ADJUSTABLE GUIDE BOW FOR AIRCRAFT TURBOMACH AND AIRCRAFT TURBOMACH

Description

technical field

[0001] The invention relates to turbomachine stator blades and more particularly to the design of variable-pitch blades with surface contouring. More specifically, the invention relates to a variable-pitch stator blade for an aircraft turbomachine and to an aircraft turbomachine. Previous art

[0002] Climate change is a major concern for many legislative and regulatory bodies worldwide. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by different countries. In particular, an ambitious standard applies to both new types of aircraft and those already in service, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been actively working for several years now to contribute to the fight against climate change.

[0003] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into account the factors impacting all phases of design and development in order to obtain less energy-intensive and more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences, with the aim of improving the energy efficiency of aircraft.

[0004] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes that minimize greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0005] This sustained research and development work focuses on new generations of aircraft engines, the weight reduction of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to provide propulsion, and, as essential complements to technological progress, aviation biofuels.

[0006] In this context, the invention relates more specifically to aspects connected with aircraft turbomachine compressors. These compressors include stator blades that straighten a primary flow, or "straightener." The airflow may exhibit unwanted secondary flows, notably due to separation at the upper and lower surfaces, which negatively impacts compressor performance and therefore the efficiency of the turbomachine.

[0007] The published patent document WO 2019 / 228897 A1 discloses blades of a straightener extending from a non-axissymmetric surface commonly referred to as "3D contouring", the latter including hollows and bumps intended to limit the occurrence of secondary flows.

[0008] However, this solution for providing trough and / or crest contours to the stator ferrules is not suitable for variable stator vanes (VSV). Indeed, while the trough and / or crest design on the ferrule is optimized for a given airflow direction, the variations in velocity and / or flow direction resulting from the pivoting of the VSV vanes render the contouring suboptimal. For example, horseshoe-shaped vortices appear with VSV vanes at high angles of attack.

[0009] The contouring proposed by document WO 2019 / 228897 A1 only concerns fixed rectifier blades and therefore does not allow for effectively limiting the occurrence of secondary flows at the level of a variable pitch rectifier. Summary of the invention Technical problem

[0010] The invention aims to solve at least one of the problems posed by the prior art. More specifically, the invention aims to provide a blade that ensures good stability in order to limit intrinsic efficiency losses and guarantee good compressor performance of a turbomachine regardless of the incidence of the VSV blades. Technical solution

[0011] The invention is the result of technological research aimed at significantly improving aircraft performance and, in this respect, contributes to reducing the environmental impact of aircraft. To this end, the present invention relates to a variable-pitch stator blade for a turbomachine, the blade comprising: an airflow guidance blade, the blade having a leading edge and a trailing edge, as well as an intrados and an extrados extending from the leading edge to the trailing edge; and one or two platform(s) disposed at one end or both ends of the blade, the platform(s) comprising an airflow guidance surface, the surface comprising a first disc portion on the intrados side and a second disc portion on the extrados side; remarkable in that the first or at least one of the first or second portions of the disc includes a bump.

[0012] Advantageously, the platform allows the blade to be fixed in a turbomachine housing, preferably by means of a trunnion ensuring a pivot connection with the housing.

[0013] Preferably, the blade comprises a single platform located at the blade tip. More preferably, the blade comprises two platforms, one at the blade tip and the other at the blade foot.

[0014] The platform(s) comprise / comprising two disk portions, preferably viewed radially. It should be noted that the disks are asymmetrical, as their airflow surface is not flat (it includes one or two bumps).

[0015] According to an advantageous embodiment of the invention, the or at least one of the first portions of the disc comprises a first bump, and the top of said first bump has a position whose projection onto the line of the blade chord at the platform is distant from the leading edge by 10% or less of the chord.

[0016] According to an advantageous embodiment of the invention, the top of the first bump is attached to the intrados or is disposed at a distance from said intrados of between 0% and 30% of a maximum thickness of the blade.

