Safran Aircraft Engines

Snecma

Snecma Moteurs

Erwan Daniel Botrel

Laurent Patrick Robert Coudert

Maxime Didier Delabriere

Sébastien Serge Francis Congratel

Patrice Eneau

Daniel Marius Man

Charlotte Marie Dujol

Léandre Ostino

Pierre Hervé Marche

Jacques Auguste Amedee Boury

Philippe Picot

Claire Dorin

Jacques Boury

Jean-Armand Marc Destouches

Regis Grohens

Thomas Michel Mervant

Régis Grohens

Gregory Pierre Lafarge

Michel Boquet

Christophe Germain Brisset

Bertrand Petot

Kevin Yannick Garles

Simon Parinet

Jacques Auguste Boury

Claire D Dorin

Jacques Auguste Amédée Boury

A turbine vane of a turbine engine such as a turboprop or a turbojet, the vane comprising a root supporting a blade, the vane including at least one air-circulation duct for cooling the air during operation, the duct including, at the vane root, an inlet portion for collecting the cooling air, the inlet portion extending away from a lower surface of the vane root opposite the blade. At least one inlet portion is provided with a helical element for rotating the cooling air in order to improve the cooling efficiency thereof.

Turbine vane including a cooling-air intake portion including a helical element for swirling the cooling air

A turbine blade, including a root, a vane including a leading edge and a trailing edge and a pressure-side wall and a suction-side wall, and including cooling vents at the trailing edge, this vane also including first and second serpentine circuits; each serpentine circuit including several ducts extending in the span direction, being connected to each other by angled portions; each serpentine circuit being supplied with air by its duct that is closest to the leading edge; and wherein the vents are supplied by the first and by the second serpentine circuit.

Turbine engine blade with improved cooling

A turbomachine turbine blade, includes a platform, a vane, a cooling cavity supplying a plurality of cooling outlets provided along the trailing edge, two radially adjacent cooling outlets defining therebetween a rib. At least one cooling hole is formed in the thickness of at least one rib and/or in the thickness of a portion of the trailing edge fillet located in the axial extension of at least one rib, so as to ensure fluid communication for a cooling flow between the inside and the outside of the blade for cooling the at least one rib.

Turbine blade comprising ribs between cooling outlets with cooling holes

A turbomachine blade including a hollow body defining a cavity, and including a downstream trailing edge. The blade further includes at least one hole which communicates with the cavity and which opens downstream onto the trailing edge. The trailing edge extends along the main radial direction of the blade, and including a downstream convex face. The blade is characterised in that each hole opens onto a planar rear face of a flat region formed in the trailing edge.

Turbine blade comprising an improved trailing-edge

A turbine vane of a turbine engine such as a turbojet. The vane includes a base supporting a blade that extends in a spanwise direction and ends in an apex. The blade includes a pressure-side wall and a suction-side wall, each ending at an end edge at the apex of the blade. The blade includes, at the apex thereof, a closed wall extending from the pressure-side wall to the suction-side wall so as to define, with the end edges, a tub shape. The pressure-side wall is curved inward so as to deviate from the spanwise direction in the region of the blade apex preceding the tub, between the base and the pressure-side end edge.

Turbine vane comprising a blade with a tub including a curved pressure side in a blade apex region

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane Psituated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

Optimized aerodynamic profile for a turbine vane, in particular for a nozzle of the fifth stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

Optimized aerodynamic profile for a turbine blade, in particular for a rotary wheel of the seventh stage of a turbine

The invention relates to a turbomachine vane comprising a root and a blade including a median main plane having a longitudinal and radial main orientation, which is carried by the root, the blade including a leading edge located at an upstream longitudinal end, a trailing edge located at a longitudinal end downstream of the leading edge with respect to a gas stream flowing around the blade, a lower surface wall and an upper surface wall which are located laterally remote from each other and each connecting the leading edge to the trailing edge, and a top located at the free outer radial end of the blade, the blade further including a fin having a longitudinal main orientation which is carried by the lower surface side, which is located at the top of the blade, wherein the fin is radially inwardly offset with respect to the top of the blade.

Turbomachine vane including an antivortex fin

The invention relates to a movable vane of a turbomachine comprising a blade () with inner cooling cavities and a root () by which means the vane can be mounted on a rotor disk, the root comprising at least two channels () communicating with said inner cavities and leading to the base () thereof, said base comprising at least two openings () through which the channels open up, and a calibrating plate () provided with calibrated perforations () corresponding to said openings being attached to the base () of the root. The vane is characterized in that a mechanical means () forming a sealing barrier between the two openings is arranged between the plate and the base of the root.

Cooled vane of a high-pressure turbine

A hollow blade including an airfoil extending along a longitudinal direction, a root, a tip, an internal cooling passage, and an open cavity defined by an end wall and a rim, together with cooling channels connecting the internal cooling passage to a pressure side. The cooling channels slope relative to the pressure side. A stack of airfoil sections of the blade at a level of the rim of the tip of the blade are offset towards the pressure side. The pressure side wall of the airfoil includes a projecting portion and cooling channels arranged in the projecting portion to open out into a terminal face of the projecting portion.

Gas turbine blade with tip sections offset towards the pressure side and with cooling channels

A turbine blade including an open cavity at its distal tip, the cavity being defined by a bottom wall and a side wall extending along the perimeter of the distal tip in an extension of the upper and lower walls of the blade, the side wall of the cavity including an opening in a vicinity of the leading edge of the blade opening into the cavity. A deflector extends at least in the middle portion of the cavity between the leading edge and the trailing edge.

Cooling of the tip of a blade

When cold and in the uncoated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X, Y, Z' given in Table 1, in which the coordinate Z&#x2032; is the quotient D/H where D is the distance from the point in question to a reference plane Psituated at the base of the nominal profile and H is the height of this profile, measured from the reference plane to the tip of the blade. The measurements D and H are taken radially relative to the axis of the turbine, while the coordinate X is measured in the axial direction of the turbine.

Optimized aerodynamic profile for a turbine blade

A turbine ring sector is disclosed. The turbine ring sector includes a wall bounding an aerodynamic flowpath in which gases flow from the upstream region toward the downstream region, and a multiperforated plate, situated on the opposite side of the aerodynamic flowpath with respect to the wall. The multiperforated plate includes a bottom and side walls. The distance between the wall of the ring sector and the bottom of the multiperforated plate defines a gap which is progressive in an axial direction, from the upstream region toward the downstream region. The value of the gap situated at the upstream region is below the value of the gap situated at the downstream region.

Turbomachine turbine ring sector

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the rectangular coordinates X,Y,Z' given in Table 1, in which the X coordinate is measured in the axial direction of the turbine, and the Z&#x2032; coordinate is the quotient D/H, where D is the distance of the point in question from a reference plane Psituated at the base of the nominal profile and H is the height of the profile measured from said reference plane to the tip of the blade, the measurements D and H being taken radially relative to the axis of the turbine.

A turbomachine blade presenting a free end having an end wall from which there projects a rim, the rim including a first fraction starting from the trailing edge of the blade and running along the suction side of the blade to a connection zone, and a second fraction of closed outline connected to the first fraction at the region of the connection zone. The second fraction extends the first fraction, running along the suction side of the blade, along the leading edge of the blade, and then returning to the connection zone.

Growing community of inventors

Alfortville, France

Erwan Daniel Botrel