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

Abstract

Disclosed is a stator vane for a compressor of an aircraft turbine engine, with a fan upstream of the compressor in the direction of flow of the air flow, the fan being driven in rotation via a speed reducer and being variable-pitch, the vane comprising a leading edge (22) and a trailing edge (24), a chord (26) extending between the leading edge and the trailing edge, the vane exhibiting an evolution in the length of the chord in cross sections of the vane in a plane transverse to the leading edge, the evolution in the length of the chord having peaks and troughs, the amplitude of the chord-length peaks relative to a mean chord length being between 5% and 20%, the wavelength between two chord-length peaks being between 10% and 50% of the mean chord length.

Description

[0001] AIRCRAFT TURBOMACHINE COMPRESSOR BLADE, COMPRESSOR, TURBOMACHINE

[0002] technical field

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

[0004] Previous art

[0005] Aircraft turbomachinery compressors are equipped with blades whose shape is crucial to the compressor's operating quality. In particular, compressor blades are expected to have a high airflow straightening capacity, as the airflow reaches the blade rows at a high angle of incidence. This is especially true in the context of compressors located directly downstream of the fan – whether these are variable-pitch fans driven by a gearbox.

[0006] There is therefore a need for aircraft turbomachine compressor blades that provide large airflow deflections, up to 50°, and operation over a wide range of angles of incidence, up to 20° and beyond.

[0007] Description of the invention

[0008] To this end, the invention proposes a stator blade for an aircraft turbomachine compressor, the blade comprising a leading edge and a trailing edge and a chord extending between the leading edge and the trailing edge, the blade exhibiting an evolution of the chord length according to sections of the blade in a plane transverse to the leading edge, the evolution of the chord length being in troughs and crests, the amplitude of the chord length crests relative to an average chord is between 5% and 20%, the wavelength between two chord length crests is between 10% and 50% of the average chord length.

[0009] According to one variant, the amplitude of the rope length peaks relative to an average rope is 10%.

[0010] According to one variant, the amplitude of the chord length peaks relative to an average chord is between 1 mm and 5 mm in absolute value. According to another variant, the wavelength between two chord length peaks is between 30% and 35% of the average chord length.

[0011] According to one variant, the wavelength between two chord length peaks is between 10% and 35% of the average chord length.

[0012] According to one variant, the wavelength between two peaks of chord length is between 30% and 50% of the length of an average chord.

[0013] According to one variant, the wavelength between two chord length peaks is between 5% and 20% of the blade height at the leading edge, preferably

[0014] 10%.

[0015] According to one variant, the wavelength between two peaks of chord length is between 2mm and 20mm in absolute value.

[0016] According to one variant, the blade further comprises an intrados, an extrados and an equidistant line between the intrados and the extrados, the blade having a skeleton angle between a tangent to the equidistant line at the leading edge and a driving axis of the turbomachine, the skeleton angle at the leading edge having an evolution according to sections of the blade in a plane transverse to the leading edge, revolution of the angle being in troughs and crests.

[0017] According to one variant, the evolution of the angle at the leading edge is proportional to the evolution of the length of the chord.

[0018] According to one variant, the amplitude of the skeleton angle ridges is a coefficient, the coefficient being between 0° per percentage of amplitude of the chord length and 2° per percentage of amplitude of the chord length, preferably between 0.5° and 1° per percentage of amplitude of the chord length.

[0019] According to one variant, the direction of evolution of the skeleton angle is to increase the angle at the leading edge when the evolution of the chord length is towards the hollows.

[0020] According to one variant, the wavelength between two skeleton angle crests is between 10% and 50% of the mean chord length, preferably between 30% and 35% of the mean chord.

[0021] According to one variant, the wavelength between two skeleton angle crests is between 5% and 20% of the blade height at the leading edge, preferably 10%. According to another variant, the blade has blade sections in a plane transverse to the leading edge, the blade comprising an evolution of an offset of one blade section relative to another, taken in a plane transverse to the leading edge, the offset being in rotation about a driving axis of the turbomachine, the revolution of the offset being in troughs and crests, the amplitude of the offset crests is between 5% and 20% of the mean chord length, preferably 10%, the wavelength between two offset crests is between 10% and 50% of the mean chord length, preferably between 30% and 35% of the mean chord length.

[0022] According to one variant, the wavelength between two offset peaks is between 5% and 20% of the blade height at the leading edge, preferably 10%.

[0023] According to one variant, the blade cut is offset towards the extrados when the revolution of the chord length is towards the hollows.

[0024] According to one variant, the blade further comprises an intrados, an extrados and a thickness between the intrados and the extrados, each section has a thickness evolution exhibiting a maximum thickness value, the blade having a maximum thickness evolution along the leading edge of the blade, this evolution exhibiting an undulation.