[0017] According to an advantageous embodiment of the invention, the or at least one of the second portions of the disc comprises a second bump.

[0018] Advantageously, the second bump helps to delay or even prevent the airflow from stalling at the level of the upper surface of the blade.

[0019] Preferably, the top of the second hump includes a position whose projection onto the blade chord at the platform is distant from the trailing edge of the blade by a distance corresponding to 10% or less of the chord.

[0020] According to an advantageous embodiment of the invention, the platform includes a peripheral edge which intersects the radial projection of the extrados on the guiding surface at a rear intersection point, the top of the second hump having a position whose projection on the blade's chord line at the platform is at a distance from the projection of the rear intersection point on the chord that is less than or equal to 10% of the chord.

[0021] According to an advantageous embodiment of the invention, apart from the bump(s), the surface of the first and second portion of the disc is substantially flat, and the top of the bump(s) rises above the substantially flat surface by a height which is between 0.5 and 5 mm, preferably between 1 and 3 mm.

[0022] According to an advantageous embodiment of the invention, the height of the top of the first bump is between 0.1 and 0.3 times the maximum thickness of the blade or between 1 and 3 times said maximum thickness, and / or the height of the top of the second bump is between 0.1 and 0.3 times the maximum thickness of the blade or between 1 and 3 times said maximum thickness.

[0023] Advantageously, this hump height minimizes secondary flows while limiting the impact of these humps on the blade's aerodynamic profile. Indeed, if the humps are too high, airflow may be blocked.

[0024] According to an advantageous embodiment of the invention, the height of the top of the hump(s) is between 1% and 10% of the chord.

[0025] According to an advantageous embodiment of the invention, the bump(s) are at a distance from the peripheral edge.

[0026] Advantageously, the first and / or second hump is / are confined to the platform and does not extend onto the fixed part of the housing, because if the hump(s) were located on a fixed part (turbomachine housing) outside the platform, the homogeneity of the aerodynamic flow would be affected. Indeed, confining the humps to the blade platforms of the invention allows said humps to act on the secondary flows regardless of the blade angle of attack.

[0027] According to an advantageous embodiment of the invention, the first hump has a principal orientation which forms an angle between 30° and 60° with the blade chord.

[0028] Preferably, the principal orientation of the first bump corresponds to a direction or principal axis of the extent of said first bump on the platform. For this purpose, the principal orientation defines the direction of the largest dimension of said first bump. The principal axis is preferably substantially parallel to the guide surface of the platform and can be viewed radially.

[0029] According to an advantageous embodiment of the invention, the second hump has a truncated cardioid shape.

[0030] The invention also relates to an aircraft turbomachine comprising a compressor, said compressor comprising a substantially axisymmetric airflow guidance surface except for an annular row of housings, the turbomachine being remarkable in that it comprises an annular row of blades according to the invention and one of its advantageous modes mentioned above, the blade platforms being received in housings of the casing such that the airflow guidance surfaces of the platforms are substantially flush with the airflow guidance surface of the casing.

[0031] Advantageously, the dawn of the invention makes it possible to avoid the appearance of secondary flows at the level of the variable-timing stator rectifier of a turbomachine compressor.

[0032] Indeed, the first hump helps to limit or even prevent unwanted secondary flow from rising towards the upper surface of a neighboring blade in a row of stator blades of the compressor. Simultaneously, the second hump helps to better guide the airflow in contact with the upper surface and prevents it from separating from said surface.

[0033] Furthermore, the invention is particularly advantageous because it improves the performance of aircraft compressors and the overall efficiency of turbomachinery, resulting in lower fuel consumption and greenhouse gas emissions, thus reducing the environmental impact of aircraft.

