TURBO MACHINE BUCKET WITH AN OPTIMIZED BAND AND METHOD FOR OPTIMIZING A BUCKET PROFILE

Description

Technical field of the invention

[0001] The present invention relates to the field of turbomachine blades for aircraft. It specifically aims at the design and / or manufacture of these turbomachine blades. Technical background

[0002] Prior art includes documents JP-A-H09 60501, JP-A-2000 018003, JP-A-H07 253001, US-A-4 165 949, and EP-A1-1 612 372.

[0003] Many mechanical components of turbomachinery are modified and / or redesigned to improve turbomachinery performance. Low-pressure turbine blades are a prime example, now featuring optimized aerodynamic profiles. Generally, a low-pressure turbine blade for a rotating runner comprises an aerodynamic blade extending along a radial axis perpendicular to the longitudinal axis and radially delimited by an inner root and an outer tail. The blade includes a leading edge and a trailing edge connected by intrados and extrados surfaces. One way to optimize the blade's aerodynamic profile is by reducing the thickness of the blade's trailing edge by approximately 0.30 mm (from 1 mm along the trailing edge, depending on the gas flow within the turbomachine), thereby minimizing aerodynamic losses.

[0004] However, this new geometry introduces specific new constraints in terms of manufacturability and design to meet lifespan objectives compared to so-called "conventional" blades. The trailing edges are also subject to high mechanical stresses along their entire radial height, particularly at the junctions with the blade roots and / or tails during blade rotation.

[0005] The present invention aims in particular to provide a turbomachine blade whose profile is optimized to reduce the stresses applied to its trailing edge during the operation of the turbomachine. Summary of the invention

[0006] We achieve this objective in accordance with the invention by means of an aircraft turbomachine blade comprising a blade extending in a radial direction, the blade having an intrados surface and an extrados surface which are connected upstream, in a direction of gas flow in the turbomachine, by a leading edge and downstream by a trailing edge, the blade having at a radially external end a heel and comprising a tip cross-section in a plane perpendicular to a radial direction of the blade, taken at the radially external end, the heel extending radially outwards from the tip cross-section which has a first center of gravity, the heel having a second center of gravity which is defined in a plane parallel to the tip cross-section, the second center of gravity being offset at least transversely from the first center of gravity,the second center of gravity being defined in a predetermined zone delimited at least in part by a first straight line and a second straight line forming substantially a V which is open towards the intrados surface and which includes a vertex whose orthogonal projection onto the transverse section of the head is located on the first center of gravity.

[0007] Thus, this solution achieves the aforementioned objective. In particular, this arrangement of the center of gravity of the blade tip reduces the stresses applied to the trailing edge by approximately 10%, taking into account the manufacturing constraints of the blade and its aerodynamic profile. This blade profile is also more efficient, and its trailing edge can be even thinner. Such a blade can also be adapted to any type of turbomachine without requiring any structural modifications for its integration.

[0008] The dawn also includes one or more of the following characteristics, taken alone or in combination: The tip cross-section is located radially just below the heel. The first center of gravity is linked to a frame of inertia comprising a first axis of inertia and a second axis of inertia that are perpendicular and pass through the first center of gravity. The first line is inclined at a first angle to the first axis of inertia, and the second line is inclined at a second angle to the first axis of inertia. The predetermined zone is delimited by a third line that is parallel to the first axis of inertia and offset towards the lower surface by a predetermined distance. The blade is solid.The blade comprises a root located at a radially inner end of the blade, opposite in the radial direction to the radially inner end. The trailing edge has a first thickening located radially between a first platform of the root and the blade, and extending at least partially transversely on either side of the trailing edge. The tail comprises a second platform defined in a plane inclined radially outwards and forming an angle between 0° and 40° with the plane of the tip cross-section. The trailing edge has a second thickening located radially between the second platform and the blade, and extending at least partially transversely on either side of the trailing edge. The first and second thickenings each comprise, respectively, an axial cross-section of generally triangular shape.The first and second thickenings each extend respectively from one of the radially inner and outer ends between 10 and 30% of the radial height of the blade. The predetermined zone is delimited by the blade skeleton line, which is intersected by the first and second lines. The predetermined zone is delimited by a curved line parallel to the blade skeleton line, the curved line being located at a maximum distance corresponding to twice the transverse thickness of the blade from the extrados surface.

