Safran Aircraft Engines

Eva Julie Lebeault

Anthony Binder

Laurent Soulat

Daniel-Ciprian Mincu

Nicolas Joseph Sirvin

Frédéric Dautreppe

Benjamin Bulot

Norman Bruno André Jodet

Mathieu Simon Paul Gruber

Fernando Gea Aguilera

Philippe Gérard Chanez

Josselin David Florian Regnard

Mathieu Patrick Jean-Louis Lallia

Pradeep Cojande

Guillaume Claude Robert Belmon

Jean-Loïc Hervé Lecordix

Jagoda Alina Worotynska

Adrien Clément Marcel Dubois

An aircraft dual-flow turbine engine assembly includes: an internal shroud for externally delimiting a primary flow path of the turbine engine gases; an external shroud for internally delimiting a secondary flow path of the turbine engine gas; and at least one air discharge duct extending between the internal shroud and the external shroud, the air discharge duct opening into the secondary flow path through an outlet orifice equipped with discharge fins. At least some of the discharge fins are movably mounted so as to be able to be incidence-control between a propulsion position, and a reverse thrust position.

Aircraft dual-flow turbine engine assembly, the assembly being equipped with incidence-controlled air discharge fins

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

The invention relates to a bypass turbomachine () for an aircraft, comprising a gas generator () and a ducted fan () comprising variable pitch blades () configured to take a reverse thrust position driving a reverse flow () of air within a secondary duct (), the gas generator () being connected to a fan casing () by a stator blade assembly () that passes through the secondary duct, first openings () for letting in air from the reverse flow being located on an outer casing () at least partially internally delimiting the outer duct, and second openings () for letting said air out being located on an inner casing () at least partially externally delimiting an inner duct (). The first openings are located within a plane that is perpendicular to a longitudinal axis (C) of the turbomachine and passes substantially through the middle of the blade assembly.

Bypass turbomachine for an aircraft

An aircraft including at least one unducted turbine engine for the propulsion of the aircraft. The turbine engine comprising: a rotor and a stator comprising a plurality of stator blades extending radially with respect to the longitudinal axis, each stator blade being defined, in a plane transverse to the longitudinal axis, by an angular position; and at least one aerodynamic obstruction positioned close to the turbine engine. The stator of the turbine engine comprises stator blades having a first chord, referred to as conventional blades, and at least one stator blade having a second chord larger than the first chord, referred to as the elongate blade, said at least one elongate blade being positioned in an interference angular range defined opposite the aerodynamic obstacle, so as to increase the straightening of the airflow from the rotor in the interference angular range.

Unducted turbine engine comprising stator blades having different chords

A process for using an air input of a turboreactor nacelle of an aircraft, comprising an air input lip which comprises at least one fixed portion and at least one portion which can be moved between a first position, in which the air input lip has an aerodynamic profile so as to guide the internal air flow over the internal wall in order to promote a thrust phase, and a second position, in which the portion is displaced in relation to the fixed portion so that the air input lip has a second radial thickness in the second position which is less than the first radial thickness in the first position so as to promote a reverse thrust phase.

Process for using an air input of a turboreactor nacelle comprising an air input lip which comprises a portion which can be moved to promote a thrust inversion phase

An air outlet for a nacelle for an aircraft turbofan having an inner wall and an outer wall connected to each other by a trailing edge, at least one radial through aperture extending over an angular portion of the air outlet and having inner and outer open faces and, for each aperture, a guiding device having inner and outer movable members movably mounted between a closed position wherein the inner and outer movable members respectively close off the inner and outer open faces, the aperture defining a closed cavity and an open position wherein the inner and outer movable members are configured to allow an external air flow to circulate in the aperture to support a reverse thrust phase.

Air outlet for a nacelle for an aircraft bypass turbojet engine comprising a guiding device to favor a reverse thrust phase

An aircraft turbojet engine extending along an X axis and comprising a blower configured to provide a reverse thrust and a nacelle comprising an air intake which comprises at least one deflection member movably mounted between a deployed position in which the deflection member projects from the inner wall or from the lip of the air intake in a radially inward direction of deployment facing the X axis or in a longitudinal direction of deployment with respect to the X axis, in order to allow a release of the reverse air flow from the inner wall to support the reverse thrust phase, and a retracted position in which the air intake has an aerodynamic profile so as to guide the internal air flow along the inner wall in order to support the thrust phase.

Turbojet engine comprising a nacelle with an air intake to promote a reversed thrust phase

The present invention concerns a turbomachine hub, intended to be mounted so as to be able to rotate about a longitudinal axis (X) of the turbomachine, the hub comprising a main body () arranged around the longitudinal axis (X), an outer surface of which has a plurality of recesses; a plurality of platforms () comprising an outer surface delimited by a circular outer edge of radius r, each platform () being arranged in a corresponding recess of the plurality of recesses of the main body (), and intended to receive a variable pitch blade (), the pitch of which is variable according to a pitch change axis (Z), the platform () being centred and able to rotate about the pitch change axis (Z); characterised in that, for each platform, at least one part of the outer surface of the platform () and the main body () of the hub comprises a curvature defined by a same sphere portion of radius R and centre C, the at least one part being situated at the outer edge of the platform (), in a junction zone between the platform () and the main body of the hub (), the centre C of the sphere being situated on the pitch change axis (Z) outside a hemispherical zone delimited between the longitudinal axis (X) and the outer surface of the platform ().

Turbomachine polyspherical hub for variable pitch blades

A turbomachine compressor module comprising an annular array of struts provided with pivotable flaps. The struts can define inter-strut spaces between two circumferentially adjacent struts, and variably oriented stator vanes are disposed at least partially in the inter-strut spaces. Also, a turbomachine having such a module and a row of rotor blades directly downstream of the module.

Compressor module for turbomachine

A discharge duct () of an intermediate housing hub for an aircraft turbojet engine, comprises an inlet end () and an outlet end (), intended to ensure the passage of air from at least one discharge inlet opening to at least one secondary outlet opening, and comprising an ejection grill () arranged at the outlet end (), the ejection grill () comprising a plurality of fins (), characterised in that the fins () comprise flow channels () for a fluid to be cooled, so as to form a heat exchange system.

Discharge duct of an intermediate housing hub for an aircraft turbojet engine comprising cooling channels

A stator wheel of a turbomachine configured to be mounted in an aircraft, the turbomachine extending along an axis, the turbomachine comprising a core flow stream for a first air flow and a bypass flow stream for a second air flow, wherein the stator wheel is configured to be mounted along the axis, the stator wheel comprising a plurality of vanes extending radially from the axis, each vane comprising a leading edge and a trailing edge that together define a chord in a plane of revolution defined relative to the axis, at least two vanes having different chords at a same radial distance, the difference in chords between two adjacent vanes being less than or equal to 25%.

Stator wheel of a turbomachine comprising vanes having different chords

An assembly for a primary flow path of a double flow turbomachine extending along a longitudinal axis including a reduction gear, configured to drive a fan in rotation, a stator including a plurality of vanes for straightening the primary flow path, and an intermediate casing, positioned downstream of the stator, including structural arms with fairings passing through the primary flow path, wherein a fairing and one straightening vane are connected together within the primary flow path by a connecting surface, the fairing and the vane being positioned longitudinally substantially at the reduction gear.

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Moissy-Cramayel, France

Eva Julie Lebeault