Turbine blade including three types of orifices for cooling the trailing edge

Abstract

A blade (18) for a turbine (7) of a turbomachine, the blade having a cooling zone (ZF) comprising three different types of cooling orifices (25, 26, 27), the cooling orifices being formed by the wall of an airfoil (30) and being radially spaced apart from one another between a platform (28) and a tip (S), a trailing edge connecting chamfer (17) having the general shape of a dihedral (DI), the dihedral being characterized by a radius (R) of the trailing edge connecting chamfer (17) and a spacing (E) between a pressure face wall (15) and a suction face wall (16) opposite to a trailing edge (BF), the spacing (E) and the radius (R) of the trailing edge connecting chamfer (17) varying radially between the platform (28) and the tip (S) of the blade (18).

Description

Technical Field

[0001] The present invention relates to the general field of cooling turbine blades of turbomachines and, more particularly, to the field of cooling moving turbine blades for turbomachines.

[0002] The invention is applicable to any type of aviation or terrestrial turbine. The invention can be particularly applicable to moving turbine blades of aircraft turbines, such as turbojets and turboprops, such as turbofans. The invention can also be applied to moving blades of industrial gas turbines. Preferably, the invention also relates to the cooling of high-pressure moving turbine blades, but is also applicable to the cooling of low-pressure and / or medium-pressure moving turbine blades. Background Art

[0003] In order to design more efficient and less consuming engines, turbine blades have developed into smaller and smaller sizes and are able to withstand higher and higher thermal and mechanical stresses such as temperature, pressure, rotation speed, etc.

[0004] As is known to all, the blades of turbomachine gas turbines (especially high pressure turbines) are subjected to high temperatures of combustion gases during engine operation. These temperatures reach values ​​far higher than the various components in contact with these gases can withstand without being damaged, which limits the service life of the components.

[0005] It is also known that an increase in the gas temperature of the high-pressure turbine makes it possible to increase the efficiency of the turbine and thus the ratio between the thrust of the engine and the weight of the aircraft propelled by this turbine. Therefore, efforts have been made to manufacture turbine blades capable of withstanding increasingly higher temperatures. Furthermore, improvements in turbine blades are aimed at reducing the fuel consumption of the engine.

[0006] One of the solutions currently available for improving blades (in particular improving the mechanical strength of blades) is to reduce the operating temperature of the blades more effectively by cooling. This cooling is obtained in particular by a cooling circuit arranged in the blades, which is intended to reduce the temperature of the blades. Due to this circuit, cooling air is usually introduced into the blades through the roots of the blades, and the cooling air passes through the blades along a path formed by a cavity made in the blades before being ejected through an orifice (or a bore), which opens on the surface of the blades, in particular at the trailing edge of the blades. Improving the cooling circuit of the blades makes it possible to reduce the air flow required to cool the blades, and also makes it possible to increase the service life of the blades and / or ensure that the service life targets are met in the event of an increase in the temperature at the inlet of the turbine stage.

[0007] Furthermore, the blades can also be modified by improving the aerodynamic profile to increase the efficiency of the blades.

[0008] Solutions aimed at improving the cooling and aerodynamic profile at the trailing edge of a moving blade of a high-pressure turbine have been described in the prior art. For example, French patent application FR 3 041 989 A1 discloses cooling the trailing edge of a high-pressure turbine blade by three different cooling zones. French patent application FR 2 864 990 A1 describes a solution for improving the cooling air discharge slots at the trailing edge of a high-pressure turbine blade.

[0009] However, there is still a need to further improve the cooling efficiency at the trailing edge of the turbine blade, in particular to ensure the mechanical strength of the turbine blade, and to provide the turbine blade with an improved aerodynamic profile so that the aerodynamic performance can be maximized. Summary of the invention

[0010] It is therefore an object of the present invention to at least partially overcome the above mentioned needs and disadvantages with respect to the embodiments of the prior art.