[0025] According to one variant, the wavelength between two peaks of maximum thickness is between 10% and 50% of the average chord length, preferably between 30% and 35% of the average chord length.

[0026] According to one variant, the wavelength between two peaks of maximum thickness is between 5% and 20% of the height of the blade at the leading edge, preferably 10%.

[0027] 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.

[0028] According to one variant, the blade or row of blades is carried by the last stator of the compressor, part of whose stators have variable pitch, or by the first fixed-pitch stator of the compressor, part of whose stators have variable pitch, or by any fixed-pitch stator of the compressor, part of whose stators have variable pitch.

[0029] According to one variant, the compressor is an axial turbomachine compressor, or a transonic variable geometry compressor, or a transonic variable geometry compressor linked to a blower, itself with variable pitch and driven by the reducer.

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

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

[0032] According to one variant, the turbomachine includes a variable-pitch fan driven by a speed reducer, at least one blade or at least one row of blades is downstream of the variable-pitch fan driven by a speed reducer.

[0033] According to one variant, the blower is unfaired.

[0034] According to one variant, at least one blade or at least one row of blades is downstream or upstream of a swan neck.

[0035] The use, in this document, of the verb "comprendre" (to understand), its variants, and its conjugations, cannot in any way exclude the presence of elements other than those mentioned. The use, in this document, of the indefinite article "un" (a), "une" (an), or the definite article "le" (the), "la" (the), or "l'" (it) to introduce an element does not exclude the presence of multiple such elements.

[0036] The terms "first >>", "second >>", "third >>, etc. are used in this document exclusively to differentiate different elements, without implying any order between these elements.

[0037] The preferred embodiments and advantages of the blade according to the invention are applicable mutatis mutandis to the aircraft compressor and turbomachine, and vice versa. The various embodiments can be considered individually or in combination. Brief description of the figures

[0038] Other features and advantages of the present invention will become apparent from the following detailed description, for the understanding of which reference will be made to the accompanying figures which show: Figure 1 illustrates a schematic view of a cross-section of an aircraft turbomachine; Figure 2 illustrates an example of a cross-section of a blade along a given streamline or radius; Figure 3 illustrates a shape of the blade; Figure 4 illustrates a shape of the blade in more detail; Figure 5 illustrates the geometric characteristics of the blade.

[0039] The drawings in the figures are not to scale. Similar features are generally denoted by similar reference numerals in the figures. Within the scope of this document, identical or analogous features may bear the same reference numerals. Furthermore, the presence of reference numerals or letters in the drawings shall not be considered limiting, even when such numerals or letters are specified in the claims.

[0040] Detailed description of embodiments of the invention

[0041] This section 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 schematic only and are not limiting.

[0042] The invention relates to a stator blade for an aircraft turbomachine compressor. The invention is applicable in the context of a turbomachine comprising a fan upstream of the compressor in the direction of airflow, the fan being driven in rotation via a speed reducer and having variable pitch. This context implies a need for blades that are more tolerant and more resistant to angle of attack and aerodynamic deviations with respect to flow separation on the blade face. In this context, the blade comprises a leading edge, a trailing edge, and a chord extending between the leading and trailing edges. The blade exhibits a chord length variation along cross-sections of the blade in a plane transverse to the leading edge. The chord length variation is characterized by troughs and crests.The invention allows for greater tolerance and resistance to angle of attack and aerodynamic deviations with respect to flow separation on the blade wall, and ensures operation over a wide angle of attack range, up to 20° (or high angle of attack). Furthermore, this allows for large airflow deviations, up to 50°.

[0043] Reference points are shown in some of these figures as abstract geometric coordinate systems primarily to quantify and / or visualize properties of embodiments of the invention. In this document, reference is made to the directions "axial," "circumferential," and "radial," corresponding respectively to directions parallel to the drive axis, essentially circular around the drive axis, and perpendicular to the drive axis. A reference point in Figure 1 specifically illustrates the direction of the referenced drive axis X, with a direction. The terms "inward" and "inward" naturally correspond to a direction toward the drive axis X along a radial direction, and the terms "outward" and "outward" to the opposite direction.The term "inlet / inlet" (and respectively "outlet / inlet"), when referring to the position of an element in an aircraft turbomachine or compressor, preferentially refers to a first (respectively, last) such element essentially upstream (respectively, essentially downstream) of the aircraft turbomachine or compressor.

