Vane for an aircraft turbine engine compressor, compressor, turbine engine

Abstract

The present invention relates to a stator vane (121) for an aircraft turbine engine compressor, the vane comprising a pressure side (16) and a suction side (18), an equidistant line (20) between the pressure side and the suction side, a leading edge (22) and a trailing edge (24), a chord (26) extending between the leading edge and the trailing edge, the vane being able to extend in an air flow path (12), the vane having an angle (29) between the tangent (28) to the equidistant line (20) and the engine axis (X), wherein an average change in the angle (29) between the leading edge and the trailing edge is smaller than the average change over 60% of the chord measured from the leading edge to the trailing edge. The invention also relates to a compressor, a turbine engine assembly and an aircraft turbine engine.

Description

[0001] Blade for aircraft turbomachine compressor, compressor, turbomachine

[0002] Technical field

[0003] The invention relates to a blade for an aircraft turbomachine compressor, a compressor and an aircraft turbomachine.

[0004] Prior art

[0005] Aircraft turbomachine compressors are equipped with blades whose shape is decisive for the operating quality of the compressors. Documents US8152473, US22016 / 169243, FR3131754, US10480532 and US2002 / 150471 describe blade shapes.

[0006] In particular, compressor blades are expected to have a strong airflow straightening capability, with the airflow reaching the blade rows at a high incidence.

[0007] There is therefore a need for aircraft turbomachine compressor blades providing large airflow deflections, up to 60°, and operation over a wide angle of incidence, up to 20°.

[0008] Summary of the invention

[0009] To this end, this invention proposes a stator blade for an aircraft turbomachine compressor, the blade comprising a lower surface and an upper surface, an equidistant line between the lower surface and the upper surface, a leading edge and a trailing edge, a chord extending between the leading edge and the trailing edge, the blade being capable of extending in an airflow vein, the blade having an angle between the tangent to the equidistant line and the engine axis, an average change in the angle between the leading edge and the trailing edge is less than the average change between 60% of the chord counting the leading edge and the trailing edge.

[0010] In particular, the average change in the angle between the leading edge and the trailing edge is less than the average change between a first point located at 60% of the chord from the leading edge and a second point located at the trailing edge (therefore in the last 40% of the chord). In the blade according to the invention: the change in the angle is at a height of the blade in the vein between 0 and 50%, at the height of the blade in the vein corresponding to the internal wall of the vein, at 1 / 5 of the height of the blade in the vein, and at mid-height of the blade in the vein, 50% of the change in the angle between the tangent to the equidistant line and the engine axis is between the first point and the second point.

[0011] According to a variant, the evolution of the angel is at a height of the blade in the vein of between 0 and 50%, preferably at the height of the blade in the vein corresponding to the internal wall of the vein and / or at 1 / 5 of the height of the blade in the vein and / or at half height of the blade in the vein.

[0012] According to a variant, at the height of the blade in the vein corresponding to the internal wall of the vein, 50% of the evolution of the angle between the engine axis and a tangent to the equidistant line is on the last 40% of its chord from the leading edge.

[0013] According to a variant, at the height of the blade in the vein corresponding to the internal wall of the vein, the angle between the engine axis and a tangent to the line equidistant from the leading edge is greater than 55°.

[0014] According to a variant, at the height of the blade in the vein corresponding to the internal wall of the vein, the total evolution of the angle between the engine axis and a tangent to the equidistant line is greater than 55°.

[0015] According to a variant, at 1 / 5 of the height of the blade in the vein, 50% of the revolution of the angle between the engine axis and a tangent to the equidistant line is on the last 40% of its chord from the leading edge. According to a variant, at 1 / 5 of the height of the blade in the vein, the angle between the engine axis and a tangent to the line equidistant from the leading edge is 40°.

[0016] According to a variant, at 1 / 5 of the height of the blade in the vein, total revolution of the angle between the engine axis and a tangent to the equidistant line is greater than 50°.

[0017] According to a variant, at mid-height of the blade in the vein, 50% of the evolution of the angle between the engine axis and a tangent to the equidistant line is on the last at most 50%, preferably 40% of its chord from the leading edge.

[0018] According to a variant, at mid-height of the blade in the vein, the angle between the engine axis and a tangent to the line equidistant from the leading edge is greater than 35° and less than 50°.