[0011] Therefore, according to one aspect of the present invention, an object of the present invention is to provide a turbine blade for a turbine, which is intended to be mounted around a rotation axis, the turbine blade comprising a platform (particularly an inner platform) and a blade, the blade extending in a radial direction relative to the rotation axis, extending from the inside to the outside away from the platform and terminating at a top, the blade comprising a leading edge and a trailing edge located downstream of the leading edge, and the blade comprising a lower surface wall and an upper surface wall, the lower surface wall and the upper surface wall each connecting the leading edge to the trailing edge, the lower surface wall and the upper surface wall being connected to each other at the trailing edge by a trailing edge chamfer,

[0012] Characterized in that the turbine blade of a turbine comprises a cooling area extending along the trailing edge, the cooling area comprising three different types of cooling apertures, the cooling apertures being formed through the wall of the blade to enable a cooling flow to circulate from the inside to the outside of the blade, the three types of cooling apertures being radially spaced apart from each other between the platform and the top,

[0013] the blade comprises, in a region at the root of the blade, close to the platform, a first type of cooling apertures in the form of first cooling apertures formed in the thickness of the trailing edge chamfer between the lower surface wall and the upper surface wall,

[0014] the blade comprises, in a region at the top of the blade, close to the top, a third type of cooling aperture in the form of a third cooling aperture, the third cooling aperture being different from the first aperture, the third cooling aperture being formed in the thickness of at least one of the lower surface wall and the upper surface wall,

[0015] the blade comprises, in a region at the middle of the blade, between a region at the root of the blade and a region at the tip of the blade, a second type of cooling aperture in the form of a second cooling aperture, the second cooling aperture being different from the first aperture and the third aperture, the second cooling aperture being formed in the thickness of the at least one of the lower surface wall and the upper surface wall,

[0016] The cooling area, lower surface wall, upper surface wall and trailing edge chamfer of the trailing edge generally have the shape of a dihedral angle, the shape of the dihedral angle is characterized by the radius of the trailing edge chamfer and the spacing between the lower surface wall and the upper surface wall opposite to the trailing edge, and the spacing between the lower surface and the upper surface and the radius of the trailing edge chamfer vary radially between the platform and the top of the blade.

[0017] In other words, the pitch and the value of the radius of the trailing edge chamfer vary in the radial direction.Thus, the pitch and the value of the radius of the trailing edge chamfer are not constant over the entire height of the blade, between the platform and the tip.

[0018] Thanks to the invention, it is possible to have a triple cooling configuration at the trailing edge of a turbine blade for a turbine and to obtain an improved aerodynamic profile of the blade. In particular, it is thus possible to reduce the fuel consumption of the turbine engine by reducing the flow required for cooling the blades, in particular the blades of the high-pressure turbine, while the blades have the same service life as turbine blades integrated with a conventional cooling circuit. In addition, by allowing a dihedral portion with dimensions adjusted according to the height to be defined at the trailing edge, the aerodynamic efficiency, in particular the aerodynamic efficiency of the high-pressure turbine stage, can be improved compared to blades incorporating conventional trailing edge cooling techniques. In addition, the service life of the blades, in particular the high-pressure turbine blades, can be increased compared to blades integrated with conventional solutions for cooling the trailing edge, such as slots at the trailing edge.

[0019] The turbine blade according to the invention may also include one or more of the following features, which are employed individually or in combination in any technically possible manner.

[0020] Preferably, the second cooling apertures and the third cooling apertures may be formed in the thickness of the lower surface wall.

[0021] Furthermore, the first cooling apertures, the second cooling apertures and the third cooling apertures may each comprise a different type of cooling aperture selected from a borehole having a circular cross section, a borehole having an elliptical cross section or a borehole in the shape of a slot.

[0022] In particular, the first cooling aperture may comprise a bore, in particular a bore having a circular cross section and / or an elliptical cross section, the second cooling aperture may comprise one of a cooling groove, in particular a cooling groove having a rectangular cross section or an elliptical cross section, and a bore, in particular a bore having a circular cross section, and the third cooling aperture may comprise the other of a cooling groove and a bore.

[0023] In particular, the first cooling aperture may comprise a bore, in particular a bore with a circular cross section and / or an elliptical cross section, the second cooling aperture may comprise a cooling slot, in particular a cooling slot with a rectangular cross section or an elliptical cross section, and the third cooling aperture may comprise a bore, in particular a bore with a circular cross section.