[0044] Figure 1 illustrates a cross-section of an aircraft turbomachine 100 into which the blade according to the invention is to be integrated. This can be a twin-flow axial turbomachine comprising several elements. In succession, along the drive 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 those skilled in the art. This is an example of an architecture, as the invention applies to other turbomachine architectures. In operation, the mechanical power of the low-pressure turbine 150 and the high-pressure turbine 140 is transmitted via shafts 101 and 102 to the low-pressure compressor 120 and the high-pressure compressor 130, respectively, as well as to the fan 110 via shaft 101.The rotors of these compressors rotate around the drive axis X, allowing them to draw in and compress air to the appropriate speeds, pressures, and temperatures, up to the inlet of the combustion chamber 160. The fan 110 generates airflows 106 and 107 upstream of the low-pressure compressor 120. The airflow 106 is primarily intended to pass axially through the aircraft turbomachine 100, thereby supplying the combustion chamber 160, while the airflow 107 is primarily intended to generate the thrust reaction necessary for the aircraft's flight. The fan 110 can be shrouded with a cowling 112 or can be an unshrouded fan 110 (without the cowling 112). The fan 110 can be considered a rotor, upstream of the low-pressure compressor 120.

[0045] The blower 110 can be a variable-pitch blower. A variable-pitch system is schematically represented by reference numeral 109 in Figure 1. In other words, the blower blades have a variable orientation relative to the direction of airflow. Put another way, the blower blades can occupy different angular pitch positions. Furthermore, a speed reducer 111 can be interposed on the shaft 101, so that the rotational speeds of the blower 110 and the low-pressure compressor 120 are proportional. The blower 110 is then driven in rotation via the speed reducer 111 around the drive shaft. The reducer 111 allows the rotational speed of the blower 110 to be reduced relative to that of the shaft 101 and the low-pressure compressor 120 to take into account the mechanical constraints and limitations of the blower.The rotational speed of the blower 110 is therefore lower than that of the low-pressure compressor 120. The blower 110 can thus have variable pitch, or it can be driven via a speed reducer, or it can have variable pitch and be driven via a speed reducer. Although not systematically referenced in Figure 1, each compressor and each turbine comprises one or more axial stages arranged in series. Each stage includes a stator (or rectifier) ​​with fixed blades and a rotor with movable blades capable of rotating around the drive shaft. For the low-pressure compressor 120 and the high-pressure compressor 130, these fixed and movable blades are respectively referenced as 121, 122 and 131, 132.Within a stage, the rotor draws in and accelerates the airflow 106 by deflecting it relative to the axis of the X-motor. The following stator, or rectifier, then directs the flow back along the X-motor axis and slows it down by converting some of its velocity into pressure. The fixed blades include vanes that can have a fixed orientation and / or variable orientation (or pitch) (the vanes occupying angular pitch positions). These adjustable blades are also known as variable-pitch vanes, or by the English acronym "VSV" for "Variable Stator Vane." Their distinguishing feature is that the inclination of their chords can vary relative to the X-motor axis, and in particular, the axes of the compressors 120 and 130. The intrados 16 and extrados 18 faces of the blades can be more or less exposed to the airflow 106. The intrados 16 extend more or less in front of the airflow 106 in order to modulate the deviation, and therefore the straightening which is imposed on the airflow 106.Compressors can have one or more stages of adjustable blades.

[0046] The angular pitch (or in other words: orientation, angular pitch position, or angular position) of the blades within the stator is controlled by a variable pitch system. In Figure 1, a variable pitch system 123 can be mounted on the housing of the low-pressure compressor 120 to control the angular pitch of the blades 121. A variable pitch system 133 can also be mounted on the housing of the high-pressure compressor 130 to control the angular pitch of the blades 131. It should be noted that this representation is in no way limiting to the number or position of variable pitch systems that the compressors 120 and 130 may include. This type of architecture is given as an example; an intermediate compressor could be present. The compressors extend along the upstream-to-downstream motor axis X, between an upstream air inlet and a downstream air outlet.The compressors include an airflow channel 12 between the upstream air inlet and the downstream air outlet. The channel 12 is a wall that directs the airflow. For example, the channel 12 directs the airflow 106. The airflow 106 enters the compressors through the upstream air inlet and exits the compressors through the downstream air outlet.

[0047] Figure 2 shows an example of a cross-section (along a constant height or streamline) of a blade 121. The blade 121 in Figure 2 is a cross-section of the blade taken in a plane transverse to the leading edge 22. The blade 121 is a type of fixed blade, possibly with angular pitch but preferably without angular pitch, and for a compressor stator, preferably a low-pressure compressor. The blade 121 may have a variable orientation relative to the stator element supporting it or, preferably, a fixed orientation relative to the stator supporting it. The blade 121 comprises a lower surface 16 and an upper surface 18. The blade 121 includes an equidistant line 20 between the lower surface 16 and the upper surface 18. The blade 121 also includes a leading edge 22 upstream in the airflow and a trailing edge 24 downstream in the airflow. The vane 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.