[0019] According to a variant, at mid-height of the blade in the vein, the total evolution of the angle between the engine axis and a tangent to the equidistant line is greater than 40°.

[0020] According to a variant, at mid-height of the blade in the vein, the angle between the engine axis and a tangent to the line equidistant from the leading edge is at least 10° less than the angle between the engine axis and a tangent to the line equidistant from the leading edge at the height in the vein corresponding to the internal wall of the vein.

[0021] According to a variant, at mid-height of the blade in the vein, the change in the angle between the engine axis and a tangent to the line equidistant from the leading edge is at least 10° less than the change in the angle between the engine axis and a tangent to the line equidistant from the leading edge at the height in the vein corresponding to the internal wall of the vein.

[0022] The invention also relates to an aircraft turbomachine compressor comprising at least one blade as described above, preferably at least one row of blades as described above. According to a variant, the row of blades is carried by a stator at the outlet of the compressor or by a first stator of the compressor.

[0023] According to a variant, the compressor being a transonic compressor with variable geometry.

[0024] The invention also relates to an aircraft turbomachine assembly comprising

[0025] At least one row of variable pitch vanes,

[0026] Downstream of the variable pitch vane row, a compressor as described previously.

[0027] According to a variant, the at least one blade of the compressor comprises a thickness between the intrados and the extrados, the blade having between the leading edge and the trailing edge a change in thickness, the blade being such that the maximum thickness is positioned before 25% of the chord from the leading edge.

[0028] According to a variant, the thickness is considered at a height of the blade in the vein corresponding to the internal wall of the vein and / or to 1 / 5 of the height of the blade in the vein and / or to 50% of the height of the blade in the vein.

[0029] Alternatively, 80% of the maximum thickness, preferably 90%, is reached before 10% of the chord from the leading edge.

[0030] The invention also relates to an aircraft turbomachine comprising a compressor as described previously or an assembly as described previously.

[0031] According to a variant, the at least one blade or the at least one row of blades is downstream of a row of variable-pitch blades, in the direction of flow of the air flow.

[0032] According to a variant, the turbomachine comprises a variable-pitch fan, the at least one blade or the at least one row of blades is downstream of the variable-pitch fan or of a rotor carrying a row of variable-pitch blades or of a stator carrying a row of variable-pitch blades.

[0033] The use in this document of the verb "to understand", its variants, as well as its conjugations, cannot in any way exclude the presence of elements other than those mentioned. The use in this document of the indefinite article "a", "an", or of the definite article "the", "the" or "I'", to introduce an element does not exclude the presence of a plurality of these elements. The terms "first", "second", "third", and so on, are, for their part, used within the framework of this document exclusively to differentiate different similar elements, and this without implying an order between these elements.

[0034] The preferred embodiments as well as the advantages of the blade according to the invention are transposed mutatis mutandis to aircraft compressors and turbomachines, and vice versa.

[0035] Brief description of the figures

[0036] Other characteristics and advantages of the present invention will appear on reading the following description for the understanding of which reference will be made to the appended figures: figure 1 illustrates a schematic view of a section of an aircraft turbomachine: figure 2 illustrates an example of blade shape; figure 3 illustrates the evolution of the shape of the blade; figure 4 illustrates the evolution of the shape of the blade.

[0037] The drawings of the figures are not to scale. Generally, like elements are denoted by like references in the figures. For the purposes of this document, identical or similar elements may bear the same references. Furthermore, the presence of reference numbers or letters in the drawings cannot be considered as limiting, including when these numbers or letters are indicated in the claims. Detailed description of particular embodiments of the invention

[0038] This part of the text describes in detail preferred embodiments of the invention. References to figures are used but the invention is not limited by them. The drawings and / or figures described below are only schematic and are not limiting.

[0039] The invention relates to a blade for an aircraft turbomachine compressor. The blade comprises a lower surface and an upper surface, an equidistant line between the lower surface and the upper surface, a leading edge and a trailing edge, a chord extending between the leading edge and the trailing edge. The blade is capable of extending in an airflow vein. The blade has an angle between the tangent to the equidistant line and the engine axis, an average change in the angle between the leading edge and the trailing edge is less than the average change between 60% of the chord counting the leading edge and the trailing edge. The shape of such a blade ensures a high airflow straightening capacity even when the airflow reaching the rows of blades is at a high incidence. In particular, such an aircraft turbomachine compressor blade shape allows large airflow deviations, up to 60°, and operation over a large incidence range, up to 20°.