[0034] It is understood that each detail of an embodiment below can be combined with each other detail of the other embodiments. Brief description of the drawings

[0035] There [ Fig.1 ] illustrates a cross-sectional view of a compressor in an axial turbomachine; The [ Fig. 2 ] represents a perspective view of a turbine blade according to the invention; The [ Fig.3 ] illustrates a radial view of the blade according to a first embodiment of the invention; The [ Fig. 4 ] illustrates a side view of the dawn of the [ Fig.3 ] ; There [ Fig. 5 ] illustrates a radial view of a blade according to a second embodiment of the invention; The [ Fig. 6 ] illustrates a side view of the dawn of the [ Fig. 5 ] ; There [ Fig. 7 ] illustrates a radial view of a blade according to a third embodiment of the invention. Detailed description of the implementation methods

[0036] In the following description, the terms "internal," "inner," "lower," "outer," and "upper" refer to positioning relative to the axis of rotation of an axial turbomachine. The axial direction corresponds to the direction along the turbomachine's axis of rotation, with lengths measured axially. Widths are measured circumferentially. The radial direction is perpendicular to the axis of rotation. Upstream and downstream refer to the main flow direction of the airflow within the turbomachine.

[0037] The dimensions of the figures are not to scale and in particular the thicknesses or radial dimensions are exaggerated to facilitate reading the figures.

[0038] There [ Fig.1 ] represents a cross-sectional view of a compressor 2 of an axial turbomachine 4.

[0039] Turbomachine 4 can correspond to a turbofan engine, turbojet, turbofan, turboprop, turboshaft engine, or any other turbofan. Alternatively, turbomachine 4 can correspond to a multi-flow turbomachine, such as a counter-rotating open rotor (CROR or unducted single fan), or any other triple-flow turbomachine.

[0040] Preferably, compressor 2 corresponds to a low-pressure compressor 2 or a high-pressure compressor (not shown). The turbomachine 4 also includes other components not shown in the [ Fig.1], such as a high-pressure compressor, a combustion chamber, and one or more turbine levels. In operation, the mechanical power of the turbine, transmitted via the central shaft to the rotor 6, sets the two compressors in motion. These compressors comprise several rows of rotor blades 8 associated with rows of stator blades 10. The rotation of the rotor 6 around its axis of rotation X thus generates an airflow and progressively compresses this air until it reaches the inlet of the combustion chamber.

[0041] A blower 12 (partially illustrated) is coupled to the rotor 6 and generates an airflow which splits into an internal radial flow F1, commonly called the primary flow F1, and an external radial flow F', which can correspond to a secondary flow F' in the context of a double-flow turbomachine, or to a tertiary flow F' of a triple-flow type turbomachine 4.

[0042] The radially external flow F' passes through an annular duct (partially shown). The secondary flow is accelerated to generate a thrust reaction useful for aircraft flight. The primary flow F1 and the radially external flow F' are both annular and separated by a separating nozzle 14.

[0043] We can see on the [ Fig.1 ] that the rotor 12 comprises several rows of rotor blades 8, in this case three, and several rows of stator blades 10, in this case three; the latter allow the conversion of the speed of the air flow F1 into pressure, in particular into static pressure.

[0044] The rotor blades 8 can extend radially from a rotor support 16 which can be a dovetail platform, a bladed one-piece drum inner ring or any other type of support for a composite rotor.

[0045] The stator blades 10 extend essentially radially from an outer casing 18. They can be fixed and immobilized there by means of fixing pins 20. They traverse radially the primary flux F1. The stator blades 10 can have fixed chords (radially constant) with respect to the outer casing 18.

[0046] The low-pressure compressor 2 includes at the inlet of a primary flow vein 22 F1, a row of 9 stator-type rectifier blades with variable pitch.

[0047] Preferably, the variable pitch of the blades 9 is ensured by means of an actuation system (not illustrated) regulating the angle formed by the blades 11 of the blades 9 around an axis 24.

