RECTIFIER BLADE FOR AN AIRCRAFT TURBOMACHINE

Abstract

A rectifier vane (14) for an aircraft turbomachine compressor, this vane (14) comprising variable-pitch vanes (16) distributed around a central axis (X), each of these vanes (16) comprising a lower surface (16a) and an upper surface (16b) which join to form a leading edge (16c) and a trailing edge (16d), and being rotatable around a radial axis (A) relative to said central axis (X) and defining with an adjacent vane (16) a channel (24) for the passage of gas between the vanes (16), each channel (24) comprising a minimum section, called the throat section (Sc), an inlet section (Se) which is located upstream of the throat section (Sc) and at a leading edge (16c), and an outlet section (Ss) which is located downstream of the throat section (Sc) and at of a trailing edge (16d). Figure for abstract: Figure 6

Description

Title of the invention: RECTIFIER BLADE FOR AN AIRCRAFT TURBOMACHINE Technical field of the invention

[0001] The present invention relates to a rectifier blade for an aircraft turbomachine compressor, as well as to a compressor and a turbomachine comprising such a blade. Technical background

[0002] The state of the art includes in particular documents FR-A1-2 853 000 and FR-Al-3 108 370.

[0003] An aircraft turbomachine compressor 10 as illustrated in [Fig.l] comprises a rotor and a stator. The rotor is formed by bladed wheels 12 and the stator is formed by stator blades 14 commonly called rectifiers, stator grids or rectifier blades. A rotor wheel 12 comprises a disk carrying rotor blades at its periphery. The stator blades 14 are interposed between the rotor wheels 12 and each comprise stator blades 16. The association of a rotor wheel 12 with a stator blade 14 forms a compression stage and a compressor can comprise several successive stages along the axis X of the compressor and the turbomachine.

[0004] The stator blades 14 are generally variable-pitch (and therefore of the VSV type, which is the acronym for Variable Stator Varies) so as to adapt the incidence of the gases on the blades 16 according to the speed of the turbomachine.

[0005] Each of the blades 16 of a stator blade 14 is thus rotatable about an axis A substantially radial relative to the central axis X of the compressor. Each blade 16 can adopt any position between an open position, also called a flag position, in which it is substantially parallel to the central axis X, and an extreme closed position in which it is inclined relative to this central axis X. There are several closed positions and the extreme closed position is the maximum closing position of the compressor stream. [Fig. 2] shows for example two positions of setting of a blade 16 about its axis A.

[0006] [Fig.3] shows the compressor vein 18 which is crossed by the blades 16. The The lines shown in this figure illustrate streamlines 20. The vein 18 can be broken down into several stream tubes 22, centered on the X axis, and which are arranged around each other. A stream tube 22 extends between two adjacent stream lines 20.

[0007] Each current tube 22 can further be decomposed into several channels 24, which are arranged circumferentially next to each other around the X axis. The vein 18 could also be considered as being broken down into several channels 24, which are arranged circumferentially next to each other around the X axis. A channel 24 extends between two adjacent blades 16.

[0008] The blades 16 thus define between them channels 24 for the passage of gases, as illustrated in [Fig. 4]. Thus, a blade 16 which is interposed between two other blades defines a first channel 24 with one of these blades 16, and a second channel 24 with the other of these blades 16.

[0009] Each blade 16 has an aerodynamic profile and comprises a lower surface 16a and an upper surface 16b which join upstream to form a leading edge 16c and downstream to form a trailing edge 16d. It is thus understood that a channel 24 is defined between the lower surface 16a of a blade 16 and the upper surface 16b of an adjacent blade 16.

[0010] Each channel 24 has a section S which varies along the central axis X and which depends on the setting position of the blades 16 ([Fig.4]). Each channel 24 has a minimum section, called the throat section Sc, an inlet section Se which is located upstream of the throat section Sc and at a leading edge 16c, and an outlet section Ss which is located downstream of the throat section Sc and at a trailing edge 16d.

[0011] In the open position, the neck section Sc of each channel 24 has a maximum value. In the extreme closed position, the neck section Sc of each channel 24 has a minimum value.

[0012] The aerodynamic objectives of a rectifier blade 14 are to ensure the deflection of the gases (change in angle of the gas flow between the inlet and the outlet of the blade), to generate a minimum of aerodynamic losses, and to be operable throughout the operating range of the engine.