[0009] The invention also relates to a wheel of a turbomachine comprising a disk centered on a longitudinal axis and a plurality of blades having any of the aforementioned characteristics, extending from the periphery of the disk and distributed regularly around the longitudinal axis.

[0010] The invention also relates to a turbomachine comprising a blade or wheel as mentioned above.

[0011] Finally, the invention relates to a method for optimizing the profile of an aircraft turbomachine blade, the blade comprising a blade extending in a radial direction, with a heel at one radially external end and a tip cross-section at the radially external end, the heel extending radially outwards from the tip cross-section, the method comprising the following steps: calculation of a first center of gravity of the blade tip cross-section, calculation of a second center of gravity of the heel; comparison of the coordinates of the first center of gravity of the tip cross-section and the second center of gravity of the heel; measurement of the distance between the first and second centers of gravity along an orthogonal projection of the plane in which the second center of gravity of the heel is defined in the plane of the tip cross-section; and compensation in which the second center of gravity of the heel is offset at least transversely from the first center of gravity of the tip cross-section in a predetermined area, the predetermined area being delimited at least in part by a first line and a second line forming substantially a V which is open towards an intrados surface of the blade and which has a vertex whose orthogonal projection is located on the first center of gravity.

[0012] According to the process, this includes a step of offsetting a trailing edge of the blade and / or a leading edge of the blade relative to a first axis of inertia of an inertia frame.

[0013] Following the process, the cross-section of the head is located radially just below the heel. Brief description of the figures

[0014] The invention will be better understood, and other objects, details, features and advantages thereof will become more apparent upon reading the detailed explanatory description that follows, of embodiments of the invention given by way of purely illustrative and non-limiting examples, with reference to the accompanying schematic drawings in which: [ Fig. 1 ] There figure 1 represents an example of a rotating wheel comprising a disk and blades extending from the periphery of the disk; Fig. 2 ] There figure 2is a side view of an example of a turbomachine blade according to the invention; [ Fig. 3 ] There figure 3 schematically represents a cross-section of the tip of a turbomachine blade at one radially external end just below a blade heel according to the invention; Fig. 4 ] There figure 4 schematically represents an example of a predetermined zone arrangement in which the center of gravity of a turbine blade heel of a turbomachine according to the invention is defined; Fig. 5 ] There figure 5 is a perspective and detail view of a blade foot with a thickening at the trailing edge according to an embodiment of the invention; [ Fig. 6 ] There figure 6 is a perspective and detail view of a blade foot with a thickening at the trailing edge according to another embodiment of the invention; [ Fig. 7 ] There figure 7is a view of the blade of a turbomachine blade at its radially external end and which includes a thickening at the trailing edge according to the invention; and [ Fig. 8 ] There figure 8 is a perspective and detail view of a turbomachine blade heel with a thickening at the trailing edge according to the invention. Detailed description of the invention

[0015] There figure 1 represents a turbine blade 1 of a turbomachine 2 with a longitudinal axis X for an aircraft, and in particular a blade for a low-pressure turbine wheel. However, the blade could be a compressor blade or any other blade intended to equip a turbomachine.

[0016] Generally, a turbine comprises one or more stages arranged successively along the longitudinal axis X of the turbomachine. Each turbine stage includes a rotating blade 3 forming a rotor and a stationary blade forming a stator. The stator blades are referred to as distributor blades. Each rotating blade 3 includes an annular disk 4 as illustrated in the figure 1 which is centered on the longitudinal axis. A plurality of movable blades 1 are mounted on the periphery of the disc and are distributed circumferentially and regularly around the disc of the moving wheel. Each moving wheel 3 is arranged downstream of a distributor wheel.

[0017] On the figure 2The moving blade 1 of the turbomachine includes an aerodynamic blade 5 which extends in a radial direction R between a radially internal end 6 and a radially external end 7. The blade 5 has a radial height which is determined between 0% at the radially internal end and 100% at the radially external end.

