an adjacent blade root and separated from each other by a hollow portion. Each tooth has, on substantially the entire longitudinal dimension thereof, a lack of symmetry in relation to a substantially radial median longitudinal plane. At least one of the circumferences of the projecting portions of the second flank is located between the circumferences of the projecting portions of the first flank and is radially shifted from the circumferences.

Rotationally symmetrical part for a turbine engine rotor, and related turbine engine rotor, turbine engine module, and turbine engine

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.

Growing community of inventors

Moissy-Cramayel, France

Nicolas Xavier Trappier