[0048] In this configuration, each blade 9 comprises a blade 11 and one or preferably two substantially circular platforms 26 arranged at the two ends of the blade 11. Each platform 26 is provided with a trunnion 28, on the side opposite the blade 11 and ensuring a pivot connection with the outer casing 18 or inner casing 30. The outer casing 18 is delimited radially internally by an upper surface 18.1 for guiding the primary flow F1, and the inner casing 30 is delimited radially externally by an inner guiding surface 30.1.

[0049] Advantageously, the 11 blades of the rectifier are identical and annularly aligned.

[0050] The housing 18, 30 includes housings 31 distributed annularly around the X axis and in which the platforms 26 are housed and pivotally mounted.

[0051] Advantageously, the platforms 26 include axisymmetric air guidance surfaces 26.1 (circular, cylindrical, conical, etc.) substantially flat and flush with the air flow guidance surfaces 18.1 and 30.1 of the housing 18, 30. Suitable seals can ensure sealing between the platforms 26 and the housing.

[0052] There [ Fig. 2 ] represents a perspective view of a blade 9 according to the invention. The blade 11 of the blade 9 comprises a leading edge 11.1 and a trailing edge 11.2, as well as an intrados 11.3 and an extrados 11.4 extending from the leading edge 11.1 to the trailing edge 11.2.

[0053] The blade 11 also includes a chord line LC, a skeleton line S and a maximum thickness e (illustrated in the [ Fig.3 ]).

[0054] The guiding surface 26.1 of the platform 26 includes a first portion of a disk 30.1 on the side of the intrados 11.3 and a second portion of a disk 30.2 on the side of the extrados 11.4.

[0055] The separation between the first 30.1 and the second portion of the disk 30.2 corresponds preferentially to the extension of the skeleton line S or the camber line S of the blade 11 on the guide surface 26.1 (viewed radially).

[0056] According to a first embodiment of the invention, the first portion 30.1 comprises a first bump 32. Details on the shape and position of the first bump 32 will be detailed later in this description.

[0057] There [ Fig.3 ] illustrates a radial view of the blade 9 according to the first embodiment of the invention.

[0058] Contour lines (iso-contour) can be seen illustrating the first bump 32 comprising a vertex 34 having a projection 34.1 onto the chord line LC of the blade 11 which is preferentially distant from the leading edge 11.1 by a distance D1 corresponding to at most 10% of the chord C. Alternatively, the distance D1 can be greater without however exceeding 15% of the chord C.

[0059] Preferably, the apex 34 of the first hump 32 is disposed adjacent (substantially attached) to the intrados 11.3. However, the apex 34 may be in the vicinity of the intrados 11.3, i.e., disposed at a distance between 0% and 50% of the maximum thickness e, and preferably between 0% and 30%.

[0060] More preferably, the apex 34 is at most 3 mm from the intrados 11.3.

[0061] Advantageously, such an arrangement of the first bump 32 allows to have a direct impact on the "horseshoe" vortex, because it makes it possible to avoid the separation of the boundary layer at the level of the blade wall on the intrados 11.3 and / or extrados 11.4 side.

[0062] The first hump 32 has a substantially circular, oval, or elliptical shape, said hump 32 being preferably substantially elliptical. However, the shape of the hump 32 may be truncated by the intrados 11.3. The first hump 32 has a principal orientation illustrated by a principal axis 32.1 which may form an angle between 30° and 60° with the chord line LC, and preferably an angle of about 45° (±10%) with the chord line LC.

[0063] Indeed, an orientation of the first hump 32 at 0° relative to the LC chord line does not obstruct secondary flows, and an orientation at 90° with the LC chord line does obstruct unwanted flows, but generates higher aerodynamic losses.

[0064] Advantageously, the orientation of the first bump 32 and the position of the apex 34 make it possible to obstruct the horseshoe vortex in order to prevent its passage towards the extrados 11.4, and this, in all possible orientations of the blade 9 around its axis 24. For this purpose, the blade 9 makes it possible to ensure optimal operation of the aircraft compressor in all flight conditions.

