Safran

Safran Aircraft Engines

Adrien Bernard Vincent Rollinger

Romain Pierre Cariou

Sylvain Paquin

Thomas Michel Flamme

Patrice Eneau

Léandre Ostino

Vianney Simon

Ba-Phuc Tang

Matthieu Jean Luc Vollebregt

Mirna Bechelany

Joseph Toussaint Tami Lizuzu

Ramzi Bohli

Michel Slusarz

Didier Maurice Marceau Guerche

Pierre Guillaume Auzillon

Thomas Olivier Michel Pierre De Rocquigny

Daniel Quach

Patrick Emilien Paul Emile Huchin

Charlotte Marie Dujol

Ngadia Taha Niane

David Grange

Jean-Claude Marcel Auguste Hanny

Vincent Marc Herb

Hugues Denis Joubert

Pascal Francis Patrick Gomez

Myriam Pelleterat De Borde

Alain Armel Le Hegarat

Filippo Pagnoni

Mathieu Jean Luc Vollebregt

Gael Philippe François Battistoni

Alice Marie Lydie Agier

Romaine Pierre Cariou

Gaël Philippe François Battistoni

The invention relates to a method for manufacturing a vane () of a turbine engine comprising a pressure-face wall () and a suction-face wall () spaced apart from one another, this vane () comprising a tip (S) with:

High-pressure turbine vane including a cavity under a squealer tip

A turbine vane includes a root carrying a blade terminated by a squealer tip, the blade having intrados and extrados walls, a leading edge, a trailing edge, and a tip wall delimiting a bottom of the squealer tip, by which the intrados wall is connected to the extrados wall. The blade also includes: a serpentine median circuit, including a first radial pipe that collects air at the root and is connected by a first bend to a second radial pipe that is connected by a second bend to a third radial pipe; a cavity under the squealer tip running along the extrados wall and extending from a central region of the squealer tip to the trailing edge; and a central radial pipe collecting air at the root and extending between at least two of the three pipes of the median circuit and directly supplying the cavity under the squealer tip.

Turbomachine blade with improved cooling

The invention relates to a turbine engine vane including a blade extending along a radial axis and a first cooling circuit arranged inside the blade, the first cooling circuit including a first cavity and a second cavity disposed downstream of the first cavity in a direction of circulation of a coolant in the blade, the first and second cavities radially extending inside the blade and being at least partly separated by a first radial partition having a radially internal free end, which at least partly demarcates a first coolant passage connecting the first and second cavities. According to the invention, the radially internal free end is enlarged by having a general transverse section substantially in the form of a keyhole.

Turbine engine vane equipped with a cooling circuit and lost-wax method for manufacturing such a vane

A method for manufacturing a ceramic core blank intended for the manufacture of hollow turbomachine blades using the technique of lost wax casting, the blank being manufactured by additive manufacture and comprising at least one first part intended to form the cavities of the hollow blades and at least one second part configured to allow the positioning of the core in a wax mold, the second part and the first part being positioned and held with respect to one another by means of a junction portion interposed between the first part and the second part, at least one through orifice extending through the second part, the junction portion and the first part, a first end of the through orifice opening to an outer face of the second part, and a second end of the through orifice opening to an outer face of the first part.

Method for manufacturing a ceramic core for manufacturing turbomachine vanes

A core for the foundry of an aeronautical part such as a turbine blade, the core being intended to be disposed in an inner housing defined by a mold, the core comprising a body intended to form the internal shape of the turbine blade, an impact portion, disposed on at least a portion of the periphery of the body so as to break a fluid jet when filling the inner housing with the fluid, the impact portion comprising a top and at least one deflection wall converging towards the top.

Core for metal casting an aeronautical part

A turbine blade including a root carrying an impeller terminated by a tip in the form of a squealer tip. This impeller also includes a serpentine median circuit, including a first radial pipe collecting air at the root and that is connected by a first bend to a second radial pipe that is connected by a second bend to a third radial pipe, a cavity under the squealer tip running along the pressure side wall, extending from a central region of the tip to the trailing edge, and a radial central pipe collecting air at the root extending between at least two of the three pipes of the median circuit and directly supplying the cavity under the squealer tip.

Turbine engine blade with improved cooling

The invention relates to a turbomachine turbine vane comprising a root having a blade () comprising a leading edge () and a trailing edge () and a lower surface wall () and an upper surface wall (), as well as: a manifold for cooling () the leading edge (); a supply line () for collecting air from the root and supplying the manifold (); a side cavity () extending along the upper surface wall () and supplied with air from the root, to form a heat shield facing the supply line (); and a circuit forming a trombone comprising a middle portion () extending along the supply line () and extending laterally to the upper surface wall (), the middle portion () being supplied by the supply line at the tip of the blade ().