[0024] Alternatively, the first cooling aperture may comprise a bore, in particular a bore with a circular cross section and / or an elliptical cross section, the second cooling aperture may comprise a bore, in particular a bore with a circular cross section, and the third cooling aperture may comprise a cooling groove, in particular a cooling groove with a rectangular cross section or an elliptical cross section.

[0025] Preferably, the first cooling aperture may comprise a bore having a circular cross section.The diameter of the first cooling aperture may be between 0.15 mm and 0.50 mm.

[0026] In addition, the radial spacing of the first orifices can be between 1.5 times and 5 times the radial size of the first orifice (in particular, the diameter of the first orifice), and the radial spacing of the first orifices corresponds to the radial size between two adjacent first orifices, and the radial spacing of the first orifices is measured from the center of the outlet cross-section of the first orifice to the center of the outlet cross-section of the adjacent first orifice.

[0027] Furthermore, a radial height of the region at the root of the blade comprising the first cooling apertures may be less than or equal to 40%, in particular less than or equal to 20%, of the radial height of the blade, the radial height of the blade corresponding to a radial dimension of the blade between the platform and the top.

[0028] Furthermore, a radial height of the region at the tip of the blade including the third cooling apertures may be less than or equal to 30% of a radial height of the bucket corresponding to a radial dimension of the bucket between the platform and the tip.

[0029] Advantageously, the spacing between the lower surface wall and the upper surface wall and the radius of the trailing edge chamfer may increase radially from the platform towards the tip of the blade.

[0030] This spacing may be defined by the distance between the lower surface wall and the upper surface wall, or the width of the dihedral, at a predetermined distance from the trailing edge, for example equal to 5 mm, or even 6 mm, or 7 mm.

[0031] Furthermore, preferably, the spacing between the lower surface wall and the upper surface wall and the radius of the trailing edge chamfer may be larger in the region at the root of the blade than in the region at the tip of the blade and in the region at the middle of the blade.

[0032] Moreover, the spacing between the lower surface and the upper surface and the radius of the trailing edge chamfer may be greater in the region at the middle of the blade than in the region at the top of the blade. Alternatively, the spacing between the lower surface wall and the upper surface wall and the radius of the trailing edge chamfer may be greater in the region at the top of the blade than in the region at the middle of the blade.

[0033] Preferably, the blade may be a moving blade for a moving turbine wheel of a turbine, in particular a high-pressure turbine.

[0034] In addition, according to another aspect of the present invention, another object of the present invention is to provide a turbine for a turbine, characterized in that the turbine comprises at least one moving wheel, at least one moving wheel comprises a plurality of moving blades such as those defined previously, and preferably, the turbine is a high-pressure turbine.

[0035] Furthermore, according to another aspect of the present invention, the object of the present invention is to provide a turbomachine, characterized in that the turbomachine comprises at least one turbine as defined above, preferably the turbomachine is a twin-shaft turbomachine.

[0036] The blade, the turbine and the turbomachine according to the invention may comprise any one of the features described in this description, taken alone or in combination with other features in any technically possible way. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] The invention can be better understood by reading the following detailed description of non-limiting examples of embodiments of the invention and by examining the schematic and partial views of the accompanying drawings, in which:

[0038] Figure 1 is a schematic axial cross-sectional view of an example of a turbofan engine suitable for implementing the present invention,

[0039] Figure 2 A first example of a trailing edge of a turbine blade according to the invention is shown according to a partial side view,

[0040] Figure 3 A second example of a trailing edge of a turbine blade according to the invention is shown according to a partial side view,

[0041] Figure 4A , Figure 4B as well as Figure 4C They are shown in Figure 5 The cross section A at the top of the blade at the dihedral angle Figure 5 The cross section B at the middle of the blade at the dihedral angle, and Figure 5 The dihedral angle of the cross section C at the root of the blade,

[0042] Figure 5 schematically shows in perspective the dihedral obtained at the trailing edge of a turbine blade according to the invention, the dihedral being formed by the planes of the lower and upper surfaces of the blade, which are joined at the trailing edge by a trailing edge chamfer,

[0043] Figure 6 A third example of a trailing edge of a turbine blade according to the invention is shown according to a partial side view, and

[0044] Figure 7 The radial evolution of the radius of the trailing edge chamfer is shown in diagram form.