[0018] In this description, the blade will be described with respect to radial (R), longitudinal (L), and transverse (T) directions, while the turbomachine will be described with respect to longitudinal (X), radial (Z), and transverse (Y) axes. These directions are perpendicular to each other. The axes are also perpendicular to each other and form an orthonormal coordinate system OXYZ, with O as the origin. The origin of the coordinate system is centered on the longitudinal axis of the turbomachine. In the installed configuration, the radial direction is parallel to the radial axis.

[0019] Furthermore, the terms "upstream," "downstream," "axial," and "axially" are defined with respect to the direction of gas flow in the turbomachine and also substantially along the longitudinal axis or direction. Similarly, the terms "radially," "internal," and "external" are defined with respect to the radial axis or radial direction.

[0020] The blade 5 comprises a leading edge 8 and a trailing edge 9 which are opposite, here along the longitudinal direction of the blade. Each blade is arranged in the aerodynamic flow so that the leading edge 8 is placed upstream of the trailing edge 9. The leading edge 8 and the trailing edge 9 are connected by an intrados surface 10 and an extrados surface 11 (cf. figure 3 ) which are opposite along the transverse direction.

[0021] The blade profile is curved, and its transverse thickness varies from the leading edge to the trailing edge. In this example, the blade is solid. That is to say, the blade has no internal cavities.

[0022] The blade is advantageously made of a metallic material or a metallic alloy such as a nickel-based alloy. An example of a nickel-based alloy is known as DS200®.

[0023] With reference to figures 1 and 2The blade includes at its radially internal end 6 a foot 12 which is designed to engage in a correspondingly shaped groove 13 in the disc 4. The disc 4, for this purpose, includes a plurality of grooves 13 regularly distributed around its periphery. The foot 12 includes, in particular, a radially internal portion having a significant thickness relative to the rest of the foot, which is called the bulb 14. The bulb is housed in the groove 13. The foot 12 also includes a stilt 15 which extends radially from the bulb 14. Typically, the stilt 15 is connected to the bulb via a neck.

[0024] The foot 12 also includes a first platform 16 that separates the blade from the foot 12. The first platform 16 defines, in particular, a radially internal wall portion delimiting a turbomachine duct in which an aerodynamic flow circulates, here a primary flow. The strut 15 extends radially between the platform 16 and the bulb 14.

[0025] The blade 5 includes at its radially external end 7 a heel 17. As we can see on the figure 2 The heel 17 typically includes a second platform 18 designed to form a radially external wall portion that also delimits the aerodynamic duct (here, the primary duct). The radially internal and external walls are radially opposed. The heel 17 is provided with blades 19 that extend radially from a radially external surface 20 of the platform. A radially internal surface 21 is opposite the radially external surface and is generally oriented towards the blade root.

[0026] In operation, each blade is subjected on the one hand to aerodynamic forces due to the circulation of the gas flow through the turbine and the blades, and on the other hand to centrifugal forces due to the rotation of the turbine disk and the blade around the longitudinal axis.

[0027] On the figure 3is represented a cross-section of the blade tip ST at 100% of the blade height (i.e., at the radially outermost tip) (and in particular where the root is connected immediately directly to the blade) as shown on the figure 2We understand that the heel extends radially outwards from the tip cross-section. The tip cross-section is defined in a first plane perpendicular to the radial direction of the blade. This tip cross-section ST has a first center of gravity G1 (associated with a mass) which has been previously determined in the O,X,Y,Z coordinate system (turbomachine coordinate system). Also shown is an inertial coordinate system of the tip cross-section comprising a first axis of inertia I1 (minor axis of inertia) and a second axis of inertia I2 (major axis of inertia) whose origin R1 is centered on the first center of gravity G1 of the tip cross-section. These first and second axes of inertia are perpendicular.

[0028] With reference to the figure 3, the first center of gravity G1 of the head cross section ST is arranged at a first predetermined distance (such as about 6 mm) from a straight line D1 defined in the first plane of the head section, tangent with the leading edge at point B1 and parallel to the first axis of inertia I1. In other words, the leading edge is at a distance VBA of the order of +6 mm from the first axis of inertia.

