Snecma

Safran Aircraft Engines

Snecma Moteurs

Societe Nationale D'etude Et De Construction De Moteurs D'aviation

Maurice Guy Judet

Fabrice Marcel Noel Garin

Michel Gilbert Roland Brault

Sébastien Serge Francis Congratel

Thomas Langevin

Fabrice Garin

Stephane Pierre Guillaume Blanchard

Jacques Boury

Ba-Phuc Tang

Patrick Claude Pasquis

Axel Sylvain Loïc Thomas

Gabrijel Radeljak

Damien Bernard Quelven

Fabrice Marcel Garin

Lucien Henri Jacques Quennehen

Nicolas Paul Tableau

Marc R Marchi

Jean-Claude Christian Taillant

Antoine Claude Michel Etienne Danis

Mathieu Dakowski

Stéphane Pierre Guillaume Blanchard

Damien Bonneau

Serge Louis Antunes

Eric Stephan Bil

Jacques Auguste Amedee Boury

Wilfried Lionel Schweblen

Alain Paul Madec

Jean-Baptiste Vincent Desforges

Michel Gilbert Brault

Clement Raphael Laroche

Franck Boisnault

Patrick Rossi

Jacky Tabardin

Fabrice Marcel Noël Garin

Michel Brault

Alexandre Xavier Bossaert

Jerome Etienne Robert Cameau

Matthieu Simon

Isabelle Monique Marie Bourriaud

Laetitia Person

David Thien-Roger Nguyen

Cecile Marie Emilienne Alirot

Christian Michel Jacques Gosselin

Clement Jarossay

Laurent Jacques Michel Coulombeau

Guillaume Michel Maurice Giliberti

Pierre Guillaume Auzillon

Jean-Baptise Vincent Desforges

Cecile Marie Alirot

Christian Michel Gosselin

Axel Sylvain Thomas

Laurent Jacques Coulombeau

Cécile Marie Emilienne Alirot

Labyrinth seal for a turbine engine, in particular of an aircraft, including a rotor element rotating about an axis of rotation, and a stator element extending around the rotor element The rotor element includes a series of annular lips extending radially outwards and surrounded by at least one abradable element carried by the stator element. Each lip includes an inner peripheral body portion, an outer peripheral body portion and an upstream annular face for impact of an air flow during operation. At least one lip has, looking from the upstream annular face, a first annular cavity with a concave rounded cross-section on its inner peripheral body portion and a second annular cavity with a concave rounded cross-section on its outer peripheral body portion.

Labyrinth seal for a turbine engine of an aircraft

Assembly () comprising an annular turbine distributor () for turbomachinery (), an annular structural element () of the turbomachinery (), coaxial with the distributor () and at least one device () for attaching the annular distributor () with the annular structural element (), characterised in that: &#x2014;the annular distributor () is crossed by at least two axial passages, and &#x2014;the or each attachment device () comprises: at least two pins (), rigidly connected to the annular structural element () and passing through one of said passages of the distributor (), a first transverse stop () for stopping the distributor () axially and passing through two consecutive pins (), and two second stops () rigidly connected to the structural element () to retain, in the pins (), the first transverse stop () for stopping the distributor () radially.

Assembly for a turbomachine comprising a distributor, a structural element of the turbomachine, and an attachment device

A support case for a turbomachine ring sector includes an annular body around a longitudinal axis of the case, and a downstream edge for attachment of the ring sector. The downstream edge extends from the body radially towards the inside of the case. The downstream edge extends circumferentially along a junction region between the body and a fixed vanes stage. The junction region includes a central portion and two lateral portions located on each side of the central portion, along a circumferential direction of the case with respect to its longitudinal axis. The central portion projects radially outwards from the case with respect to the lateral portions.

Turbomachine case comprising a central part projecting from two lateral portions in a junction region

A mobile vane for a turbine engine, including a root designed to be inserted into a receiving element of a rotor disk for a turbine engine, a platform carried by the root, and a blade extending from the platform. The platform includes an upstream edge. The upstream edge includes a lug for engaging in a locking notch of the disk in such a way as to hold the vane axially in relation to the disk, according to the longitudinal direction of the receiving element.

Mobile vane for a turbine engine, comprising a lug engaging in a locking notch of a rotor disk

The present invention relates to an assembly () comprising:

Assembly for attaching a nozzle to a structural element of a turbine engine

A turbine blade for an aircraft turbomachine including a root, an airfoil and a platform inserted between the airfoil and the root and delimiting a gas circulation flowpath, the platform having two axial ends each forming an angel wing of which at least one has an internal cavity that will be supplied with air from the root of the blade. At least one of the two angel wings is drilled with at least one bleed hole for passage of a bleed flow from the internal cavity that will limit/prevent gas recirculation outside the flowpath.

Mobile turbine blade with an improved design for an aircraft turbomachine

The present invention relates to a turbine assembly () of a turbine engine (), comprising at least: a first bladed rotor (), a bladed stator () and a second bladed rotor () arranged in series, the rotors () being mounted on a shaft (); a sealing plate () extending between the stator () and the shaft () and separating a first recess (C) arranged between the first rotor () and the stator (), from a second recess (C) arranged between the stator () and the second rotor (); and pressure-reducing means () positioned inside the first recess (C), the assembly being characterized in that said pressure-reducing means () comprise a plurality of substantially radial recompression fins () extending into the first recess (C).