[0048] The blade is able to extend along its height within the airflow channel 12. The blade's height is substantially radial, substantially along the Z direction. The airflow channel 12 has an annular structure. The airflow channel 12 has a rotary structure. The channel 12 can thus have an inner wall, along the radial direction inward. The channel 12 can have an outer wall, along the radial direction outward. In cross-section, the channel 12 has a height, along which the blades 121 extend. The height of the channel is between the inner and outer walls. The height extends along the radial Z direction, which is transverse to the axis of the X-axis. In the following, the shape of the blade varies depending on the cross-section of the blade along the axis of the X-axis, at different heights along the direction transverse to the axis of the X-axis.The invention deals with the shape of the blade, and particularly with its profiles corresponding to sections of the blade at different vein heights by a streamline.

[0049] Figure 3 illustrates a shape of the blade 121. Figure 3 shows the blade 121 extending in the radial direction. The blade 121 extends in the Z direction. In particular, the blade 121 is corrugated, at least on a portion of its leading edge. The blade 121 exhibits corrugations according to one or more parameters that vary according to laws, as described below. The parameters and laws defining the corrugations are considered individually or in combination. The blade itself is corrugated and comprises one or more types of corrugation. The corrugation of the blade exhibits troughs and crests. For the shape of the dawn itself and in what follows, the notion of 'troughs and ridges' (or equivalent formulations) refers to a shape or values ​​that increase (or rise) and decrease (or fall) in an alternate, sinusoidal manner (like a curve resembling a sinusoid with arcs similar to those of a sinusoid).This can be a sinusoidal evolution, following a regular or irregular sinusoid. In other words, it can be a shape or values ​​describing curves that are alternately concave and convex.

[0050] Figure 4 illustrates a blade shape in more detail. Figure 4 shows cross-sections of the blade 121 taken in a transverse plane 56 at the leading edge 22. In other words, Figure 4 shows cross-sections in the transverse plane 56 at the leading edge at different positions or heights along the leading edge, as shown in Figure 3. These are different cross-sections taken at different positions or heights of the blade, along the radial Z direction. Section 121a is the average cross-section. Figure 4 omits all the reference symbols from Figure 2 (such as the chord 26, angle 29, thickness 50, etc.) for the sake of clarity. For each section 121 a, 121 b, 121 c, ... we will refer to figure 2 which shows one of these sections.

[0051] Figure 5 illustrates the blade parameters, along the x-coordinate of the driving axis X and the h-coordinate of the radial direction Z. The coordinates h1 and h2 show the blade waviness or the waviness of a blade parameter as a function of its height. The x1 coordinate shows the position of the blade 121 along the driving axis. The blade and its parameters (such as chord length, angle 29, cutting offset, or thickness) vary according to the laws described below. The amplitude A shows the peak-to-average amplitude 52 of a blade parameter. The wavelength λ shows the peak-to-peak wavelength.

[0052] The angle of attack (or angle of incidence) is the angle formed by the direction of the airflow at the leading edge 22 with respect to the tangent 28 to the equidistant line 20. In Figure 2, it is the angle between one or the other of the arrows 30 representing the direction of the airflow and the tangent 28 to the equidistant line 20 (the airflow can be "above" or "below" the tangent 28). The arrows 30 show different (maximum) angles of attack. The distance between the two arrows 30 corresponds to the range of angles of attack. Within the framework of the invention, the blade is able to ensure a large air deflection for operation over a large range of air incidence, 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, for example [-10°; +10°] or [-5°; +15°], etc.).

[0053] The chord law describes the evolution of the chord length 26 along sections of the blade in a plane transverse to the leading edge (over at least a portion of the leading edge). The chord law describes the evolution of the chord length along sections in a plane transverse to the leading edge at different positions or heights along the leading edge, in the radial direction Z. In other words, the chord law describes the evolution of the chord length 26 along sections of the blade at different heights along the blade's height, in the radial direction. Put another way, the blade exhibits a chord length evolution 26 along its height. Between two successive, spaced sections, the chord length changes. Figures 3 and 4 show that sections 121a, 121b, 121c, ... have chords whose lengths vary from one section to the next.

[0054] The evolution of the skeleton angle 29 along blade sections in a plane transverse to the leading edge is called the skeleton angle law. The skeleton angle law describes the evolution of the skeleton angle 29 at the leading edge as a function of blade sections in a plane transverse to the leading edge at different positions or heights along the leading edge, in the radial direction. The waviness is applied to the value of the skeleton angle at the leading edge. Between two successive, spaced sections, the angle 29 changes. Figures 3 and 4 show that sections 121a, 121b, 121c, ... have an angle 29 that varies along the chord and from one section to the next.

