TURBOMACHINES BLADE WITH A COOLING CIRCUIT AND METHOD FOR PRODUCING SUCH A BLADE FROM LOST WAX

Description

Field of invention

[0001] The present invention relates to the field of turbomachines and in particular to a turbomachine blade equipped with a cooling circuit intended to cool it. It also relates to a lost wax manufacturing method for such a blade. Technical background

[0002] The technical background includes in particular documents US-A1-2019 / 316472, US-A1-2019 / 112942 and CN-A-101 007 337.

[0003] Turbomachine blades, especially high-pressure turbine blades, are subjected to very high temperatures that can reduce their lifespan and degrade the performance of the turbomachine. Indeed, the turbomachine turbines are arranged downstream of the combustion chamber of the turbomachine, which ejects a hot gas flow that is expanded by the turbines and allows them to rotate for the operation of the turbomachine. The high-pressure turbine, which is placed directly at the outlet of the combustion chamber, undergoes the highest temperatures.

[0004] In order to enable the turbine blades to withstand these severe thermal stresses, it is known to provide a cooling circuit in which circulates relatively colder air which is taken from the compressors, the latter being located upstream of the combustion chamber. More precisely, each turbine blade comprises a blade with a pressure side wall and an extrados side wall which are connected upstream by a leading edge and downstream by a trailing edge.

[0005] The cooling circuit generally comprises several cavities inside the blade, some of which communicate with each other and which are supplied with cooling air from the root of the blade, part of this cooling air opening into outlet orifices which are placed near the trailing edge. These orifices deliver jets of cooling air onto the walls of the blade.

[0006] It is known that the cooling circuit comprises several partitions extending radially in the blade so as to form "rising" and "descending" cavities arranged successively according to the direction of circulation of the cooling air and which communicate with each other by curved passages. These cavities and passages are known under the expression "trombone" circuit.

[0007] The cooling circuit cavities are generally made by at least one foundry core which is exploited in a blade manufacturing process using the lost wax casting technique.

[0008] Cooling circuit cavities are often equipped with turbulence promoters to increase heat exchange. A turbulence promoter is an element projecting into the cavity whose function is to generate disturbances and turbulence in the air flow circulating in the cavity, in order to increase heat exchange between this air flow and the walls of the cavity (see for example document FR-A1-3 065 985).

[0009] Common turbulence promoters are of different types but historically constrained by the demolding capacity required by the cores used for cavity formation and turbulence promoters.

[0010] Common shapes include simple disruptors, straight or inclined, in chevrons, simple protrusions or hollows in a cylindrical or teardrop shape, or even bridges that cross cavities and connect two opposite walls (typically intrados and extrados).

[0011] These disruptors, depending on their type, spacing and geometric dimensions, have different thermal efficiencies, but this efficiency remains proportional to the pressure losses generated: in a pressure loss / thermal efficiency diagram, all these geometries are essentially aligned. However, since the overpressure rate available to circulate air in the circuit is limited, there is a strong interest in increasing the heat exchange / pressure loss ratio.

[0012] The heat exchanges obtained are also not very homogeneous.

[0013] It is, moreover, difficult to increase heat exchanges on one face (typically a hot wall), while limiting heat exchanges on the opposite face (typically a cold internal partition), because these exchanges unnecessarily increase pressure losses and increase the temperature gradient between the cold internal partition and the hot external wall, which is detrimental to the mechanical strength of the blade.

[0014] There is therefore a need to define a technology that provides a solution to at least some of these problems. Summary of the invention

[0015] The objective of the present invention is to improve the heat exchange ratio on the hot walls / pressure losses in the cavity, by using a new configuration of turbulence promoters.

[0016] We achieve this objective in accordance with the invention by means of an aircraft turbomachine blade, this blade comprising a blade and a cooling circuit inside the blade, this cooling circuit comprising at least one longitudinal cavity for the flow of a cooling air flow, the cooling circuit further comprising elements which project into said cavity and which are configured to disturb said air flow, characterized in that each of said projecting elements has a general arch shape and comprises two lateral legs and a median roof, the legs extending between a first wall of the cavity and said roof, this roof connecting said legs together, and in that each of said elements defines internally, with said first wall, a first passage section, and externally, with a second wall opposite the first wall, a second passage section,and each of said elements being configured so that said first or second passage section is reduced from upstream to downstream relative to the direction of said air flow.,

[0017] The present invention thus proposes to use arch-shaped projecting elements. These projecting elements have the particularity of allowing a reduction in the air passage section inside or outside them, so as to accelerate the air which passes through this passage section. Increasing the air flow speed on the wall to which a projecting element is connected, or on the opposite wall, makes it possible to increase the heat exchanges between the air and this wall and therefore to optimize the cooling of the blade.

