Safran Aircraft Engines

Institutul National De Cercetare-dez-voltare Turbomotoare - Comoti

Norman Bruno André Jodet

Jéremy Paul Francisco Gonzalez

Jérémy Paul Francisco Gonzalez

Jacky Novi Mardjono

Georges Jean Xavier Riou

Josselin David Florian Regnard

Laurent Louis Robert Baudoin

Stéphane Orcel

Francis Couillard

Kevin Morgane Lemarchand

Mathieu Simon Paul Gruber

Anthony Binder

Eva Julie Lebeault

Fernando Gea Aguilera

Laurent Soulat

Guillaume Martin

Laurence Francine Vion

Arnaud Dubourg

Edith-Roland Fotsing

Annie Ross

Jean-Michel Boiteux

Guillaume Bodard

Hervé Rolland

Jean-Michel Daniel Paul Boiteux

Adrien Clément Marcel Dubois

Cristian Mihail Stanica

Rasika Fernando

Pascal Jérôme Romano

Daniel Therriault

Radu Mihalache

Romulus Petcu

Valentin Silivestru

Jean Michel Roux

Lucia Raluca Voicu

An aeronautical thruster having a longitudinal axis and having a hub and at least two annular rows of unducted blades including an upstream annular row and a downstream annular row which are spaced apart from one another along the longitudinal axis, the upstream annular row being rotatable around the longitudinal axis, and the downstream annular row comprising a series of blades including a first blade and a second blade each extending in a radial direction from the hub to define a radial dimension between the hub and a radially outer end of the corresponding blade, wherein the radial dimension of the first blade is greater than the radial dimension of the second blade.

Aeronautical thruster

A cascade for a thrust reverser device intended to be mounted on a turbomachine of an aircraft, the cascade including first partitions extending in a first direction, seconds partitions extending in a second direction orthogonal to the first direction, the second partitions extending in a third direction secant to a plane comprising first and second directions between a first and a second end. Each second partition forms, with reference planes parallel to the second direction and to the third direction, a plurality of different angles at positions distinct from the height of the second partitions separating the first end from the second end of the second partitions, each angle being formed between a reference plane and the tangent to the second partition taken at the intersection between said corresponding reference plane and the second partition at the position of the height of the corresponding second partition.

Thrust reverser cascade including acoustic treatment

The invention relates to acoustic management, on an aircraft turbomachine or on a nacelle, via a panel (). On a support () is reserved a recess (), recessed with respect to a surrounding general surface () for contact with moving air. The recess () is adapted to receive the panel, as another so-called surface for contact with moving air. The support () and/or the panel comprise removable connecting elements for, in the recess, mounting the panel removably with respect to the support, the panel being an acoustic panel () or a non-acoustic panel ().

Acoustic management, on a turbomachine or a nacelle

A manufacturing method of a turbomachine airfoil, such as an outlet guide vane airfoil, comprising positioning a first fibrous wall preform on a first mold portion, placing at least one core on the first wall preform, positioning a second fibrous wall preform on the core, assembling a second mold portion to the first mold portion so as to form a mold around the first and second wall preforms, applying a hardening treatment to the first and second wall preforms, removing the core, and positioning a reinforcing structure between the first wall preform and the second wall preform.

Blade for a turbomachine and method for the manufacture thereof

A cascade type thrust reverser device for a turbomachine of an aircraft, comprising a cascade including first partitions, second partitions intersecting the first partitions, and cavities, and a casing including a housing into which the cascade can be inserted in a first direction, the casing and the cascade being in relative translation with respect to one another in the first direction. The casing comprises a perforated wall intended to be in contact with an air flow and including orifices and wall strips with no orifices and intended to face the first walls of the cascade when the device is in a first position in which the cascade is entirely positioned in the housing.

Acoustic attenuation on a turbine engine wall

A cascade type thrust reverser device for a turbomachine of an aircraft, comprising a thrust reverser cascade and a casing, the cascade including first cavities, and the casing comprising an opening defining a housing wherein said cascade can be inserted in a first direction, and the casing and said cascade being in relative translation with respect to one another in the first direction between a first position of the device in which the cascade is entirely positioned in the housing and a second position of the device in which said cascade is at least partially outside said housing. The casing comprises an acoustic treatment panel including second cavities extending in a second plane parallel to the first plane, each first cavity facing a second cavity when the device is in the first position to form an acoustic treatment cell.

Internal structure for a turbomachine nacelle, the internal structure being intended to surround at least one part of a compartment capable of receiving a gas generator of the turbomachine, the internal structure comprising a ventilation cavity of said compartment, the ventilation cavity being provided with a main ventilation outlet and an auxiliary ventilation outlet separate to the main ventilation outlet, the internal structure comprising a shut-off member movable relative to the ventilation cavity between a flow position and a shut-off position in which the shut-off member shuts off the auxiliary ventilation outlet to a greater extent than in the flow position.

