US-12624640-B2 - Turbine engine element comprising at least one blade obtained by additive manufacturing

Abstract

The present invention relates to a turbomachine element ( 1 ), comprising at least one blade ( 2 ) obtained by additive manufacturing, the blade ( 2 ) having a skin ( 4 ) and an internal lattice ( 6 ) allowing air circulation in the blade ( 2 ) and having an additive manufacturing support function for the skin ( 4 ).

Inventors

• Sylvain Pierre Votie
• Denis Daniel Jean BOISSELEAU
• Xavier Roger BETBEDER-LAÜQUE

Assignees

• SAFRAN HELICOPTER ENGINES

Dates

Publication Date

20260512

Application Date

20220916

Priority Date

20210917

Claims (11)

1. 1 . A turbine engine element, the turbine engine element comprising at least one blade obtained by additive manufacturing, the at least one blade comprising: a skin; and an internal lattice comprising a mesh formed by a network of solid structures and empty regions connected together and arranged in a closed pattern to form a grid, the lattice being configured for allowing air circulation in the blade and for supporting the skin during the additive manufacturing, wherein the at least one blade further comprises at least one insert positioned in the lattice, and the lattice further comprises an inner portion and an outer portion, separated by the insert.
2. 2 . The turbine engine element according to claim 1 , wherein the lattice comprises a variable density.
3. 3 . The turbine engine element according to claim 2 , wherein the density is higher close to the skin.
4. 4 . The turbine engine element according to claim 1 , wherein the at least one insert further comprises at least one opening configured for allowing air circulation towards the skin.
5. 5 . The turbine engine element according to claim 1 , further comprising two circumferential duct walls, wherein the at least one blade extends in a radial direction between the two circumferential duct walls, the radial direction being perpendicular to a main axis of a turbine engine and intersecting the main axis, the skin forming two tangential walls of the at least one blade.
6. 6 . The turbine engine element according to claim 5 , wherein the blade further comprises two openings, each opening extending in a plane perpendicular to the radial direction.
7. 7 . The turbine engine element according to claim 6 , wherein the blade does not have a wall extending in a plane perpendicular to the radial direction.
8. 8 . The turbine engine element according to claim 1 , wherein the turbine engine element being at least one of: a high-pressure nozzle; an inlet guide vane; and a variable stator vane.
9. 9 . An aircraft turbine engine, the aircraft turbine engine comprising a plurality of turbine engine elements according to claim 1 .
10. 10 . An aircraft, the aircraft comprising the aircraft turbine engine according to claim 9 .
11. 11 . A method for manufacturing the turbine engine element according to claim 1 , the method comprising the additive manufacturing of the skin and the lattice of the least one blade of the turbine engine element, the lattice supporting the skin during the additive manufacturing, the method being performed so that the lattice extends in the at least one blade at the end of the additive manufacturing.

Description

FIELD OF THE INVENTION The present invention relates to the field of turbine engines and, more particularly, to cooling systems of turbine engine elements. Prior Art The parts inside a turbine engine are subjected to very high thermal stresses. In order to prevent certain elements from breaking or wearing out prematurely, it is necessary to cool them during the operation of the turbine engine. This is particularly the case for high-pressure nozzles, inlet guide vanes and variable stator vanes. Currently, several methods are known for cooling these elements, including: added multiply-pierced insert (impact): with this technology (shown in FIGS. 1 and 2), the ventilating air is guided at high-speed by the piercings (impact holes produced in an insert) onto the part to be ventilated.Internal coil type circuit: with this technology (shown in FIG. 3), the ventilating air is guided through the part to be cooled in order to produce a convective heat exchange. It is discharged into the main duct (hot gas) via vents.Disruptors (bridges/fins): with this technology (shown in FIG. 4), the path of the ventilating air may be partially obstructed by disruptors in order to produce local heat exchange between the disruptors and the ventilating air (+acceleration of the air in order to increase convective heat exchange). These disruptors can also be used to thermally connect the pressure side and suction side of a blade. In addition, it is known from document FR3085713 to incorporate cooling capillaries in a vane (or a blade). These capillaries pass through the vane and thus enable air circulation through the vane. This system enables efficient cooling but is not optimal in the case of producing the element (vane, blade or nozzle) by additive manufacturing. More specifically, since a vane (or blade) may be a hollow part, during its production by additive manufacturing, it may be necessary to incorporate a manufacturing support in the vane. As a function of the direction of manufacture, the support can support the material deposited, for example in order to produce one of the outer faces of the vane. However, for many elements of a turbine engine, additive manufacturing by laser powder bed fusion imposes a direction of manufacture along the axis of the turbine. In this context, a hollowed out blade with many deflections cannot be produced in a single piece without supports. It is therefore essential to add manufacturing supports in the blade. However, these supports, which are not initially provided, can interfere with the capillaries, and potentially degrade the intrinsic performance of the part. Moreover, certain unsupported capillaries may be obstructed and inaccessible at the end of the additive manufacturing method. In this context, it is necessary to supply a turbine engine element comprising a blade having a structure that is suitable to be cooled and to be produced by additive manufacturing. DISCLOSURE OF THE INVENTION According to a first aspect, the invention proposes a turbine engine element, comprising at least one blade obtained by additive manufacturing, the blade having a skin and an internal lattice allowing air circulation in the blade and having an additive manufacturing support function for the skin. The lattice can have a variable density. The lattice can have a higher density close to the skin. The blade can have at least one insert positioned in the lattice. The insert can have at least one opening allowing air circulation towards the skin. The lattice can comprise an inner portion and an outer portion, separated by the insert. The element can have two circumferential duct walls between which said at least one blade extends in a direction radial to a main axis of the turbine engine, the skin forming two tangential walls of said at least one blade. The blade can have openings extending in a plane perpendicular to the radial direction. The blade may not have a wall extending in a plane perpendicular to the radial direction. The element can be chosen from a high-pressure nozzle, an inlet guide vane and a variable stator vane. According to a second aspect, the invention proposes an aircraft turbine engine, the turbine engine comprising an element according to the first aspect. According to a third aspect, the invention proposes an aircraft comprising a turbine engine according to the second aspect. According to a fourth aspect, the invention proposes a method for manufacturing a turbine engine element according to the first aspect, the method comprising additive manufacturing of a skin of a blade of the element using a lattice as support, the lattice extending in the blade at the end of manufacturing. DESCRIPTION OF THE FIGURES Other features, aims and advantages of the invention will emerge from the following description, which is given purely by way of illustration and not being limiting and which should be read with reference to the attached drawings, in which: FIG. 1 is a representation of a device of th