Single-crystal component for aircraft turbine

Abstract

A single-crystal component for an aircraft turbine (8), made of a superalloy (31), said component comprising at least one internal cavity (C1) lined at least partially on the inside with a thermally conductive coating (28) having a thermal conductivity greater than that of the substrate forming said component. Figure for the abstract: Fig. 3.

Description

Title of the invention: Single-crystal component for aircraft turbine technical field

[0001] The present exposition relates generally to the field of single-crystal components for gas turbines, and more particularly to the cooling of such components. Prior art

[0002] An aircraft generally comprises a plurality of turbomachines, each employing a gas turbine (also called a "combustion turbine"). Among the different turbomachine configurations, a distinction is made between turbojets, whose thrust, intended to propel the aircraft, is generated by reaction to the ejection of hot gases.

[0003] Such a turbojet 1, as illustrated in [Fig.1], is housed in a nacelle 2 of tubular structure along a longitudinal axis X-X', and includes an air inlet 3 and a fan 4 intended to be driven in rotation so as to ensure the compression of the air and its division into a primary airflow and a secondary airflow.

[0004] The primary airflow (or hot airflow) circulates successively through a low-pressure compressor 5, a high-pressure compressor 6, combustion chambers 7 whose combustion gases (or hot gases) drive in rotation one or more high-pressure turbines 8 and one or more low-pressure turbines 9. As for the secondary airflow (or cold airflow), it bypasses the entire part of the turbojet 1 which is traversed by the primary airflow to join it and mix with it in an ejection nozzle 10 before being expelled.

[0005] In the remainder of this presentation and unless otherwise specified, upstream and downstream are defined with respect to the normal flow direction of the primary airflow through the turbojet 1, i.e. from upstream to downstream, or generally from the air inlet 3 to the ejection nozzle 10.

[0006] The constant improvement of these propulsion systems, generally mandated by aeronautical regulations, leads manufacturers to improve the aircraft's propulsion efficiency and therefore its design. This propulsion efficiency also depends on the meteorological conditions in which the aircraft operates, which are increasingly severe.

[0007] To this end, in order to improve engine efficiency, it has been proposed to increase the temperature of the combustion gases intended to pass through the turbines, in particular the high-pressure turbine 8. This temperature increase, however, exceeds by several hundred degrees the melting point of the material that constitutes the blades of the turbine, the said blades being single-crystal parts made of a superalloy (Thus, when reference is made in the text of this presentation to a blade, it is also a reference, by substitution, to a single-crystal part. Similarly, when reference is made in the text of this presentation to a single-crystal part, it is also a reference, by substitution, to a blade).

[0008] Among the possible superalloy materials, one can cite a nickel-based or cobalt-based superalloy, for example.

[0009] The lifespan of these blades is then likely to decrease drastically when they are in regular contact with hot gases exhibiting such temperatures.

[0010] One solution consists of modifying the internal structure of said blades to integrate cooling circuits and thus reduce their temperature when they operate in this environment. To do this, channels are cut into the blade to circulate a cooling fluid inside. The cooling fluid is generally drawn upstream of the low-pressure compressor 5 because it has a temperature 500 to 1000 degrees Celsius lower than the temperature of the hot gases.

[0011] Such a cooling circuit consists of a single cavity in direct contact with the outer walls of the blade, in this case the upper and lower surfaces. In other words, the cavity extends through the thickness of the blade in a direction perpendicular to the longitudinal mean centerline (generally known as the "blade core") between the upper and lower surfaces. The cavity extends to the upper and lower surfaces without any structural obstruction.

[0012] However, the temperature at the level of said mean line is lower than that of the intrados and extrados walls. The latter expand much more than the blade core and therefore subject the blade structure to mechanical stresses likely to alter its integrity and consequently impact its lifespan.

[0013] Other designers have chosen to integrate a row of cooling circuits (and therefore a plurality of cavities) within the thickness of the blade profile, thus forming a profile called an "H-cavity". This type of design is distinguished by the alternating arrangement of a single cavity and double cavities (i.e., superimposed cavities) within the thickness of the blade.

[0014] However, this complex geometry requires particular attention to manage the differential expansions caused by temperature differences between the hot outer walls, which undergo significant expansion, and the cold core of the blade, which expands less. Such differential expansions lead to deformation, notably a bending towards the extrados of the wall. Indeed, the partition located at the The level of the double cavities in the thickness of the blade is very cold, compared to the said external walls, which are warmer and therefore expand more.