Cooled turbine vane

Aviation turbine blade () extending in the radial direction from a blade root () as far as an upper partition wall (), said blade () comprising a plurality of inner cavities (C-C) defining at least one cooling circuit, each of said inner cavities (C-C) being defined by walls among inner walls (), a lower surface wall (), an upper surface wall (), the blade root () and the upper partition wall (), said blade () being characterized in that it comprises at least one reinforcing beam () disposed inside one of the inner cavities (C, C), and connecting the blade root () to the upper partition wall (), said reinforcing beam () is not connected to the inner walls (), the lower surface wall () and the upper surface wall ().

Turbine blade having an improved structure

Blade for a turbine, comprising a blade root and an airfoil () extending radially outwards from the blade root (), the airfoil () comprising a first internal cooling circuit including an intrados cavity () extending radially along the intrados wall () and a first inner wall () arranged between the intrados wall () and the extrados wall (), an extrados cavity () extending radially along the extrados wall () and a second inner wall () arranged between the intrados wall () and the extrados wall (). The first cooling circuit includes one inner through cavity () defined between two through walls () each extending between the intrados wall () and the extrados wall (). The intrados cavity (), the extrados cavity () and the inner through cavity () are fluidly connected in series.

Blade comprising an improved cooling circuit

An aviation turbine blade () includes at least a first lower surface cavity (C) and a first upper surface cavity (C), each adjacent to a first through-cavity (C) and a second through-cavity (C), the first upper surface cavity (C) being adjacent to the upper surface wall (), the first lower surface cavity (C) being adjacent to the lower surface wall (), each of the first and second through-cavities (C, C) extending from the lower surface wall () as far as the upper surface wall (), the second through-cavity (C) including a first inner wall (P) extending from the upper surface wall () as far as the first through-cavity (C), and a second inner wall (P) extending from the lower surface wall () as far as the first through-cavity (C). The first (P) and second (P) inner walls are not connected.

A ceramic core used for fabricating a hollow turbine blade for a turbine engine by using the lost-wax casting technique and shaped to constitute the cavities of the blade as a single element, includes, in order to feed the insides of these cavities jointly with cooling air, core portions that are to form first and second lateral cavities and that are connected to a core portion that is to form at least one central cavity, firstly in the core root via at least two ceramic junctions, and secondly at various heights up the core via a plurality of other ceramic junctions of positioning that defines the thickness of the internal partitions of the blade, while also ensuring additional cooling air for predetermined critical zones of the first and second lateral cavities.

Ceramic core for a multi-cavity turbine blade

An assembly for manufacturing a wax moulding of a turbine engine blade has a wax injection mould in which a core is able to be mounted in a predetermined moulding position, the core including a main element and at least a first secondary element each including at least one functional part and a non-functional part, wherein the non-functional part of the first secondary element includes a rod portion extending in a longitudinal direction of the blade and housed in a first slot of the non-functional part of the main member, the mould including a first internal boss clamping said rod portion at the bottom of the first slot.

Assembly for manufacturing a turbine engine blade

An aviation turbine blade extending in the radial direction and presenting a pressure side and a suction side, including a plurality of pressure side cavities extending radially at the pressure side of the blade, a plurality of suction side cavities extending radially at the suction side of the blade, and at least one central cavity located in the central portion of the blade and surrounded by pressure side cavities and by suction side cavities, the blade further including a plurality of cooling circuits, in which at least a first cooling circuit comprises: a first cavity and a second cavity, the first and second cavities communicating with each other at a radially inner end and at a radially outer end of the blade.

Turbine blade comprising a cooling circuit

The invention provides a method of assembling together a first core () and a second core () in order to make a non-permanent model configured for use in lost wax molding to form a part having a first cavity and a second cavity corresponding respectively to the first core and to the second core. The invention is characterized by the fact that the first and second cores () are assembled together with a first spacer (), the first spacer () being arranged between the first and second cores.

Method for creating a nonpermanent model

A core for casting a blade of a turbomachine, the core including a first part for defining a first cavity and a second part of which at least one portion defines a second cavity located between the first cavity and the tip of the blade. The portion of the second part includes a through-hole that ends opposite the first part so as to define, in the cast blade, an outer face of a duct for removing dust from the first cavity. An alumina pin which is secured to the end face of the first part extends into the through-hole to define the inner face of the duct. A centering sheath placed between the pin and the through-hole centers the pin in relation to the through-hole. The centering sheath is made of a material that is intended to dissolve before the blade is cast around the core.

Growing community of inventors

Moissy-Cramayel, France

Adrien Bernard Vincent Rollinger