[0040] In particular, the average change in the angle between the leading edge and the trailing edge is less than the average change between a first point located at 60% of the chord from the leading edge and a second point located at the trailing edge (therefore in the last 40% of the chord). In the blade according to the invention: the change in the angle is at a height of the blade in the vein between 0 and 50%, at the height of the blade in the vein corresponding to the internal wall of the vein, at 1 / 5 of the height of the blade in the vein, and at mid-height of the blade in the vein, 50% of the change in the angle between the tangent to the equidistant line and the engine axis is between the first point and the second point.

[0041] Such a blade allows to seek the best compromise between performance and stability. Indeed, the greater the skeleton angle law (average evolution of the angle between the tangent to the equidistant line and the engine axis) downstream of the chord (beyond half starting from the leading edge), the more stable the cut at the height considered, but the more it generates pressure loss. Thus, such an average evolution of the chord as a function of the height makes it possible to penalize more severely in pressure loss than the critical height (20%), to penalize moderately the surrounding heights (0-50%) for which the difficulty in stability is less but effective, and not at all at stable heights (above 50% in height).

[0042] References are shown in some of these figures as abstract geometric references essentially in order to quantify and / or visualize properties of embodiments of the invention. In the context of this document, reference is made to the directions "axial", "circumferential" and "radial" corresponding respectively to directions parallel to the motor axis, essentially circular around the motor axis, and perpendicular to the motor axis. A reference in Figure 1 illustrates in particular the direction of the motor axis referenced X, with a direction. The terms "inner" and "inner" naturally correspond to a direction towards the motor axis X in a radial direction, and the terms "outer" and "outer" to the opposite direction in this direction.The term "inlet" (and respectively "outlet"), to refer to the position of an element in an aircraft turbomachine or in a compressor, preferably refers to a first (respectively, last) such element essentially upstream (respectively, essentially downstream) of the aircraft turbomachine or the compressor. Figure 1 illustrates a section of an aircraft turbomachine 100 on which it is intended to integrate the blade according to the invention. It may be a double-flow axial turbomachine comprising a certain number of elements. In succession, along the engine axis X, there is a fan 110, a low-pressure compressor 120, a high-pressure compressor 130, a combustion chamber 160, a high-pressure turbine 140 and a low-pressure turbine 150. These elements are known to a person skilled in the art. The blower 110 may be a variable-pitch blower, represented schematically by the reference 109 in FIG. 1.This is an example of architecture because the invention applies to other turbomachine architectures. In operation, the mechanical power of the low 150 and high 140 pressure turbines is transmitted via shafts 101 and 102 to the low 120 and high 130 pressure compressors respectively, as well as to the fan 110 via the shaft 101. A reducer 111 can be interposed on the shaft 101, so that the rotational speeds of the fan 110 and the low pressure compressor 120 are proportional. The rotors of these compressors rotate around the engine axis X, allowing them to suck in and compress air to bring it to suitable speeds, pressures and temperatures, up to the inlet of the combustion chamber 160. The blower 110 makes it possible to generate primary 106 and secondary 107 air flows upstream of the low-pressure compressor 120.The primary air flow 106 is mainly intended to axially pass through the aircraft turbomachine 100, thereby supplying the combustion chamber 160, while the secondary air flow 107 is mainly intended to generate a thrust reaction necessary for the flight of the aircraft.

[0043] Although not systematically referenced in Figure 1, each compressor and each turbine comprises one or more axial stages arranged in series. Each stage comprises a stator (or rectifier) ​​with fixed blades (or vanes) and a rotor with moving blades (or vanes) capable of being rotated around the engine axis. For the low pressure 120 and high pressure 130 compressors, these fixed and moving blades are respectively referenced by 121, 122 and 131, 132. Within a stage, the rotor sucks in and accelerates the primary air flow 106 by deflecting it relative to the engine axis X and the stator or rectifier which follows straightens the flow in the engine axis X and slows it down by transforming part of its speed into pressure. Fixed vanes include blades that may be fixed in orientation and / or variable in orientation (or pitch) (the blades occupying angular pitch positions).These steerable vanes are also known as variable-pitch vanes, or according to the English acronym "VSV" for "Variable Stator Vane". Their particularity is that the inclination of their chords can vary relative to the axis of the engine X, and the axis of the compressors 120, 130 in particular. The intrados 16 and extrados 18 faces of the vanes can be more or less exposed to the primary air flow 106. The intrados 16 extend more or less facing the primary air flow 106 in order to modulate the deflection, and therefore the straightening which is imposed on the primary air flow 106. The compressors can comprise one or more stages of steerable vanes.