[0015] This results in a line of overstress at the level of the cavities which are located at the level of the extrados, and which manifest themselves by radial cracks, indicating a potential failure of the structure under the effect of mechanical stresses.

[0016] In this context, the lifespan of the blades can drastically decrease when they are regularly exposed to hot, high-temperature gases.

[0017] There is therefore a need to maintain the mechanical integrity of the blade during operation while equipping it with an efficient cooling system that allows it to operate continuously and repeatedly within the desired temperatures. Description of the invention

[0018] The present description relates to a single-crystal aircraft turbine part, made of a superalloy, said part comprising at least one internal cavity lined at least partially inside with a thermally conductive coating whose thermal conductivity is greater than that of the substrate forming said part.

[0019] As mentioned above, gas turbines are subjected to extreme operating conditions, with high temperatures and pressures, which imposes stringent requirements regarding materials and design. Single-crystal components, made from superalloys, are widely used in this field due to their superior mechanical properties and heat resistance.

[0020] Such a single-crystal part can be a high-pressure turbine blade, a turbine disc, or a turbine cone, for example.

[0021] The single-crystal component includes at least one internal cavity, which is designed to optimize the cooling and thermal management of the turbine. Such a cavity allows the passage of a cooling fluid, thereby reducing the temperature of the surfaces subjected to hot gas flows.

[0022] Inside this cavity, a thermally conductive coating is applied whose thermal conductivity is greater than that of the substrate of the part, i.e. the superalloy. By being thermally conductive, the coating allows better heat dissipation, thus facilitating heat transfer between the internal cavity and the sensitive areas of the part (in the case of a blade: the leading edge, the trailing edge, the lower surface wall and the upper surface wall; in a general case, these are the external walls of the single-crystal part).

[0023] This thermally conductive coating can be partially applied in the cavity, which means that it is not necessarily applied over the entire surface internal cavity. Some areas may be covered, while others may remain bare, exposed to the base material (the superalloy).

[0024] According to one embodiment of the invention, the thermally conductive coating is composed of a CuNiAg, and / or AgNiCu, and / or CuNiAl alloy.

[0025] The designations CuNiAg, AgNiCu, and CuNiAl refer to specific metallic alloys composed of copper (Cu), silver (Ag), and aluminum (Al). Each combination exhibits unique characteristics in terms of thermal conductivity, corrosion resistance, and mechanical properties known to those skilled in the art.

[0026] It should be noted that the use of the term "and / or" between the different alloy combinations indicates that the coating can be formulated from one or more of these options, thus offering flexibility in design and manufacture. This allows a person skilled in the art to select the most appropriate alloy according to specific performance and production cost requirements.

[0027] By integrating these specific alloys into the thermally conductive coating, the invention aims to improve the thermal performance of turbines while taking into account the associated challenges and constraints.

[0028] According to one embodiment of the invention, the thermally conductive coating is present in the material of the part over a non-zero depth of less than 200 qm.

[0029] The thermally conductive coating is therefore not superficial because it penetrates slightly into the structure of the material, which facilitates the dissipation of heat generated by the hot gases in the turbine.

[0030] Furthermore, a coating depth of less than 200 sq m helps to maintain the mechanical integrity of the part's material. Indeed, if the coating were too thick, it could create internal stresses or alter the mechanical properties of the part, particularly in terms of fatigue resistance.

[0031] According to one embodiment of the invention, said at least one internal cavity is totally lined with said thermally conductive coating.

[0032] The internal surface of the cavity is therefore entirely covered by the thermally conductive coating. This means that no part of the cavity's internal surface is left uncoated, thus allowing for homogeneity in thermal conductivity within the part, if desired. This is also advantageous for reducing, or even eliminating, hot spots that could result from uncoated areas.

[0033] According to one embodiment of the invention, the part is a blade extending radially about an axis of rotation of the turbine and comprising: - a blade root comprising a circulation circuit for a blade cooling fluid and provided with a platform defining an external radial end of the blade root and, - a blade extending from said platform towards an external radial end of the blade and comprising a leading edge and a trailing edge, an intrados wall and an extrados wall each connecting the leading edge and the trailing edge, and transverse partitions each extending between the intrados wall and the extrados wall so as to define a plurality of internal cavities, defined above, within the blade, the blade comprising a plurality of internal cooling circuits, each being disposed in an internal cavity and connected to said circulation circuit, said internal cooling circuits being arranged radially between the intrados wall and the extrados wall and extending successively between the leading edge and the trailing edge.