[0013] The aerodynamic profiles of the blades 16 are therefore designed to meet these needs. For this, the profiles are characterized under several aerodynamic conditions (angle, flow speed, etc.) corresponding to the different operating points.

[0014] The rectifier vanes 14 have the constraint of being evaluated with a different pitch angle depending on the engine speeds considered. Thus, the profile of a vane 16 of this type of vane 14 is produced by considering several aerodynamic profiles: the profile at nominal pitch and the profile(s) of the other specification point(s).

[0015] A rectifier vane 14 has the difficulty of requiring optimization not only in several operating points but also for several profiles, because depending on the engine speeds, the timing law will be controlled, and therefore the timing of the blades 16 will be modified. Thus, a blade 16 which would be optimized for a timing and triangulation of a high speed point (open timing, zero and positive incidence) can be strongly mismatched when the blade 16 is in closed timing and in very high incidence negative. This negative incidence can even lead to significant separations on the intrados 14a of the blade 16. In addition, the separation topology will significantly reorient the outlet angle of the flow, this angle will be significantly different from the skeleton angle of the blade. Failure to control this angle of the flow at the outlet of the blade 14 can have considerable impacts on the compressor and the supply of the blades located downstream, and therefore an impact on the performance and the pumping margin of the compressor.

[0016] The present invention provides an improvement to the current technique, which is simple, effective and economical. Summary of the invention

[0017] The invention relates to a stator vane for an aircraft turbomachine compressor, this vane comprising variable-pitch blades distributed around a central axis, each of these blades comprising a lower surface and an upper surface which join to form a leading edge and a trailing edge, each of these blades being rotatable around a radial axis relative to said central axis and defining with an adjacent blade a gas passage channel between the blades, each channel comprising a minimum section, called the throat section, an inlet section which is located upstream of the throat section and at a leading edge, and an outlet section which is located downstream of the throat section and at a trailing edge, the blades being movable around their radial axes from an open position in which the throat section of each channel has a maximum value,up to an extreme closed position in which the throat section of each channel has a minimum value, characterized in that the blades are profiled so that, in the extreme closed position, the section of each channel decreases between the inlet section and the throat section, then increases continuously between the throat section and the outlet section.

[0018] The present invention thus proposes a particular evolution of the section of each channel when the blades are in the extreme closed position. The inventors have thus chosen to optimize this parameter in this extreme position, which means that this parameter can evolve differently when the blades are in another position. Advantageously, this optimization is made for a reference point such as for example a ground idle point of the engine, a point at which the compressor must not make too many losses under penalty of generating penalizing efficiencies and therefore excessively high temperatures in the turbine of the turbomachine.

[0019] The blading according to the invention may comprise one or more of the following characteristics, taken in isolation from one another, or in combination with one another:

[0020] - the blades are profiled so that a curve of evolution of the section of each channel along the central axis has a tangent at the neck section;

[0021] - in the extreme closed position, the outlet section of each channel defines the value maximum of the canal section;

[0022] - each of the blades includes a bump on its lower surface which extends substantially over the entire extent of the blade along the radial axis and which defines said section at the neck;

[0023] - the boss of each blade is located at a distance from the leading edge of the blade cor corresponding, which represents 3 to 10% of a length of the channel defined by this blade measured along said central axis.

[0024] The present invention also relates to a compressor, in particular a high pressure compressor, for an aircraft turbomachine, comprising at least one blade as described above.

[0025] The invention further relates to a turbomachine, in particular for an aircraft, comprising at least one blade or compressor as described above. Brief description of the figures

[0026] Other characteristics and advantages will emerge from the following description of a non-limiting embodiment of the invention with reference to the appended drawings in which:

[0027] [Fig-1] [Fig.l] is a partial schematic view in axial section of a compressor aircraft turbomachine,

[0028] [Fig.2] [Fig.2] is a schematic view of a variable pitch vane of a rectifier vane,

[0029] [Fig.3] [Fig.3] is a partial schematic view in axial section of a rectifier blade,

[0030] [Fig.4] [Fig.4] is a partial schematic perspective view of two blades of a rectifier blade,

[0031] [Fig.5a-5c] Figures 5a-5c are graphs showing the evolution of the section of a channel along the central axis of the blading, according to three possible configurations; figure 5c represents the configuration according to the invention and figures 5a and 5b represent other configurations,

[0032] [Fig.6] [Fig.6] is a partial schematic view of a rectifier blade showing the profiles of the blades according to the three configurations of Figures 5a-5c, and

[0033] [Fig.7a-7c] Figures 7a are schematic perspective views of a blade of a stator blade according to the three configurations of figures 5a-5c, and show the influence of these configurations on the separations. Detailed description of the invention

[0034] Figures 1 to 4 have been described in the above.