[0044] The angular setting (or in other words: orientation, angular setting position or even angular position) of the blades within the fixed-blade stator is controlled by a variable setting system. In Figure 1, a variable setting system 123 may be mounted on the casing of the low-pressure compressor 120, to control the angular setting of the blades 121. A variable setting system 133 may also be mounted on the casing of the high-pressure compressor 130, to control the angular setting of the blades 131. It should be noted that this representation is in no way limiting the number or position of the variable setting systems that the compressors 120, 130 may comprise. This type of architecture is given as an example; an intermediate compressor could be present. The compressors extend along the engine axis X directed from upstream to downstream, between an upstream air inlet end and a downstream air outlet end.The compressors comprise an air flow vein 12 between the upstream air inlet end and the downstream air outlet end. The vein 12 is a wall channeling the air flow. For example, the vein 12 channels the first air flow 106. The first air flow 106 enters the compressors through the upstream air inlet end and leaves the compressors through the downstream air outlet end.

[0045] Figure 2 shows an example of a blade shape 121, which is a type of fixed blade, possibly with angular pitch but preferably without angular pitch, and for a low pressure compressor stator. The blade 121 may have a variable orientation relative to the stator element supporting it or preferably be of fixed orientation relative to the stator supporting it. The blade 121 comprises a lower surface 16 and an upper surface 18. The blade 121 comprises an equidistant line 20 between the lower surface 16 and the upper surface 18. The blade 121 also comprises a leading edge 22 upstream in the air flow and a trailing edge 24 downstream in the air flow. The blade 121 also includes a chord 26 extending between the leading edge 22 and the trailing edge 24. A tangent 28 to the equidistant line 20 is shown.

[0046] The blade is able to extend into the air flow vein 12. The air flow vein 12 has an annular structure. The air flow vein 12 has a structure of revolution. The vein 12 may thus comprise an inner wall, in the radial direction towards the inside. The vein 12 may comprise an outer wall, in the radial direction towards the outside. In section, the vein 12 has a height, along which the blades 121 extend. The height of the vein is between the inner wall and the outer wall. The height extends in the radial direction which is transverse to the axis of the engine X. In the following, the shape of the blade varies depending on the cut of the blade along the axis of the engine X, at different heights in the direction transverse to the axis of the engine X. The invention deals with the shape of the blade, and particularly its profiles corresponding to cuts of the blade at different heights of vein by a streamline.

[0047] The angle of incidence (or angle of incidence or angle of incidence) is the angle formed by the direction of the airflow at the leading edge 22 relative to the tangent 28 to the equidistant line 20. In Figure 2, this is the angle between one or other of the arrows 30 representing the direction of the airflow and the tangent 28 to the equidistant line 20 (the airflow being able to be "above" or "below" the tangent 28). The arrows 30 show different angles of incidence (maximum). The difference between the two arrows 30 corresponds to the angle of incidence range. In the context of the invention, the blade is capable of ensuring a large air deflection for operation over a large air incidence range, for example a range of 20° between the two arrows 30 (the distribution of the range on either side of the tangent 28 can be variable).Furthermore, the skeleton angle law is the evolution, along the chord 26, of the angle 29 between the tangent 28 to the equidistant line 20 and the engine axis X. The thickness law is the evolution, along the chord 26, of the thickness 50 of the blade.