[0034] As stated above, the radial direction corresponds to the direction in which the blade extends relative to the axis of rotation of the turbine carrying the blade. The radial and longitudinal directions of the blade are therefore considered synonymous in the remainder of this description.

[0035] In each cooling circuit, the internal cavity (the blade core or "cold zone") extends on either side of the mean line which extends radially along the intrados and extrados walls.

[0036] According to one embodiment of the invention, the platform includes a cooling cavity connected to said fluid circulation circuit, and includes a plurality of holes, opening from the platform opposite the blade and / or on the edge of the platform, so as to at least partially evacuate the cooling fluid from the platform.

[0037] The cooling fluid first circulates in the blade foot and then in said cooling cavity hollowed out in the thickness of the platform.

[0038] The cooling fluid is then at least partially released through said holes in the ejection nozzle. The platform is then further cooled, which helps to lower the average temperature of the blade.

[0039] By way of example, said holes may have a diameter between 0.3 and 0.35 mm.

[0040] According to certain embodiments of the invention, the blade comprises a plurality of slots arranged along the trailing edge, and a plurality of holes positioned along the leading edge, on the surface of the intrados wall and on the surface of the extrados wall, each slot and hole connecting one of the internal cooling circuits to an external volume of the blade.

[0041] The cooling fluid is therefore discharged into the ejection nozzle after circulating through the various cavities designed to cool the blade. Such holes are located on the outer surface of the sensitive areas – the volume external to the blade – here the intrados wall and the extrados wall.

[0042] These holes are also positioned along the leading edge, which is also a sensitive area, just like the intrados wall and the extrados wall, because it is likely to exhibit greater thermal expansion than the blade core following regular contact with hot gases.

[0043] As with the leading edge, the trailing edge is a sensitive area with slots because it also changes in contact with hot gases.

[0044] The holes may have a diameter between 0.3 and 0.35 mm and / or be inclined so as to generate a film of air (“film-cooling” according to the Anglo-Saxon term) directed downstream and intended to further cool the blade.

[0045] The present exposition further relates to a gas turbine comprising a plurality of single-crystal parts as defined above.

[0046] The present description further relates to a turbomachine comprising a gas turbine as defined above.

[0047] The present description further relates to a method for manufacturing a single-crystal part as defined above, comprising the following successive steps: 1) a step of depositing said thermally conductive coating at least partially on the external surface of at least one ceramic core, each ceramic core being intended to allow the formation of an internal cavity as defined above, after its detachment;

[0048] 2) a step of injecting wax into a first mold so as to create a model in wax of the single-crystal part to be manufactured, the first mold containing said at least one ceramic core lined with the thermally conductive coating;

[0049] 3) a step of covering the wax model with a refractory layer to create a second mold;

[0050] 4) a step of removing the wax from the second mold;

[0051] 5) a step of casting the superalloy into the second mold and around said mold minus a ceramic core;

[0052] 6) a step of removing the second mold; and

[0053] 7) a step of unseating said at least one ceramic core after the solidification of the superalloy, so as to obtain the single-crystal part resulting from steps 1) to 6).

[0054] In other words, in step 1), the thermally conductive coating is applied to the external surface of a ceramic core or a plurality of ceramic cores. This core is intended to form an internal cavity that will be present in the final single-crystal piece after its deburring.

[0055] Once the coating has been applied, the next step is to inject wax to create the wax model. This model determines the final shape of the part and will therefore serve as the basis for the subsequent steps.

[0056] After the creation of the wax model, it is covered with a refractory layer that withstands the high temperatures encountered during the casting of the superalloy. Thus, by hardening around the wax model, the refractory coating creates a second robust mold that will maintain the shape of the part during the casting process.

[0057] Once the refractory layer has hardened, the wax is removed from the second mold. This step is generally carried out by heating, which melts the wax and allows it to be removed without damaging the refractory coating. Naturally, the blade shell is fired before the superalloy is poured into it.

