Variable pitch stator blade including an aerodynamic fin

Abstract

The present invention provides a variable pitch stator blade (14) of a compressor stator (44) of a turbine, which extends along a radial longitudinal plane and includes a radially inner plate (16) and a radially outer plate (18), at least one airfoil (12) extending between the radially inner plate and the radially outer plate, each plate (16, 18) including a transverse longitudinal wall (22) facing the other plate (16, 18), characterized in that the 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).

Description

Technical Field

[0001] The present invention relates to a stator blade of a turbine engine, and more particularly to a compressor stator blade, which is configured to reduce the formation of vortices at the blade root or tip. Background Art

[0002] A turbine engine compressor stage consists of a plurality of moving blades mounted on a rotor and a plurality of fixed blades distributed around the main axis of the compressor, and the plurality of fixed blades form a stator.

[0003] The stator is used to correct the airflow circulating in the main nozzle before it comes into contact with the impeller.

[0004] To improve the performance of the compressor, it is very important to control the aerodynamic losses generated in the airflow passing through the compressor.

[0005] Some aerodynamic losses are generated by secondary airflows and take the form of vortices.

[0006] As can be seen in FIG. 3 for example, some vortices 10, called corner vortices, result from the interaction between the boundary layer flowing along the blade 12 of the blade 14 and the boundary layer flowing along the plate 16 of the blade 14.

[0007] Similarly, some stator stages include variable pitch stator blades, commonly referred to by the acronym "VSV" of "Variable Stator Vane", for which the orientation of each blade 14 of the stage about its main axis can be modified by a suitable mechanism.

[0008] The object of the present invention is to propose a stator blade for a turbine engine compressor, the production of which limits the formation of these corner vortices. Summary of the Invention

[0009] The present invention relates to a variable pitch stator blade of a turbine engine compressor stator, which extends along a longitudinal radial plane including a radially inner plate and a radially outer plate, and at least one blade extends between the radially inner plate and the radially outer plate.

[0010] Each plate includes a transverse longitudinal wall facing the other plate.

[0011] Characterized in that the transverse longitudinal wall of at least one plate includes at least one fin protruding radially towards the other plate.

[0012] The presence of these fins helps to break the vortices formed at the connection between the blade and each plate.

[0013] Preferably, the transverse longitudinal wall includes a plurality of fins distributed transversely on either side of the blade.

[0014] Preferably, each fin is substantially parallel to the vane and is laterally positioned at a distance from the vane.

[0015] Preferably, the vane includes a first upstream longitudinal end edge called the leading edge and a second downstream longitudinal end edge called the trailing edge, and each fin includes an upstream longitudinal end that is longitudinally offset downstream relative to the leading edge.

[0016] 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.

[0017] Preferably, each fin includes a radial end face that is flush with the transverse longitudinal wall of the plate at the upstream longitudinal end of the fin.

[0018] Preferably, the plate includes an outer peripheral surface that radially extends from the outer peripheral edge of the transverse longitudinal wall, and each fin includes a radial end face that is flush with the outer peripheral surface of the plate at the downstream longitudinal end of the fin.

[0019] Preferably, the cross-section of each fin is linear along the transverse longitudinal plane.

[0020] Preferably, the cross-section of each fin is curved along the transverse longitudinal plane and has a curvature similar to that of the vane.

[0021] Preferably, the maximum radial dimension of each fin is less than 25% of the radial distance between the transverse longitudinal walls of the plate.

[0022] The present invention also relates to an aircraft turboengine compressor, which includes a stator formed by a plurality of blades according to the present invention distributed around the main axis A of the compressor, wherein the vanes of the blades are radially oriented relative to the main axis A of the compressor, and is characterized in that each blade is rotatably and movably mounted about its radial main axis. Description of the Drawings

[0023] FIG. 1 is a schematic perspective view of a turboengine compressor including variable pitch stator blades.

[0024] FIG. 2 is a more detailed view at a larger scale of a part of the compressor shown in FIG. 1, which shows certain variable pitch blades.

[0025] FIG. 3 is a detailed view of one end of a variable pitch vane according to the prior art, showing the formation of eddy currents at the connection between the vane and the plate.

[0026] FIG. 4 is a view similar to FIG. 3, showing a blade according to the present invention.

[0027] FIG. 5 is a top view of the blade component shown in FIG. 4, showing two fins on the plate.

[0028] Figure 6 is a side view of the blade shown in Figures 4 and 5. DETAILED DESCRIPTION

[0029] For the purposes of describing the present invention, the radial, longitudinal and transverse orientations according to the markers R, L, T indicated in the figures will be used in a non - limiting manner.

[0030] Figure 1 shows a part of a turbine engine, in particular a part of a high - pressure or low - pressure compressor. This compressor 40 has a main axis A extending along the main longitudinal direction of the turbine engine.

[0031] The compressor 40 includes a number of stages, each of which includes a set of moving blades 42 supported by a rotor disc and a plurality of stator blades 14 forming a stator 44, the stator 44 being designed to redirect the air flow along the main axis direction of the compressor.