[0055] The tangential stacking law, for blade sections taken in a plane transverse to the leading edge 22, describes the evolution of the offset of one section relative to another, the offset being rotational around the turbomachine's X-axis (or along a Y-direction transverse to the X-axis and the Z-direction). It is the rotational offset of the position of a particular point on the blade relative to the same particular point on another section—such as a point on the leading edge 22, the trailing edge 24, or the center of gravity of a section. Between two successive, spaced-apart sections, one section is offset relative to the other. Figures 3 and 4 show that sections 121a, 121b, 121c, ... are stacked along the blade height with offsets relative to each other. We can see that a point taken at the leading edge 22 or at the center of gravity is offset from one cut to the other.

[0056] Furthermore, each cross-section has a thickness law with a maximum thickness value. The blade exhibits a thickness evolution corresponding to the revolution of the maximum thickness along the leading edge of the blade, this evolution exhibiting undulation. The appropriate shape of these undulations allows the generation of a vortex system on the blade surface, which stabilizes the flow at high angles of attack. According to Figures 3 and 4, cross-sections 121a, 121b, 121c,... have a thickness 50 that varies along the chord and from one cross-section to the next.

[0057] The blade 121 is capable of straightening the trailing edge airflow so that the straightened airflow is as close as possible to the motor axis X or at an angle suitable for the downstream rotors, but as constant as possible and independent of upstream conditions. The blade 121 is capable of providing a large airflow deflection. In Figure 2, the straightened airflow is represented by arrow 25. The airflow deflection is the angle between arrow 25 and arrow 30. Thus, the airflow deflection varies with the variation of arrow 30. It is possible that, depending on the angle of arrow 30, the angle of arrow 25 may vary more slightly, due to the stator's straightening efficiency under incidence. The blade 121 is capable of providing a large deflection, up to 50°. In addition, the 121 blade is capable of withstanding significant variations in air incidence, as indicated by arrows 30. The incidence range is typically up to 20° and beyond.

[0058] The blade 121 can exhibit a change in chord length 26 according to blade sections in the plane 56 transverse to the leading edge. The change in chord length can be in troughs and crests. In other words, in relation to the chord law, the blade exhibits an undulation that is a sinusoid around a so-called "smooth" chord law (without the undulations). The change in chord length in troughs and crests refers to a shape or values ​​that increase (or rise) and decrease (or fall) alternately in a sinusoidal manner (like a curve resembling a sine wave with arcs similar to those of a sine wave). It can be a sinusoidal change. Such a sinusoidal change is along a regular or irregular sinusoid. In other words, it can be a shape or values ​​describing alternately concave and convex curves.In Figure 4, chord 26 (visible in Figure 2 but not in Figure 4 for clarity) varies in length along troughs and crests, depending on the cross-section. The amplitude A (peak to mean) of the chord length crests relative to a mean chord can be between 5% and 20%, preferably 10% (throughout the document, the mean chord is the average of the chord over its height. Mathematically, this is Qi). dh / H). The wavelength A between two chord length crests. The peak-to-peak (peak-to-peak) distance can be between 10% and 50% of the mean chord length. Such a blade allows for greater tolerance and resistance to angle of attack and aerodynamic deviations with respect to flow separation on the blade face, and ensures operation over a wide angle of attack range, up to 20° (or high angle of attack). Furthermore, such a blade allows for large airflow deviations, up to 50°. The amplitude A (peak-to-mean) of the chord length peaks relative to a mean chord can be between 1 mm and 5 mm in absolute value, as an example only. The wavelength A between two chord length peaks (peak-to-peak) can be between 30% and 35% of the mean chord length. The wavelength Δ between two chord length peaks (peak to peak) can be between 10% and 35% of the average chord length.The wavelength λ between two chord length crests (peak to crest) can be between 30% and 50% of the average chord length. The wavelength λ between two chord length crests (peak to crest) can be between 20% and 40% of the average chord length. These wavelength ranges allow operation over a wide angle of attack, up to 20° (or high angle of attack). These wavelength ranges allow for large airflow deflections, up to 50°. The wavelength λ between two chord length crests (peak to crest) can be between 5% and 20% of the blade height at the leading edge 22, preferably 10%. The wavelength λ between two chord length crests (peak to crest) can be between 2 mm and 20 mm in absolute value, as an example only. These characteristics are applied section by section, according to figure 4, resulting in a blade of the shape shown in figure 3.

