Single-crystal component for an aircraft turbine

Abstract

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

Description

Single-crystal component for aircraft turbine Technical Field

[0001] This presentation generally concerns the field of single-crystal parts for gas turbines, and more specifically the cooling of such parts. Previous technique

[0002] An aircraft typically comprises a plurality of turbomachines, each employing a gas turbine (also known as a "combustion turbine"). Among the various turbomachine configurations, turbojets are distinguished, in which the thrust used to propel the aircraft is generated by reaction to the ejection of hot gases.

[0003] Such a turbojet 1, as illustrated in Figure 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 in relation 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 The aircraft's propulsive efficiency, and therefore its architecture, is also affected. This propulsive efficiency also depends on the weather conditions in which the aircraft operates, which are now 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 temperature of the material that constitutes the turbine blades, 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] Examples of possible superalloy materials include nickel-based or cobalt-based superalloys.

[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 involves modifying the internal structure of the turbine blades to integrate cooling circuits, thereby reducing their temperature when operating in this environment. To achieve this, channels are machined into the blade to circulate a cooling fluid. This cooling fluid is generally drawn upstream of the low-pressure compressor 5 because its temperature is 500 to 1000 degrees Celsius lower than that of the hot gases.

[0011] Such a cooling circuit consists of a single cavity in direct contact with the outer surfaces 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 extrados and intrados walls without any structural obstruction.

[0012] However, the temperature at the mean surface line is lower than that of the intrados and extrados surfaces. The latter expand much more than the blade core and therefore subject the blade structure to mechanical stresses that can compromise 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 careful attention to manage the differential expansion 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 expansion leads to deformation, notably a bending of the wall towards the extrados. This is because the partition located at the double cavities within the blade's thickness is very cold compared to the hotter outer walls, which 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 as 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] Therefore, there is a need to maintain the mechanical integrity of the blade in operation while equipping it with an efficient cooling system that allows it to operate in a prolonged and recurring manner within the desired temperatures. Description of the invention

[0018] The present presentation relates to a single-crystal aircraft turbine component made of a superalloy, said component 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 component.

[0019] As mentioned above, gas turbines are subjected to extreme operating conditions, with high temperatures and pressures, which imposes stringent requirements on 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 could 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 turbine cooling and thermal management. Such a cavity allows the passage of a cooling fluid, thereby reducing the temperature of 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 for 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 and the upper surface; in a general case, these are the external surfaces of the single-crystal part).

[0023] This thermally conductive coating can be partially applied within the cavity, meaning it is not necessarily applied to the entire internal surface of the 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 metal 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 is worth noting 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 manufacturing. This allows the person skilled in the art to select the most appropriate alloy based on 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 pm.

[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 µm helps maintain the mechanical integrity of the part's material. Indeed, if the coating were too thick, it could create internal stresses or alter the part's mechanical properties, 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 avoiding, hotspots that could result from uncovered 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, as defined in claim 1, 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 mentioned 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. Therefore, the radial and longitudinal directions of the blade are considered synonymous in the remainder of this discussion.

[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 the said cooling cavity carved into the thickness of the platform.

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

[0039] As an example, these holes can have a diameter between 0.3 and 0.35 mm.

[0040] According to some 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 – in this case, the intrados and extrados walls.

[0042] These holes are also positioned along the leading edge, which is also a sensitive area, just like the intrados and extrados walls, 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 slits because it also changes in contact with hot gases.

[0044] The holes can 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] According to some embodiments of the invention, the thermal conductivity is greater than 70 W / m / K.

[0046] According to certain embodiments of the invention, the CuNiAg alloy has a thermal conductivity ranging from 110 to 160 W / m / K at 800°C, the AgNiCu alloy having a thermal conductivity ranging from 120 to 180 W / m / K at 800°C, and the CuNIAI alloy exhibiting a thermal conductivity ranging from 70 to 130 W / m / K at 800°C.

[0047] The present presentation also relates to a gas turbine comprising a plurality of single-crystal parts as defined above.

[0048] This presentation also relates to a turbomachine comprising a gas turbine as defined above.

[0049] This presentation also concerns a manufacturing process for 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; 2) a step of injecting wax into a first mold so as to create a wax model of the single-crystal part to be manufactured, the first mold containing said at least one ceramic core lined with the thermally conductive coating; 3) a step of covering the wax model with a refractory layer to create a second mold; 4) a step of removing the wax from the second mold; 5) a step of pouring the superalloy into the second mold and around said at least one ceramic core; 6) a step of removing the second mold; and 7) a step of detaching 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).