[0045] The same reference numbers may be used throughout the drawings to identify the same or similar elements.

[0046] Furthermore, in order to make the drawings easier to read, the various parts shown in the drawings are not necessarily in consistent proportion. DETAILED DESCRIPTION

[0047] Throughout the description, it should be noted that the axis 2 of the turbine 1 is referred to as the radial symmetry axis of the turbine (see Figure 1 ). The axial direction of the turbine 1 corresponds to the axis of rotation 2 of the turbine 1. The radial direction of the turbine 1 is the direction perpendicular to the axis 2 of the turbine 1. Furthermore, unless otherwise stated, the adjectives and adverbs axial, radial, axially and radially are used with reference to the aforementioned axial and radial directions, and the terms inner (or internal) and outer (or external) are used with reference to the radial direction, such that an inner part of an element is closer to the axis 2 of the turbine 1 than an outer part of the same element. Furthermore, it should be noted that the terms upstream and downstream are considered relative to the normal gas flow main direction 5 of the turbine 1 (from upstream to downstream).

[0048] Figure 1 An aircraft turbomachine 1 is shown, for example a turbofan and a two-shaft turbojet engine, having a central longitudinal axis 2, about which various components of the aircraft turbomachine extend. The turbomachine comprises, from upstream to downstream along a main direction 5 of gas flow through the turbomachine, a fan 3, a low-pressure compressor 4, a high-pressure compressor 6, a combustion chamber 11, a high-pressure turbine 7 and a low-pressure turbine 8.

[0049] Typically, after passing through the fan, the air is divided into a central main flow 12a and a secondary flow 12b surrounding the main flow. The main flow 12a flows in a primary gas flow path 14a that passes through the compressor 4, 6, the combustion chamber 11 and the turbine 7, 8. The secondary flow 12b flows in a secondary flow path 14b that is radially outwardly bounded by the engine casing and surrounded by the nacelle 9.

[0050] Typically, the high-pressure turbine 7 has alternating moving wheels and distributors. The distributor includes a plurality of fixed blades, and the moving wheel includes a plurality of moving blades 18, such as in Figure 2 Shown in.

[0051] The moving blades 18 of the high-pressure turbine 7 can be cooled at the trailing edge primarily by means of drilled holes produced in the thickness of the trailing edge chamfer or by means of cooling grooves.

[0052] Both cooling techniques have advantages and disadvantages. Therefore, from a thermomechanical point of view, the drilled holes made in the thickness of the trailing edge chamfer represent a good compromise. The drilled holes ensure that the heat level at the trailing edge is controlled. However, considering the minimum thickness of the material in this area, this requires defining a relatively thick trailing edge with an aerodynamic profile, which has an adverse effect on the aerodynamic efficiency of the aerodynamic profile. In addition, the slots at the trailing edge make it possible to define a relatively good aerodynamic dihedral. However, from a thermomechanical point of view, the efficiency of the slots is lower than that of the drilled holes occurring in the thickness of the trailing edge chamfer. In fact, the air injected at the inlet of the trailing edge slot, which cools the slot bottom in the form of a film, heats up when passing through the slot bottom in contact with the air of the main flow path.

[0053] In order to improve the cooling at the trailing edge and obtain an improved aerodynamic profile of the moving blade 18, the invention proposes defining a three-fold configuration for injecting cooling air at the trailing edge, associated with a dihedral portion at the trailing edge with an aerodynamic profile, which three-fold configuration can be adjusted according to the radial height, according to the local thermomechanical needs, in order to maximize the aerodynamic efficiency.

[0054] therefore, Figure 2 A first example of cooling of the trailing edge BF of a moving blade 18 of a turbine according to the invention is shown according to a partial side view, the moving blade being equipped, for example, with Figure 1 The high-pressure turbine 7 of the turbojet engine 1 .

[0055] The blade 18 comprises an inner platform 28 and vanes 30 extending in the radial direction 23 , extending from the inside outwards away from the inner platform 28 and terminating at a tip S of the blade.