[0059] The blade 121 can also exhibit an evolution of the skeleton angle 29, according to blade sections in the plane 56 transverse to the leading edge 22. The evolution of the skeleton angle 29 can be trough-shaped and crest-shaped. In other words, in relation to the skeleton angle law, the blade exhibits an undulation of the skeleton angle 29 value at the leading edge that is a sinusoid around the so-called "smooth" leading edge angle law (without the undulations). The trough-shaped and crest-shaped evolution of the skeleton angle 29 refers to a shape or values ​​that increase (or rise) and decrease (or fall) alternately in a sinusoidal manner (like a curve resembling a sinusoid with arcs similar to those of a sinusoid). It can be a sinusoidal evolution. Such a sinusoidal evolution follows a regular or irregular sinusoid.In other words, it can be a shape or values ​​describing alternately concave and convex curves. In Figure 4, angle 29 varies along troughs and crests, depending on the cross-section. The evolution (or variation) of angle 29 can be proportional to the evolution (such as the elongation) of the chord length. The amplitude A (peak to mean) of the crests of skeleton angle 29 can be a coefficient K, the coefficient K being between 0° per percent of the chord length amplitude and 2° per percent of the chord length amplitude, preferably the coefficient K is between 0.5° and 1°, preferably 0.75°, per percent of the chord length amplitude. Preferably, the direction of evolution (or modification) of angle 29 is to increase the angle at the leading edge when the evolution of the chord length is towards the hollows (or in other words, when the chord is shortened).The wavelength λ between two crests at an angle of 29° (peak to crest) can be between 10% and 50% of the mean chord, preferably between 30% and 35% of the mean chord length. The wavelength λ between two crests at an angle of 29° (peak to crest) can be between 5% and 20% of the blade height at the leading edge 22°, preferably 10%. The wavelength λ between two crests at an angle of 29° (peak to crest) can be between 2 mm and 20 mm in absolute value. These characteristics apply section by section, as shown in Figure 4, resulting in a blade with the shape shown in Figure 3.

[0060] The blade 121 allows for a combined large deflection of the airflow and operation with a high angle of attack. As already mentioned, this can be achieved through the chord law as described previously. This can be further enhanced by combining the chord law with the skeleton angle law, whereby the blade exhibits a waviness at the leading edge with a skeleton angle value of 29, which is a sinusoid around the so-called "smooth" (unwaviness-free) leading edge angle law, or, in the terms previously used to describe the evolution of skeleton angle 29.

[0061] The blade 121 can also exhibit an evolution of the offset between two blade sections, measured in plane 56 transverse to the leading edge, the offset being in rotation around a driving axis (X) of the turbomachine. The evolution of the offset can be in troughs and crests. In other words, in relation to the tangential stacking law (rotational offset around the driving axis of one section or section relative to another), the blade exhibits an offset undulation that is a sinusoid around the so-called "smooth" stacking axis (without the undulations). In Figure 4, the offset between two sections varies according to troughs and crests, depending on the sections. The evolution of the displacement in troughs and peaks refers to a shape or values ​​that increase (or rise) and decrease (or fall) in an alternating, sinusoidal manner (like a curve resembling a sine wave with arcs similar to those of a sine wave). It can be a sinusoidal evolution.Such a sinusoidal evolution follows a regular or irregular sinusoid. In other words, it can be a shape or values ​​describing alternately concave and convex curves. The amplitude A (peak-to-mean) of the offset peaks can be between 5% and 20% of the mean chord length, preferably 10%. The absolute value of the amplitude A (peak-to-mean) of the offset peaks can be between 1 mm and 5 mm. The wavelength A between two offset peaks (peak-to-peak) can be between 10% and 50% of the mean chord length, preferably between 30% and 35% of the mean chord length. The wavelength A between two offset peaks can be between 5% and 20% of the blade height at the leading edge 22, preferably 10%. The wavelength λ between two offset peaks can be between 2 mm and 20 mm in absolute value. These characteristics apply section by section, as shown in Figure 4, resulting in a blade with the shape represented in Figure 3.In stacking, the shorter section is preferably offset towards the upper surface. The direction of thickness variation (or modification) is to offset the blade cut towards the upper surface when the chord length variation is towards the hollows (or in other words, when the chord is shortened).

[0062] The blade 121 can also exhibit a change in maximum thickness along its leading edge. This change in maximum thickness can be in troughs and crests. In other words, in relation to the thickness law, the blade exhibits a wave of the maximum thickness, which is a sinusoid around the so-called "smooth" thickness (without the undulations). In Figure 4, the maximum thickness varies in troughs and crests, depending on the cross-section. The amplitude A (peak to mean) of the maximum thickness crests relative to an initial thickness (or the so-called "smooth" thickness law, i.e., without undulations) can be, as a percentage of the initial thickness, from 0 to 2 times the amplitude of the chord change, preferably 1 time. The wavelength Δ between two peaks of maximum thickness (peak to peak) can be between 10% and 50% of the mean chord, preferably between 30% and 35% of the mean chord length.The wavelength λ between two peaks of maximum thickness (peak to peak) can be between 5% and 20% of the blade height at the leading edge 22, preferably 10%. The wavelength λ between two peaks of maximum thickness (peak to peak) can be between 2 mm and 20 mm in absolute value. These characteristics apply section by section, as shown in Figure 4, resulting in a blade with the shape shown in Figure 3.