[0048] The blade 121 is capable of straightening the air flow at the trailing edge so that the straightened air flow is as much as possible along the engine axis X or at an angle adapted for the downstream rotors but as constant as possible and independent of what is happening upstream. The blade 121 is capable of ensuring a large deflection of the air flow. In Figure 2, the straightened air flow is shown diagrammatically by the arrow 25. The air deflection is the angle between the arrow 25 and the arrow 30. Thus the air deflection varies with the variation of the arrow 30. It is possible that depending on the angle of the arrow 30 the angle of the arrow 25 varies more slightly, due to the efficiency in straightening the stator under incidence. The blade 121 is capable of ensuring a large deflection, up to 60°. Furthermore, the blade 121 is capable of withstanding significant variations in air incidence, according to the arrows 30.Figure 3 illustrates the evolution of the shape of the blade 121 (or evolution of the curvature or even, evolution of the camber). Figure 3 shows the dimensionless skeleton angle law (“A”) on the ordinate following the position along the chord 26 of the profile (“C”) on the abscissa (from the leading edge 22 - dimensionless position). It shows the evolution of the angle 29 of the leading edge 22 to the angle 29 of the trailing edge 24. Thus the dimensionless value of the angle 29 is 1 at the leading edge and 0 at the trailing edge. Figure 3 shows how the evolution of the angle 29 is distributed and in particular how the 50% of the evolution of the angle 29 are positioned. The particularity of the blade is that the shape (or the profile) varies according to the height of the vein (or in other words, according to the height of the blade). Figure 3 shows the evolution of the shape of the blade in the lower part of the blade, between 0 and 50% of its height in the vein.Curve 36 shows the blade shape or profile according to a section at a height of the blade in the vein corresponding to the inner wall of the vein. The section is at the root of the blade, in a radial direction towards the outside. Curve 38 shows the blade shape or profile according to a section at a height of the blade in the vein corresponding to 1 / 5 of the height in the vein, counting from the inner wall of the vein. The section is at 1 / 5 of the height of the blade, in a radial direction towards the outside. Curve 40 shows the blade shape or profile according to a section at a height of the blade in the vein corresponding to the mid-height of the vein counting from the inner wall of the vein. The section is at the mid-height of the blade height.

[0049] The blade 121 has an average change in the angle 29 between the leading edge and the trailing edge which is less than the average change between 60% of the chord counting the leading edge and the trailing edge. In other words, the average revolution of the angle 29 between the leading edge 22 and the trailing edge 24 is less than the average change between a first point located at 60% of the chord counting from the leading edge 22 and a second point located at the trailing edge 24. In other words, the average change in the angle 29 is greater over the last 40% of the chord, counting from the leading edge. In Figure 3, it can be seen that the curves 36, 38, 40 are steeper beyond 0.6 on the abscissa than before 0.6. The curves are steepest between 60% of the chord from the leading edge to the trailing edge. The blade includes an angle of 29 at the leading edge and an angle of 29 at the trailing edge.The angle 29 follows an evolution (or “variation” or “deviation”) which is distributed (or “distributed”) from the leading edge 22 to the trailing edge 24. This evolution is more or less accentuated from the leading edge to the trailing edge. The average distribution (therefore the average evolution) from the leading edge 22 to the trailing edge 24 is less than the average distribution (therefore the average evolution) from said first point to said second point. The dimensionless average evolution (or “variation” or “deviation”) of the angle 29 from the leading edge 22 to the trailing edge 24 is according to a slope less than the dimensionless average revolution slope (or “variation” or “deviation”) of the angle 29 from said first point to said second point. The total change in angle 29 is the difference between angle 29 at the leading edge 22 and angle 29 at the trailing edge.

[0050] Thus, concerning the law of curvature of the blade, the angle of attack increasing the deviation of the air flow close to the leading edge, the blade 121 makes it possible to limit the deviation of the air flow on the upstream part of the blade, while controlling that the residual deviation of the air flow on the downstream part of the blade is still compatible with its stability. Beyond 60% of the chord from the leading edge, the evolution of the angle between the engine axis X and the tangent 28 depends on the position of the blade cut according to the height of the vein. In this document, the evolution of the angle 29 mentioned relates to the average evolution of the angle 29.

[0051] Curve 36, at a height of the blade in the vein corresponding to the internal wall of the vein, shows that the average evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is less than the average revolution between 60% of the chord counting the leading edge and the trailing edge. Beyond 60% of the chord counting the leading edge 22, revolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is steeper. The slope of curve 36 accelerates. According to the profile of the blade 121 according to the curve 36, 50% of the evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is on the last 40% of its chord from the leading edge 22. The angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 at the leading edge 22 is greater than 50°. The total evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is greater than 55°.In other words, the value of the angle 29 at the leading edge 22 is 55° greater than the value of the angle 29 at the trailing edge 24. The root section has a total evolution greater than 55°, distributed 50% over the last 40% of chord 26 from the leading edge.