[0065] There [ Fig. 4 ] illustrates a side view of dawn 9 of the [ Fig.3Here, dawn 9 is partially illustrated; in particular, the lower half of said dawn 9 is represented, and the dotted line may possibly correspond to an axis of symmetry of dawn 9.

[0066] We see that the summit 34 of the first bump 32 rises above the substantially flat surface 26.1 by a height H1.

[0067] Preferably, the height H1 of the apex 34 is between 1% and 10% of the chord of the blade 11.

[0068] The height H1 is determined based on the maximum thickness e (illustrated in the [ Fig.3]) of the blade 11 at the platform 26 (that is to say, in cases where the thickness of the blade 9 is not constant along its radial height, the height of the hump is defined relative to the maximum thickness). Indeed, it is preferable that the height H1 be between 0.1 and 3 times the maximum thickness e. More preferably, the height H1 is between 0.1 and 0.3 times the maximum thickness e when the first hump 32 is at the tip of the blade, and between 1 and 3 times the thickness e when said hump 32 is at the tip of the blade.

[0069] For this purpose, the height H1 can, for example, correspond to 1 mm when the thickness e is equal to 10 mm, in particular in the case where the first bump 32 is located on the platform at the top of the blade, or said height H1 can be equal to 3 mm if the thickness e is 1 mm, in particular in the case where the first bump 32 is located on the platform at the bottom of the blade.

[0070] Preferably, the height H1 is between 0.5 and 5 mm, and more preferably between 1 and 3 mm. Below this range (height H1 less than 1 mm), tolerances in the shape of the first hump 32 may affect its effectiveness in damping secondary flows. Conversely, above this range (height H1 greater than 3 mm), the compressor's performance at nominal angle of incidence may be degraded.

[0071] The first hump 32 can generally have a convex and / or concave shape. Preferably, the first hump 32 has a convex shape at the upstream part corresponding to about half of the extent of said hump 32 (here the term about corresponds to ±20% of the extent), said first hump 32 has a concave shape on a downstream part corresponding to another half, and more preferably, the convex half is disposed downstream of the hump 32 (closer to the trailing edge 11.2) so as to place the convex part in direct contact with the secondary flows to block them and allow to better guide the flow downstream with the concave half.

[0072] There [ Fig. 5 ] illustrates a radial view of a blade 109 according to a second embodiment of the invention, in which the blade 109 includes a second bump 132 at the right of the second portion of disk 30.2 of the surface 26.1.

[0073] Identical elements between the first and second embodiments have the same reference signs, while additional or substantially similar elements are incremented by 100.

[0074] Preferably, the dawn 109 includes the first bump 32 at the right of the first portion of disc 30.1 and the second bump 132 arranged at the right of the second portion of disc 30.2.

[0075] Platform 26 includes a peripheral edge 26.2, and the first 32 and second bump 132 are at a distance from said peripheral edge 26.2. This allows the 3D contouring to rotate entirely with the blade 109.

[0076] The upper surface 11.4 has a rear intersection point 11.5 with the platform 26, and specifically with the peripheral edge 26.2, said point 11.5 corresponds to an intersection between said peripheral edge 26.2 and the radial projection of the upper surface 11.4 on the guide surface 26.1. For this purpose, the rear intersection point 11.5 is located downstream of the platform 26 (near the trailing edge 11.2 of the blade 11), and the second hump 132 includes a vertex 134 whose projection 134.1 on the chord line LC is a distance D2 from a projection 11.6 of the rear intersection point 11.5 on the chord line LC.

[0077] Preferably, the distance D2 corresponds to at most 15% of the chord C, and more preferably, the distance D2 corresponds to 10% or less of the chord C.

[0078] In an alternative not shown, the vertex 134 can be positioned so that its projection 134.1 onto the chord line LC is distant from the leading edge 11.1 by a distance D1 corresponding to at least 10% of the chord C.