[0029] The first center of gravity G1 of the head cross-section ST is also located at a second predetermined distance (such as approximately 5 mm) from a straight line D2 defined in the first plane of the head cross-section, tangent to the trailing edge 9 at point B2 and parallel to the first axis of inertia I1. The trailing edge is at a distance VBF of approximately 5 mm from the first axis of inertia I1. Similarly, the trailing edge is located at a distance VSI of approximately 16 mm from the first axis of inertia I1.

[0030] On the figure 3We can also see that a straight line TE tangent to the extrados surface 11 at a point B3, parallel to the first axis of inertia I1, is located at a distance DE of approximately 3 mm from the first center of gravity G1. The second axis of inertia I2 intersects the intrados surface 10 at a point B4 located at a distance DI of approximately 0.6 mm from the first center of gravity G1.

[0031] Blade 5 is subjected to stress, particularly at its trailing edge 9, primarily due to bending moments (M1 and M2) generated by aerodynamic and centrifugal forces, as well as tensile stress (Fz) due to centrifugal force. This stress can be expressed by the following material strength equation: σ = Fz S − M 1 I 1 X 2 − M 2 I 2 X 1

[0032] With Fz: corresponding to the tensile force induced by the rotation of the blade. M1 and M2: corresponding respectively to the bending moments due to aerodynamic and centrifugal forces in the inertial frame of the tip cross-section at 100% of the blade height. S: is the cross-sectional area of ​​the blade at the tip. X1: corresponds to the distance (VBF) between a point B2 on the trailing edge 9 and the first axis I1. X2: corresponds to the distance (VSI) between the trailing edge 9 and the second axis I2.

[0033] This equation is advantageously applied at the leading and trailing edges of the blade head cross-section.

[0034] The material distribution of the blade in the ST head cross-section is greater around the second axis of inertia than around the first axis of inertia.

[0035] To limit, or even eliminate or transfer elsewhere, the stresses applied to the trailing edge (which is relatively thin compared to the trailing edge of a conventional blade), the blade has an optimized profile or geometry. For example, the trailing edge 9 has a transverse thickness of 0.30 mm, or even 0.20 mm from an axial distance of 1 mm from the trailing edge.

[0036] To that end and with reference to the figure 4The heel (associated with a mass) is configured to present a second center of gravity G2, which is defined in a second plane (approximately perpendicular to the radial direction) parallel to the first plane of the head cross-section. In particular, the second center of gravity G2 is offset at least transversely from the first center of gravity G1 of the head cross-section ST, which is located at the radially outermost end (i.e., at 100% of the blade height along the radial direction).

[0037] Advantageously, but not exclusively, the second center of gravity of the wing root is defined in a predetermined zone ZG located upstream of the blade and near the lower surface 10 of the tip cross-section. This predetermined zone is also defined in the plane of the second center of gravity. More precisely, the predetermined zone is located upstream of a median plane of the tip cross-section that is parallel to the second axis of inertia and includes the radial direction. Placing the center of gravity of the wing root in such a zone reduces the stresses on the trailing edge of the blade.

[0038] As can be seen precisely on the figure 4, the predetermined zone ZG is represented hatched is delimited at least in part, following an orthogonal projection of the plane including the center of gravity G2 of the heel in the plane of the cross section of the head ST, by a first line L1 and a second line L2 forming substantially a V open towards the intrados surface 10 and whose apex is located on the center of gravity of the cross section of the head ST.

[0039] The first straight line L1 is inclined at a first angle β (beta) with respect to the first axis of inertia I1 and passes through the first center of gravity G1 the head cross section ST. This first angle is between 1 and 10°.

[0040] The second line L2 is inclined at a second angle y (gamma) with respect to the second axis of inertia I2 and passes through the center of gravity G1 of the head cross-section ST. The second line L2 is also parallel to the longitudinal axis X, as we can see on the figure 4The second angle γ that the second line L2 forms with the second axis of inertia is advantageously between 35° and 50°.

