Safran Aircraft Engines

Snecma

Safran Nacelles

Kevin Morgane Lemarchand

Nathalie Nowakowski

Gilles Alain Marie Charier

Gilles Alain Charier

Tewfik Boudebiza

Laurent Soulat

Michel Gilbert Roland Brault

Guillaume Patrice Kubiak

Augustin Marc Michel Curlier

Pierre-Alain Jean Philippe Reigner

Nils Bordoni

Guillaume Olivier Vartan Martin

Adrien Pierre Jean Pertat

Bruno Albert Beutin

Olivier Kerbler

Norman Bruno André Jodet

Augustin Curlier

Nicolas Joseph Sirvin

Nicolas Jérôme Jean Tantot

Philippe Gerard Chanez

Nicolas Alain Bader

Carmen Gina Ancuta

Christian Sylvain Vessot

Guillaume Martin

Dominique Gerhardt Mayhew

Matthieu Pierre Michel Dubosc

Nicolas Maurice Herve Aussedat

Michel Gilbert Brault

Mélody Seriset

Melody Seriset

Ludovic Michael Laurent Toupet

Lionel Jean Leon Lefranc

Mathieu Lallia

Thomas Alain Vincent

A fan module with variable pitch blades for a longitudinal axis propulsion assembly. The module includes a rotor through which a rotor shaft runs along the longitudinal axis X and carries the blades of the fan. A system for changing the pitch of the blades includes a mechanism connected to the blades of the fan and a controller acting on the mechanism. The controller has a fixed body and a movable body translatable along the longitudinal axis X relative to the fixed body, and a load transfer bearing arranged between the mechanism and the controller. The fan module includes a power shaft driving the rotor shaft via an epicyclic train speed reducer, the speed reducer including a sun gear connected to the power shaft and a satellite carrier including a fixed annular ferrule secured to a fixed casing. The fixed body of the controller is mounted on the annular ferrule.

Fan module with variable pitch blades

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

A turbomachine comprising a ducted fan, a low-pressure turbine shaft and a reduction gearbox housed in a casing between the fan and the low-pressure turbine shaft, the fan rotor supplying airflow to a primary stream and a secondary stream and comprising a hub of diameter D, wherein&#x2014;the diameter Dof the fan rotor is greater than 82 inches (2.08 metres), &#x2014;the pressure ratio of the fan is between 1.10 and 1.35, the turbomachine comprises a low-pressure compressor separate from the fan, the reduction gearbox being interposed between the fan rotor and a turbine shaft of the low-pressure compressor, and wherein the reduction gearbox casing has an outside diameter Dgreater than the diameter Dof the hub, the pitch diameter Dof the reduction gearbox ring being between 0.15 and 0.35 times the fan rotor diameter.

Turbomachine with fan rotor and reduction gearbox driving a low-pressure decompressor shaft

The present invention relates to a powerplant of an aircraft including at least one twin-spool gas generator of longitudinal axis (XX), at least one main fan arranged upstream of the gas generator on the longitudinal axis (XX) and driven in rotation by the gas generator, the main fan being shrouded with a main fan housing; and at least one auxiliary fan with axis (XY, XY') offset relative to the longitudinal axis (XX) and driven by the gas generator, the auxiliary fan being shrouded with an auxiliary fan housing, the gas generator including at least one low-pressure compressor and a low-pressure turbine connected by a low-pressure shaft, the low-pressure turbine driving in rotation the main fan and the auxiliary fan via at least one power transmission system including a differential gear system incorporating a bevel gear.

Aircraft propulsion assembly equipped with a main fan and with a least one offset fan

An aircraft including a fuselage, a lateral support wing () and a propulsion assembly () mounted under the wing. The wing includes at least two structural spars () extending from the fuselage toward the tip of the wing, one of these () being upstream and the other () downstream. The propulsion assembly includes a gas generator () and at least two offset fans () arranged on either side of the axis of the gas generator. The offset fans () are attached directly to one of the spars () and the gas generator () is attached directly to the two spars. The leading edge of the wing forms a given sweep of angle (&#x3b1;) with the axis of the fuselage. The two offset fans () are axially offset from one another.

Aircraft with multiple fan propulsion assembly fixed under the wing

Aircraft turbine engine comprising a low-pressure spool that comprises a low-pressure shaft (), means () for taking off power from said low-pressure shaft, and a fan () that is driven by said low-pressure shaft by means of a reduction gear (), said reduction gear comprising at least one first element () that is connected to said low-pressure shaft for conjoint rotation, at least one second element () that is connected to said fan for conjoint rotation, and at least one third element () that is connected to a stator casing of the turbine engine, characterised in that said at least one third element is connected to said stator casing by disengageable connection means (), and comprising at least one member that can move from a first position in which said at least one third element is fixedly connected to said stator casing into a second position in which said at least one third element is separated from said stator casing and is free to rotate about said longitudinal axis.

