Aircelle Limited

Safran Nacelles

Safran Ceramics

Aircella

Jean-Philippe Joret

Guy Bernard Vauchel

Xavier Cazuc

Pierre Caruel

Vincent Devanlay

Hazem Kioua

Vincent Jean-François Peyron

Sébastien Laurent Marie Pascal

Peter Segat

Jérôme Corfa

Xavier Bouteiller

Fabien Bravin

Thomas Vandellos

Emmanuel Lesteven

Charles-Antoine Lampaert

Eric Conete

Patrick Boileau

Laurent Georges Valleroy

Vincent Peyron

Pierre Andre Baudu

Denis Guillois

Patrick André Boileau

Guillaume Lefort

Benoit Carrere

Stephane Beilliard

Gina Ferrier

Clément Marie Benoît Roussille

François Bellet

Serge Bunel

Andre Baillard

Thomas Marlay

Jerome Collier

Damien Lemoine

Matthieu Menielle

Georges Alain Bouret

Jérémie Rabineau

Romain Gardel

Pascal Soulier

Vincent Dutot

Pierre-Alain Reboul

Pierre-Alain Chouard

Philippe Avenel

Alexis Yves-Marie Loncle

Boualem Merabet

Anne-Laure Gibouin

Regis Fassier

André Baillard

Loïc Le Boulicaut

Christophe Guenadou

Vincent Taddei

Hélène Malot

Franck Laurant

Marine Ollivier

Francois Bellet

Guillaume Levieux

Pierre Baudu

Stéphane Riquier

A thrust reverser of a nacelle of a turbojet engine of an aircraft extending about a longitudinal axis and including a fixed structure and at least one door which is rotatable between a closed position and an open position and including a downstream frame spaced axially from a downstream end edge of the door by an axial clearance parallel to the longitudinal axis. The thrust reverser includes a system for diverting fluids to the exterior including at least one projection fixed onto the door or the other fixed structure and extending into the axial clearance and towards one of the fixed structure or the door, while leaving an axial gap between a free end of said projection and one of the fixed structure or the door.

Thrust reverser with door of a nacelle of a turbojet engine of an aircraft with a system for diverting fluids to the exterior

An assembly for a turbomachine with a longitudinal axis having an exhaust cone with an outer annular wall for the flow of a primary airflow and an annular box arranged radially inside the outer annular wall, an exhaust casing arranged upstream of and connected to the exhaust cone. One end of the outer annular wall or one end of the annular box may be free to move relative to and have no mechanical connection with the exhaust cone or the exhaust casing.

Fastening of an exhaust cone in a turbomachine turbine

A nacelle for an aircraft propulsion assembly includes an actuator, a secondary door subject to the actuator and contributing to defining an external aerodynamic surface of the nacelle in a first position, a main door contributing to defining the external aerodynamic surface and including a recess that the actuator passes through in a deployed position but not in a retracted position, and a flap connected to the main door. The flap is movable between a closed position wherein the flap closes off at least a first region of the recess through which the actuator passes in the deployed position whereby the flap contributes to defining the external aerodynamic surface, and an open position wherein the flap is moved away from the recess so as to clear the passage for the actuator through the first region of the recess.

Nacelle for an aircraft propulsion assembly comprising a flap for closing a recess for the passage of an actuator

The invention relates to an assembly () which comprises a front frame () of a thrust reverser structure of an aircraft nacelle and at least one cascade (). The assembly is characterized in that, when the cascade is fixed to the frame, the axial distance (L) that separates the first blading () upstream of the cascade from a covering area () between the extension of the cascade and the front frame is greater than or equal to once the mean height (h) of the cascade.

Front frame and deflection grid of a thrust reverser of an aircraft nacelle

The invention relates to an assembly for a turbine of a turbomachine with a longitudinal axis (X) comprising:

Fastening of an exhaust cone in a turbomachine nozzle

A thrust reverser for an aircraft bypass turbojet engine nacelle has a generally annular shape around an axis and includes an annular frame for securing deflection grids. The frame has a first frusto-conical wall widening in the downstream direction and including an upstream peripheral edge configured to be attached to a casing of the turbojet engine. A downstream peripheral edge of the wall extends in the continuation of the wall and secures the upstream ends of the grids. A second annular wall extends radially outwards from an outer frusto-conical face of the first wall. The first and second walls are integrally formed and the second wall has axial openings through which actuators pass.

Thrust reverser for an aircraft bypass turbojet engine nacelle

Disclosed is an air intake of a turbine engine nacelle extending along an axis in which a flow of air circulates from upstream to downstream, the air intake extending annularly about the axis, the air intake having a movable upstream portion and a stationary downstream portion, at least one controllable movement member for moving the movable upstream portion between a retracted position and an extended position in order to enable circulation of an air flow passing between the movable upstream portion and the stationary downstream portion, at least one de-icing circuit having at least one circulation channel fluidically connecting the movable upstream portion and a hot air source, the circulation channel having at least one non-telescopic upstream pipe and at least one telescopic downstream member.