[0035] As illustrated in these figures, the invention relates to a rectifier vane 14 for an aircraft turbomachine compressor 10.

[0036] The blading 14 comprises variable-pitch blades 16 distributed around a central axis X.

[0037] Each blade 16 has a lower surface 16a and an upper surface 16b which join to form a leading edge 16c and a trailing edge 16d.

[0038] Each blade 16 is rotatable about a radial axis A relative to the central axis X and defines with an adjacent blade 16 a channel 24 for the passage of gas between the blades 16.

[0039] Each channel 24 comprises a minimum section, called the throat section Sc, an inlet section Se which is located upstream of the throat section Sc and at a leading edge 16c, and an outlet section Ss which is located downstream of the throat section Sc and at a trailing edge 16d.

[0040] The expressions "upstream" and "downstream" refer to the flow of gases through the blades 16 in operation, this flow taking place along the central axis X from the leading edges 16c of the blades 16 to their trailing edges 16d.

[0041] The blades 16 are movable around their axes A from an open position (not shown) in which the throat section Sc of each channel 24 has a maximum value, to an extreme closed position (illustrated in [Fig.4]) in which the throat section Sc of each channel 24 has a minimum value.

[0042] Each channel 24 has a length L measured along the X axis ([Fig.4]). This length L is the distance between the inlet and outlet sections Se, Ss.

[0043] Several configurations exist for the evolution of the section S of each channel 24 along the axis X and this evolution can also be different depending on the setting position of the blades 16.

[0044] In the present invention, we are only interested in the evolution of the section S of the channel 24 when the blades 16 are in the extreme closed position. There are several configurations, three of which are shown in Figures 5a to 5c.

[0045] Each of these figures shows the evolution of the section S along the X axis.

[0046] Figure 5c illustrates the present invention and Figures 5a and 5b illustrate confi figures outside the scope of protection of the invention and are provided as comparative examples.

[0047] In Figure 5c, the blades 16 are profiled so that, in the extreme closed position, the section S of each channel 24 decreases between the inlet section Se and the throat section Sc, then increases continuously between the throat section Sc and the outlet section Ss. It is thus understood that the channel 24 first comprises a convergent part C1 (linked to a reduction in the section), then a divergent part C2 (linked to an increase in the section).

[0048] The blades 16 are thus profiled so that the evolution curve of the section S of each channel 24 along the X axis has a general U shape.

[0049] Advantageously, the evolution curve of the section S of each channel 24 has a tangent T1 at the level of the neck section Sc.

[0050] Advantageously, the evolution curve of the section S of each channel 24 has another tangent T2 at the level of the output section Ss.

[0051] A tangent of the curve means that the section S evolves gradually and therefore without sudden change. A sudden change in the section would cause a peak to appear on the curve. On the contrary, the lower end of the curve is curved.

[0052] The or each tangent T1, T2 is preferably horizontal, i.e. parallel to the abscissa axis X.

[0053] [Fig.6] illustrates an embodiment of the profiles of the blades 16 which makes it possible to obtain the evolution of the section S illustrated in figure 5c. The blades 16 are shown in the extreme closed position in this [Fig.6].

[0054] The contour line H1 designates the profiles of the blades 16 according to the configuration of FIG. 5c.

[0055] Preferably, each blade 16 comprises a boss 30 on its intrados 16a which extends substantially over the entire extent of the blade 16 along the axis A and which defines the neck section Sc (figures 6 and 7c). In other words, the neck section Sc is measured between this boss 30 and the extrados 16b opposite the adjacent blade 16.