[0063] Thus, the invention relates to the shape of the blade, and particularly to the evolution of its parameters or profiles (especially in 2D) in terms of thickness, chord, tangential stacking, and angle along its height, through wave or undulation patterns. As shown in Figure 4, the above characteristics vary and combine. The appropriate shape of these undulations allows for the generation of a vortex or eddy system on the blade surface, which stabilizes the airflow at high angles of attack but generates additional losses at the adaptation angle of attack. The parameters described correspond to an optimization between the formation of eddies (which generate aerodynamic losses) while limiting their occurrence to stabilize the flow (and energize the boundary layer).Thus, a person skilled in the art would not apply the teachings of the invention in a prior art context where the compressor is not connected to a variable-pitch blower driven by a gearbox. Indeed, the performance cost is not justified by the minimal need to increase angle of attack. In the case of a compressor connected to a variable-pitch blower driven by a gearbox (which is the context of the present invention), it turns out that fixed-pitch compressors can withstand a very high angle of attack under certain operating conditions. Since a prior art blade is incapable of stably withstanding such an angle of attack, the designer is forced to use such a blade in a way that severely penalizes its performance (which is therefore not incentivized).Indeed, such a blade in a state-of-the-art structure, capable of only 10° of angle of attack but required to maintain a stable 20°, will have to be used at its performance point at -10° of angle of attack, resulting in numerous aerodynamic losses (intrados separation, shock, vortices). These additional losses are far greater than those caused by the skin vortices due to the described undulations, and would allow the blade to be used closer to its nominal angle of attack (for example, while a state-of-the-art blade is used at -10° in performance to be able to maintain its maximum 10° angle of attack, a wavy blade capable of 15° of angle of attack could be used at only -5° of its nominal angle of attack in performance).

[0064] The invention also relates to an aircraft turbomachine compressor 120, comprising at least one blade 121, preferably at least one row of blades 121, comprising all or part of the described features. The invention also relates to a turbomachine comprising the compressor. This is an axial compressor for a turbomachine. More particularly, it is a transonic variable-geometry compressor. Even more particularly, it is a transonic variable-geometry compressor connected to a fan, itself with variable pitch or variable pitch and driven by the gearbox. The blade 121 or the row of blades 121 may be mounted on the last stator of the compressor, some of whose stators have variable pitch. The blade 121 or the row of blades 121 may be mounted on the last stator 124 at the compressor outlet, which must direct the flow to a purely axial velocity.The blade 121 or the row of blades 121 can be supported by the first fixed-pitch stator of the compressor, some of whose stators have variable pitch. The blade 121 or the row of blades 121 can be supported by any fixed-pitch stator of the compressor, some of whose stators have variable pitch.

[0065] The compressor incorporating the blade 121 can be located upstream or downstream of a gooseneck. The compressor can also be located between two goosenecks. The compressor can be located downstream of the gearbox, which is itself downstream of a variable-pitch fan (or variable-pitch propeller) – shrouded or unshrouded. The blade 121 allows for a combined large airflow deflection and operation with a large angle of attack. This can be achieved through the chord law or a combination of the chord law with one or more of the following laws: the skeleton angle law, the tangential stacking law, and the thickness law, or with a tangential stacking law independently of the other laws. This can also be achieved through any one of these laws considered individually, or through a combination of two or more of them.

[0066] The 121 blade described previously can be classified as a stator blade.

[0067] The present invention has been described in relation to specific embodiments, which are purely illustrative and should not be considered limiting. Generally, it will be obvious to a person skilled in the art that the present invention is not limited to the examples illustrated and / or described above.

Claims

Demands 1. Stator blade for aircraft turbomachine compressor, the blade comprising - a leading edge (22) and a trailing edge (24) - a chord (26) extending between the leading edge and the trailing edge, the blade exhibiting an evolution of the chord length according to sections of the blade in a plane transverse to the leading edge, the evolution of the chord length being in troughs and crests, the amplitude of the chord length crests relative to an average chord is between 5% and 20%, the wavelength between two chord length crests is between 10% and 50% of the length of an average chord.

2. Blade according to claim 1, wherein the amplitude of the chord length crests relative to an average chord is 10%.

3. Blade according to claim 1 or 2, wherein the amplitude of the chord length crests relative to an average chord is between 1 mm and 5 mm in absolute value.

4. Blade according to any one of the preceding claims, wherein the wavelength between two chord length crests is between 30% and 35% of the average chord length.