Internal structure for a turbomachine nacelle

A cascade for a thrust reversal device intended to be mounted on a turbomachine of an aircraft, the cascade including first partitions extending in a first direction, second fixed partitions extending in a second direction orthogonal to the first direction, and a frame inside which the first and second partitions extend, the frame including at least two fixed walls extending according to the first direction, and at least one part of each first partition extending between two second partitions. At least one first partition is mobile according to the second direction between a first position wherein the first partition is distant, in the second direction, from the fixed walls to form a plurality of resonating cavities with the first partitions and/or the fixed walls, and a second position wherein the one first partition is in contact with a fixed wall or another first partition.

The invention relates to acoustic management, on an aircraft turbomachine () or on a nacelle (), via a panel (). On a support () is reserved a recess (), recessed with respect to a surrounding general surface () for contact with moving air. The recess () is adapted to receive the panel, as another so-called surface for contact with moving air. The support () and/or the panel comprise removable connecting elements for, in the recess (), mounting it removably with respect to the support, the panel being an acoustic panel or a non-acoustic panel.

Method for preparing a support and for acoustic management on a turbine engine or a nacelle

An acoustic panel for an aircraft propulsion unit, comprises a core and an entry layer in contact with the core, the entry layer comprising lower elements and upper elements, each comprising an inner face, an outer face and at least one through-hole, the inner faces of the lower elements and upper elements being in contact with each other and forming a channel defining a baffle and opening into a cell of the core.

Acoustic panel and associated propulsion unit

A nacelle for a turbomachine is provided and includes an inner wall that is annular about a longitudinal axis of the nacelle, the annular inner wall being designed to surround part of the turbomachine. The nacelle further includes an annular outer wall surrounding the annular inner wall. The annular outer wall includes a first acoustically porous part, and the annular inner wall including a second acoustically porous part. The first and second porous parts are arranged facing one another so as to allow soundwaves, emitted by the turbomachine housed in the nacelle, to pass through the annular inner wall and then the annular outer wall so as to escape from the nacelle.

Turbomachine nacelle having acoustically porous walls

A flow-straightener vane of a bypass turbomachine includes a plurality of vane sections stacked radially with respect to a longitudinal axis (X) along a stacking line (L) between a root end and a tip end. Each vane section has a pressure-face surface and a suction-face surface extending axially between an upstream leading edge and a downstream trailing edge. Between the leading and trailing edges of each vane section there is formed a profile chord (CA) the length of which is substantially constant between the tip end and the root end, and the stacking line (L) exhibits a curvature in a plane passing more or less through the axis (X) and through the stacking line (L), situated in the vicinity of the tip end and oriented from downstream towards upstream.

Turbomachine fan flow-straightener vane, turbomachine assembly comprising such a vane and turbomachine equipped with said vane or said assembly

An acoustic treatment panel intended to be disposed on at least one wall of a turbojet engine in contact with a fluid flow, the panel including a first acoustically reflective plate, a second plate and a plurality of cavities mounted between the first plate and the second plate and including a plurality of cells. The second plate is a one-piece plate through which a plurality of channels pass, each opening out on the one hand onto a first orifice formed on a first face of the second plate and on the other hand onto a second orifice formed on a second face of the second plate, the length of each channel extending between its first orifice and its second orifice being greater than the thickness of the second plate.

Acoustic treatment panel for a turbojet engine

An assembly for the rear of a bypass turbomachine () comprises a primary nozzle () comprising a trailing edge defining a primary flow path portion and a secondary nozzle () defining a secondary flow path portion, defined about a longitudinal axis (X), said secondary nozzle being configured to eject a mixture of the flows coming from a secondary flow path (Vs) and from a primary flow path (Vp) of the turbomachine (), the secondary nozzle being of convergent-divergent shape with a throat () corresponding to a minimum cross section of the nozzle (), the secondary nozzle () comprising, at the throat (), a periodic succession of lobes () which are situated along the internal circumference of the secondary nozzle (). The assembly also comprises a lobed mixer () at the downstream end of the primary nozzle (), this having an alternation of hot lobes () extending inside the secondary flow path and of cold lobes () extending inside the primary flow path. The lobes of the nozzle () which are concave (), which is to say radially towards the outside, and respectively which are convex (), which is to say radically towards the inside, if the longitudinal offset is disregarded, radically face the respectively hot lobes () and cold lobes () of the mixer ().

Growing community of inventors

Moissy-Cramayel, France

Norman Bruno André Jodet