[0052] Curve 38, at a height of the blade in the vein corresponding to 1 / 5 of the height in the vein, shows that the average evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is less than the average evolution between 60% of the chord counting the leading edge and the trailing edge. Beyond 60% of the chord counting the leading edge 22, the evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is greater. The slope of curve 38 accelerates. According to the profile of the blade 121 according to the curve 38, 50% of the evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is on the last 40% of its chord from the leading edge 22. The angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 at the leading edge 22 is 40°. The total evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is greater than 50°.The 1 / 5 height cut presents a total evolution greater than 50°, distributed 50% over the last 40% of chord 26.

[0053] Curve 40, at a height of the blade in the vein corresponding to the mid-height of the vein, shows that the average evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is less than the average evolution between 60% of the chord counting the leading edge and the trailing edge. Beyond 60% of the chord counting the leading edge 22, the evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is greater. The slope of curve 40 accelerates. According to the profile of the blade 121 according to the curve 40, 50% of the evolution of the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 is on the last at most 50%, preferably 40% of its chord from the leading edge 22. Thus, the mid-height cut advances its evolution. The angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 at the leading edge 22 is greater than 35° and less than 50°.At the mid-height of the vein, the total change in the angle 29 between the motor axis X and the tangent 28 to the equidistant line 20 is greater than 40°. The point at which 50% of the change is reached is advanced by 10% of the chord at mid-height relative to its position at the inner wall of the vein. At the mid-height of the vein, the angle 29 between the motor axis X and the tangent 28 to the equidistant line 20 at the leading edge is at least 10° less than the angle 29 between the motor axis X and the tangent 28 to the equidistant line 20 at the leading edge at the height in the vein corresponding to the section in the inner wall of the vein. The change in the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 at the leading edge is at least 10° less than the change in the angle 29 between the engine axis X and the tangent 28 to the equidistant line 20 at the leading edge at the height in the vein corresponding to the internal wall of the vein.

[0054] Figure 4 illustrates the evolution of the shape of the blade 121. Figure 4 shows the thickness law (“E”) on the ordinate following the position along the chord 26 of the profile (“C”) on the abscissa (from the leading edge 22 - dimensionless position). It shows the evolution of the thickness between the intrados 16 and the extrados 18 from the leading edge 22 to the trailing edge 24. Figure 4 shows how the evolution of the thickness of the blade 121 is distributed. The particularity of the blade is that the shape (or the profile) varies according to the height of the vein (or in other words, according to the height of the blade). Figure 4 shows the evolution of the shape of the blade in the lower part of the blade, between 0 and 50% of its height in the vein. Figure 4 shows the evolution of the thickness of the blade 121 in cross section at the height of the blade in the vein. Curve 42 shows the shape or profile of the blade according to a section at a height in the vein corresponding to the internal wall of the vein.The section is at the base of the blade, in a radial direction towards the outside. Curve 44 shows the shape or profile of the blade according to a section at a height in the vein corresponding to 1 / 5 of the height in the vein, counting from the internal wall of the vein. The section is at 1 / 5 of the height of the blade, in a radial direction towards the outside. Curve 46 shows the shape or profile of the blade according to a section at a height in the vein corresponding to the mid-height of the vein counting from the internal wall of the vein. The section is at the mid-height of the height of the blade.

[0055] The blade 121 has a thickness change between the leading edge and the trailing edge. The blade is such that the maximum thickness is positioned before 25% of the chord from the leading edge. In other words, the thickness change of the blade is greater over the first 25% of the chord, starting from the leading edge. In Figure 2, we see that the curves 42, 44, 46 are steeper before 0.25 on the abscissa than after 0.25. The curves are steeper between the leading edge and 25% of the chord from the leading edge.