[0079] The summit 134 is attached to the extrados 11.4 or disposed at a distance from said extrados 11.4 between 0% and 50%, and preferably between 0% and 30% of the maximum thickness e, similarly to the disposition of the summit 34 of the first hump 32 with respect to the intrados 11.3.

[0080] Advantageously, such an arrangement of the second hump 132 makes it possible to delay or even avoid the separation of the boundary layer at the level of the blade wall on the extrados side 11.4.

[0081] In this configuration, the first bump 32 acts directly on the "horseshoe" vortex phenomenon in order to limit its impact, and the second bump 132 allows in a curative way to compensate for stalls that may be caused at the right of the extrados 11.4 by the phenomenon.

[0082] Preferably, the second hump 132 comprises a truncated cardioid shape, or a substantially circular, oval, or elliptical shape, said second hump 132 being preferably in the shape of a truncated cardioid with a principal orientation substantially parallel to the extrados 11.4, so that the second hump 132 extends along its principal direction which follows the curvature of the extrados 11.4. Indeed, the second hump 132 follows the contour of the extrados 11.4 so as to be able to better guide the airflow and to prevent its separation from said extrados 11.4.

[0083] There [ Fig. 6 ] illustrates a side view of dawn 109 of the [ Fig. 5]. We see that the summit 134 of the second bump 132 rises above the substantially flat surface 26.1 by a height H2.

[0084] Preferably, the height H2 of the top 134 is determined as a function of the maximum thickness e similarly to the height H1 of the first bump 32. For this purpose, the height H2 is between 0.1 and 3 times the thickness e, and more preferably between 0.1 and 0.3 times the maximum thickness e when the second bump 132 is at the top of the blade, and between 1 and 3 times the thickness e when said bump 132 is at the bottom of the blade.

[0085] The height H2 can, for example, correspond to 1 mm when the thickness e is equal to 10 mm, in particular in the case where the second bump 132 is located on the platform at the top of the blade, or said height H2 can be equal to 3 mm if the thickness e is 1 mm, in particular in the case where the second bump 132 is located on the platform at the bottom of the blade.

[0086] However, it should be noted that on the same substantially flat surface 26.1, the height H2 is preferentially less than the height H1.

[0087] Preferably, the height H2 is between 1% and 10% of the chord of blade 11. More preferably, the height H2 is between 0.5 and 5 mm, and more preferably between 1 and 3 mm.

[0088] The second hump 132 can generally have a convex and / or concave shape. Preferably, the second hump 132 has a convex shape at the upstream part corresponding to about half of the extent of said second hump 132 (here the term about corresponds to ±20% of the extent), said second hump 132 has a concave shape on a downstream part corresponding to another half, and more preferably, the convex half is disposed downstream of the hump 132 (closer to the trailing edge 11.2) so as to allow better guidance of the airflow downstream.

[0089] There [ Fig. 7 ] illustrates a radial view of a blade 209 according to a third embodiment of the invention, in which the substantially flat surface 26.1 of the blade 209 comprises only the second bump 132 at the right of the second portion of disk 30.2.

[0090] Indeed, dawn 209, according to the third embodiment, lacks the first hump 32 (illustrated on the figures 1-6 ). In this configuration, during high angles of attack, the second hump 132 helps to maintain the flow of air over the upper surface 11.4 in order to prevent its separation which may be caused by the horseshoe vortex.

[0091] It is understood that the three modes of implementation of figures 3 , 5 And 7can be combined on the same blade: a blade can have a hump on the lower surface and a hump on the upper surface (or vice versa), or have two humps on one of its platforms and none on another.

[0092] Also, different blades of the same row of blades may exhibit one and / or another of the configurations of the three embodiments illustrated.