Turbine assembly of an aircraft turbine engine

The turbine () has a rotor () with blades () arranged on a periphery of disks () in radial planes (), and a tie rod () extending along slots () in the disks. An annular spacer (') fixes a position of the tie rod relative to a center line (M) of the disks, and is arranged in a channel (). The spacer has through-openings, which are arranged radially relative to the tie rod or to its center line, and extend coaxially. The channel carries a cooling medium and is limited by a separating pipe () radially outwards. The through-openings serve for flow of the medium.

Turbomachine comprising a shaft sleeve and associated sleeve tube

The invention relates to a trunnion () for a high-pressure turbine (), to be arranged between a shaft of a low-pressure turbine () and an inner surface () of a seal mounting () of the low-pressure turbine (), the trunnion () being characterized in that it includes a drop-launching extension () arranged such as to extend opposite a flared portion () of the inner surface () of the seal mounting (), such that when the trunnion () is rotated about the shaft of the low-pressure turbine (), oil (H), which tends to penetrate between the trunnion () and the seal mounting (), is thrown by centrifugal effect from the drop-launching extension () toward the flared portion () of the inner surface () of the seal mounting ().

Trunnion for high-pressure turbine, and turbojet engine including such a trunnion

A turbine wheel for a turbine engine, comprising a disk carrying blades each having a platform, carrying an impeller connected by a stilt to a root. Each blade comprises an upstream radial wall and a downstream radial wall extending inwards from the platform of the blade. The roots of the blades are engaged into slots on the periphery of the disk, so that the radial walls of the blades and the platforms of the blades are circumferentially arranged end-to-end, and define inter-blade cavities radially inside the platforms, each accommodating an inter-blade sealing and vibration damping member. At least one member comprises an elastically deformable element configured so as to be elastically prestressed against the walls of a cavity in the member mounting position.

Turbine wheel for a turbine engine

A rotating assembly for a turbine engine, comprising a disc having an outer periphery having alternating slots and teeth, blades radially extending from the disc and roots of which are axially engaged in the slots, with spaces called slot cavities being provided between the roots of the blades and the slots, platforms laterally extending from the blades and circumferentially arranged end-to-end, so as to form spaces called inter-blade cavities, and a downstream annular shroud, comprising an outer annular sealing lip opposite the downstream ends of the platforms. The downstream shroud further comprises an intermediate annular sealing lip opposite the downstream faces of the teeth of the disc, radially between the slot cavities and the inter-blade cavities.

The present disclosure aims at achieving a mounting of a turbomachine stator blading while avoiding a delicate machining process of the engine casing. It also aims at simplifying the mounting as well as at reducing the costs, and at decreasing the number of item references for the engine. To achieve these objects, the turbomachine uses an extension of the engine casing. A turbomachine comprises at least one stator blading for air or gas stream flowing on the internal engine casing. The stator blading comprises a rim which is assembled on a radial extension of engine casing with an axial fastener. In a preferred embodiment of the turbomachine, the engine casing presents an annular wall provided with fingers originating axially from this wall and extending according to a curvature so that they present a radial extension tip against the rim of the stator blading.

Method for mounting a stator blading of a turbomachine, an engine casing and a turbomachine comprising at least one stator blading mounted on this engine casing

An assembly with a turbine disk for a gas turbine engine and a bearing-supporting journal, the turbine disk having a radial annular fastening flange secured to a radial annular part of the journal by bolts, the bolts passing successively through fastening bores formed in the radial annular fastening flange of the turbine disk and in the radial annular part of the journal, the radial annular part of the journal having air circulation openings, the openings being formed between the fastening bores of the journal.

Assembly including a turbine disk for a gas turbine engine and a bearing-supporting journal, and cooling circuit for the turbine disk of such an assembly

A stationary vane assembly for a turbine engine including an inner casing and at least two angular sectors forming a high-pressure guide vane assembly, each angular sector including a fastening tab for fastening the angular sector on the casing. The stationary vane assembly including a ring of abradable material fixed on the casing, on the side of a mobile vane assembly using a plurality of screw-nut assemblies, the stationary vane assembly also including radial retaining pins for the angular sectors introduced into first bores formed in the fastening tabs and in the mounting clip, the bores having axes substantially parallel to the longitudinal axis, the mounting clip being fixed on the casing by the same screw-nut assemblies as those fixing the ring of abradable material support on the casing.

Fixed vane assembly for a turbine engine having a reduced weight, and turbine engine comprising at least one such fixed vane assembly

A turbine engine including a channel for injecting a cooling air flow of a high-pressure turbine disk, opening into a cavity that is substantially isolated, upstream, from a cavity in which an air flow sampled at the output of a high-pressure compressor circulates, by a first labyrinth seal, and downstream, from a cavity communicating with the primary flow of the turbine engine, by a second labyrinth seal. The turbine engine includes channels communicating with the injection channel and opening through a static part of the first labyrinth seal between two lips of that seal, so as to allow an air flow coming from the injection channel to be injected between the lips.

Turbine engine including an improved means for adjusting the flow rate of a cooling air flow sampled at the output of a high-pressure compressor using an annular air injection channel

Growing community of inventors

Dammarie les Lys, France

Maurice Guy Judet