Snecma

Safran Aircraft Engines

Jean Bertucchi

Matthieu Leyko

Mathieu Gaillot

Damien Joseph Cellier

Julien Szydlowski

Ghislain Maxime Romuald Madiot

Irwin Pierre Robert Kernemp

Stephane Emmanuel Daniel Bensilum

Aurélien Mazzella

Irwin Kernemp

Arthur Droit

Stéphane Emmanuel Daniel Bensilum

Gaétan Jean Mabboux

Stéphane Bensilum

A part or an assembly of parts of a splitter stage of a turbofan, including a plurality of obstacles including at least one a blade and at least one arm, and a platform from which the obstacles extend, wherein the arm is defined by a skeleton angle having a trailing value on the trailing edge of between 3&#xb0; and 15&#xb0;.

Dimensioning of the skeleton angle of the trailing edge of the arms crossing the by-pass flow of a turbofan

A discharge system of a separated twin-flow turbojet for an aircraft, supported by a suspension mast, is disclosed. The system includes a main nozzle delimited by an annular cowl with a slot of annular shape defining upstream and downstream portions of the cowl and which is traversed by the suspension mast. The downstream portion of the cowl of the main nozzle includes a first part extending downstream from the upstream portion of the cowl to a trailing edge of the main nozzle, on either side of the suspension mast along two predefined angular sectors; and a second part formed from an internal contour of the slot and having a trailing edge with a diameter smaller than that of the trailing edge associated with the first part of the downstream portion of the cowl. Connecting walls laterally connect the first and second parts of the downstream portion.

Discharge system of a separated twin-flow turbojet for an aircraft, corresponding turbojet and associated design method

A turbine engine convergent-divergent nozzle including an annular central element and an annular cap arranged coaxially around the central element to co-operate with the central element to define an annular flow channel for a gas stream from the engine. Between a throat section and an ejection section of the nozzle, the central element and the cap present respective internal profiles in longitudinal section that are modeled by curves having respective radii of curvature that are identical in absolute value.

Convergent-divergent nozzle for a turbine engine

A pylon for fixing an aircraft engine having unducted pusher propellers, the pylon ensuring the fixation of a propulsive system on the boattail of the aircraft, the pylon having a trailing edge, with an upper face and a lower face, for an airflow encountered by the pylon, wherein at least one of the two faces of the upper face and the lower face of the trailing edge is inclinable, at least in part.

Pylon for fixing an aircraft engine having unducted pusher propellers

A turbomachine including a nozzle extending along an axial direction and configured to convey a hot gas stream; and a mounting pylon for mounting the turbomachme, in which the pylon extends radially from a base arranged in the vicinity of the nozzle, is provided. The turbomachine includes a pylon cooler device having at least one auxiliary gas duct having a gas inlet configured to take cold gas from the outside of the nozzle, and a gas outlet opening out in the vicinity of the base of the pylon.

Turbomachine comprising a device for the cooling of a pylon

In an assembly including a bypass turbine engine and an attachment system attaching the turbine engine to an aircraft structure, the attachment system includes a mast, at least two rear attachments connecting the mast to two rear attachment points of the turbine engine, and a rear covering covering the rear attachments and providing a streamlined fairing. The rear covering includes two side walls, each including a first portion extending along and in proximity of the mast and extending rearwards, and a second portion projecting sideways relative to the first portion towards its outside, to cover the rear attachments. Each second portion is substantially a bulge and is connected to the first portion along a curved connection line extending between front and rear points situated substantially level with an inner longitudinal edge of the first portion to minimize the impact on a flow of bypass stream leaving the bypass passage.

Assembly formed by a turbine engine and a system for attaching it to an aircraft structure

A turbine engine pipe having a longitudinal axis, which, at an upstream end thereof, includes an outer wall and a core, only an upstream portion of which is surrounded by the outer wall. The boundary between the upstream portion and a downstream portion is defined by the intersection between the core and the surface formed by bisectors of acute angles formed by straight lines perpendicular to the outer wall and straight lines perpendicular to an outer surface of the core passing through the downstream end of the outer wall. The outer surface of the downstream portion of the core has a sharply curved convex region close to the boundary, the region having a radius of curvature in a radial plane of less than 30 times the radius of the core at the boundary, and a first slightly curved region between the boundary and the sharply curved convex region.

Growing community of inventors

Thiais, France

Jean Bertucchi