[0032] The main direction of each blade 14 of the stator 44 is radial with respect to the main axis A of the compressor 40, i.e., each blade extends along a radial direction R perpendicular to the longitudinal direction L.

[0033] It will be understood that due to the local inclination of the air flow jet in which the blade 14 is located, the blade may be slightly inclined with respect to the radial direction R.

[0034] The stator 44 is of the variable pitch type, i.e., each stator blade 14 is pivotally movable about its main axis, which here extends along the radial direction R, to change its inclination with respect to the plane defined by the longitudinal direction L and the radial direction R of the blade 14, i.e., with respect to the overall flow direction of the air flow in the compressor 40. It will be understood that not all of the stator blades 14 of a turbine engine are of variable pitch. Thus, only the stator blades 14 of the first stage of the compressor 40 are movable, and the blades forming the other stators are fixed in the stator.

[0035] In the following description, only the moving blades 14, i.e., the variable pitch blades, will be referred to. Thus, the term blade 14 will be used only to describe this movable blade 14.

[0036] It will be understood that when the blade 14 pivots in the compressor 40, the associated markers R, L, T pivot in the same way about the radial direction R.

[0037] As can be seen in more detail in Figure 2, each blade 14 includes a radially inner plate 16, which together with a fixed wall 17 of the compressor 40 defines the radial interior of the air flow jet.

[0038] The blade 14 also includes a radially outer plate 18 (see Figure 1), which together with a fixed wall 19 of the compressor 40 defines the radial exterior of the air flow jet.

[0039] Since the blade 14 is a variable pitch blade, the main form of each plate 16, 18 is a disk coaxial with the radial rotation axis of the blade 14 relative to the fixed platform of the compressor.

[0040] Each plate 16, 18 includes an annular outer peripheral wall 20, which is centered on this radial rotation axis and is positioned in a complementary circular hole 21 formed in the associated fixed platform.

[0041] Each plate 16, 18 also includes a transverse longitudinal wall 22 that extends generally longitudinally and transversely and faces the other plate 16, 18.

[0042] Each transverse longitudinal wall 22 extends generally along the extension of the annular wall of the associated platform to reconstitute the air flow jet.

[0043] The blade 14 includes a blade 12 that extends along the main axis of the blade 14, that is, here it extends in the radial direction R between the two plates 16, 18.

[0044] The blade 12 includes an upstream longitudinal end edge 24 commonly referred to as the leading edge, a downstream longitudinal end edge 26 commonly referred to as the trailing edge, a lower surface wall 28, and an upper surface wall 30. The lower surface wall 28 and the upper surface wall 30 laterally define the blade 12 and 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.

[0045] During the operation of the turbomachine, and as can be seen in FIG. 4, the air flow in the jet generates a perturbation 32 at the junction between each radial end of the blade 12 and the transverse longitudinal wall 22 of the plates 16, 18, that is, at the junction between the lower surface wall 28 or the upper surface wall 30 and the transverse longitudinal wall 22 of the plates 16, 18. Here, only the perturbation between the upper surface wall 30 and the transverse longitudinal wall 22 of the radially inner plate 16 is shown.

[0046] These perturbations 32 in the form of vortices tend to increase in the flow direction, that is, from upstream to downstream, that is, they are larger when approaching the trailing edge of the blade 12.

[0047] To limit the spread of the perturbation 32 in the downstream direction, and as can be seen from FIGS. 4 to 6, each transverse longitudinal wall 22 of the plates 16, 18 includes at least one fin 34 that extends by protruding radially with respect to the transverse longitudinal wall 22.

[0048] The fin 34 forms an obstacle to the vortices formed at the plate 16, preventing the perturbation 32 from further increasing or even causing the formation of multiple perturbations 33 with smaller amplitudes.

[0049] According to a preferred embodiment shown in the figures, the transverse longitudinal wall 22 of the plate 16 includes two fins 34 that are laterally distributed on either side of the vane 12. It will be understood that the present invention is not limited to this embodiment, and different numbers of fins 34 can be arranged on either side of the vane 12.

[0050] Each fin 34 extends generally parallel to the vane 12, i.e., in a plane parallel to the longitudinal direction L and the perpendicular direction V in the embodiment shown in the figures.

[0051] The fins are laterally positioned facing the lower surface wall 28 or the upper surface wall 30 and extend laterally away from that wall 28, 30 of the vane 12.

[0052] The cross-section of each fin 34 along the transverse longitudinal plane can be linear or curved. In this case, the curvature of the fin 34 is similar to the curvature of the lower surface wall 28 or the upper surface wall 30 located beside it. Thus, in the latter case, the fin 34 follows the flow direction of the air flow in the stator.

[0053] Each fin 34 is defined by a radial end face 36 and two side surfaces 38 separated from each other by this radial end face 36. Thus, the cross-section of the fin 34 along the transverse radial plane, i.e., along the directions R, T, is generally rectangular.

[0054] According to an alternative embodiment, the acute angle of the ridge of each fin 34 is replaced by a rounded corner, thus forming a fin 34 with a rounded edge.