[0050] In other words, in step 1), the thermally conductive coating is applied to the external surface of a ceramic core or a plurality of cores. This core is intended to form an internal cavity that will be present in the final single-crystal part after its debossing.

[0051] Once the coating is applied, the next step involves injecting wax to create the wax model. This model determines the final shape of the piece and will serve as the basis for the subsequent steps.

[0052] After the wax model is created, it is covered with a refractory layer that withstands the high temperatures encountered during the casting of the superalloy. As it hardens around the wax model, the refractory coating creates a robust second mold that will maintain the shape of the part during the casting process.

[0053] Once the refractory layer has hardened, the wax is removed from the second mold. This step is usually done by heating, which melts the wax and allows it to be removed without damaging the refractory coating. Naturally, the blade's shell is fired before the superalloy is poured into it.

[0054] When the second mold is ready, the molten superalloy is poured into it, surrounding the ceramic core(s). Specifically, as the superalloy is poured, it comes into contact with the applied coating, enabling chemical interaction between the superalloy, the thermally conductive coating, and the core surface during the melting phase. It is worth noting that once the superalloy has cooled, the part is protected by the combination of elements diffused throughout its structure. In other words, the local chemistry of the part has been favorably modified, contributing to improved resistance to the high temperatures of the blade in its overall structure. Thus, no post-casting deposition process is necessary to protect the resulting single-crystal part.

[0055] The demolding of the core(s) and the second mold is then carried out after the superalloy has solidified (and cooled), which, among other things, exposes the internal cavity(ies). Once the core(s) are removed, the finished part can be inspected and tested to ensure it meets the required specifications. This final process yields a functional single-crystal part, ready for integration into applications found in aircraft turbines.

[0056] According to one embodiment of the invention, the thermally conductive coating is deposited by implementing a chemical vapor deposition process.

[0057] 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 this invention, the thermally conductive coating is deposited using this method.

[0058] Applying the thermally conductive coating via CVD allows access to 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

[0059] Other objects, features, and advantages of the invention will be better understood upon reading the detailed description below of various embodiments of the invention given by way of non-limiting example. This description refers to the accompanying figure pages, on which: - [Fig. 1] Figure 1 schematically presents a cross-sectional view of an aircraft turbomachine; - [Fig. 2] Figure 2 illustrates a perspective view of a single-crystal part according to the present invention; - [Fig. 3] Figure 3 illustrates a cross-sectional view of an embodiment of the internal structure of a blade as a single-crystal part, according to plane P of figure 2; - [Fig. 4] Figure 4 illustrates the different manufacturing stages of said single-crystal part according to the invention.

[0060] It should be noted that across all figures, common elements are identified by identical numerical references. Description of the implementation methods

[0061] To make the explanation more concrete, an example of a blade 11 (as a single-crystal component) is illustrated in Figure 2 and 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.

[0062] More specifically, the blade 11 is generally made of a nickel-based superalloy to withstand high thermal and mechanical stresses when in operation.

[0063] This material typically constitutes 80%, or even 90%, or even 99% of the structure of blade 11.

[0064] Of course, the nickel-based superalloy can be replaced by any other material deemed suitable by a person skilled in the art for manufacturing such a blade 11.

[0065] Furthermore, the blade 11 can be manufactured 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.

[0066] This manufacturing process makes it possible to obtain a blade with an aerodynamic profile designed to provide air passages between adjacent blades 11 without causing airflow turbulence detrimental to the efficiency of the turbine 8.

[0067] The blade 11 can be manufactured entirely or partially by additive manufacturing. By "partially," we mean 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.

[0068] The blade 11 includes a blade foot 12 which generally has a fir tree-shaped profile intended to be inserted into a rotor disc of the turbine 8 and thus be driven around its axis of rotation.

[0069] The blade foot 12 is provided with a platform 13 defining an external radial end of the blade foot 12 and supporting a blade 14 which extends radially between the blade foot 12 and an external radial end of the blade which is not visible in Figure 2. In other words, the blade 14 extends along a radial direction Y-Y' relative to the axis X-X' which also corresponds to the axis of rotation of the turbine 8 as illustrated in figure 1. The radial direction and the longitudinal direction of the blade 14 are then coincident.

[0070] 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.

[0071] 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.

[0072] Such areas are considered sensitive zones because their mechanical properties decrease as the temperature of the hot gases increases. They are then likely to expand more than the relatively cooler core of the turbine.

[0073] To overcome these disadvantages, the blade base 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.

[0074] The cooling fluid is usually 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.

[0075] 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 figure 3.

[0076] 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.

[0077] 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.

[0078] 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.

[0079] As an 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.

[0080] Furthermore, the cooling fluid circulation circuit 19 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 designed to pass through the blade platform 13 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.