[0056] Furthermore, typically, the blade 30 includes a leading edge BA and a trailing edge BF located downstream of the leading edge BA. The blade 30 also includes a lower surface wall 15 and an upper surface wall 16, each connecting the leading edge BA to the trailing edge BF, and the lower surface wall 15 and the upper surface wall 16 are connected to each other at the trailing edge BF by a trailing edge chamfer 17.

[0057] According to the invention, the blade 18 comprises a cooling zone ZF extending along the trailing edge BF, the cooling zone comprising three different types of cooling orifices 25, 26, 27 formed through the wall of the blade 30 to enable the circulation of a cooling flow from the inside to the outside of the blade 30. These three types of cooling orifices 25, 26, 27 are radially spaced apart from one another between the inner platform 28 and the top S.

[0058] Therefore, in the area at the root ZP of the blade, close to the inner platform 28, which is a particularly critical mechanical area, the blade 18 comprises a first type of cooling holes 25, which are in the form of first cooling holes 25, which are formed in the thickness of the trailing edge chamfer 17 between the lower surface wall 15 and the upper surface wall 16.

[0059] These first orifices 25 are here made in the shape of a drilled hole, preferably the orifices have a circular cross section, so preferably the orifices 25 are cylindrical, or the orifices have an elliptical cross section, as in Figure 2 These first orifices 25 may have a larger radial dimension, in particular a diameter D, which is between 0.15 mm and 0.50 mm. In addition, the radial spacing p of these first orifices 25 is between 1.5 and 5 times the radial dimension D of the first orifices 25 (in particular, the diameter of the first orifices), and the radial spacing of the first orifices corresponds to the radial dimension between two adjacent first orifices 25, and the radial spacing of the first orifices is measured from the center of the outlet cross section of the first orifice 25 to the center of the outlet cross section of the adjacent first orifice 25.

[0060] The manufacture of these first orifices 25 makes it possible, among other things, to ensure the presence of a large amount of metal material. Therefore, the radius R of the trailing edge chamfer 17 is large and the spacing E of the dihedral angle DI is relatively thick, as will be described later with reference to Figure 4A and Figure 5 to describe.

[0061] Furthermore, in a region at the top ZS of the blade, close to the top S, the blade 18 comprises a third type of cooling apertures 27 in the form of third cooling apertures 27 which are different from the first apertures 25 and which are formed in the thickness of at least one of the lower surface wall 15 and the upper surface wall 16, preferably in the thickness of the lower surface wall 15.

[0062] Here Figure 2 In the example of FIG. 1 , these third cooling holes 27 are made in the shape of a drilled hole, in particular with a circular cross section. The third cooling holes make it possible to define a smallest possible radius R of the trailing edge chamfer 17 and a dihedral portion DI with a thin spacing E to improve aerodynamic efficiency, as will be described later with reference to FIG. Figure 4C and Figure 5 Explained.

[0063] Finally, in the zone at the middle ZM of the blade, situated between the zone at the root ZP of the blade and the zone at the top ZS of the blade (this zone is also a sensitive zone from a thermomechanical point of view), the blade 18 comprises a second type of cooling orifices 26 in the form of second cooling orifices 26, which are different from the first orifices 25 and the third orifices 27 and are formed in the thickness of the lower surface wall 15.

[0064] Here Figure 2 In the example of , these second cooling orifices 26 are in the shape of cooling slots, in particular with a rectangular or elliptical cross section. These second cooling orifices 26 are located axially between the third orifices 27 upstream and the first orifices 25 downstream. The second cooling orifices make it possible to obtain a radius R of the trailing edge chamfer 17 that is smaller than the radius of the trailing edge chamfer at the root of the blade, and to obtain a dihedral angle DI with a pitch E that is thinner than at the root of the blade.

[0065] Thanks to the invention, in addition to making it possible to define an improved aerodynamic profile at the trailing edge chamfer 17 and the dihedral angle DI of the trailing edge, it is also possible to minimize the amount of cooling used to cool the trailing edge BF. Compared to a conventional trailing edge cooled by slots over the entire radial height of the blade, the solution of the invention uses lower cooling flows, in particular due to the presence of simple drilled holes in the lower surface wall 15 in the region of the trailing edge BF, these drilled holes being minimally restrictive from a thermomechanical point of view.

