Stator blade with variable pitch, featuring aerodynamic fins

Description

TECHNICAL FIELD

[0001] The invention relates to a turbomachine stator blade, in particular a compressor stator blade, which is shaped to reduce the formation of vortices at the root or tip of the blade. STATE OF THE PRIOR ART

[0002] A turbomachine compressor stage is composed of a plurality of moving blades mounted on a rotor and a plurality of fixed blades, distributed around the main axis of the compressor, which form a rectifier.

[0003] The rectifier is intended to straighten an airflow circulating in the primary vein, before it comes into contact with an impeller. An example is disclosed in US patent application 6,283,705 B1.

[0004] In order to improve compressor performance, it is important to control the aerodynamic losses that are generated in the airflow through the compressor.

[0005] Among the aerodynamic losses, some are generated by secondary air flows, taking the form of vortices.

[0006] As can be seen as an example in the figure 3 , certain vortices 10, called corner vortices, result from the interaction between the boundary layer flowing along the blade 12 of a vane 14 with the boundary layers flowing along a plate 16 of the vane 14.

[0007] Also, some rectifier stages include variable pitch vanes, commonly referred to by the acronym "VSV" for "Variable Statoric Vane", for which the orientation of each vane 14 of this stage, around its main axis, is capable of being modified by an appropriate mechanism.

[0008] The aim of the invention is to propose a rectifier blade for a turbomachine compressor which is designed to limit the formation of these corner vortices. STATEMENT OF THE INVENTION

[0009] The invention relates to a variable-pitch stator vane for a turbomachine compressor rectifier, which extends in a longitudinal radial plane comprising a radially inner plate and a radially outer plate between which at least one blade extends, each plate comprising a transverse longitudinal wall facing the other plate, said transverse longitudinal wall of at least one plate comprises at least one fin projecting radially in the direction of the other plate in which the blade comprises a first upstream longitudinal end edge called the leading edge and a second downstream longitudinal end edge called the trailing edge, characterized in that each fin comprises an upstream longitudinal end offset longitudinally downstream relative to the leading edge.

[0010] The presence of the fins helps to break up the vortices forming at the connection between the blade and each plate.

[0011] Preferably, said transverse longitudinal wall comprises several fins distributed transversely on either side of the blade.

[0012] Preferably, each fin is substantially parallel to the blade and is located transversely away from the blade.

[0013] Preferably, the longitudinal distance between the upstream longitudinal end of each fin and the leading edge is at least equal to 30% of the longitudinal length of the plate.

[0014] Preferably, each fin has a radial end face which is flush with the transverse longitudinal wall of the plate at the upstream longitudinal end of the fin.

[0015] Preferably, the plate has a peripheral face extending radially from the peripheral edge of the transverse longitudinal wall, and each fin has a radial end face which is flush with the peripheral face of the plate at the downstream longitudinal end of the fin.

[0016] Preferably, the section of each fin, along a transverse longitudinal plane, is rectilinear.

[0017] Preferably, the section of each fin, along a transverse longitudinal plane, is curved and has a curvature similar to the curvature of the blade.

[0018] Preferably, the maximum radial dimension of each fin is less than 25% of the radial distance between the transverse longitudinal walls of the plates.

[0019] The invention also proposes an aircraft turbomachine compressor comprising a rectifier formed from a plurality of blades according to the invention, distributed around a main axis A of the compressor, with their blades oriented radially relative to the main axis A of the compressor, characterized in that each blade is mounted to rotate about its main radial axis. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] [ Fig. 1 ] is a schematic perspective representation of a turbomachine compressor incorporating variable-pitch stator blades. [ Fig. 2 ] is a larger-scale detail of part of the compressor shown in the figure 1 , showing some variable pitch blades. [ Fig. 3 ] is a detail of one end of a variable-pitch blade according to the prior art, showing the formation of vortices at the connection between the blade and a plate. [ Fig. 4 ] is a view similar to that of the figure 3, representing a blade according to the invention. [ Fig. 5 ] is a top view of the blade portion shown in figure 4 , showing two fins on the plate. [ Fig. 6 ] is a side view of the dawn shown in figures 4 And 5 . DESCRIPTION OF EMBODIMENTS

[0021] For the description of the invention, the radial, longitudinal and transverse orientations according to the reference R, L, T indicated in the figures will be adopted without limitation.

[0022] We represented at the figure 1 a part of a turbomachine, in particular a high or low pressure compressor. This compressor 40 has a main axis A extending in the main longitudinal direction of the turbomachine.

[0023] The compressor 40 comprises several stages, each of which comprises a set of moving blades 42 carried by a rotor disk and a plurality of stator blades 14 forming a rectifier 44 intended to reorient the air flow in the main axial direction of the compressor.

[0024] The main orientation of each blade 14 of the rectifier 44 is radial relative to the main axis A of the compressor 40, that is to say that each blade extends in the radial direction R perpendicular to the longitudinal direction L.

[0025] It will be understood that the blade may be slightly inclined relative to the radial direction R, due to the local inclination of the air flow vein in which the blade 14 is located.