[0058] When the second mold is ready, the molten superalloy is poured into it, surrounding the ceramic core(s). More specifically, when the superalloy is poured, it comes into contact with the applied coating and allows for chemical interaction between the superalloy, the thermally conductive coating, and the core surface during the melting phase. It should be noted that once the superalloy has cooled, the part is protected by the combination of elements diffused within its structure. In other words, the local chemistry of the part has been favorably modified, which contributes to improved resistance to high temperatures of the blade in its overall structure. Thus, no post-casting deposition process is necessary to protect the resulting single-crystal part.

[0059] The demolding of the core(s), as well as the second mold, is then carried out after the superalloy has solidified (and cooled), which allows, among other things, the internal cavity(ies) to be exposed. Once the core(s) have been removed, the finished part can be inspected and tested to ensure that it meets the required specifications. This final process yields a functional single-crystal part, ready for integration into applications encountered in aircraft turbines.

[0060] According to one embodiment of the invention, the thermally conductive coating is deposited by implementing a chemical vapor phase deposition.

[0061] Chemical Vapor Deposition (CVD) is a technique used to apply thin films of materials to substrates, which is advantageous when manufacturing parts with specific thermal properties. Thus, in the context of the invention, the thermally conductive coating is deposited using this method.

[0062] The application of the thermally conductive coating via CVD makes it possible to reach the areas to be coated, ensuring that the coating is applied homogeneously and continuously over the entire surface in a targeted manner. Brief description of the drawings

[0063] Other objects, features, and advantages of the invention will be better understood upon reading the following detailed description of various embodiments of the invention given by way of non-limiting examples. This description refers to the accompanying figure pages, on which: - [Fig.1] Fig.1 schematically presents a cross-sectional view of an aircraft turbomachine; - [Fig. 2] Fig. 2 illustrates a perspective view of a room monocrystalline according to the present invention; - [Fig. 3] Figure 3 illustrates a cross-sectional view of one embodiment of the internal structure of a blade as a single-crystal piece, according to plane P of [Fig.2]; - [Fig. 4] Fig. 4 illustrates the different stages of manufacturing said part monocrystalline according to the invention.

[0064] It should be noted that across all the figures, the common elements are identified by identical numerical references. Description of the implementation methods

[0065] To make the explanation more concrete, an example of a blade 11 (as a single-crystal component) is illustrated in [Fig. 2] and is described in detail below. It should be noted that the invention is not limited to this example and can be applied to any single-crystal turbine component made of a superalloy.

[0066] More specifically, the blade 11 is generally made of a nickel-based superalloy so as to withstand high thermal and mechanical stresses when in operation.

[0067] This material generally constitutes 80%, or even 90%, or even 99% of the structure of the blade 11.

[0068] Of course, the nickel-based superalloy can be replaced by any other material deemed suitable by those skilled in the art for manufacturing such a blade 11.

[0069] Furthermore, the blade 11 can be produced by lost-wax casting, a process known to those skilled in the art. The manufacturing process for the blade 11 according to the invention will be detailed below.

[0070] This manufacturing process makes it possible to obtain a blade having an aerodynamic profile designed to provide air passages between adjacent blades 11 without causing airflow turbulence detrimental to the efficiency of the turbine 8.

[0071] The blade 11 can be completely or partially manufactured by additive manufacturing. By "partially" is meant that certain specific parts of the blade 11 such as the tub are manufactured by additive manufacturing, while other parts of the blade 11 are manufactured by another blade manufacturing process.

[0072] The blade 11 includes a blade foot 12 which generally has a fir tree-shaped profile intended to be inserted into a disk of the turbine rotor 8 and thus be driven around its axis of rotation.

[0073] The blade root 12 is provided with a platform 13 defining an external radial end of the blade root 12 and supporting a blade 14 that extends radially between the blade root 12 and an external radial end of the blade that is not visible in [Fig. 2]. In other words, the blade 14 extends in a radial direction Y-Y' with respect to the axis X-X', which also corresponds to the axis of rotation of the turbine 8 as illustrated in [Fig. 1]. The radial and longitudinal directions of the blade 14 are thus identical.

[0074] More specifically, the blade 14 comprises four distinct parts: a leading edge 15 and a trailing edge 16, as well as an intrados wall 17 and an extrados wall 18, each connecting the leading edge 15 and the trailing edge 16.