[0066] exist Figure 2 In a first example, the region at the top ZS of the blade comprises two third cooling orifices 27. However, this number is variable. For example, Figure 3 Shown according to Figure 2 The view is similar to the view of the second embodiment, in which three third cooling openings 27 are present.

[0067] An optimal compromise between aerodynamic performance and thermomechanical behavior can be obtained, for example, by integrating first cooling orifices 25, preferably in the shape of drilled holes, for example, the radial height HP of the zone at the root ZP of the blade including these first cooling orifices 25 being less than or equal to 20% of the radial height H of the blade 30 corresponding to the radial dimension of the blade 30 between the inner platform 28 and the tip S, as shown in Figure 2 Furthermore, the radial height HS of the region at the top ZS of the blade including the third cooling apertures 27 is less than or equal to 30% of the radial height H of the bucket 30, and the third cooling apertures are preferably in the shape of drilled holes.

[0068] However, it is possible to modify this distribution according to the desired purpose. Thus, in order to significantly increase the useful life of the blade 20 , the extent of the first orifices 25 can be increased to 40% of the total height H of the blade 30 , thereby compromising aerodynamic performance.

[0069] Furthermore, it should be noted that, if the thickness of the material at the top ZS of the blade is sufficient, the third orifice 27 may comprise an orifice comprising a cylindrical first portion and a trumpet-shaped second portion comprising diverging walls, thus resembling a cooling orifice of the “fan-shaped hole” type in order to maximize the cooling efficiency in this area.

[0070] Figure 4A , Figure 4B , Figure 4C as well as Figure 5 It also makes it possible to illustrate the configuration of the optimized aerodynamic profile obtained by the invention at the trailing edge BF of a moving blade 18 of the turbine.

[0071] Thus, the cooling zone ZF of the trailing edge BF, the lower surface wall 15, the upper surface wall 16 and the trailing edge chamfer 17 together define a dihedral angle DI having a substantially Figure 5 The shape of the dihedral angle DI can be seen in FIG. 1 , which is characterized by the radius R of the trailing edge chamfer 17 and the spacing E between the lower surface wall 15 and the upper surface wall 16 opposite the trailing edge BF, as shown in FIG. Figure 4A , Figure 4B as well as Figure 4C Advantageously, the spacing E between the lower surface wall 15 and the upper surface wall 16 and the radius R of the trailing edge chamfer 17 vary radially between the inner platform 28 and the top S of the blade 18 .

[0072] More specifically, in the region at the root ZP of the blade, such as at Figure 4AAs can be seen in FIG. 1 , the spacing E is thick and the value of the radius R is high. In particular, this spacing E between the lower surface wall 15 and the upper surface wall 16 and the radius R of the trailing edge chamfer 17 are greater in the region at the root ZP of the blade than in the region at the top ZS of the blade and in the region at the middle ZM of the blade.

[0073] In the region at the middle ZM of the blade, as in Figure 4B As can be seen in FIG. 1 , the dihedral portion DI is thinner, with a radius R having a value lower than that at the root of the blade. In particular, the spacing E between the lower surface wall 15 and the upper surface wall 16 and the radius R of the trailing edge chamfer 17 are greater in the region at the middle ZM of the blade than in the region at the top ZS of the blade.

[0074] Finally, in the region at the top ZS of the blade, as in Figure 4C As can be seen in the figure, the dihedral angle DI is thin and the radius R is small.

[0075] Thus, the solution of the invention increases the radius R of the trailing edge chamfer 17 by 30% to 100% in the region of the root ZP of the blade and reduces the radius R of the trailing edge chamfer 17 by 10% to 50% in the region of the top ZS of the blade, compared to a reference aerodynamic profile drawn for a high-pressure turbine blade cooled by a slot at the trailing edge.

[0076] The solution of the invention can also be adjusted according to the local thermomechanical situation of the blade 18 to which the invention is applied. For example, the number of second orifices 26 (in particular, the shape of the slots) in the middle of the blade can be adjusted according to the value of the inlet temperature of the turbine of the engine to which the technology is applied, or during a new conception or design after thermal shading performance tests or durability tests of the turbine blades, which led to the identification of local hot spots at the trailing edge.

