Safran Aircraft Engines

Snecma

Jonathan Evert Vlastuin

Mathieu Simon Paul Gruber

Gilles Alain Marie Charier

Gilles Alain Charier

Clement Marcel Maurice Dejeu

Nathalie Nowakowski

Nicolas Jérôme Jean Tantot

Adrien Louis Nicolas Laurenceau

Cyril Verbrugge

Henri Yesilcimen

Caroline Marie Frantz

Adrien Jacques Philippe Fabre

Vivien Mickaël Courtier

Sébastien Emile Philippe Tajan

Adrien Philippe Fabre

Adrien Laurenceau

Laurence Francine Vion

Damien Bernard Emeric Guegan

Pierre-Hugues Ambroise Maxime Victor Retiveau

Kévin Lemarchand

Dominique Gerhadt Mayhew

Prasaad Cojande

Anthony Pascal Eloi Louet

Anthony Louet

Clement Dejeu

Kévin Morgane Lemarchand

Caroline Marie Franz

Clement Marcel Dejeu

Nicolas Jerôme Jean Tantot

Clément Dejeu

Clément Marcel Maurice Dejeu

A turbofan includes a fan, a casing positioned downstream of the fan and separating a primary flowpath from a secondary flowpath, a compressor, a combustion chamber and a turbine being arranged in the primary flowpath, the turbofan having a differential transmission coupled to the turbine, and a power supply device configured to provide additional power to the one provided by the turbine to drive the compressor.

Turbofan comprising a power supply device to drive the compressor

A variable pitch bladed disc including a plurality of blades, each being of variable pitch about a blade axis of rotation and having a root, the plurality of blades including at least one first blade and at least one second blade, a plurality of rotor connecting shafts, each shaft having a root and a tip, the root of each blade being mounted on the tip of a corresponding rotor connecting shaft via a pivot so as to allow each blade to be rotated about the blade axis of rotation, the first blade having a first rotation axis inclination such that the rotation axis thereof is inclined in a fixed manner with respect to a radial axis passing through the root of the corresponding shaft, and the second blade has a second rotation axis inclination different from the first rotation axis inclination.

Variable pitch bladed disc

A variable-pitch bladed disc including a plurality of blades, each at a variable pitch in relation to a rotation axis of the blade and each having a root, the plurality of blades including at least one first blade and at least one second blade, a plurality of rotor connecting shafts, each having a root and a tip, each root being mounted at the tip of a corresponding rotor connecting shaft by way of a pivot so as to allow each blade to rotate about the blade rotation axis, the first blade having a first blade inclination, such that the first blade is inclined in a fixed manner with respect to the blade rotation axis of the first blade, and the second blade having a second blade inclination different from the first blade inclination.

A method of altering the twisting relationship for the aerodynamic surface of a fan blade of a gas turbine engine, wherein the following steps are performed: establishing, for a portion of the aerodynamic surface of the fan blade, an alteration relationship defined by variation of a pitch angle of the blade as a function of radial height along the blade, the alteration relationship including alterations that are each defined by a height along with the radial height of the fan blade and by an amplitude; and applying the alteration relationship as established in this way to an initial twisting relationship of the fan blade so as to obtain an altered twisting relationship for the fan blade, the initial twisting relationship being defined by a polynomial for the radial height of the fan blade as a function of its pitch angle.

Method for altering the law of twist of the aerodynamic surface of a gas turbine engine fan blade

Turbojet engine includes a fan, a compressor, a combustion chamber, and a turbine configured to rotatably drive the fan via a first reduction gear and the compressor via a second reduction gear, in which an outlet of the first reduction gear is rotatably coupled to an inlet of the second reduction gear by a reduction shaft, the reduction shaft being supported by a bearing arranged between the first reduction gear and the second reduction gear.

Turbojet with step-down mechanism

A vane configured to be placed with a plurality of identical vanes so as to form a vane wheel for an aeroengine, the vane wheel defining an axis, the vane having an airfoil presenting a leading edge and a trailing edge, the leading-edge curve describing the shape of the leading edge of the airfoil in a view perpendicular to the airfoil presenting at least one leading-edge undulation, said at least one leading-edge undulation extending over less than 30% of a length of the airfoil from the first end of the airfoil.