[0075] These four parts are subjected to regular contact with hot gases having temperatures that are several hundred degrees higher than the melting temperature of the material constituting the blade 11.

[0076] Such parts are considered to be sensitive areas because their mechanical properties decrease when the temperature of the hot gases increases. They are then likely to expand more than the relatively cooler core of the turbine blades.

[0077] To overcome these disadvantages, the blade foot 12 includes a circulation circuit 19 of a blade cooling fluid 11 intended to circulate inside the blade 14 of the blade 11 and thus cool it.

[0078] The cooling fluid is generally taken upstream of the low-pressure compressor 5 and has a temperature 500 to 1000 degrees Celsius or more lower than the temperature of the hot gases flowing around the blade 14.

[0079] The interior of the blade 14 then consists of a plurality of internal cooling circuits for the blade 11 which are connected to said circulation circuit 19 and which will be described in [Fig.3].

[0080] More specifically, the cooling fluid first circulates in the blade foot 12 and then in the cooling circuits arranged inside the blade 14 before being released into an ejection nozzle 10.

[0081] For this purpose, the blade has a plurality of slots arranged along the trailing edge 16, and a plurality of holes positioned along the leading edge 15, on the outer surface of the intrados wall 17 and on the outer surface of the extrados wall 18.

[0082] The holes and slots are then positioned at the sensitive areas and connect one, several or all of the internal cooling circuits to an external volume of the blade 11.

[0083] By way of example, the holes have a diameter between 0.3 and 0.35 mm and / or are inclined so as to generate a film of air ("film-cooling" according to the Anglo-Saxon term) directed downstream and intended to cool the blade 11 further.

[0084] Furthermore, the circulation circuit 19 of the cooling fluid for the blade 11 can be connected directly or indirectly to said cooling circuits. The connection is considered "direct" when the circulation circuit 19 is connected to the cooling circuits by continuous channels intended to pass through the platform 13 of the blade 11 before coming into contact with the cooling circuits. The connection is considered "indirect" when a cooling cavity 22 is cut into the thickness of the platform 13 and connects the circulation circuit 19 to the cooling circuits.

[0085] Thus, when the connection is indirect, the cooling fluid circulates first in the blade foot 12 and then in said cooling cavity 22 hollowed out in the platform 13, which makes it possible to cool the platform 13, consequently contributing to lowering the average temperature of the blade 11.

[0086] Since the cooling cavity 22 is wider than the channels, it has a better cooling capacity. It therefore has two roles: to cool the platform 13 and to circulate the cooling fluid towards the blade 14.

[0087] To further cool said platform 13, the latter has a plurality of holes, opening from the platform 13 opposite the blade 14. The holes can be arranged alternately or additionally on the edge 25 of the platform 13.

[0088] The cooling fluid can therefore escape through the holes made on the blade 14 but also through the holes opening from the platform 13.

[0089] The holes drilled on the surface of the platform 13 can have a diameter between 0.3 and 0.35 mm for example.

[0090] Fig. 3 represents a section along plane P visible in Fig. 2 and illustrates an embodiment of a cooling circuit through which the cooling fluid circulates before being released into the ejection nozzle 10. The section plane P is perpendicular to the radial direction Y-Y'.

[0091] Figure 3 shows a plurality of cooling circuits 26, here five, arranged radially between the intrados wall 17 and the extrados wall 18, and extending successively between the leading edge 15 and the trailing edge 16 and thus along a The mean longitudinal line Z-Z' (or axis Z-Z') between the lower surface wall 17 and the upper surface wall 18. The mean longitudinal line Z-Z' is defined as a median line between the lower surface wall 17 and the upper surface wall 18, in a plane perpendicular to the radial direction Y-Y'. In the case of a blade 14 with symmetrical geometry, the mean longitudinal line Z-Z' corresponds to the chord of the airfoil of the blade 14.

[0092] The five cooling circuits 26 are separated from each other within the blade 14 by through-partitions 27, each extending between the lower surface wall 17 and the upper surface wall 18. It is possible to connect these through-partitions by other through-partitions 27 extending between the leading edge 15 and the trailing edge 16.

[0093] In other words, the through-partitions 27 define a plurality of internal cavities within the blade 14 so that each internal cooling circuit 26 is housed in a separate internal cavity.