[0041] The predetermined zone ZG is also delimited, at least in part, by a third line L3 which is parallel to the first axis of inertia. This third line L3 is located at a distance from the first axis of inertia (towards the extrados surface) which is between 0 and 1 mm depending on the geometry of the blade and the turbine stages.

[0042] We can also see on the figure 4 that the predetermined zone ZG is delimited upstream by the skeleton line S of the curved blade which intersects the first straight line L1 and the second straight line L2.

[0043] More generally, the predetermined zone ZG is delimited by: The first straight line L1, the second straight line L2 which forms approximately a V open towards the intrados surface 10, the orthogonal projection of the apex of the V onto the cross-section of the head located on the first center of gravity G1, and a curved line parallel to the skeleton line which is located at a maximum distance corresponding to twice the transverse thickness of the blade from the extrados surface. This curved line intersects the first and second straight lines and is a maximum boundary of the predetermined zone ZG beyond the extrados surface.

[0044] The platform 18 of the heel 17 is defined in a plane inclined radially outwards. This plane forms an angle α (alpha) between 0° and 40° with the plane of the head cross-section. This angle simplifies the modeling by calculation, thus avoiding the need to work with non-integer and / or inclined sections. This angle depends on the shape of the turbine runner as well as the geometry of the blade.

[0045] Advantageously, for a heel weighing between 10 and 20 g, the third line L3 is offset from the first axis of inertia by a distance of between 0 and 0.3 mm with a platform inclined between 20 and 30° relative to the plane of the head cross-section. Alternatively, the third line is offset by a distance of between 0 and 1 mm with a platform inclined between 0 and 20° relative to the plane of the head cross-section.

[0046] Another possibility or complement to reduce stress at the trailing edge is to optimize the trailing edge profile locally. With reference to the figure 5 The trailing edge 9 has a first thickening 30 located between the root platform and the blade, extending at least partially transversely on either side of the trailing edge. This thickening 30 also extends over 10% of the blade height from the radially inner tip 6. The thickening or additional thickness is on the order of 0.05 mm (in the transverse direction). The thickening advantageously has a substantially triangular axial cross-section.

[0047] According to another embodiment illustrated on the figure 6The first thickening 30 is located between the platform 16 of the foot and the trailing edge 9 of the blade. The first thickening 30 extends transversely on either side of the trailing edge and radially up to approximately 30% of the blade height.

[0048] Adding this thickening to the trailing edge and the radially inner tip of the blade reduces stress by 2 to 7%. This thickening also facilitates the manufacturing of the blade with a thin trailing edge. Furthermore, this geometry limits the risk of material deflection. Deflection is then likely to occur in a less stressed area than the original, allowing for a 10 to 20% margin of error on certain materials.

[0049] As we can see precisely on the figures 7 and 8The trailing edge also has a second thickening 31 located at the radially external end of the trailing edge, near the heel platform. This thickening 31 has the same configuration as the first thickening, namely a height between 10% and 30% from the radially external end and a transverse extension on either side of the trailing edge.

[0050] The center of gravity of the heel is determined using a process for optimizing the profile of an aircraft turbomachine blade. The various steps of the process are implemented using CAD / CAM and / or calculation software.

[0051] Initially, the characteristics of the blade, such as its mass, material, dimensions, etc., are referenced in the software.

[0052] In a first step, blade 5 is divided into several cross-sections (horizontal) according to its radial height.

[0053] The cross-section of the blade tip at the outer radial end (and located just before the radius of the heel, at 100% of the blade height) is selected.

[0054] The process then includes a step of calculating the center of gravity G1 of the head cross-section ST. Prior to this calculation step, a mass is associated with the head cross-section ST. The center of gravity G1 is defined as the geometric center of gravity of the head cross-section.

[0055] The coordinates of the center of gravity G1 are defined, in the plane of the head cross-section, with respect to the inertial frame comprising the first axis of inertia I1 and the second axis of inertia I2 whose origin R1 is located on the center of gravity G1 of the head cross-section.

[0056] The process includes a further step of calculating the center of gravity of the blade's heel. Prior to this step, the mass of the blade's heel is measured, which allows its center of gravity to be determined.

