Safran

Safran Aircraft Engines

Romain Pierre Cariou

Adrien Bernard Vincent Rollinger

Sylvain Paquin

Ba-Phuc Tang

Thomas Michel Flamme

Léandre Ostino

Vianney Simon

Thomas Olivier Michel Pierre De Rocquigny

Patrice Eneau

Michel Slusarz

Pierre Guillaume Auzillon

Ngadia Taha Niane

Ramzi Bohli

David Grange

Didier Maurice Marceau Guerche

Alain Armel Le Hegarat

Filippo Pagnoni

Myriam Pelleterat De Borde

Jeremy Jacques Attilio Fanelli

Jérémy Jacques Attilio Fanelli

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

A turbine engine blade includes an airfoil with a pressure-side wall and a suction-side wall which are connected upstream by a leading edge and downstream by a trailing edge. A cooling circuit has an internal cavity extending inside the airfoil and a plurality of outlet openings, each oriented substantially along a longitudinal axis X. Each outlet opening communicates with the cavity and is arranged in the vicinity of the trailing edge. A calibration device is arranged in the cavity and provided with calibration conduits arranged substantially opposite the outlet openings. The calibration conduits each include an oblong transverse section which is substantially perpendicular to the longitudinal axis.

Turbine engine blade provided with an optimized cooling circuit

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

A turbine engine vane including a blade extending following a radial axis and a cooling circuit arranged inside the blade, the cooling circuit including a first cavity and a second cavity arranged downstream from the first cavity following a direction of circulation of a cooling fluid, the first and second cavities extend radially inside the blade and being separated at least partially by a radial partition having a radially external free end which delimits at least partially a passage connecting the first and second cavities, the radial partition connecting a first wall in contact with the outside environment of the blade to a second opposite wall substantially following a transversal axis, perpendicular to the radial axis, respectively in a connection zone. According to the invention, at least one connection zone presents a thickening having a substantially triangular general transversal cross-section.

Turbine engine blade equipped with a cooling circuit with optimized connection zone

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.

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.

Growing community of inventors

Moissy-Cramayel, France

Romain Pierre Cariou