Safran Helicopter Engines

Turbomeca

Laurent Pierre Tarnowski

Camel Serghine

Loïs Pierre Denis Vive

Thomas Drouin

Alphonse Puerto

Bernard Claude Pons

Youssef Bouchia

Mathieu Laurent Herran

Jérôme Yves Félix Gilbert Porodo

Hubert Hippolyte Vignau

Nicolas Bulot

Christophe Michel Georges Marcel Brillet

Pierre-Luc Regaud

Jérôme Porodo

Sébastien Combebias

Fabien Artus

Rainer Thomas

Ludovic André Joël Naudot

Nicolas Rochuon

Vincent Duqueine

Michael Derderian

Didier Claude Chartrain

Jerome Porodo

Jerome Yves Felix Gilbert Porodo

A method for manufacturing a compressor impeller or rotor including a hub that carries blades, involving a step of manufacturing a hub which includes all or some of the blades, and a step of additive manufacture by adding localised material using a method such as the LMD process to form or finish each blade.

Method for manufacturing a compressor impeller for a turbomachine

Aircraft turbomachine including a centrifugal compressor, a combustion chamber, the combustion chamber being supplied by the compressor via a diffuser and via a straightener, and a heat exchanger, the exchanger including a first circuit, supplied with exhaust gas from the turbomachine, and a second circuit, which are connected by volutes on the one hand to an outlet of the diffuser and on the other hand to an inlet of the straightener, the volutes having reversed winding directions such that their connection ports to the exchanger are independent of one another and are substantially diametrically opposed, and such that the minimum cross section of each duct is situated at a larger cross section of the other duct.

Recovered-cycle aircraft turbomachine

An aircraft turbomachine having a centrifugal compressor, an annular combustion chamber, an annular casing extending around the chamber and delimiting an annular space (E) in which the chamber is situated, and a heat exchanger. The heat exchanger can include a first circuit supplied with exhaust gas from the turbomachine, and a second circuit connected by first and second volutes respectively to an outlet of the compressor and to the annular space. The first and second volutes can be positioned at an axial distance from one another, and the second volute can be connected to the annular space by a straightener which is situated at least in part outside the casing and which is integrated into an annular connecting pipe which connects the second volute to this casing.

The invention relates to a rotor blade () for a turbomachine, in particular of an aircraft, comprising an airfoil () comprising a pressure face () and a suction face () extending from a leading edge () to a trailing edge (), the airfoil () comprising an axis of elongation extending substantially along the leading () and trailing () edges, the airfoil () comprising a radially firmer end for connection to a rotor and a free radially outer end. According to the invention, the airfoil () further comprises at least one series of fins () situated on said free end, each of these fins () comprising a pressure face () situated on the suction face () side of the airfoil (), and a suction face () situated on the pressure face () side of the airfoil ().

Rotor blade for a turbomachine

A wheel blade for a turbomachine compressor. The blade comprises a body extending between the root of the blade in the direction of the height of the blade. The body extends from a leading-edge to a trailing edge in a direction of the camber line of the blade. The blade comprises a winglet located in the extension of the body in the direction of the height. The winglet is located at the tip of the blade and at the leading edge of the blade. The body has a curvature oriented in a first direction of rotation with respect to the direction of the camber line, the winglet having a curvature in a second direction of rotation counter to the first direction of rotation.

Wheel blade for turbomachine, comprising a winglet at its tip and at the leading edge

The invention relates to a vane () of a diffuser () for a radial or mixed-flow compressor () of an engine (), including a leading edge () arranged facing a flow of gas, a trailing edge () being opposite the leading edge (), a side upper surface wall () and a side lower surface wall () which connect the leading edge () to the trailing edge (), and a profile including a curved line () having at least two points of inflection (I, I) between the leading edge () and the trailing edge (). The invention also relates to a corresponding radial diffuser ().

Radial or mixed-flow compressor diffuser having vanes

An impeller of a centrifugal compressor, the impeller including a web and blades secured to the web on a front face of the web. A point of intersection between a trailing edge and a blade root is at least one half-thickness of the web further forward than the blade root at an intermediate diameter of the impeller, and a point of intersection between the trailing edge and the blade tip is also further forward than the blade tip at an intermediate diameter of the impeller.

Centrifugal compressor impeller

A diffuser of a compressor of centrifugal or mixed type includes two end plates which enclose a plurality of regularly distributed circumferential blades, and at least one transverse upstream passage produced in lower or upper surfaces of the blades. An injection/withdrawal coupling is achieved by a recirculation of a stream in an air passage of the diffuser on the basis of injection of air from at least one point in a leading edge zone of an upstream side of the diffuser. Blowing of air is then effected in at least one groove formed along a lateral flank of each blade by withdrawal of the air stream in a region of a trailing edge. Thereby, effectively separation of the air in a boundary layer in a gas turbine compressor diffuser is realized by re-energizing the boundary layer with air at a higher pressure by a suction/re-injection coupling.

Gas turbine diffuser blowing method and corresponding diffuser

The invention relates to a secondary air system for a compressor of centrifugal or mixed type including a rotor presenting an axis of rotation, said compressor being adapted to compress an oxidizer gas. The secondary air system includes an oxidizer gas bleed system arranged in the rotor.

Air bleed having an inertial filter in the tandem rotor of a compressor

A diffuser for a mixed flow or centrifugal compressor of a gas turbine. The diffuser includes at least one vane presenting a pressure side, a suction side, and a first flank. The vane includes a plurality of orifices opening out into the suction side and/or the pressure side and communicating with at least one cavity formed in the vane, the cavity extending transversely relative to the vane and opening out into its first flank. The cross section of the cavity varies in the transverse direction of the vane, the cross section increasing towards the first flank.

Growing community of inventors

Moissy-Cramayel, France

Laurent Pierre Tarnowski