[0057] A comparison of the coordinates between the center of gravity G1 of the cross section of head ST and the center of gravity G2 of heel 17 is carried out.

[0058] To do this, a measurement of the distance between the first and second centers of gravity G1, G2 is taken. For this purpose, an orthogonal projection of the heel plane, including the center of gravity G2, is made onto the plane of the head cross-section ST.

[0059] Finally, the process includes a compensation step in which the second center of gravity of the tail is offset from the first center of gravity of the leading edge within a predetermined area. This compensation step includes at least a transverse and axial offset of the tail's center of gravity G2 within the predetermined area to reduce the aerodynamic stresses acting on the relatively thin trailing edge.

[0060] Typically, once compared, if the center of gravity G2 of the heel is not within the predetermined zone ZG, at least partially V-shaped, a modification of the heel's mass distribution is made to move the heel's center of gravity G2 forward or backward. This operation is advantageously performed manually (in the design software) and naturally depends on the blade's manufacturing and integration criteria.

[0061] To compensate for the stresses at the relatively thin trailing edge of the blade, we can also manually adjust the distance of the leading edge and / or trailing edge relative to the first axis of inertia. To do this, we position the blade cross-sections relative to each other along the radial direction to limit the moment between them. By manually adjusting the values ​​of δAx and δTg (from the formula below) relative to the first axis of inertia, the offset is created, and thus the moment is generated. In the case of the thin trailing edge, we aim to modify the offset to reduce the stresses at the trailing edge.

[0062] The formula below characterizes the trailing edge stress σBF along the first axis of inertia I1. This formula characterizes the compensation at the leading edge and / or trailing edge as stated above. σ BF = Fz S − V BF I 1 × cos θ × M aers Ax + δ Ax × F u − sin θ × M aeso T a + δ T S × F

[0063] With: Fz: Traction force induced by centrifugal force. F c<: centrifugal force. S: head cross-section. V BF: distance from the trailing edge point to the first inertial axis I1. M Ax< aero: M1 in the inertial frame. Pitch angle θ (theta): platform inclination angle between the turbomachine frame and the inertial frame. M Tg< aero: M2 in the inertial frame. δ Ax<: relates to the axial offset of the trailing edge. δ Tg<: relates to the tangential offset of the trailing edge.

[0064] When the distance VBA from the leading edge to the first axis of inertia I1 is less than the distance VBF from the trailing edge to the first axis of inertia I1, we apply, for example, an axial compensation (or axial displacement value (δ Ax< )) of the leading edge upstream by approximately 1 mm and of the leading edge downstream by approximately 5 mm relative to the axis of inertia I1. That is, we reduce or increase the distance of the trailing or leading edges from the first axis of inertia I1. To achieve greater axial moment compensation, we can implement an axial offset of the leading edge upstream by approximately 2 mm and an axial offset of the trailing edge downstream by approximately 6 mm.

[0065] When the leading edge distance (VBA) is greater than the trailing edge distance (VBF), we apply a relatively large axial displacement value to the leading and trailing edges and / or a tangential displacement. In this case, it is possible to minimize the trailing edge stress.

Claims

1. A vane (1) for a turbine engine of an aircraft comprising a blade (5) extending in a radial direction, the blade having a pressure side surface (10) and a suction side surface (12) which are connected upstream, in a direction of circulation of a gas in the turbine engine, by a leading edge (8) and downstream by a trailing edge (9), the blade having, at a radially outer end (7), a root (17) and comprising a transverse head section (ST), in a first plane perpendicular to the radial direction of the blade, taken at the radially outer end (7), the root extending radially towards the outside from the transverse head section which has a first centre of gravity (G1), the root (17) having a second centre of gravity (G2) which is defined in a second plane parallel to the first plane of the transverse head section (S1), characterized in that the second centre of gravity is offset at least transversely from the first centre of gravity, the second centre of gravity (G2) being defined in a predetermined zone (ZG) delimited at least in part by a first straight line (L1) and a second straight line (L2) substantially forming a V which is open towards the pressure side surface (10) and which comprises an apex whose orthogonal projection on the transverse head section is located on the first centre of gravity (G1).

