Safran Aircraft Engines

Snecma

Philippe Gérard Edmond Joly

Laurent Jablonski

Jean-Marc Claude Perrollaz

Romain Nicolas Lagarde

Eric Jacques Boston

Charles Jean-Pierre Douguet

François Jean Comin

Damien Merlot

Christophe Perdrigeon

Michel Andre Bouru

Francois Jean Comin

Michel André Bouru

Edouard Antoine Dominique Marie De Jaeghere

Hervé Pohier

Adrien Jacques Philippe Fabre

Rémi Roland Robert Mercier

Anthony Lafitte

Romain Jean-Louis Robert Thory

Pierre-Guillaume Bardin

Marie-Oceane Parent

Etienne Ide

Jean Marc Claude Perrollaz

Hervé Gignoux

An assembly for a turbine engine extending along an axis, includes a stator, a low-pressure compressor shaft, a low-pressure compressor comprising a rotor comprising a drum rotatably coupled to the shaft of the low-pressure compressor, and a fan comprising a disk rotatably coupled to the shaft of the low-pressure compressor. The drum has a radially internal part rotatably coupled to the shaft of the low-pressure compressor and a radially external part fixed to the radially internal part by means of detachable fixing means. A first axial retention means is configured to axially and detachably retain the disk of the fan relative to the shaft of the low-pressure compressor. A second axial retention means is configured to axially and detachably retain the radially internal part of the drum relative to the shaft of the low-pressure compressor.

Assembly for a turbine engine

A turbomachine assembly includes a casing, first and second rotors, and a damper. The first rotor includes a disk and blades and is movable in rotation relative to the casing. The second rotor is movable relative to the casing around a longitudinal axis. The damper damps a movement of the first rotor relative to the second rotor. The damper includes first to third parts. The first part bears on the first rotor in a first area extending over a first angular sector around the longitudinal axis and applies a first centrifugal force on the first rotor. The second part bears on the first rotor in a second area that is smaller than the first angular sector and extends over a second angular sector around the longitudinal axis. The third part bears on the second rotor and applies a second centrifugal force on the second rotor.

Assembly for turbomachine

The present invention relates to an assembly for a turbomachine, comprising: a first rotor; a second rotor, and a damper () configured to damp a displacement of the first rotor relative to the second rotor, the damper comprising: a first portion () bearing against the first rotor and having a first thickness, a second portion () bearing against the second rotor and having a second radial thickness, and a third portion () connecting the first portion () to the second portion () and having a third radial thickness, wherein the third radial thickness is greater than at least one of the first radial thickness and the second radial thickness.

Assembly for a turbomachine

The present invention relates to a turbomachine assembly, comprising: a casing (), a first rotor () which is movable in rotation with respect to the casing () about a longitudinal axis (X-X), and comprising: *a disk (), and *a plurality of blades () capable of flapping with respect to the disk () during a rotation of the first rotor () with respect to the casing (), a second rotor () which is movable in rotation with respect to the casing () about the longitudinal axis (X-X), and a damper () which is configured to damp a displacement of the first rotor () with respect to the second rotor () in a plane orthogonal to the longitudinal axis (X-X), the displacement being caused by a flapping of at least one blade () among the plurality of blades (), the damper () comprising: o a first bearing part (): *bearing against the first rotor (), and *being configured to apply a first centrifugal force (C) to the first rotor (), o a second bearing part (): *bearing against the second rotor (), and *being configured to apply a second centrifugal force (C) on the second rotor (), and o a linking part (): *connecting the first bearing part () to the second bearing part (), and being thinned relative to the first bearing part () and the second bearing part (), and o a flyweight () which is fixedly mounted on the damper ().

Turbomachine assembly having a damper

The invention relates to a bladed disk () of a fan, comprising: a hub () comprising an outer radial platform () designed so as to define an inner gas flow stream in the turbomachine, a plurality of blades () comprising a root () connected to the platform (), a leading edge () and a trailing edge (), a groove formed in the platform () around part of the root () of each blade () in an area adjacent to the leading edge () and/or the trailing edge (), and a joint () placed in the groove () in such a way that it extends in the extension of the radially outer face () of the platform () in order to ensure a continuity of the inner flow stream.

Bladed disk flexible in the lower part of the blades

The invention relates to an assembly () for a turbomachine comprising: characterized in that the damping device () is annular while extending circumferentially around the turbomachine longitudinal axis (X-X) and in that the damping device () comprises a first radial external surface () supported with friction against the first module () as well as a second radial external surface () supported with friction against the second module (), so as to couple the modules () in order to damp their respective vibrational movements during operation.

Damping device

A turbomachine assembly is provided. The assembly includes: a first rotor module including a first blade, a second rotor module, connected to the first rotor module, and including a second blade of smaller length than the first blade, and a damping device including a plurality of first radial external surfaces supported with friction against the first module. The damping device is stepped, and includes a second radial external surface supported with friction against the second module, so as to couple the modules for the purpose of damping their respective vibrational movements during operation.

An assembly for a turbomachine including a first rotor module () with a first blade (), and a second rotor module (), connected to the first rotor module (). The second rotor module includes a second blade with a smaller length than the first blade (). A damping device () is attached to the second rotor module () and includes a radial external surface () supported with friction against the first module (), so as to couple the modules () for the purpose of damping their respective vibrational movements during operation.

The invention relates to a turbomachine assembly () comprising:

The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream at the radially inner ends of the leading edges of the fan blades, divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.

Fan for a turbomachine

The invention relates to a blade, the platform () of which has at least one bumper () which rises, on the inner face, from a connection to an intermediate area of the blade located along a longitudinal axis of the blade, between the blade root and the outer face (), towards a free side edge () of the platform. The platform bumper () locally has a thinned portion at a distance from said free side edge of the platform. The invention is applicable to turboshaft engine fans.

Blade with a platform and a hollow bumper

A fan blisk for a turbomachine, the blisk comprising a hub delimited by an upstream face and a downstream face in addition to an outer peripheral face and a revolving inner peripheral face delimiting an inner opening, the hub carrying blades each having a leading edge and a trailing edge, the hub and the blades forming a one-piece assembly. The upstream face and/or the downstream face is offset, being located along the axis of rotation (AX) between the leading edges and the trailing edges of the blades, the blisk (bladed disk) comprising a filling member mounted on the offset face, the filling member comprising an inner centring ferrule engaging in the inner face of the hub, a radial portion resting against the offset face and an outer ferrule extending the outer peripheral face of the hub.

Blisk comprising a hub having a recessed face on which a filling member is mounted

Relates to a vane for a turbomachine. The vane (') has a blade (&#x2032;) and a root (&#x2032;) to be engaged in an axial groove in a disc of the turbomachine. The upstream end (&#x2032;) of the root is connected to a radially internal end (&#x2032;) of the leading edge (&#x2032;) of the blade by the upstream end of a connecting zone having a discontinuity towards the downstream end, so that said radially internal end of the leading edge of the blade is situated further downstream than the upstream end of the root.

Vane with spoiler

The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream () at the radially inner ends of the leading edges of the fan blades (), divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.

Growing community of inventors

Vaux Le Penil, France

Philippe Gérard Edmond Joly