Leading edge profile of vanes

The invention relates to an unducted-fan aircraft engine comprising a generally cylindrical nacelle through which a primary jet flows, the nacelle bearing a fan rotor comprising variable-pitch vanes () located radially outside the nacelle in order to be traversed by a secondary jet () flowing longitudinally around the nacelle. The rotor comprises a hub bearing variable-pitch vane supports each carrying one vane (), each vane () comprising a blade extending from a root that is used to removably attach same to a base of the associated support (). Each base () is located radially outside the nacelle so that it is in the secondary jet, and each assembly formed by a base () and the vane root supported by the base () is surrounded and covered by a detachable aerodynamic cover ().

Unducted-fan aircraft engine including a propeller comprising vanes with roots outside the nacelle and covered by detachable covers

A method for modeling at least a part of a blade of a non-streamlined propeller, part of the blade having an offset. The method includes (a) Parameterization of at least one Bezier curve representing a deformation of the blade characterizing the offset, defined by: a. First and second end control points (PCU1, PCUK); b. At least one intermediate control point (PCUi, i&#x2208;[[2,K&#x2212;1]]) disposed between the end points (PCU1, PCUK). The parameterization being performed according to at least one deformation parameter and the cutting height in the blade, on the basis of which the abscissa of the intermediate control point (PCUi) and the ordinate of the second end point (PCUK) are expressed. Optimized values of the deformation parameter or parameters are determined and then output.

Method for modeling a non-streamlined propeller blade

A variable-pitch vane including a plurality of propeller blades (). Each blade has a variable pitch according to a blade rotational axis (A) and a root (). A plurality of rotor connecting shafts (), each shaft having a foot () and a head (). The root () of each blade is mounted on the head () of a rotor connecting shaft via a pivot () in such a way as to allow each blade () to rotate according to the blade rotational axis (A), in which each blade () has a blade pitch, such that the blade rotational axis (A) is inclined relative to a radial axis (A) passing through the foot () of the corresponding shaft.

Variable-pitch vane

A method for modeling an offset portion of a blade of a non-ducted propeller is provided. The method includes: parametrizing a C1-class curve representing a deformation of the blade characterizing the offset, according to a position along a section at a given height in the blade, the curve intersecting consecutively through a first bend control point, a central control point, and a second bend control point, the first and second bend control points defining the extent of the blade section, the parametrization being implemented according to a first deformation parameter defining the abscissa of the central control point, a second parameter of deformation defining the ordinate of the second bend point, and a third deformation parameter defining the angle of the tangent to the curve at the second bend control point; optimizing one of the deformation parameters; and plotting the values of the optimized parameters on an interface.

Method for modeling a blade of a non-ducted propeller

A propeller blade extending radially between a blade root and a blade tip and having a leading edge and a trailing edge, together with a pressure side face and a suction side face. Along the leading edge, the angle (&#x3b1;) between the radial direction and a tangent to the leading edge passes through a local maximum zsituated in the range 60% to 80% of the height of the blade. Furthermore, the suction side face of the blade includes an indentation extending axially from the leading edge to the trailing edge and located at a height of the blade between the local maximum zand the blade tip.

Propeller blade for a turbomachine

An aircraft turbopropeller includes two coaxial contra-rotating unducted propellers, upstream and downstream respectively, each propeller including an annular row of blades. The blades of the downstream propeller have a reverse sweep in relation to that of the blades of the upstream propeller. The leading edges of the blades of the downstream propeller extend radially outwards from downstream to upstream, at least over a radially outer portion of the blades.

Aircraft turbopropeller

The invention relates to an aircraft turboprop engine having two unducted propellers, coaxial about a rotation axis (LL), respectively upstream () and downstream (), each propeller comprising an annular row of blades, the blades of the downstream propeller () each having a truncated head and each comprising a leading edge () and a trailing edge (), the radially external ends (AB) of which are distant from each other and are connected by a terminal edge (), characterised in that the terminal edge () of each blade of the downstream propeller () has, at at least two points, tangents in a meridian plane having different slopes, the upstream end (A) of the terminal edge defining a maximum radial dimension (R) of the blade and its downstream end (B) defining a minimum radial dimension (R) of the terminal edge.

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Moissy-Cramayel, France

Jonathan Evert Vlastuin