[0077] If the turbine inlet temperature is relatively low, the number of second orifices 26 (in particular the shape of the slots) can be reduced in favour of third orifices 27 , in particular in the shape of drill holes, which make it possible to define a better aerodynamic dihedral angle DI favouring efficiency.

[0078] Conversely, if the turbine inlet temperature is relatively high, the number of second orifices 26 (in particular, the shape of the slots) may be increased to ensure safe thermomechanical conditions in the region, thereby compromising aerodynamic efficiency.

[0079] In this case, the aerodynamic profile at the trailing edge BF is thinner in the middle ZM of the blade than in the tip ZS of the blade, in terms of the thickness of the dihedral angle DI and the value of the radius R of the trailing edge chamfer 17. Figure 6A third embodiment of a blade 18 is shown, in which the second opening 26 is in the form of a circular bore formed in the lower surface wall 15 and the third opening 27 is in the form of a cooling slot. In this case, the spacing E between the lower surface wall 15 and the upper surface wall 16 and the radius R of the trailing edge chamfer 17 are greater in the region at the top ZS of the blade than in the region at the middle ZM of the blade.

[0080] In addition, from the previous reference Figure 2 In the described embodiment, it is possible to determine the mean value V of the radial evolution of the radius R of the trailing edge chamfer 17 平均 , minimum value V 最小 And the maximum value V 最大 These values ​​are determined for the case in which the height HP represents 20% of the total height H of the blade and the height HS represents 30% of the total height H of the blade.

[0081] Figure 7 The ratio h / H is shown as a function of the radius R of the trailing edge chamfer 17 , wherein h corresponds to the radial height taken into account for observing the value of the radius R, starting from the inner platform 28 .

[0082] In addition, the value V 最小 、V 平均 and V 最大 This is shown in Table 1 below.

[0083] h / H <![CDATA[V 最小 (mm)]]> <![CDATA[V 平均 (mm)]]> <![CDATA[V 最大 (mm)]]> 0 0.5 0.75 1 0.1 0.5 0.75 1 0.2 0.5 0.75 1 0.3 0.3 0.45 0.6 0.4 0.3 0.45 0.6 0.5 0.3 0.45 0.6 0.6 0.3 0.45 0.6 0.7 0.2 0.3 0.4 0.8 0.2 0.3 0.4 0.9 0.2 0.3 0.4 1 0.2 0.3 0.4

[0084] Table 1: Radial evolution of the values ​​of the radius R of the trailing edge chamfer

[0085] Therefore, it can be seen that in Figure 2 In the example of FIG. 2 , the radius R decreases in the radial direction from the inner platform 28 towards the tip S of the blade, which makes it possible to improve the efficiency.

[0086] Of course, the present invention is not limited to the embodiments just described, and those skilled in the art may make various modifications to the embodiments.

