Safran Aircraft Engines

Laurent Louis Robert Baudoin

Josselin David Florian Regnard

Norman Bruno André Jodet

Caroline Marie Frantz

Georges Jean Xavier Riou

Nicolas Joseph Sirvin

Jéremy Paul Francisco Gonzalez

Norman Bruno Andre Jodet

Daniel-Ciprian Mincu

Frédéric Dautreppe

Jagoda Alina Worotynska

Jeremy Paul Gonzalez

Hélène Monique Orsi

Jean Michel Roux

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

Method for using an air intake of a turbojet engine nacelle comprising at least one elastically deformable portion, at least one connecting member mounted in an annular cavity integrally with the elastically deformable portion, and at least one controllable displacement member, in which method: during a thrust phase of the turbojet engine, the controllable displacement member moves the connecting member into a first position in which the elastically deformable portion of the air intake lip has an aerodynamic profile, and during a thrust reversal phase of the turbojet engine the controllable displacement member moves the connecting member into a second position in which the elastically deformable portion of the air intake lip has an irregular profile so as to allow a release of the reverse air flow from the elastically deformable portion.

Method for operating an air intake of a turbojet engine nacelle during a thrust phase and during a reverse thrust phase

Device for perforating an acoustic annular panel for a turbine engine, wherein it includes a perforation needle support ring, the ring having an axis A of revolution and being intended to be inserted substantially coaxially inside the panel, the ring being sectored and including ring sectors regularly distributed around the axis and each one radially mobile opposite the axis from a radially inner rest position up to a perforation position of the radially outer panel, and at least one rotating actuator of rotation axis A and mechanically connected to the ring sectors such that a rotation of the actuator leads to a radial movement of the ring sectors.

Device for perforating an acoustic annular panel for a turbine engine

A ground testing device for a turbomachine, including an aircraft engine () extending axially from an air inlet to a nozzle, a shroud () having an axis of revolution (A) and surrounding at least partially the engine () so as to create an annular air flow between the engine () and the shroud () when the engine () is in operation, and a framework () allowing the shroud () to be supported.

Ground testing device for a turbomachine

An aircraft propulsion unit including two distinct fan spools, and including a first duct and a second duct extending downstream of the two fan spools, the propulsion unit including a control ring disposed on one of the ducts, downstream of one of the fan spools, with at least one annular internal wall extending in an interior space of the duct, the control unit of the control ring being configured to modify the shape of the internal wall, and to cause an air passage cross section in the duct to vary at the ring, a device for acquiring acoustic signals generated by acoustic waves propagating downstream of the fan spools, and a device for processing the acoustic signals, configured to measure a dephasing between the acoustic waves and to control the control ring depending on the dephasing.

Active system generating destructive acoustic interference for an aircraft engine with multiple fan spools

The invention relates to an assembly for a turbine engine (), the assembly including: a casing () extending along a longitudinal axis (X-X') of the turbine engine (), a surface heat exchanger () arranged along an inner edge () of said casing () and including a plurality of cooling fins () extending radially inwards and longitudinally, and a central body () partially concentrically arranged within said casing (), said casing () and the central body () defining between one another an annular air channel (Va), the width (h) of which is substantially constant in the region of the heat exchanger (). The assembly is characterized in that the fins () are distributed over more than 180&#xb0; of the inner edge () of the casing () and in that the mean height ((h) of the fins () is more than 5% of the width (h) of the annular air channel (Va) in the region the heat exchanger ().

Surface heat exchanger and acoustic treatment

An assembly including a compressor for a turbine engine, and a discharge system including a supply duct, one end of which is connected to the compressor, the duct being configured to collect therefrom a flow of air compressed by the compressor; and a pressure-reduction device including an inlet and an outlet, the inlet being connected to the other end of the supply duct, wherein the pressure reduction device includes a casing forming a volume between the inlet and the outlet, a porous material occupying the volume, and a device for holding the porous material within the casing, wherein the outlet has an open section allowing the upper air to pass through the open section of the inlet.

System for discharging a compressor flow of a turbine engine

The invention relates to an aircraft turbofan type engine () comprising an air intake duct () having an inner face provided with an acoustic attenuation shroud () located upstream from a fan () of this engine (). This acoustic attenuation shroud () comprises a regular alternation of first portions and second portions with different first and second thicknesses along the radial direction, the number of alternations being greater than twice the number of passing azimuth acoustic vibration modes occurring in the engine () without the acoustic attenuation shroud () operating under conditions corresponding to a landing phase and a take-off phase.

Growing community of inventors

Moissy-Cramayel, France

Laurent Louis Robert Baudoin