[0094] It is understood that the illustrated example is not limiting, and that the number of cooling circuits 26 is defined according to the geometry of the blade 14, and in particular its dimensions. More generally, the blade 14 comprises at least two cooling circuits 26 or at least three cooling circuits 26.

[0095] In order to cool the blade 11, each cooling circuit 26 has its internal cavity Cl lined at least partially on the inside with a thermally conductive coating 28 whose thermal conductivity is greater than that of the substrate forming said part 11, i.e., the superalloy. By being thermally conductive, the coating 28 allows for better heat dissipation from the blade, thus facilitating heat transfer between the internal cavity Cl and the sensitive areas of the part. The mechanical stresses that could be generated by differential expansion between the sensitive areas and the blade core are then limited, or even eliminated.

[0096] This thermally conductive coating 28 can be partially applied within the cavity Cl, meaning that it is not necessarily applied over the entire internal surface of the internal cavity Cl. Some areas may be covered, while others may remain bare, exposed to the base material (the superalloy). Alternatively, the internal cavity Cl can be completely lined with said thermally conductive coating 28, as illustrated in [Fig. 3]. The internal surface of the internal cavity Cl is thus entirely covered by the thermally conductive coating 28. This means that no part of the internal surface of the cavity is left uncoated, thereby ensuring homogeneity in thermal conductivity within the cavity.

[0097] Such a thermally conductive coating 28 may be composed of a CuNiAg, and / or AgNiCu, and / or CuNiAl alloy as defined above. Furthermore, it may also be present in the material of the part to a non-zero depth of less than 200 sq m.

[0098] The thermally conductive coating 28 is therefore not superficial because it penetrates slightly into the structure of the material, which facilitates the dissipation of heat generated by the hot gases in the turbine.

[0099] Furthermore, a coating depth of less than 200 µm helps to maintain the mechanical integrity of the part's material. Indeed, if the coating were too thick, it could create internal stresses or alter the mechanical properties of the part, particularly in terms of fatigue resistance.

[0100] It should be noted that the internal cavity Cl may include flow disruptors for the cooling fluid of the blade 11 and which may be sized in the form of protrusions intended to promote the flow of said fluid towards sensitive areas, in particular the trailing edge 16, the intrados wall 17 and the extrados wall 18.

[0101] Figure 4 illustrates the different stages of a process S0 for manufacturing said single-crystal part according to the invention, here a blade.

[0102] The process S0 begins with a first step SI of depositing said thermally conductive coating 28 at least partially, or even totally, on the external surface of at least one ceramic core 29, each ceramic core 29 being intended to allow the formation of an internal cavity Cl, as defined above, after its detachment.

[0103] The process S0 continues with a second step S2 of injecting wax 30 into a first mold Ml so as to create a wax model of the single-crystal part to be manufactured, the first mold Ml containing said at least one ceramic core 29 lined with the thermally conductive coating 28.

[0104] The process S0 continues with a third step S3, not visible in the figure, of covering the wax model M1 with a refractory layer to create a second mold M2.

[0105] The process S0 continues with a fourth step S4, not visible in the figure, of removing the wax 30 from the second mold M2. The shell of the developing blade is typically baked to obtain a stable structure that will withstand the high-temperature casting of the superalloy. This ensures that the shell and core are well consolidated to withstand the casting process in the next step S5, without deforming or breaking.

[0106] The process S0 then continues with a fifth step S5 of pouring the superalloy 31 into the second mold M2 and around it at least one ceramic core 29. More specifically, when the superalloy 30 is poured, it comes into contact with the applied coating 28 and allows a chemical interaction between the superalloy 31, the thermally conductive coating 28, and the surface of the core 29 during the melting phase. It should be noted that once the superalloy 31 has cooled, the part is protected thanks to the combination of elements diffused within its structure. In other words, the local chemistry of the part has been favorably modified, contributing to improved high-temperature resistance of the blade in its overall structure through a modification of the thermal expansion differential. Thus, no post-casting deposition process is necessary to protect the resulting single-crystal part.

[0107] Finally, the process S0 ends with a sixth step S6 of detachment of said at least one ceramic core 29 after the solidification of the superalloy 31, so as to obtain the single-crystal part resulting from the succession of steps from the first step SI to the fifth step S5.

[0108] The process S0 may further include an optional seventh step S7 comprising an inspection of the resulting single-crystal part to ensure that it meets the required specifications. This final process yields a functional single-crystal part, ready for integration into applications encountered in aircraft turbines.