[0018] The principle of the invention is therefore to increase heat exchanges by acting on the air flow speed rather than on turbulence. Indeed, turbulence increases pressure losses at the same time as heat exchanges while acceleration increases heat exchanges while refining the boundary layer and thus reducing pressure losses. In addition, the establishment of turbulence requires the implementation of several promoters before reaching an optimal level for increasing heat exchanges and cannot, therefore, be used for very local cooling.

[0019] The dawn also includes one or more of the following characteristics, taken alone or in combination: the roof comprising a first face which faces the first wall, and a second face which faces the second wall, at least one of the first and second faces being inclined relative to the facing wall, from upstream to downstream relative to the direction of said air flow, so as to reduce the corresponding passage section; said roof of each of said projecting elements is inclined relative to said first and second walls; said first and second faces are substantially parallel to each other; said projecting elements are arranged one behind the other inside said cavity and form a row of projecting elements extending along said direction of said air flow, the row comprising for example between 5 and 10 projecting elements; said legs of each of said projecting elements are separated from each other by a distance representing more than 40% of a width of said cavity;said legs of each of said projecting elements are oriented convergently from upstream to downstream; said legs and said roof of each of said projecting elements have upstream edges which are convexly rounded; said roof of each of said projecting elements is located substantially at half height in said cavity, this height being measured between said first and second walls; each of said projecting elements defines an internal air passage section which has a generally rectangular, circular, oblong or trapezoidal shape; the reduction in the passage section from the inlet to the outlet of each of the projecting elements is of the order of at least 10%; the passage section inside and at the inlet of each of the projecting elements may represent at least 25% of the total passage section of the cavity;each of the projecting elements has a length noted L3 and has an entrance having a height noted h, the length L3 being between 0.5.h and 3.h.;

[0020] The present invention also relates to a method for the lost wax manufacturing of a blade as described above, characterized in that it comprises the following steps: providing a refractory ceramic or metallic core for the formation of said cavity, this core having an elongated shape and comprising internal U-shaped conduits which are configured to form by molding said projecting elements, each of these conduits each comprising two ends which open onto the same face of the core configured to form by molding said first wall, injecting wax so as to coat the core and to form a model, manufacturing a shell enveloping the model, pouring a molten metal inside the shell so as to form the blade, and unsticking the shell and the core so as to release the blade and to form the cooling circuit with its cavity and its projecting elements.

[0021] The invention further relates to a turbomachine turbine comprising at least one turbomachine blade having the aforementioned characteristics.

[0022] The invention further relates to a turbomachine comprising at least one turbomachine turbine as mentioned above. Brief description of the figures

[0023] The invention will be better understood, and other aims, details, characteristics and advantages thereof will appear more clearly on reading the detailed explanatory description which follows, of embodiments of the invention given as purely illustrative and non-limiting examples, with reference to the appended schematic drawings in which: [ Fig. 1 ] There figure 1 is a schematic view in axial and partial section of an example of a turbomachine to which the invention applies; [ Fig. 2 ] There figure 2 is a schematic view of an axial section of a turbomachine blade with a cooling circuit according to the invention; [ Fig. 3 ] There figure 3is a schematic cross-sectional view of a blade of a turbomachine blade comprising a cooling circuit with various cavities; [ Fig. 4 ] There figure 4 is a perspective view of projecting elements which are located in a cavity of a cooling circuit of a blade according to the invention; [ Fig. 5A-5B ] THE Figures 5A and 5B are respectively schematic views of a projecting element, seen from above in a cavity, and seen in section along line BB of the Figure 5A ; [ Fig. 5C-5D ] THE Figures 5C and 5D are respectively schematic views of the projecting element, seen in section along line CC of the Figure 5A , seen in section along line DD of the Figure 5A ; [ Fig. 6A-6B ] There Figure 6A is a schematic sectional view of a cavity of a cooling circuit of a blade according to the invention, the sectional plane passing through legs of the projecting elements located in this cavity, and the Figure 6Bbeing a larger scale view of a detail of the Figure 6A ; [ Fig. 7A-7B ] There Figure 7A is a schematic sectional view of a cavity of a cooling circuit of a blade according to the invention, the sectional plane passing through roofs of the projecting elements located in this cavity, and the Figure 7B being a larger scale view of a detail of the Figure 7A ; [ Fig. 8 ] There figure 8 is a schematic view of a ceramic core for molding the cavity of a blade according to the invention; and [ Fig. 9 ] There figure 9 is a schematic sectional view along line IX-IX of the figure 8 . Detailed description of the invention