5. Blade according to any one of the preceding claims, wherein the wavelength between two chord length crests is between 10% and 35% of the average chord length.

6. Blade according to any one of the preceding claims, wherein the wavelength between two chord length crests is between 30% and 50% of the average chord length.

7. Blade according to any one of the preceding claims, wherein the wavelength between two chord length crests is between 5% and 20% of the blade height at the leading edge (22), preferably 10%.

8. Blade according to any one of the preceding claims, wherein the wavelength between two crests of chord length is between 2mm and 20mm in absolute value.

9. Blade according to any one of the preceding claims, further comprising an intrados (16), an extrados (18) and an equidistant line (20) between the intrados and the extrados, the blade having a skeleton angle (29) between a tangent (28) to the equidistant line (20) at the leading edge and a driving axis (X) of the turbomachine, the skeleton angle (29) at the leading edge having an evolution along sections of the blade in a plane transverse to the leading edge, the evolution of the angle (29) being in troughs and crests.

10. Blade according to the preceding claim, wherein the evolution of the angle (29) at the leading edge is proportional to the evolution of the chord length.

11. Blade according to the preceding claim, wherein the amplitude of the skeleton angle crests (29) is a coefficient (K), the coefficient (K) being between 0° per percent of amplitude of the chord length and 2° per percent of amplitude of the chord length, preferably between 0.5° and 1° per percent of amplitude of the chord length.

12. Blade according to one of the two preceding claims, wherein the direction of evolution of the skeleton angle (29) is to increase the angle at the leading edge when the evolution of the chord length is towards the hollows.

13. Blade according to one of the three preceding claims, wherein the wavelength between two skeleton angle crests (29) is between 10% and 50% of the mean chord length, preferably between 30% and 35% of the mean chord.

14. Blade according to any one of the four preceding claims, wherein the wavelength between two skeleton angle crests (29) is between 5% and 20% of the blade height at the leading edge (22), preferably 10%.

15. Blade according to any one of the preceding claims, the blade having blade sections in a plane transverse to the leading edge, the blade comprising an evolution of an offset of one blade section relative to another, taken in a plane transverse to the leading edge, the offset being in rotation about a driving axis (X) of the turbomachine, the evolution of the offset being in troughs and crests, the amplitude of the offset crests being between 5% and 20% of the mean chord length, preferably 10%, the wavelength between two offset crests being between 10% and 50% of the mean chord length, preferably between 30% and 35% of the mean chord length.

16. Blade according to the preceding claim, wherein the wavelength between two offset crests is between 5% and 20% of the blade height at the leading edge (22), preferably 10%.

17. Blade according to one of the two preceding claims, wherein the blade cut is offset towards the extrados when the evolution of the chord length is towards the hollows.

18. Blade according to any one of the preceding claims, further comprising an intrados (16), an extrados (18) and a thickness between the intrados (16) and the extrados (18), each cross-section having a thickness evolution exhibiting a maximum thickness value, the blade having a maximum thickness evolution along the leading edge of the blade, this evolution exhibiting a waviness.

19. Blade according to the preceding claim, wherein the wavelength between two peaks of maximum thickness is between 10% and 50% of the length of an average chord, preferably between 30% and 35% of the length of an average chord.

20. Blade according to one of the two preceding claims, wherein the wavelength between two peaks of maximum thickness is between 5% and 20% of the height of the blade at the leading edge (22), preferably 10%.

21. Aircraft turbomachine compressor (120) comprising at least one blade (121) according to any one of the preceding claims, preferably at least one row of blades (121) according to any one of the preceding claims.

22. Compressor according to the preceding claim, in which the blade or row of blades is carried by the last stator of the compressor, part of whose stators have variable pitch, or by the first fixed-pitch stator of the compressor, part of whose stators have variable pitch, or by any fixed-pitch stator of the compressor, part of whose stators have variable pitch.

23. Compressor according to one of the two preceding claims, the compressor being an axial compressor for turbomachinery, or a transonic variable geometry compressor, or a transonic variable geometry compressor linked to a blower, itself with variable pitch and driven by the reducer.

24. Aircraft turbomachine (100) comprising a compressor according to any one of the preceding claims.

25. Turbomachine according to the preceding claim, wherein at least one blade (121) or at least one row of blades (121) is upstream or downstream of a row of variable pitch blades, in the direction of airflow.

26. Turbomachine according to one of two preceding claims, comprising a fan (110) with variable pitch driven by a speed reducer, at least one blade (121) or at least one row of blades (121) is downstream of the fan with variable pitch driven by a speed reducer.

27. Turbomachine according to one of the three preceding claims, wherein the fan is unshod.

28. Turbomachine according to any one of the four preceding claims, wherein at least one blade (121) or at least one row of blades (121) is downstream or upstream of a gooseneck.

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