[0109] Although the present invention has been described with reference to specific embodiments, it is evident that modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the claims. In particular, the number of cooling circuits and cooling cavities is not limited to those shown in this example. The present invention can also be applied to other turbine blades besides high-pressure turbine blades for turbomachinery. Furthermore, the invention applies more generally to any single-crystal turbine component made of a superalloy. Therefore, the description and drawings should be considered illustrative rather than restrictive.

[0110] It is also evident that all the characteristics described with reference to a device are transposable, alone or in combination, to a process and vice versa.

Claims

Demands

1. Single-crystal part for an aircraft turbine (8), made of a superalloy (31), said part comprising at least one internal cavity (Cl) lined inside at least partially with a thermally conductive coating (28) having a thermal conductivity greater than that of the substrate forming said part.

2. Single crystal part according to claim 1, wherein the thermally conductive coating (28) is composed of a CuNiAg, and / or AgNiCu, and / or CuNiAl alloy.

3. Single crystal part according to claim 1 or 2, wherein the thermally conductive coating (28) is present in the material of the part over a non-zero depth of less than 200 pm.

4. Single-crystal part according to any one of the preceding claims, wherein said at least one internal cavity (Cl) is totally lined with said thermally conductive coating (28).

5. A single-crystal component according to any one of the preceding claims, the component being a blade (11) extending radially about an axis of rotation of the turbine (8) and comprising: - a blade root (12) comprising a circulation circuit (19) for a blade cooling fluid (11) and provided with a platform (13) defining an external radial end of the blade root (12), and - a blade (14) extending from said platform (13) to an external radial end of the blade (11) and comprising a leading edge (15) and a trailing edge (16), an intrados wall (17) and an extrados wall (18) each connecting the leading edge (15) and the trailing edge (16), and transverse partitions (27) each extending between the intrados wall (17) and the extrados wall (18) so as to define a plurality of internal cavities (Cl), as defined according to any one of claims 1 to 4, within the blade (14), the blade (14) comprising a plurality of internal cooling circuits (26), each being disposed in an internal cavity (Cl) and connected to said circulation circuit (19),said internal cooling circuits (26) being arranged radially between the intrados wall (17) and the extrados wall (18) and extending successively between the leading edge (15) and the trailing edge (16).

6. A single-crystal part according to claim 5, in which the platform (13) has a cooling cavity (22) connected to said fluid circulation circuit (19), and has a plurality of holes, opening from the platform (13) opposite the blade (14) and / or on the edge (25) of the platform (13), so as to at least partially evacuate the cooling fluid (19) from the platform (13).

7. Single-crystal part according to claim 5 or 6, wherein the blade (14) comprises a plurality of slots arranged along the trailing edge (16), and a plurality of holes positioned along the leading edge (15), on the surface of the lower surface wall (17) and on the surface of the upper surface wall (18), each slot and hole connecting one of the internal cooling circuits (26) to an external volume of the blade.

8. Gas turbine (8) comprising a plurality of single-crystal parts according to any one of claims 1 to 7.

9. Turbomachine (1) comprising a gas turbine (8) according to claim 8.

10. A method for manufacturing a single-crystal part according to any one of claims 1 to 7, the method comprising the following successive steps: 1) a step (S1) of depositing said thermally conductive coating (28) at least partially on the external surface of at least one ceramic core (29), each ceramic core (29) being intended to allow the formation of an internal cavity (C1), according to any one of claims 1 to 7, after its demolding; 2) a step (S2) of injecting wax (30) into a first mold (M1) so as to create a wax model of the single-crystal part to be manufactured, the first mold (M1) containing said at least one ceramic core (29) lined with the thermally conductive coating (28); 3) a step (S3) of covering the wax model with a refractory layer to create a second mold (M2); 4) a step (S4) of removing the wax from the second mold (M2);5) a step (S5) of casting the superalloy (31) into the second mold (M2) and around said at least one ceramic core (28); 6) a step of unmolding the second mold; and 7) a step (S6) of unmolding said at least one ceramic core (28) after the solidification of the superalloy (31), so as to obtain the single-crystal part resulting from steps 1) to A'»;

11. 0). Manufacturing method according to claim 10, wherein the thermally conductive coating (28) is deposited by implementing a chemical vapor deposition.