[0055] The side surfaces are preferably parallel to each other and parallel to the radial direction R of the blade 14.

[0056] The radial dimension of the fin 34 is variable and increases along the longitudinal direction. The radial end face 36 is preferably circular along the radial direction.

[0057] Preferably, the maximum radial dimension of each fin is less than 25% of the radial dimension of the vane 12, i.e., less than 25% of the radial distance between the transverse longitudinal walls 22 of the plates 16, 18.

[0058] This radial dimension is sufficient to prevent vortices having dimensions of the same order of magnitude as this radial dimension without unduly affecting the flow elsewhere.

[0059] 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 can avoid generating an "upward step" effect that is harmful to the performance of the turbomachine.

[0060] Moreover, as can be seen in FIG. 6, at the downstream longitudinal end of the fin 34, i.e., at the downstream end of the fin 34, the radial end face 36 extends substantially radially and is flush with the outer peripheral wall 20 of the plate 16.

[0061] Thus, at each longitudinal end thereof, the radial end face 36 is flush with, or extends from, the transverse longitudinal wall 22 or the outer peripheral wall 20 of the plate 16.

[0062] As a result, there is no detachment between the radial end face 36 of the fin 34 and the transverse longitudinal wall 22 or the outer peripheral wall 20 of the plate 16, which can prevent the formation of additional vortices that may be harmful to the performance of the blade 14.

[0063] To limit the disturbances in the air that may be generated by the fins, the upstream longitudinal end of each fin 34 is longitudinally offset downstream relative to the leading edge 24 of the blade 12.

[0064] Preferably, this downstream longitudinal offset is at least equal to 30% of the longitudinal length of the plate 16.

[0065] The region located at the leading edge 24 of the blade 12 is a region where the air flow is not fully oriented parallel to the blade. For certain operating points, an aerodynamic blockage region (with a high Mach number) may occur near the leading edge 24. Therefore, protrusions should be avoided in this region.

[0066] As described above, the blade 14 rotates and moves relative to the compressor stator about a radial axis.

[0067] Since the fins 34 are mounted on the plate 16, the fins 34 can also rotate and move integrally with the blade 14, ensuring good rectification efficiency of the air flow in the compressor and reducing the disturbances 32 commonly referred to as "corner vortices".

[0068] According to a preferred embodiment, each fin 34 is integrally formed with the plate 16, and the plate 16 is associated with the fin.

[0069] This embodiment can be achieved by molding or any other embodiment, such as by additive techniques.

[0070] The above description of one or two fins applies to the fins supported by the plate 16 located at the radially inner end of the blade 12. It will be understood that this specification will apply in the same manner to each fin 34 supported by the plate 18 located at the radially outer end of the blade 12.

Claims

1. A variable pitch stator vane (14) for a turbine engine compressor stator (44), which extends along a longitudinal radial plane including a radially inner plate (16) and a radially outer plate (18), and at least one vane (12) extends between the radially inner plate and the radially outer plate, each plate (16, 18) includes a transverse longitudinal wall (22) facing the other plate (16, 18), characterized in that, the 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), the vane (12) includes a first upstream longitudinal end edge (24) called the leading edge and a second downstream longitudinal end edge (26) called the trailing edge, and each fin (34) includes an upstream longitudinal end longitudinally offset downstream relative to the leading edge (24).

2. The vane (14) according to claim 1, characterized in that, the transverse longitudinal wall (22) includes a plurality of fins (34) distributed transversely on either side of the vane (12).

3. The vane (14) according to claim 1 or 2, characterized in that, each fin (34) is substantially parallel to the vane (12) and is transversely positioned at a distance from the vane (12).

4. The vane (14) according to claim 3, characterized 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 plates (16, 18).

5. The vane (14) according to claim 1 or 2, characterized in that, each fin (34) includes a radial end face (36), and the radial end face is flush with the transverse longitudinal wall (22) of the plates (16, 18) at the upstream longitudinal end of the fin (34).

6. The vane (14) according to claim 1 or 2, wherein, the plates (16, 18) include an outer peripheral surface (20) extending radially from the outer peripheral edge of the transverse longitudinal wall (22), characterized in that each fin (34) includes a radial end face (36), and the radial end face is flush with the outer peripheral surface of the plates (16, 18) at the downstream longitudinal end of the fin (34).

7. The vane (14) according to claim 1 or 2, characterized in that, the cross-section of each fin (34) is linear along a transverse longitudinal plane.

8. The vane (14) according to claim 1 or 2, characterized in that, the cross-section of each fin (34) is curved along a transverse longitudinal plane and has a curvature similar to that of the vane (12).

9. The vane (14) according to claim 1 or 2, characterized 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. An aircraft turbine engine compressor, including a stator (44), the stator being formed by a plurality of vanes (14) according to any one of claims 1-9 distributed around the main axis (A) of the compressor, wherein, The blade (12) of the paddle is radially oriented relative to the main axis (A) of the compressor. It is characterized in that each of the blades (14) is rotatably mounted about its radial main axis.

Images