Claims

1. A blade (18) of a turbine (7) of a turbomachine, the blade being intended to be mounted about an axis of rotation (2), the blade comprising a platform (28) and a vane (30), the vane extending in a radial direction (23) relative to the axis of rotation (2), extending from the inside outwards away from the platform (28) and terminating at a tip (S), the vane (30) comprising a leading edge (BA) and a trailing edge (BF) located downstream of the leading edge (BA), and the vane (30) comprising a lower surface wall (15) and an upper surface wall (16), the lower surface wall and the upper surface wall each connecting the leading edge (BA) to the trailing edge (BF), the lower surface wall (15) and the upper surface wall (16) being connected to each other at the trailing edge (BF) by a trailing edge chamfer (17), It is characterized in that The blade comprises a cooling zone (ZF) extending along the trailing edge (BF), the cooling zone (ZF) comprising three different types of cooling orifices (25, 26, 27) formed through the wall of the blade (30) to enable a cooling flow to circulate from the inside to the outside of the blade (30), the three types of cooling orifices (25, 26, 27) being radially spaced apart from one another between the platform (28) and the top (S), The blade (18) comprises, in a region at the root (ZP) of the blade close to the platform (28), a first type of cooling aperture (25) in the form of a first cooling aperture (25) formed in the thickness of a trailing edge chamfer (17) between the lower surface wall (15) and the upper surface wall (16), The blade (18) comprises, at the blade tip (ZS), in a region close to said tip (S), a third type of cooling aperture (27) in the form of a third cooling aperture (27) which is different from said first cooling aperture (25) and which is formed in the thickness of at least one of said lower surface wall (15) and said upper surface wall (16), The blade (18) comprises, in a region at the middle (ZM) of the blade, between a region at the root (ZP) of the blade and a region at the tip (ZS) of the blade, a second type of cooling aperture (26) in the form of a second cooling aperture (26) which is different from the first cooling aperture (25) and the third cooling aperture (27), the second cooling aperture being formed in the thickness of the at least one of the lower surface wall (15) and the upper surface wall (16), The cooling zone (ZF) of the trailing edge (BF), the lower surface wall (15), the upper surface wall (16) and the trailing edge chamfer (17) generally have the shape of a dihedral angle (DI), the shape of the dihedral angle being characterized by a radius (R) of the trailing edge chamfer (17) and a spacing (E) between the lower surface wall (15) and the upper surface wall (16) opposite to the trailing edge (BF), the spacing (E) being defined by a distance between the lower surface wall (15) and the upper surface wall (16) at a predetermined distance from the trailing edge (BF), and the spacing (E) between the lower surface wall (15) and the upper surface wall (16) and the radius (R) of the trailing edge chamfer (17) varying radially between the platform (28) and the top (S) of the blade (18), the spacing (E) and the radius (R) of the trailing edge chamfer (17) varying in value in the radial direction.

2. The blade according to claim 1, characterized in that The second cooling aperture (26) and the third cooling aperture (27) are formed in the thickness of the lower surface wall (15).

3. The blade according to claim 1 or 2, characterized in that: The first cooling aperture (25), the second cooling aperture (26) and the third cooling aperture (27) each comprise a different type of cooling aperture selected from a borehole having a circular cross section, a borehole having an elliptical cross section or a borehole in the shape of a slot.

4. The blade according to claim 1 or 2, characterized in that: The radial height (HP) of the region including the first cooling aperture (25) at the root (ZP) of the blade is less than or equal to 40% of the radial height (H) of the bucket (30), the radial height of the bucket corresponding to the radial dimension of the bucket (30) between the platform (28) and the tip (S).

5. The blade according to claim 1 or 2, characterized in that: The radial height (HS) of the region including the third cooling apertures (27) at the blade tip (ZS) is less than or equal to 30% of the radial height (H) of the blade (30), the radial height of the blade corresponding to the radial dimension of the blade (30) between the platform (28) and the tip (S).

6. The blade according to claim 1 or 2, characterized in that: The spacing (E) between the lower surface wall (15) and the upper surface wall (16) and the radius (R) of the trailing edge chamfer (17) increase radially from the platform (28) to the tip (S) of the blade (18).

7. The blade according to claim 1 or 2, characterized in that: The spacing (E) between the lower surface wall (15) and the upper surface wall (16) and the radius (R) of the trailing edge chamfer (17) are greater in the region of the root (ZP) of the blade than in the region of the blade tip (ZS) and in the region of the middle (ZM) of the blade.

8. The blade according to claim 1 or 2, characterized in that: The blades are moving blades (18) for a moving wheel of a turbine (7) of a turbomachine.

9. The blade according to claim 4, characterized in that The radial height (HP) of the region at the root (ZP) of the blade including the first cooling aperture (25) is less than or equal to 20% of the radial height of the bucket.

10. The blade according to claim 8, characterized in that The turbine is a high-pressure turbine.

11. A turbine (7) for a turbomachine, characterized in that The turbine comprises at least one moving wheel comprising a plurality of blades (18) according to any one of the preceding claims.

12. The turbine (7) for a turbomachine according to claim 11, characterized in that The turbine is a high-pressure turbine.

13. A turbine (1), characterized in that: The turbomachine comprises at least one turbine according to claim 11 or 12 .

14. The turbine (1) according to claim 13, characterized in that The turbine is a twin-shaft turbine.

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