Claims

1. A variable-pitch stator blade (9; 109) of an aircraft turbomachine (4), the blade (9; 109) comprising: - a airfoil (11) for directing an airflow, the airfoil (11) having a leading edge (11.1) and a trailing edge (11.2), and a pressure side (11.3) and a suction side (11.4) extending from the leading edge (11.1) to the trailing edge (11.2); and - one or two platforms (26) disposed at one end or both ends of the airfoil (11), the one or two platforms (26) comprising a guide surface (26.1) for the airflow, the surface (26.1) comprising a first disk portion (30.1) on the pressure side (11.3) and a second disk portion (30.2) on the suction side (11.4); characterized in that at least one of the first (30.1) or second disk portions (30.2) includes a bump (32; 132).

2. The blade (9; 109) according to claim 1, characterized in that at least one of the first disk portions (30.1) includes a first bump (32), and the peak (34) of said first bump (32) has a position where the projection (34.1) on the chord line (LC) of the blade (9; 109) at the platform (26) is at a distance from the leading edge (11.1) of 10% or less of the chord (C).

3. The blade (9; 109) according to one of the preceding claims, characterized in that the peak (34) of the first bump (32) is adjacent to the pressure side (11.3) or is disposed at a distance from said pressure side (11.3) ranging between 0% and 30% of a maximum thickness (e) of the airfoil (11).

4. The blade (109) according to one of the preceding claims, characterized in that at least one of the second disk portions (30.2) includes a second bump (132).

5. The blade (109) according to claim 4, characterized in that the platform (26) includes a peripheral edge (26.2) that intersects the radial projection of the suction side (11.4) on the guide surface (26.1) at a rear intersection point (11.5), and the peak (134) of the second bump (132) has a position where the projection (134.1) on the chord line (LC) of the blade (109) at the platform (26) is at a distance (D2) from the projection (11.6) of the rear intersection point (11.5) on the chord line (LC), which is less than or equal to 10% of the chord (C).

6. The blade (9; 109) according to one of claims 4 or 5, characterized in that, apart from the bump(s) (32; 132), the surface (26.1) of the first (30.1) and second disk portions (30.2) is substantially flat, and the peak (34; 134) of the bump(s) (32; 132) rises above the substantially flat surface (26.1) by a height (H1; H2) that ranges between 0.5 and 5 mm, preferably between 1 and 3 mm.

7. The blade (9; 109) according to the previous claim, characterized in that the height (H1) of the peak (34; 134) of the first bump (32) is between 0.1 and 0.3 times the maximum thickness (e) of the airfoil (11) or between 1 to 3 times said maximum thickness (e), and / or the height (H2) of the peak (134) of the second bump (132) is between 0.1 and 0.3 times the maximum thickness (e) of the airfoil (11) or between 1 to 3 times said maximum thickness (e).

8. The blade (9; 109) according to claim 6 or 7, characterized in that the height (H1; H2) of the peak (34; 134) of the bump(s) (32; 132) is between 1% and 10% of the chord (C).

9. The blade (9; 109) according to one of the preceding claims, characterized in that the bump(s) (32; 132) are at a distance from the peripheral edge (26.2).

10. The blade (9; 109) according to one of the preceding claims, characterized in that the first bump (32) has a main orientation (32.1) that forms an angle between 30° and 60° with the chord line (LC) of the blade (9; 109).

11. The blade (9; 109) according to one of claims 4 to 10, characterized in that the second bump (132) has a truncated cardioid shape.

12. An aircraft turbomachine (4) comprising a compressor (2), said compressor (2) comprising a guide surface (18.1, 30.1) for a substantially axisymmetric airflow (F1), the turbomachine (4) being characterized in that it comprises an annular row of blades (9; 109) according to one of claims 1 to 11, the platforms (26) of the blades (9; 109) being received in housings (31) of the casing (18, 30) such that the airflow guide surfaces (26.1) of the platforms (26) are substantially flush with the airflow guide surface (18.1, 30.1) of the casing (18, 30).