[0056] Thus, concerning the blade thickness law, strong variations in air incidence constrain the blade, particularly in its upstream part, the thickness of the blade upstream of the chord allows the blade to better resist on the upstream part of the blade. Concerning the thickness law, the curvature of the leading edge generates a significant acceleration when it is bypassed by the air, acceleration which destabilizes the air flow. Thus, advancing the maximum thickness of the profile towards the leading edge allows its curvature to be minimized. The thickness is considered at a height of the blade in the vein between 0 and 50%. The thickness close to the leading edge stabilizes the behavior of the blade but increases the pressure losses.Thus, the invention aims to stabilize the blade profiles at a height between 0 and 50% of the blade, where the blade is the least stable (a pressure loss between 50-100% of height being less critical because the blade does not show any difficulty in stability). The maximum thickness positioned before 25% of the chord from the leading edge is considered in section, at a height of the blade in the vein corresponding to the internal wall of the vein (curve 42) and / or at 1 / 5 of the height of the blade in the vein (curve 44) and / or at 50% of the height of the blade in the vein (curve 46). Preferably, 80% of the maximum thickness, preferably 90%, is reached before 10% of the chord from the leading edge. This makes it possible to further strengthen the blade, and to withstand significant variations in incidence on the leading edge of the blade. In relative magnitude, the maximum thickness is greater than 6% of the chord.Thickening towards the leading edge improves stability at flow incidence by reducing its radius of curvature up to a certain point. Indeed, too much thickness too close to the leading edge results in a very small local radius (in the extreme, with 100% maximum thickness at 0% of the leading edge, we have a rectangle with two very unstable corners at incidence, despite an infinite radius of curvature at the leading edge since we have a segment). The 90% maximum thickness reached at 10% of the chord is the best robustness to guarantee the largest radius of curvature at the leading edge without reducing the radius of curvature too much between 10% and 25% of the chord.

[0057] All these characteristics of the blade 121 can be considered separately or in combination. In particular, the average evolution of the angle 29 relative to the chord and the evolution of the thickness relative to the chord can be considered separately or in combination. The blade ensures a uniform flow along the blade. The blade makes it possible to avoid separation of the air before reaching the trailing edge. The blade allows in a combined manner a large deflection of the air flow and a large incidence of the air flow. The blade makes it possible to withstand large variations in incidence of the air flow. This makes it possible to improve the performance of a compressor using such a blade. Also, this makes it possible to improve the performance of a turbomachine assembly according to the invention comprising at least one row of variable-pitch blades, and downstream of the row of variable-pitch blades, comprising a compressor with at least one blade 121 as described above.Indeed, blades 121 located downstream of variable-pitch blades in the direction of flow of the air flow undergo significant variations in the incidence of the air flow due to the presence of the variable-pitch blades upstream. The blades 121 as described are capable of withstanding large variations in the incidence of the air flow.

[0058] The invention also relates to a compressor 120 of an aircraft turbomachine, comprising at least one blade 121, preferably at least one row of blades 121, comprising all or part of the characteristics described. The invention also relates to a turbomachine assembly comprising at least one row of blades (of a rotor or a stator) with variable pitch, and, downstream of the row of blades with variable pitch, the compressor 120. The invention also relates to a turbomachine comprising the compressor or the assembly with the compressor. It is in particular an axial compressor for a turbomachine. More particularly, it is a transonic compressor with variable geometry (with blades moving according to a pitch on the stators). The compressor may be downstream of the fan 110 of the turbomachine, the fan itself being with variable pitch.The row of blades 121 can be carried by the last stator 124 at the outlet of this compressor (called OGV for "outlet guide vane"), which must straighten the flow towards a purely axial speed and / or the first fixed-pitch stator of this compressor, in the direction of flow of the air flow. The row of blades 121 is downstream (directly (i.e. immediately after) or indirectly (i.e. separated by other row(s) of blades) of a row of variable-pitch blades (of a rotor, a stator or the fan), in the direction of flow of the air flow. The blade 121 allows in a combined manner a large deflection of the air flow and operation with a large incidence intake - preferably by the combination of the thickness law and the skeleton curvature law.

[0059] The blade 121 described previously with the evolution of the angle 29 and / or the evolution of the thickness can be described as a stator blade, or a shaped blade, a shaped stator blade. The present invention has been described in relation to specific embodiments, which have a purely illustrative value and should not be considered as limiting. In general, it will be obvious to a person skilled in the art that the present invention is in no way limited to the examples illustrated and / or described above. For example, the invention is applicable at the blade head, in combination or independently of the evolution of the angle 29 between 0 and 50% of height.

Claims

Claims 1. Stator blade (121) for an aircraft turbomachine compressor, the blade comprising a lower surface (16) and an upper surface (18), an equidistant line (20) between the lower surface and the upper surface, a leading edge (22) and a trailing edge (24), a chord (26) extending between the leading edge and the trailing edge, the blade being able to extend in an airflow vein (12), the blade having an angle (29) between the tangent (28) to the equidistant line (20) and the engine axis (X), an average change in the angle (29) between the leading edge and the trailing edge is less than the average change between a first point located at 60% of the chord from the leading edge and a second point located at the trailing edge and in which - revolution of the angle (29) is at a height of the blade in the vein between 0 and 50%, - at the height of the blade in the vein corresponding to the internal wall of the vein, at 1 / 5 of the height of the blade in the vein, and at mid-height of the blade in the vein, 50% of the change in the angle (29) between the tangent (28) to the equidistant line (20) and the engine axis (X) is between the first point and the second point.