[0081] Thus, when the connection is indirect, the cooling fluid first circulates in the blade foot 12 and then in said cooling cavity 22 hollowed out in the platform 13, which allows the platform 13 to be cooled, consequently contributing to lowering the average temperature of the blade 11.

[0082] Since the cooling cavity 22 is wider than the channels, it has a greater cooling capacity. It therefore serves two purposes: cooling the platform 13 and circulating the cooling fluid towards the blade 14.

[0083] 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.

[0084] The cooling fluid can therefore escape through the holes made on the blade 14 but also through the holes opening from the platform 13.

[0085] The holes drilled on the surface of platform 13 can have a diameter between 0.3 and 0.35 mm for example.

[0086] Figure 3 represents a section along plane P visible in Figure 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'.

[0087] Figure 3 shows a plurality of cooling circuits 26, here five, arranged radially between the lower surface wall 17 and the upper surface wall 18, and extending successively between the leading edge 15 and the trailing edge 16, and thus along a mean line Z-Z' (or axis Z-Z') between the lower surface wall 17 and the upper surface wall 18. The longitudinal mean 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 a symmetrical geometry, the longitudinal mean line Z-Z' corresponds to the chord of the blade's airfoil.

[0088] 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.

[0089] 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.

[0090] It is understood that the illustrated example is not exhaustive, 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.

[0091] To cool the blade 11, each cooling circuit 26 has its internal cavity C1 lined at least partially 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 within the internal cavity C1 and the sensitive areas of the part. The mechanical stresses that could be generated by differential expansion between the sensitive areas and the core of the blade are then limited, or even eliminated.

[0092] This thermally conductive coating 28 can be partially applied within cavity C1, meaning that it is not necessarily applied to the entire internal surface of the cavity. Some areas may be covered, while others may remain bare, exposed to the base material (the superalloy). Alternatively, the internal cavity C1 can be completely lined with the thermally conductive coating 28, as illustrated in Figure 3. The internal surface of the cavity C1 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.

[0093] 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 within the material of the part to a non-zero depth of less than 200 pm.

[0094] 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.

[0095] Furthermore, a coating depth of less than 200 µm helps maintain the mechanical integrity of the part's material. Indeed, if the coating were too thick, it could create internal stresses or alter the part's mechanical properties, particularly in terms of fatigue resistance.

[0096] It should be noted that the internal cavity C1 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.

[0097] Figure 4 illustrates the different stages of a process S0 for manufacturing said single-crystal part according to the invention, here a blade.

[0098] The process S0 begins with a first step S1 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 C1, as defined above, after its detachment.

[0099] The process S0 continues with a second 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.

[0100] 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.

[0101] The S0 process continues with a fourth step, S4 (not shown in the figure), which involves removing wax 30 from the second mold, M2. The shell of the developing blade is typically baked to achieve 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 subsequent step, S5, without deforming or breaking.

[0102] The S0 process 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, enabling 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 by 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 deposition process is required. post-casting is not necessary to protect the resulting single-crystal part.

[0103] Finally, the S0 process 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 S1 to the fifth step S5.

[0104] The S0 process may further include an optional seventh step, S7, comprising an inspection of the resulting single-crystal part to ensure it meets the required specifications. This final process yields a functional single-crystal part, ready for integration into applications encountered in aircraft turbines.

[0105] 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 turbine blades other than high-pressure turbomachinery turbine blades. 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.

[0106] 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 (C1) 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 (C1) 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 foot (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 foot (12) and, - a blade (14) extending from said platform (13) towards 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 (C1), 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 (C1) 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. A single-crystal part according to any one of the preceding claims, having a thermal conductivity greater than 70 W.m- 1 .K- 1 .

9. A single-crystal part according to any one of the preceding claims in combination with claim 2, wherein the CuNiAg alloy has a thermal conductivity of 110 to 160 W / m / K at 800°C, the AgNiCu alloy having a thermal conductivity of 120 to 180 W / m / K at 800°C, and the CuNIAI alloy exhibiting a thermal conductivity ranging from 70 to 130 W / m / K at 800°C.

10. Gas turbine (8) comprising a plurality of single-crystal parts according to any one of claims 1 to 9.

11. Turbomachine (1) comprising a gas turbine (8) according to claim 10.

12. A method for manufacturing a single-crystal part, according to any one of claims 1 to 9, 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 9, after its detachment; 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 removing the second mold; and 7) a step (S6) of uncoating 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 6).

13. A manufacturing method according to claim 12, wherein the thermally conductive coating (28) is deposited by implementing a chemical vapor deposition.

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