[0024] There figure 1shows a partial axial sectional view of a turbomachine 1 with longitudinal axis X to which the invention applies. The turbomachine shown is a dual-flow, dual-spool turbomachine intended to be mounted on an aircraft according to the invention. Of course, the invention is not limited to this type of turbomachine.

[0025] This double-flow turbomachine 1 generally comprises a fan 2 mounted upstream of a gas generator 3. In the present invention, and generally, the terms “upstream” and “downstream” are defined in relation to the circulation of gases in the turbomachine and here along the longitudinal axis X (and even from left to right on the figure 1). The terms "axial" and "axially" are defined with respect to the longitudinal axis X. Similarly, the terms "radial", "internal" and "external" are defined with respect to a radial axis Z perpendicular to the longitudinal axis X and with respect to the distance from the longitudinal axis X.

[0026] The gas generator 3 comprises, from upstream to downstream, a low pressure compressor 4a, a high pressure compressor 4b, a combustion chamber 5, a high pressure turbine 6a and a low pressure turbine 6b.

[0027] The fan 2, which is surrounded by a fan casing 7 carried by a nacelle 8, divides the air entering the turbomachine into a primary air flow which passes through the gas generator 3 and in particular in a primary vein 9, and into a secondary air flow which circulates around the gas generator in a secondary vein 10.

[0028] The secondary air flow is ejected by a secondary nozzle 11 terminating the nacelle while the primary air flow is ejected outside the turbomachine via an ejection nozzle 12 located downstream of the gas generator 3.

[0029] The high-pressure turbine 6a, like the low-pressure turbine 6b, comprises one or more stages. Each stage comprises a stator blade mounted upstream of a moving blade. The stator blade comprises a plurality of stator or fixed blades, called distributors, which are distributed circumferentially around the longitudinal axis X. The moving blade comprises a plurality of moving blades which are equally distributed circumferentially around a disc centered on the longitudinal axis X. The distributors deflect and accelerate the aerodynamic flow leaving the combustion chamber towards the moving blades so that they are driven in rotation.

[0030] In reference to the figures 2 And 3, each turbine blade (and here a high-pressure turbine moving blade 20) comprises a blade 21 rising radially from a platform 22. The latter is carried by a root 23 which is intended to be implanted in one of the corresponding grooves of the turbine disk. Each blade 21 comprises a lower surface wall 24 and an upper surface wall 25 which are connected upstream by a leading edge 26 and downstream by a trailing edge 27. The lower surface 24 and upper surface 25 walls are opposite along a transverse axis T which is perpendicular to the longitudinal X and radial Z axes.

[0031] Increased performance requirements and evolving aeronautical regulations are pushing current engine manufacturers to design engines operating in increasingly harsh environments (temperature, pressure, rotational speed, emissions, etc.). This leads to the need to define "new generation" high-pressure turbine blades that can withstand this type of stress.

[0032] Increasing the temperature of the gas driving the blade improves the efficiency of the turbomachine. This temperature increasingly exceeds the melting temperature of the superalloy constituting the blade by several hundred degrees. The blade must therefore be cooled more and more efficiently.

[0033] For this, the blade 20 comprises a cooling circuit 28 which is arranged inside the blade 21 and which is intended to cool the walls of the blade undergoing the high temperatures of the primary air flow leaving the combustion chamber 5 and passing through it. This cooling circuit 28 comprises several cavities which communicate with each other so as to form a "trombone" type duct. The latter comprises several passages or turns so that a cooling fluid, here cooling air, sweeps the entire blade and from top to bottom along the radial axis.

[0034] The root 23 comprises a supply channel 30 which comprises a cooling air inlet 31 taken upstream of the combustion chamber 5 such as on the low pressure compressor 4a, and which opens into the trombone-type duct. The channel 30 also opens onto a radially internal face 32 of the root of the blade which comprises the cooling air inlet 31. The cooling circuit 28 also comprises outlet orifices 33 which are arranged in the vicinity of the trailing edge 27 of the blade. The outlet orifices 33 are oriented substantially along the longitudinal axis X and are aligned and regularly distributed substantially along the radial axis. In this way, the cooling air RF which circulates from the root of the blade, passes through the cavities inside the blade and opens into the outlet orifices 33.