[0026] The rectifier 44 is of the variable pitch type, that is to say that each stator blade 14 is pivotally movable about its main axis, which here extends in the radial direction R, to modify its inclination relative to a plane defined by the longitudinal direction L and the radial direction R of the blade 14, that is to say relative to the overall flow direction of the air flow in the compressor 40. It will be understood that not all the stator blades 14 of the turbomachine are variable pitch. Thus, only the stator blades 14 of the first stages of the compressor 40 are movable, the blades constituting the other rectifiers are fixed in the rectifier.

[0027] In the following description, reference will only be made to a movable blade 14, i.e. a variable-pitch blade. Consequently, only the term blade 14 will be used to describe this movable blade 14.

[0028] It will be understood that during a pivoting of the blade 14 in the compressor 40, the reference R, L, T which is associated with it pivots identically around the radial direction R.

[0029] As can be seen in more detail in the figure 2 , each blade 14 comprises a radially internal plate 16 which delimits, with a fixed wall 17 of the compressor 40, the radially internal part of the air flow vein.

[0030] The blade 14 also includes a radially external plate 18 (visible at the figure 1 ), which delimits, with a fixed wall 19 of the compressor 40, the radially external part of the air flow vein.

[0031] Since the blade 14 is a variable-pitch blade, each plate 16, 18 is in the main form of a disc coaxial with the radial axis of rotation of the blade 14 relative to the fixed platforms of the compressor.

[0032] Each plate 16, 18 comprises an annular peripheral wall 20 which is centered on the radial axis of rotation and which is located in a complementary circular orifice 21 formed in the associated fixed platform.

[0033] Each plate 16, 18 also comprises a transverse longitudinal wall 22 which extends mainly longitudinally and transversely and which faces the other plate 16, 18.

[0034] Each transverse longitudinal wall 22 extends mainly in the extension of the annular wall of the associated platform, to reconstitute the air circulation vein.

[0035] The blade 14 comprises a vane 12 which extends along the main axis of the blade 14, that is to say here in the radial direction R between the two plates 16, 18.

[0036] The blade 12 comprises an upstream longitudinal end edge 24 commonly called the leading edge, a downstream longitudinal end edge 26 commonly called the trailing edge, a lower surface wall 28 and an upper surface wall 30 which transversely delimit the blade 12 and which extend longitudinally between the leading edge 24 and the trailing edge 26 and radially between the two transverse longitudinal walls 22 of the plates 16, 18.

[0037] During the operation of the turbomachine, and as can be seen in the figure 4, the circulation of air in the vein produces disturbances 32 at the connection between each radial end of the blade 12 and the transverse longitudinal wall 22 of a plate 16, 18, that is to say at the connection between the intrados wall 28 or the extrados wall 30 with the transverse longitudinal wall 22 of a plate 16, 18. Here, only the disturbances between the extrados wall 30 and the transverse longitudinal wall 22 of the radially internal plate 16 have been represented.

[0038] These disturbances 32, which take the form of vortices, tend to amplify according to the direction of flow, that is to say from upstream to downstream, that is to say they are more significant as they approach the trailing edge of the blade 12.

[0039] To limit the expansion of disturbances 32 downstream, and as can be seen in figures 4 to 6, each transverse longitudinal wall 22 of a plate 16, 18, comprises at least one fin 34 which extends by projecting relative to said transverse longitudinal wall 22 in the radial direction.

[0040] The fin 34 forms an obstacle to the vortex flow forming at the level of the plate 16, preventing the disturbance 32 from developing further, or even causing the formation of several disturbances 33 of lesser amplitude.

[0041] According to a preferred embodiment shown in the figures, the transverse longitudinal wall 22 of the plate 16 comprises two fins 34 which are distributed transversely on either side of the blade 12. It will be understood that the invention is not limited to this embodiment, and that a different number of fins 34 can be arranged on one or the other side of the blade 12.

[0042] Each fin 34 extends substantially parallel to the blade 12, that is to say in a plane parallel to the longitudinal L and vertical V directions in the mode shown in the figures.

[0043] The fin is located transversely opposite the intrados wall 28 or the extrados wall 30 and it extends transversely at a distance from this wall 28, 30 of the blade 12.

[0044] The section of each fin 34 along a transverse longitudinal plane may be rectilinear, or curved, and in this case, the curvature of the fin 34 is similar to the curvature of the intrados wall 28 or the extrados wall 30, next to which the fin 34 is located. Thus, in the latter case, the fin 34 follows the direction of flow of the air flow in the rectifier.

[0045] Each fin 34 is delimited by a radial end face 36 and two lateral faces 38 which are separated from each other by the radial end face 36. The section of the fin 34 along a transverse radial plane, that is to say along the directions R, T is thus substantially rectangular.

[0046] According to an alternative embodiment, the sharp angles of the edges of each fin 34 are replaced by fillets, thus forming a fin 34 with rounded edges.

[0047] The side faces are preferably parallel to each other and parallel to the radial direction R of the blade 14.

[0048] The radial dimension of the fin 34 is variable and increases in the longitudinal direction. The radial end face 36 is preferably curved in the radial direction.

