Safran Aircraft Engines

Paul Antoine Foresto

Adrien Jacques Philippe Fabre

Vivien Mickaël Courtier

Stéphane Roger Mahias

Christophe Paul Jacquemard

Régis Eugène Henri Servant

Christophe Marcel Lucien Perdrigeon

Guillaume Pascal Jean-Charles Gondre

Anthony Binder

William Henri Joseph Riera

Sébastien Emile Philippe Tajan

Norman Bruno Andre Jodet

Adrien Philippe Fabre

Cedric Zaccardi

Suzanne Madeleine Coustillas

Fabrice Michel Francois Rene Aubert

Stéphane Ibarzo

Didier Pierre Mozer

Alexandre Gerard Francois Couilleaux

Fabrice Michel François René Aubert

An intermediate casing includes an inner wall having an outer surface extending along a longitudinal axis, the outer surface having a first profile in a first plane, an outer wall having an inner surface a second profile in the first plane, and first through fourth arms each extending radially. The outer and inner surfaces and the first and second arms delimit a first cavity, the first cavity having a first area and the outer and inner surfaces being separated by a first distance. The outer and inner surfaces and the third and fourth arms delimit a second cavity, the second cavity having a second area and the outer and inner surfaces being separated by a second distance. The first and second profiles are such that the first area is substantially identical to the second area and the first distance is different from the second distance.

Intermediate casing for a gas turbine engine

The invention relates to a turbomachine assembly () comprising a fan () and a booster drum-type part (), the fan () comprising:&#x2014;blades () comprising an airfoil () and an extension () mounted on and attached to the trailing edge () of the airfoil (),&#x2014;a fan () disc () and&#x2014;a series of inter-blade platforms (), the extension () of each blade () extending beyond the downstream face () of the fan () disc () in the direction of the upstream edge () of the part () and covering, at least partially, the cavity () between the fan () and the part ().

Turbomachine assembly comprising fan blades with an extended trailing edge

The invention relates to a module of an aircraft bypass engine including an annular flow path of a secondary flow of gas wherein is mounted an annular row of stator blades configured to straighten the gas flow, an arm extending in the flow path downstream from said annular row of blades and being connected physically to one of the stator blades so as to form a unitary aerodynamic assembly, wherein at least one of the two blades adjacent to the blade of the unitary aerodynamic assembly is of a variable pitch type and capable of being rotationally moved about a substantially radial axis with respect to a longitudinal axis of the engine.

Module of an aircraft bypass engine of which one arm integrates a stator blade

A turbine engine fan having an axis of revolution wherein a fan disk having an upstream face and a radial face is configured to receive a series of fan blades. An inter-blade platform wherein an upstream end, a blocking ferrule, are applied and attached on the one hand to the upstream end of the platform and on the other hand to the upstream face of the fan disk, and an inlet cone. The blocking ferrule includes a radial ring and a lug, extending axially from the radial ring and configured to form a radial abutment for the inlet cone.

Fan comprising an inter-blade platform attached upstream by a ferrule

The invention relates to a fan blink () for an aircraft turbomachine, comprising a hub, an annular platform () and fan blades () arranged projecting from the annual platform. It also comprises a mechanical discharge slit () from a trailing edge () of the fan blade, associated with at least one of the fan blades (), for the case of ingestion of a bird, the slit being made on the annular platform () going around the trailing edge ().

Fan for aircraft turbomachine

A guide vane located in a fan air flow in an aircraft twin-spool turbomachine, the vane being made with an extrados body and an intrados body between which there is a thermal conduction matrix. Furthermore, the attachment devices between the two spools are arranged outside the aerodynamic part of the vane.

Outlet guide vane for turbomachine, made from several parts assembled together by attachment means outside the flow stream

A hub for a propeller having variable-pitch blades for a turbomachine with longitudinal rotation axis, including a plurality of blade platforms, each one being designed to receive a blade root, the platforms being distributed around the annular outer periphery of the hub concentrically with the longitudinal rotation axis, wherein each blade platform has the general geometric shape of a spherical cap, and wherein, between two adjacent blade platforms around the longitudinal rotation axis, the transverse dimension, in particular the diameter, of the hub can vary axially and radially.

Hub for propeller having variable-pitch blades, with radial and axial dimensioning variation

A system is configured to change the pitch of blades of at least one turbine engine propeller provided with multiple blades. The system includes a control means acting on a connecting mechanism connected to the blades of the propeller. The control means includes a stationary body and a movable body translatable along a longitudinal axis (X) relative to the stationary body and defining two chambers. A fluid supply means is configured to be connected to a fluid supply source and to supply the chambers. The fluid supply means includes at least one channel. The channel is arranged and extends along an axis (Q) parallel to the longitudinal axis in an overthickness of a wall of the stationary body, the overthickness extending radially inside relative to the longitudinal axis.

Pitch change system equipped with means for supplying fluid to a control means and corresponding turbine engine

A turboprop including a propeller including a blade extending in a direction, which also includes a root, a leading edge, a trailing edge, and a wing tip, and an inner air stream channel, wherein the inner air stream channel includes an inlet located at the root of the blade and an outlet leading to the trailing edge of the blade transversely directed in relation to the main elongation direction, such that an inner stream of air flowing in the inner air stream channel by entering via the inlet adjacent to the root of the blade is discharged via the outlet adjacent to the trailing edge of the blade by forming a stream of blown air that moves away from the trailing edge in a direction which is transverse to the main elongation direction and which has a component in the direction of a skeleton line of the blade at the trailing edge.

Turboprop

The invention relates to a method for manufacturing a blade made of composite material extending in a main elongation direction, the composite material of said blade comprising a fiber reinforcement densified by a matrix, the method comprising the following steps:

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Moissy-Cramayel, France

Paul Antoine Foresto