[0035] As illustrated in detail on the figure 3, the cooling circuit 28 comprises several cavities arranged successively from upstream to downstream of the blade 21. In particular, a first cavity 34 and a second cavity 35 each extend along the radial axis in the blade. The second cavity 35 is arranged downstream of the first cavity 34 in the direction of circulation of the cooling air (and from upstream to downstream along the longitudinal axis X). The first cavity 34 and the second cavity 35 are separated, at least in part, by a first radial partition 36 which has a radially internal free end 37. The latter is located at the level of the connecting end 38 of the root of the blade (radially opposite the free end 39 of the blade). The free end 39 of the blade further comprises a closing wall (not shown) which makes it possible to contain the cooling air inside the blade for its cooling.The first partition 36 is connected to the closing wall at its radially outer end (opposite its radially inner free end 37).

[0036] The cooling circuit 28 also comprises a third cavity 42 which extends radially inside the blade. The third cavity 42 is arranged upstream of the first cavity 34 in the direction of circulation of the cooling air RF. The third cavity 42 is separated at least in part from the first cavity 34 by a second radial partition 43 which comprises a radially external free end 44. The third cavity 42 and the first cavity 34 are connected by a second passage 45 for cooling fluid which is delimited at least in part by the radially external free end 44. The closing wall also delimits the second passage 45.

[0037] The 34, 35 and 42 arranged successively according to the direction of circulation of the cooling fluid form the trombone type conduit.

[0038] The blade 21 may include another cooling circuit 46 which also allows the blade to be cooled. The cooling circuit 46 includes a lower surface cavity 47 which extends radially inside the blade. The lower surface cavity 47 serves to specifically cool the lower surface wall and the upper part of the blade along the radial axis. The air which is injected into this cavity can exit the blade through the outlet orifices 33 or through other orifices which would be located on the lower surface wall for example. As we can see on the figure 3, the intrados cavity 47 extends transversely between the internal wall 41 and the intrados wall 24. The latter also extends longitudinally in the direction of circulation of the air flow between the cavities 35, 42. In other words, the cavity 35 transversely covers the cavities 34 and 47. Its length is substantially identical to that of the first cavity 34 in the direction of circulation of the cooling air (axial direction).

[0039] Upstream of the third cavity 42 is arranged an upstream cavity 48 of another cooling circuit of the blade 21, this cavity 48 extending radially along the leading edge 26.

[0040] It is clear from the above that a turbomachine blade may comprise one or more internal cooling circuits and that each of these circuits may comprise one or more cavities for circulating a cooling air flow.

[0041] It is known to provide turbulence promoters in at least one of these cavities. These turbulence promoters are elements projecting from a wall of a cavity for circulating an air flow, which have the function of generating disturbances and turbulence in this air flow in order to increase the heat exchanges between this air flow and the wall of the cavity.

[0042] The present invention proposes a new advantageous configuration of these projecting elements which are in the form of arches 50, as can be seen in the figure 4 . Each arch 50 or projecting element comprises two lateral legs 52 and a medial roof 54. The legs 52 extend between a first wall 56 of the cavity 58 and the roof 54, and this roof 54 extends between the legs 52 and connects them together.

[0043] As can be seen in the Figures 5A to 5D, each arch 50 defines internally, with the first wall 56 to which it is connected, a first passage section S1-S1', and externally, with a second wall 60 opposite the first wall 56, a second passage section S2-S2'. The passage section is denoted S1 inside and at the inlet or upstream of the arch 50, S1' inside and at the outlet or downstream of the arch, S2 outside and at the inlet of the arch, and S2' outside and at the outlet of the arch. Each arch 50 is configured so that the first or second passage section S1-S1', S2-S2' is reduced from upstream to downstream relative to the direction of the air flow (arrow RF). In other words, each arch 50 is configured so that the passage section S1' at the exit of the arch is less than the passage section S1 at the entrance of the arch, or that the passage section S2' at the exit of the arch is less than the passage section S2 at the entrance of the arch.Each passage section has a general shape: rectangular, circular, oblong, trapezoidal, etc.