2. Blade (121) according to one of the preceding claims, at the height of the blade in the vein corresponding to the internal wall of the vein, the angle (29) between the engine axis (X) and a tangent (28) to the equidistant line (20) at the leading edge is greater than 55°.

3. Blade (121) according to one of the preceding claims, at the height of the blade in the vein corresponding to the internal wall of the vein, total revolution of the angle (29) between the engine axis (X) and a tangent (28) to the equidistant line (20) is greater than 55°.

4. Blade (121) according to one of the preceding claims, in which, at 1 / 5 of the height of the blade in the vein, the angle (29) between the engine axis (X) and a tangent (28) to the equidistant line (20) at the leading edge is 40°.

5. Blade (121) according to one of the preceding claims, in which, at 1 / 5 of the height of the blade in the vein, the total evolution of the angle (29) between the engine axis (X) and a tangent (28) to the equidistant line (20) is greater than 50°.

6. Blade (121) according to one of the preceding claims, in which, at mid-height of the blade in the vein, the angle (29) between the engine axis (X) and a tangent (28) to the equidistant line (20) at the leading edge is greater than 35° and less than 50°.

7. Blade (121) according to one of the preceding claims, in which, at mid-height of the blade in the vein, the total evolution of the angle (29) between the engine axis (X) and a tangent (28) to the equidistant line (20) is greater than 40°.

8. Blade (121) according to one of the preceding claims, in which, at mid-height of the blade in the vein, the angle (29) between the motor axis (X) and a tangent (28) to the equidistant line (20) at the leading edge is at least 10° less than the angle (29) between the motor axis (X) and a tangent (28) to the equidistant line (20) at the leading edge at the height in the vein corresponding to the internal wall of the vein.

9. Blade (121) according to one of the preceding claims, in which, at mid-height of the blade in the vein, the evolution of the angle (29) between the engine axis (X) and a tangent (28) to the equidistant line (20) at leading edge is at least 10° less than the change in the angle (29) between the engine axis (X) and a tangent (28) to the equidistant line (20) at the leading edge at the height in the vein corresponding to the internal wall of the vein.

10. Aircraft turbomachine compressor (120) comprising at least one blade (121) according to one of claims 1 to 13, preferably at least one row of blades (121) according to one of claims 1 to 13. 1 1. Compressor according to the preceding claim, in which the row of blades is carried by a stator at the outlet of the compressor or by a first stator of the compressor.

12. Compressor according to the preceding claim, the compressor being a transonic compressor with variable geometry.

13. Aircraft turbomachine assembly comprising - At least one row of variable pitch blades, - Downstream of the row of variable-pitch blades, a compressor according to one of claims 14 to 16.

14. Assembly according to the preceding claim, in which the at least one blade (121) of the compressor comprises a thickness between the intrados (16) and the extrados (18), the blade having between the leading edge and the trailing edge a change in thickness, the blade being such that the maximum thickness is positioned before 25% of the chord from the leading edge.

15. Assembly according to one of the two preceding claims, in which the thickness is considered at a height of the blade in the vein corresponding to the internal wall of the vein and / or to 1 / 5 of the height of the blade in the vein and / or to 50% of the height of the blade in the vein.

16. Assembly according to one of the three preceding claims, in which 80% of the maximum thickness, preferably 90%, is reached before 10% of the chord from the leading edge.

17. Aircraft turbomachine (100) comprising a compressor according to one of claims 14 to 16 or an assembly according to one of claims 17 to 20.

18. Turbomachine according to the preceding claim, in which the at least one blade (121) or the at least one row of blades (121) is downstream of a row of variable-pitch blades, in the direction of flow of the air flow.

19. Turbomachine according to one of two preceding claims, comprising a variable-pitch fan (110), the at least one blade (121) or the at least one row of blades (121) is downstream of the variable-pitch fan or of a rotor carrying a row of variable-pitch blades or of a stator carrying a row of variable-pitch blades.

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