[0049] Preferably, the maximum radial dimension of each fin is less than 25% of the radial dimension of the blade 12, that is to say less than 25% of the radial distance between the transverse longitudinal walls 22 of the plates 16, 18.

[0050] This radial dimension is sufficient to combat vortices, which are of the same order of size as this radial dimension, without affecting the flow too much elsewhere.

[0051] The radial end face 36 is flush with the transverse longitudinal wall 22 of the plate 16 at the upstream longitudinal end of the fin 34. This makes it possible to avoid having an "upward step" effect, which would be detrimental to the efficiency of the turbomachine.

[0052] Moreover, as can be seen in the figure 6, at the downstream longitudinal end of the fin 34, that is to say at the downstream end of the fin (34), the radial end face 36 is curved to extend mainly radially and to be flush with the peripheral wall 20 of the plate 16.

[0053] Thus, at each of its longitudinal ends, the radial end face 36 is flush with, or in the extension of, the transverse longitudinal wall 22 or the peripheral wall 20 of the plate 16.

[0054] There is then no detachment between the radial end face 36 of the fin 34 and the transverse longitudinal wall 22 or the peripheral wall 20 of the plate 16, which makes it possible to avoid the formation of additional disturbances which could harm the performance of the blade 14.

[0055] In order to limit the disturbances in the air flow that a fin can produce, the upstream longitudinal end of each fin 34 is offset longitudinally downstream relative to the leading edge 24 of the blade 12.

[0056] Preferably, this longitudinal offset downstream is at least equal to 30% of the longitudinal length of the plate 16.

[0057] The area located at the leading edge 24 of the blade 12 is an area in which the airflow is not yet completely oriented parallel to the blade. An aerodynamic blocking area (at high Mach) may appear near the leading edge 24 for certain operating points, so projections in this area must be avoided. As previously stated, the blade 14 is rotatable relative to the compressor stator around a radial axis.

[0058] Since the vanes 34 are mounted on the plate 16, the vanes 34 are also movable in rotation in a manner integral with the blade 14, guaranteeing good efficiency in the straightening of the air flow in the compressor and the reduction of disturbances 32 commonly called "corner vortices".

[0059] According to a preferred embodiment, each fin 34 is made integrally with the plate 16 associated with it.

[0060] This production can be by molding or any other method, such as additive technology.

[0061] The description which has just been given for one or two fins applies to the fins which are carried by the plate 16 located at the internal radial end of the blade 12. It will be understood that this description will apply identically to each fin 34 which is carried by the plate 18 located at the external radial end of the blade 12.

Claims

1. Variable-pitch stator vane (14) for a turbine engine compressor stator (44), which extends in a longitudinal radial plane including a radially inner plate (16) and a radially outer plate (18) between which at least one blade (12) extends, each plate (16, 18) including a transverse longitudinal wall (22) facing the other plate (16, 18), wherein said transverse longitudinal wall (22) of at least one plate (16, 18) includes at least one fin (34) protruding radially towards the other plate (16, 18), wherein the blade (12) includes a first upstream longitudinal end edge (24) referred to as the leading edge and a second downstream longitudinal end edge (26) referred to as the trailing edge, characterised in that each fin (34) includes an upstream longitudinal end longitudinally offset downstream in relation to the leading edge (24).

2. Vane (14) according to the preceding claim, characterised in that said transverse longitudinal wall (22) includes several fins (34) distributed transversally on either side of the blade (12).

3. Vane (14) according to any one of the preceding claims, characterised in that each fin (34) is substantially parallel with the blade (12) and is located transversally at a distance from the blade (12).

4. Vane (14) according to any one of the preceding claims, characterised in that the longitudinal distance between the upstream longitudinal end of each fin (34) and the leading edge (24) is at least equal to 30% of the longitudinal length of the plate (16, 18).

5. Vane (14) according to any one of the preceding claims, characterised in that each fin (34) includes a radial end face (36) which is flush with the transverse longitudinal wall (22) of the plate (16, 18) at the upstream longitudinal end of the fin (34).

6. Vane (14) according to any one of the preceding claims, wherein the plate (16, 18) includes a peripheral face (20) extending radially from the peripheral edge of the transverse longitudinal wall (22), characterised in that each fin (34) includes a radial end face (36) which is flush with the peripheral face of the plate (16, 18) at the downstream longitudinal end of the fin (34).

7. Vane (14) according to any one of the preceding claims, characterised in that the cross-section of each fin (34), along a transverse longitudinal plane is rectilinear.

8. Vane (14) according to any one of claims 1 to 6, characterised in that the cross-section of each fin (34), along a transverse longitudinal plane is curved and has a similar curvature to the curvature of the blade (12).

9. Vane (14) according to any one of the preceding claims, characterised in that the maximum radial dimension of each fin (34) is less than 25% of the radial distance between the transverse longitudinal walls of the plates (16, 18).

10. Aircraft turbine engine compressor including a stator (44) formed from a plurality of vanes (14) according to the preceding claims, distributed about a main axis (A) of the compressor, with the blades (12) thereof radially oriented in relation to the main axis (A) of the compressor, characterised in that each blade (14) is movably mounted in rotation about the radial main axis thereof.