[0044] Reducing the passage section makes it possible to increase the speed of the air flow. Thus, when the passage section is reduced inside the arch 50 and therefore on the side of the first wall 56, the air flow which flows over this first wall 56 is accelerated, which makes it possible to increase the heat exchanges between this air flow and the wall 56.

[0045] When the passage section is reduced outside the arch 50, on the side of the second wall 60, the air flow which flows over this second wall is accelerated, which makes it possible to increase the thermal exchanges between this air flow and the wall 60.

[0046] The reduction of the passage section can be obtained by varying the dimensions of the arch 50, and in particular the widths and / or heights of the entry and exit of the arch, the length as well as the thickness of the legs 52 and roof 54 of the arch 50, etc. The distance between two consecutive arches 50 can also make it possible to modulate the level of exchanges.

[0047] The roof 54 comprises a first face 54a which faces the first wall 56 on which the arch 50 or the element projects. The roof 54 comprises a second face 54b which faces the second wall 60 and opposite the wall 54a.

[0048] Depending on the cooling requirements and manufacturing constraints, the arch 50 can therefore be positioned on the wall 56 located on the intrados or extrados side of the blade 21. The wall 56 is then cooled by adjusting the speed, and the opposite wall 60 is also cooled but by optimizing the vortices and therefore the turbulence in the air flow RF. Depending on the cooling requirements of the wall 60, the intensity of the vortices generated at the upper corners 70 of the arch 50 can for example be modulated by adjusting for example the slope of the arch 50 (difference in inlet / outlet height relative to its length) or, more generally, on the shape of its generator. In the figures, the arch 50 has in axial section ( Figure 5D) a rectangular shape whose corners are radiated and whose dimensions decrease linearly to form a convergent, but this section could be an arc of a circle, an ellipse, sinusoidal or any other shape depending on the type of flow desired on the wall 60. The inlet and outlet sections S1, S1', S2, S2' can thus vary, as mentioned above, and the convergence can be non-linear, so as to minimize pressure losses for example.

[0049] Alternatively, the arch 50 may be configured to accelerate the air flow at the wall 60 and not at the wall 56. In this case, the arch 50 still forms a convergent in the direction of flow but on the side of this wall 60.

[0050] According to the embodiment of the invention illustrated in the figure 4and in Figures 6A to 6D, at least one of the faces 54a, 54b is inclined relative to the facing wall 56, 60, from upstream to downstream relative to the direction of the air flow RF, so as to reduce the corresponding passage section S1-S1', S2-S2'.

[0051] The inclination of one of the faces 54a, 54b or both faces of the roof 54 may be achieved by an inclination of the roof 54, a variation in thickness of the roof, or a combination of both.

[0052] In the example shown, the faces 54a, 54b are substantially parallel to each other. The face 54a is inclined from upstream to downstream towards the wall 56 and therefore reduces the first passage section S1-S1' inside the arch 50, between the inlet and the outlet of the arch, which causes an acceleration of the air flow on the wall 56.

[0053] The legs 52 and the roof 54 of each arch 50 have upstream edges 62 which are rounded convex in the example shown in figure 4. The legs 52 may also have downstream edges 64 which are convexly rounded. The ends of the legs 52, opposite the roof 54, may be connected to the wall 56 by fillets 66. The legs 52 are connected to the roof 54 by the corners 70 which are curved and devoid of sharp edges. All of these characteristics make it possible to improve the aerodynamics of the arch 50 so as to ensure its function of accelerating the air flow passing inside the arch while generating the least possible disturbance in this air flow.

[0054] The arches 50 are preferably arranged one behind the other inside the cavity 58 and form a row of arches. Each row comprises, for example, between 5 and 10 projecting elements, as illustrated in Figures 6A And 7A .

[0055] THE Figures 6A and 6Ballow it to be seen that the legs 52 can also be inclined relative to each other as well as relative to the side walls 72 of the cavity 58. The legs 52 converge, for example, towards each other from upstream to downstream so as to accentuate the reduction in the passage section.

[0056] The legs 52 are located in the immediate vicinity of the side walls 68 in the example shown. The legs 52 are separated from each other by a minimum distance L1 representing at least 40%, or even 60%, of a width L2 of the cavity 58.

[0057] THE Figures 7A and 7B allow the inclination of the roofs 54 of the arches 50 of this embodiment to be observed. The roofs 54 are located substantially halfway up in the cavity 58 in the example shown.