2. The vane (1) according to the preceding claim, characterised in that the first centre of gravity (G1) is linked to a reference frame of inertia comprising a first axis of inertia (I1) and a second axis of inertia (I2) which are perpendicular and which pass through the first centre of gravity, the first straight line (L1) being inclined at a first angle to the first axis of inertia (I1) and the second straight line (L2) being inclined at a second angle to the first axis of inertia (I2).

3. The vane (1) according to the preceding claim, characterised in that the first centre of gravity (G1) is arranged at: - a first predetermined distance from a third straight line (D1), defined in the first plane of the transverse head section, which is tangent with the leading edge (8) at a point (B1) and parallel to the first axis of inertia (I1) and a second predetermined distance, and - a second predetermined distance from a fourth straight line (D2) in the first plane of the transverse head section which is tangent to the trailing edge (9) at a point (B2) and parallel to the first axis of inertia (I1).

4. The vane (1) according to one of the preceding claims, characterised in that the predetermined zone (ZG) is delimited by a third straight line (L3) which is parallel to the first axis of inertia (I1) and offset towards the pressure side surface (10) by a predetermined distance.

5. The vane (1) according to any one of the preceding claims, characterised in that the vane is movable and the blade (5) is full.

6. The vane (1) according to any one of the preceding claims, characterised in that the vane comprises a base (12) located at a radially inner end (6) of the blade, opposite in the radial direction to the radially outer end (7), the trailing edge (9) having a first thickening (30) located radially between a first platform (16) of the base and the blade, and extending transversely on either side of the trailing edge (9).

7. The vane (1) according to any one of the preceding claims, characterised in that the root (17) comprises a second platform (18) which is defined in a plane inclined radially towards the outside and forming an angle of between 0° and 40° with the plane of the transverse head section.

8. The vane (1) according to the preceding claim, characterised in that the trailing edge (9) has a second thickening (31) located between the second platform (18) and the blade and extending at least partially transversely on either side of the trailing edge.

9. The vane (1) according to one of claims 6 to 8, characterised in that the first thickening (30) and the second thickening (31) each comprise an axial section of generally triangular shape.

10. The vane (1) according to one of claims 6 to 8, characterised in that the first thickening (30) and the second thickening (31) each extend respectively from one of the radially inner and outer ends between 10 and 30% of the radial height of the blade.

11. The vane (1) according to any one of the preceding claims, characterised in that the predetermined zone is delimited by the median line of the blade which is intersected by the first straight line and the second straight line.

12. The vane (1) according to one of the preceding claims, characterised in that the predetermined zone is delimited by a curved line parallel to the median line of the blade, the curved line being located at a maximum distance corresponding to twice the transverse thickness of the blade from the suction side surface.

13. A mobile wheel of a turbine engine comprising a disc (4) centred on a longitudinal axis X and a plurality of vanes (1), according to any one of claims 1 to 12, extending from the periphery of the disc and evenly distributed about the longitudinal axis.

14. A turbine engine (2) comprising a vane (1) according to any one of claims 1 to 12 or a wheel according to the preceding claim.

15. A method for optimising a profile of a vane for a turbine engine of an aircraft, the vane comprising a blade (5) with a root (17), extending in a radial direction, at a radially outer end (7) and a transverse head section (ST) at the radially outer end, the root (17) extending radially towards the outside from the transverse head section (ST), the method comprising the following steps of: - calculating a first centre of gravity (G1) of the transverse head section of the blade, - calculating a second centre of gravity (G2) of the root (17); - comparing the coordinates of the first centre of gravity (G1) of transverse head section and the second centre of gravity (G2) of the root; - measuring the distance between the first and second centre of gravity according to an orthogonal projection of the plane in which the second centre of gravity of the root is defined in the plane of the transverse head section; and - compensating in which the second centre of gravity of the root is offset at least transversely from the first centre of the gravity of the transverse head section in a predetermined zone (ZG), the predetermined zone (ZG) being delimited at least partially by a first straight line and a second straight line substantially forming a V which is open towards a pressure side surface of the blade and which comprises an apex, the orthogonal projection of which is located on the first centre of gravity.