Aircraft turbine engine with planetary or epicyclic gear train

An aircraft propulsion assembly comprising at least a main turbine mounted along a longitudinal axis, at least one main fan arranged upstream of the main turbine along the longitudinal axis and driven in rotation by the said main turbine, the said main fan being ducted by a main fan casing, an auxiliary turbine mounted along the longitudinal axis, the auxiliary turbine being independent of the main turbine, an auxiliary fan of axis offset with respect to the longitudinal axis and driven by the auxiliary turbine, the auxiliary fan being ducted by an auxiliary fan casing, the main casing being separate and distinct from the auxiliary casing so as respectively to generate a main secondary flow and an auxiliary secondary flow which remain independent of one another until they are discharged into the atmosphere.

Aircraft propulsion assembly equipped with a main fan and with at least one offset fan

Fan module with variable-pitch blades for a propulsion system that includes:

An aircraft propulsion assembly including a bypass turbojet having a stationary inter-compressor casing positioned upstream from a space between passages, a nacelle including at its downstream end an inner wall that defines the outside of the space between passages and the inside of the bypass stream flow passage, and an outer wall arranged around the inner wall and that defines the outside of the bypass stream flow passage, at least a portion of the inner wall being suitable for moving between a maintenance position and a working position, and a thrust-reverser including a downstream element of the outer wall that is movable in translation between a retracted position and a thrust-reversal position in which it allows bypass stream deflector elements to be deployed in a radial direction, the deployment of the stream deflector elements being controlled by guide rods.

Aircraft propulsion unit having thrust reverser flaps connected to an inter-compressor casing located between the engine compressors

An aircraft includes a wing having a first upstream spar and a second downstream spar extending in the direction of the span of said wing, and at least one propulsion unit supported by the wing. The propulsion unit includes a turboprop engine and a propeller. The propeller includes an external annular casing fixed to a suction surface of the wing, and at least to the first upstream spar via at least one first and second fastener.

Aircraft with a propulsion unit with offset fan

A front part of an aircraft turbomachine of the bypass type including a single fan, a gas generator positioned downstream of the fan, a reducer interposed between the gas generator and the fan, a flow splitter and a structure with structural outlet guide vanes. The vanes each have a root arranged upstream of the flow splitter and the reducer is also arranged at least half upstream of the flow splitter.

Aircraft turbomachine front part

The invention relates to an aviation turbine engine fan module including a de-icing system () for de-icing an inlet cone () and comprising a sheath () placed inside an inside space defined upstream by the inlet cone, said sheath comprising a first duct () having at least one hot air admission orifice (), said first duct being configured to convey hot air from a bearing enclosure () of the engine towards a wall of the inlet cone in order to heat it from the inside, the sheath further comprising a second duct () having at least one outlet situated downstream from the admission orifice of the first duct, said second duct being configured to discharge air from the first duct towards the downstream end of the engine. The invention also provides a method of de-icing a turbine engine inlet cone.

Turbine engine fan module including a turbine engine inlet cone de-icing system, and a de-icing method

A variable-orientation stator vane of a turbine engine, comprising a wall that extends between a leading edge and a trailing edge forming a chord line. The chord line increases from a root end towards a tip end of said vane. This vane is fitted pivotably in a turbine engine casing so as to divert a flow of cold air circulating in an annular duct thereof.

Turbine engine stator vane, fan casing comprising such a vane, thrust reverser system of a turbine engine equipped with such a vane and turbine engine equipped with said vane, said casing or said system

Disclosed herein is a fan module with variable pitch blades for a propulsion unit, the fan module comprising a rotor carrying the blades and comprising an inner shaft and an outer casing defining between them a space, a control device for controlling the pitch of the blades, and a feathering device for feathering the blades. The rotation of the rotor is guided by a bearing, and the module further comprises means for recovering and guiding a liquid lubricant of the bearing, with the recovery and guidance means being configured to recover and guide said lubricant from an axial upstream end of the casing, axially from upstream to downstream and radially from the inside to the outside, under the centrifugal effect.

A modular assembly for a turbine engine has an annular bearing support that is configured to be connected to a first anti-friction bearing, which is configured to be fitted around a first shaft of the turbine engine. The modular assembly further includes a second anti-friction bearing configured to be fitted around a second shaft of the turbine engine, wherein the second shaft is not parallel to said first shaft.

Growing community of inventors

Moissy-Cramayel, France

Kevin Morgane Lemarchand