[0058] THE Figures 6A to 7Balso allow us to observe the zones Z1 of acceleration of the air flow inside the arches 50 and the overall acceleration of the air flow in the cavity 58 (arrow RF). It is also noted that the equilibrium level of heat exchange is reached from the first arch 50. Furthermore, the level of exchange is much more homogeneous over the width L2 of the cavity 58. The inventors have noted that with this geometry, the heat flow increases by 5.1% while the pressure drop only increases by 2.7% compared to conventional turbulence promoters. The arches 50 generate controlled vortices Z2 which also cool the opposite wall 60 much more efficiently than conventional promoters.

[0059] The reduction in the passage section inside or outside the arch 50, as mentioned above, may be of the order of at least 10%. In the case, for example, where it is the internal passage section S1-S1' of the arch 50 which is reduced, this means that the ratio S1' / S1 is less than or equal to 0.9.

[0060] The passage section inside and at the entrance of the arch 50 may represent at least 25% of the total passage section of the cavity. In other words, the ratio S1 / (S1+S2) is at least equal to 0.25.

[0061] The arch has a length noted L3 and its entrance has a height noted h. Preferably, the length L3 is between 0.5.h and 3.h (cf. Figure 5D ).

[0062] These different parameters can be optimized according to the other dimensions of the cavity 58 and the arch 50 (thickness, etc.) in order to control the variation of the section outside the arch and therefore the vortices which can be created there. Thus, if the wall 60 opposite that of the arch 50 does not require additional cooling, we will seek to limit the section expansion outside the arch to less than 30% in order to limit separations and therefore pressure losses. Conversely, if the cooling of the opposite wall must be improved, a greater expansion outside the arch could be interesting, the section convergence under the arch could then be increased (while avoiding too strong an acceleration leading to reaching Mach 1 for example in the cavity, which would generate high pressure losses).

[0063] The thickness of the walls of the arch 50 must be as thin as possible while allowing manufacturability and ensuring the mechanical strength of the blade 20 in operation. Preferably, the inlet material section of the arch 50 must not represent more than 40% of the total section of the cavity.

[0064] Advantageously, but not limited to, the blade 20 is made from a metal alloy and according to a manufacturing process using the lost wax casting technique or lost model. The metal alloy is preferably nickel-based and may be monocrystalline.

[0065] This method comprises a first step of manufacturing one or more foundry cores. In the present example, the blade comprising a blade provided with at least one cooling air circulation cavity is produced from a foundry core, one embodiment of which is shown in figures 8 and 9 .

[0066] A foundry core is conventionally obtained by ceramic injection (ceramic which is then debinded and sintered). The core used in the context of the present invention cannot, however, be produced by injection because it is not easily demolded.

[0067] This core can, however, be obtained by additive manufacturing of ceramic or any other suitable material such as refractory metals.

[0068] In the case of figures 8 and 9, the ceramic core 74 has an elongated shape and comprises internal U-shaped conduits 76 which are configured to form by molding the projecting elements, i.e. the arches 50. Each of these conduits 76 comprises two ends 76a which open onto the same face 78 of the core configured to form by molding the aforementioned first wall 56. The core 74 comprises a face 80 opposite the face 78 and configured to form by molding the aforementioned wall 60. The lateral faces 82 of the core 74 are configured to form by molding the aforementioned lateral walls 72.

[0069] In another step of the process, wax or an equivalent material is injected around the core 74 or even a set of cores, which are previously arranged, advantageously, but not limited to, in a press. Once the wax has cooled, we obtain a model comprising the cores embedded in the wax.

[0070] The model is arranged in a column with other similar models to form a cluster.

[0071] The method further comprises producing a shell in a refractory material around the cluster and which acts as a mold. The refractory material in this example is a ceramic. The shell is produced by immersing the cluster several times in a ceramic slip.

[0072] In another step of the process, molten metal is poured or cast inside the shell to fill the cavities created when the wax was removed from the models and intended to form the metal parts, in this case the turbine blades. In fact, prior to this metal pouring step, a wax removal step is carried out.

[0073] When the shell is cooled and solidified, a knockout step destroys the shell and the cores in the metal parts (blade) so as to reveal the final blade and the cooling fluid circulation cavities.