Air inlet of a turbomachine nacelle including a circulation conduit of a hot air flow between a movable upstream part and a fixed downstream part

A supersonic air intake of an aircraft propulsion assembly comprises: •⋅ pipe with a longitudinal axis configured to guide the flow of air into the turbine engine, •⋅ a speed reducer fixedly mounted inside the pipe along the longitudinal axis, •⋅ at least one support arm linking the speed reducer to the pipe and having an upstream outer wall and a downstream outer wall together defining a cavity, and •⋅ at least one de-icing device having an inner wall mounted in the cavity opposite the upstream outer wall in order to together define a calibrated de-icing space, as well as a member for supplying and a member for discharging a de-icing air flow in the de-icing space.

Supersonic air inlet of an aircraft propelling assembly comprising a support arm de-icing device and de-icing method

Disclosed is a supersonic speed attenuator for an air inlet of an aircraft power plant, having a conical external wall with a tapered end upstream and an internal dividing wall that delimits, with the conical external wall, a cavity; the supersonic speed attenuator further having a de-icing device having: an internal wall mounted in the cavity opposite the conical external wall so that, together, the two walls delimit a calibrated de-icing volume; at least one member for supplying a de-icing air flow, opening into the de-icing volume; and at least one member for discharging the de-icing air flow from the de-icing space.

Supersonic speed attenuator of an air intake of an aircraft propulsion assembly comprising a de-icing device and de-icing method

The invention relates to an assembly for a turbomachine, comprising an exhaust cone () and an exhaust case having an annular inner shell, the exhaust cone () and the exhaust case extending about an axis (X), the exhaust cone () having a radially outer skin () extending in the projection of said inner shell and a radially inner skin () defining a chamber () with the outer skin (), the upstream end () of the cone () being fastened to the inner shell () via connecting lugs (), the assembly comprising an annular fairing () located radially outside said lugs () and extending axially between the inner shell () of the exhaust case () and the outer skin () of the exhaust cone (), said fairing () extending in the projection of said inner shell () and of the outer skin (), the assembly being characterized in that the fairing () is segmented and consists of multiple angular segments () arranged circumferentially end to end.

Exhaust cone for a turbomachine

An assembly for a turbomachine nozzle having a longitudinal axis (X) includes an exhaust cone with an annular outer wall for guiding a primary air flow and an annular chamber placed radially inside said annular outer wall. The annular chamber includes an annular inner wall placed radially inside the annular outer wall of the exhaust cone. An exhaust case is placed upstream of the ejection cone. A connecting member is interposed longitudinally between the exhaust case and the exhaust cone and fastened to the exhaust case and to the annular inner wall. The assembly further includes elastic means configured to deform radially and prestressed radially between the annular outer wall and the connecting member.

A nacelle for a propulsion assembly, the nacelle including a front internal structure having a framework formed of longitudinal beams and including removable, dismountable or retractable panels which are provided to facilitate the access for maintenance to the parts of the motor shrouded by these panels.

Nacelle for a propulsion assembly with a very high bypass ratio, comprising a removable and structural front internal structure

The present disclosure relates to a rear assembly of a turbojet engine nacelle that includes a thrust reverser fixed structure having at least one half-cowl mounted on a beam pivotally hinged by at least one hinge, a suspension mast on which the fixed structure is movably mounted, a hatch for access to the hinge of the half-cowl, and a fireproof sealing device which includes a sealing gasket adapted to provide sealing between a median section of the nacelle and the beam area. The sealing gasket includes a tubular gasket type lower portion which extends at least partially over a seal bearing edge of the half-cowl that is adapted to cooperate with a seal bearing wall of the mast and an upper portion including an elastically deformable sealing sheet extending from the seal bearing wall of the mast to the rear face of the hatch for access to the hinge.

Rear assembly of a turbojet engine nacelle including a fireproof sealing device

A propulsion unit for an aircraft is provided. The propulsion unit includes a stationary unit having at least one cowl pivotally mounted on an axis collinear with a longitudinal axis of the propulsion unit between an operating position or a maintenance position allowing access to the turbojet engine for regular maintenance operations and irregular maintenance operations. The maintenance position for irregular maintenance operations includes an opening angle of the cowl that is greater than the opening angle of the cowl to carry out regular maintenance operations. The propulsion unit further includes at least one cylinder and at least one safety connecting rod secured to the cowl and stationary unit.

Propulsion unit for an aircraft and method for opening a movable cowl of said propulsion unit

An air intake structure for a turbojet engine nacelle includes an inner panel to be attached to a fan casing, and an external panel capable of a translational movement in a longitudinal direction of the nacelle. The external panel incorporates a portion of an air inlet lip able to provide a junction between the inner panel and the external panel, and each of the air inlet lip portion and the inner panel is equipped, at least in the region of a joining end, with an acoustic attenuation structure. In particular, at least one of the joining ends is equipped with a radial buffer able to come into contact with a corresponding joining flange exhibited by the other joining end when the external panel is in a closed position.

Air inlet structure for a turbojet engine nacelle of laminar type

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Beuzeville, France

Jean-Philippe Joret