[0056] The hump 30 of a blade 16 or the neck section Sc defined by this blade 16, is preferably located between 3 and 10% of the length L of the channel 24 defined by this blade 16, measured from the leading edge of the blade 16.

[0057] Figure 7c shows the gas separation zones at the level of the intrados 16a of the blade 16. It can be seen that these zones ZI are relatively small.

[0058] In Figure 5a, the blades 16 are profiled so that, in the extreme closed position, the section S of each channel 24 decreases between the inlet section Se and the throat section Sc, then increases and decreases between the throat section Sc and the outlet section Ss. The channel 24 is therefore convergent (part Cl of the curve), divergent (part C2) then convergent again (part C3).

[0059] In Figure 5b, the blades 16 are profiled so that, in the closed extreme position, the section S of each channel 24 increases continuously between the inlet section Se and the outlet section Ss. The channel 24 is therefore entirely divergent (part C2).

[0060] The contour lines H2 and H3 in [Fig.6] respectively designate the profiles of the blades 16 according to the configurations of FIGS. 5a and 5b.

[0061] Figure 7a and 7b show the gas separation zones Z2, Z3 at the level of the intrados of the blade 16, respectively according to the configurations of Figures 5a and 5b. It can be seen that these zones Z2, Z3 are larger than the zone of Figure 7c.

[0062] Indeed, in the case of figures 5a and 7a, the diverging part is too short which generates more detachments. In the case of figures 5b and 7b, the neck section Sc is at the level of the leading edge 16c which creates poor incidence resistance and detachments.

[0063] The three configurations were evaluated by aerodynamic calculations and showed:

[0064] - in the extreme closed position and at a typical slow-motion point, a clear reduction in losses for the configuration according to the invention,

[0065] - in open position and at a high speed type point: a small increase in losses, a priori negligible given the gains made at low speed.

[0066] A positive effect of the configuration according to the invention has also been observed on the exit angles of the blades 16 in the extreme closed position.

[0067] The invention makes it possible to significantly reduce separations, losses and flow angle disturbances when the blades 16 are in the extremely closed position, without significantly penalizing losses in another position.

Claims

Claims

1. A rectifier vane (14) for an aircraft turbomachine compressor, this vane (14) comprising variable-pitch vanes (16) distributed around a central axis (X), each of these vanes (16) comprising a lower surface (16a) and an upper surface (16b) which join to form a leading edge (16c) and a trailing edge (16d), each of these vanes (16) being rotatable around a radial axis (A) relative to said central axis (X) and defining with an adjacent vane (16) a channel (24) for the passage of gas between the vanes (16), each channel (24) comprising a minimum section, called the throat section (Sc), an inlet section (Se) which is located upstream of the throat section (Sc) and at a leading edge (16c), and an outlet section (Ss) which is located downstream of the section at the neck (Sc) and at the level of a trailing edge (16d),the blades (16) being movable about their radial axes (A) from an open position in which the neck section (Sc) of each channel (24) has a maximum value, to an extreme closed position in which the neck section (Sc) of each channel (24) has a minimum value, the blades (16) being profiled so that, in the extreme closed position, the section (S) of each channel (24) decreases between the inlet section (Se) and the neck section (Sc), then increases continuously between the neck section (Sc) and the outlet section (Ss), characterized in that each of the blades (16) comprises a boss (30) on its intrados (16a) which extends substantially over the entire extent of the blade (16) along the radial axis (A) and which defines said section at the neck (Sc).,

2. Blading (14) according to claim 1, in which the blades (16) are profiled so that a curve of evolution of the section (S) of each channel (24) along the central axis (X) has a tangent (Tl) at the level of the neck section (Sc).

3. Blading (14) according to claim 1 or 2, in which, in the closed extreme position, the outlet section (S) of each channel (24) defines the maximum value of the section (S) of the channel (24).

4. Blading (14) according to one of the preceding claims, in which the boss (30) of each blade (16) is located at a distance from the leading edge (16c) of the corresponding blade (16), which represents 3 to 10% of a length of the channel (24) defined by this blade (16) measured along said central axis (X).

5. Compressor (10), in particular high pressure, for a turbomachine aircraft, comprising at least one blade (14) according to one of the preceding claims.

6. Aircraft turbomachine, comprising at least one blade (14) according to one of claims 1 to 4 or a compressor (10) according to claim 5.

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