[0074] The present invention provides several advantages including: an improvement in the heat exchange / pressure drop ratio in the cooling cavity of the blade, an improvement in the volume ratio of the projecting element and therefore of the arch (and therefore of the mass of the blade) / exchange surface, better control of the exchange improvement zone and better homogeneity of the exchanges in this zone, a potential reduction in the number of projecting elements (and therefore of the mass of the blade) to cool the critical zones, possible cooling of the opposite wall if necessary by adjusting the shape of the projecting element (in particular the slope of the external face of its roof to increase or reduce the intensity of the vortices released at the corners of the arch), etc.

Claims

1. A vane (20) for an aircraft turbomachine, this vane comprising a blade (21) and a cooling circuit (28) inside the blade, this cooling circuit (28) comprising at least one longitudinal cavity (58) for the flow of a cooling air stream (RF), the cooling circuit further comprising projecting elements into said cavity and which are configured to disrupt said air stream, characterised in that each of said projecting elements has the general shape of an arch (50) and comprises two lateral legs (52) and a median roof (54), the legs (52) extending between a first wall (56) of the cavity (58) and said roof (54), this roof (54) interconnecting said legs (52), and in that each of said elements defines internally, with said first wall (56), a first passage cross-section (S1-S1'), and externally, with a second wall (60) opposite the first wall (56), a second passage cross-section (S2-S2'), and each of said elements being configured so that said first or second passage cross-section is reduced from upstream to downstream with respect to the direction of said air stream.

2. The vane (20) according to claim 1, characterised in that the roof (54) comprises a first face (54a) which faces the first wall (56), and a second face (54b) which faces the second wall (60), at least one of the first and second faces (54a, 54b) being inclined relative to the facing wall (56, 60), from upstream to downstream relative to the direction of said air stream (RF), so as to reduce the corresponding passage cross-section (S1-S1', S2-S2').

3. The vane (20) according to the preceding claim, characterised in that said first and second faces (54a, 54b) are substantially parallel to each other.

4. The vane (20) according to one of the preceding claims, characterised in that said roof (54) of each of said projecting elements is inclined with respect to said first and second walls (56, 60).

5. The vane (20) according to any one of the preceding claims, characterised in that said projecting elements are arranged one behind the other inside said cavity (58) and form a row of projecting elements extending along said direction of said air stream, the row comprising for example between 5 and 10 projecting elements.

6. The vane (20) according to any one of the preceding claims, characterised in that said legs (52) of each of said projecting elements are separated from each other by a distance (L1) representing more than 40% of a width (L2) of said cavity (58).

7. The vane (20) according to any one of the preceding claims, characterised in that said legs (52) of each of said projecting elements are oriented in a convergent manner from upstream to downstream.

8. The vane (20) according to any one of the preceding claims, characterised in that said legs (52) and said roof (54) of each of said projecting elements comprise upstream edges (62) which are convexly rounded.

9. The vane (20) according to any one of the preceding claims, characterised in that said roof (54) of each of said projecting elements is located substantially at half a height in said cavity (58), this height being measured between said first and second walls (56, 60).

10. The vane (20) according to any one of the preceding claims, characterised in that each of said projecting elements defines an internal air passage cross-section (S1-S1') which has a generally rectangular, circular, oblong or trapezoidal shape.

11. The vane (20) according to any one of the preceding claims, characterised in that the reduction in the passage cross-section from the inlet to the outlet of each of the projecting elements is of the order of at least 10%.

12. The vane (20) according to any one of the preceding claims, characterised in that the passage cross-section inside and at the inlet of each of the projecting elements can represent at least 25% of the total passage cross-section of the cavity.

13. The vane (20) according to any one of the preceding claims, characterised in that each of the projecting elements has a length noted L3 and has an inlet with a height noted h, the length L3 being between 0.5.h and 3.h.

14. A turbomachine turbine comprising at least one vane (20) according to any one of the preceding claims.

15. A method for manufacturing lost-wax of a vane (20) according to any one of claims 1 to 13, characterized in that it comprises the following steps: - providing a refractory ceramic or metallic core (74) for forming said cavity (58), this core having an elongate shape and comprising internal U-shaped ducts (76) which are configured to form said projecting elements by moulding, each of said ducts (76) each comprising two ends (76a) which open onto a same face (78) of the core configured to form said first wall (56) by moulding, - injecting wax so as to coat the core (74) and form a model, - manufacturing a shell enveloping the model, - pouring molten metal into the shell to form the vane, and - stripping the shell and core (74) so as to release the vane (20) and form the cooling circuit (28) with its cavity (58) and projecting elements.