Turbine engine assembly

Abstract

The invention relates to an aircraft turbine engine assembly comprising at least one row of variable-pitch vanes, a compressor having at least one vane (121) downstream of the row of variable-pitch vanes, the vane comprising a pressure side (16) and a suction side (18), a thickness between the pressure side (16) and the suction side (18), a leading edge (22) and a trailing edge (24), a chord (26) extending between the leading edge and the trailing edge, the vane being extendable in an air flow passage (12), the thickness of the vane varying between the leading edge and the trailing edge, the vane being such that the maximum thickness is positioned before 25% of the chord measured from the leading edge. The invention also relates to an aircraft turbine engine.

Description

Technical Field

[0001] This invention relates to turbine engine components and turbine engines. Background Technology

[0002] Aircraft turbine engines are equipped with blades that are crucial to their operational quality. Document US6264429 describes turbine engine blades.

[0003] Some of the blades can be adjustable. These blades have variable pitch. Downstream of these variable-pitch blades along the direction of airflow, the turbine engine may also include a compressor equipped with blades. Due to the presence of variable-pitch blades upstream, these blades are subjected to significant changes in the angle of attack of the airflow.

[0004] Therefore, a turbine engine assembly is needed, in which at least some of the blades are capable of withstanding large changes in the angle of attack of the airflow. Summary of the Invention

[0005] Therefore, the present invention proposes an aircraft turbine engine assembly, comprising: -At least one row of variable pitch blades - A compressor having at least one blade downstream of a row of variable pitch blades, the blade including a pressure side and a suction side, a thickness between the pressure side and the suction side, a leading edge and a trailing edge, a chord extending between the leading edge and the trailing edge, the blade being able to extend into an airflow passage, the thickness of the blade varying between the leading edge and the trailing edge, the blade such that the maximum thickness is positioned before 25% of the chord from the leading edge.

[0006] According to one variation, the thickness is considered at a height between 0% and 50% of the blade height in the channel.

[0007] According to a variation, the thickness is considered at the following height of the impeller in the channel, which corresponds to the inner wall of the channel and / or 1 / 5 of the height of the impeller in the channel and / or 50% of the height of the impeller in the channel.

[0008] According to one variation, 80% of the maximum thickness, preferably 90%, is reached before 10% of the chord from the leading edge.

[0009] In one embodiment, the component further includes a rotor, which includes at least one row of variable pitch blades.

[0010] According to one variation, at least one blade also includes an equidistant line between the pressure side and the suction side, the blade having an angle between the tangent of the equidistant line and the engine axis, the average variation of which between the leading and trailing edges is less than the average variation between the chord at 60% from the leading edge and the trailing edge. Specifically, the average variation of the angle between the leading and trailing edges is less than the average variation between a first point (located at 60% of the chord from the leading edge) and a second point located at the trailing edge (i.e., in the last 40% of the chord).

[0011] According to a variation, at the height corresponding to the inner wall of the channel of at least one impeller, - 50% of the change in angle between the engine axis and the tangent of the equidistant line occurs in the last 40% of the chord of the blade from the leading edge, and / or - The angle between the engine axis and the tangent of the equidistant line at the leading edge is greater than 55°, and / or - The total change in the angle between the engine axis and the tangent of the equidistant line is greater than 55°.

[0012] According to one variation, at 1 / 5 of the height of at least one blade in the channel... - 50% of the change in angle between the engine axis and the tangent of the equidistant line occurs in the last 40% of the chord of the blade from the leading edge, and / or - The angle between the engine axis and the tangent of the equidistant line at the leading edge is greater than 40°, and / or - The total change in the angle between the engine axis and the tangent of the equidistant line is greater than 50°.

[0013] According to a variation, at the midpoint height of at least one blade in the channel, - 50% of the change in angle between the engine axis and the tangent of the equidistant line occurs at most 50% of the last chord of the blade from the leading edge, preferably 40%, and / or - The angle between the engine axis and the tangent of the equidistant line at the leading edge is greater than 35° and less than 50°, and / or - The total change in the angle between the engine axis and the tangent of the equidistant line is greater than 40°, and / or - The angle between the tangent of the engine axis and the equidistant line at the leading edge is at least 10° smaller than the angle between the tangent of the engine axis and the equidistant line at the height corresponding to the inner wall of the channel, and / or - The change in angle between the engine axis and the tangent of the equidistant line at the leading edge is at least 10° smaller than the change in angle between the engine axis and the tangent of the equidistant line at the leading edge at the height corresponding to the inner wall of the channel.

[0014] The present invention also relates to an aircraft turbine engine comprising the components described above.

[0015] According to one variant, the turbine engine includes a variable pitch fan, with at least one blade or at least one row of blades located downstream of the variable pitch fan, or downstream of a rotor carrying a row of variable pitch blades, or downstream of a stator carrying a row of variable pitch blades.

[0016] In one embodiment, the turbine engine further includes a low-pressure turbine and a high-pressure turbine, as well as a high-pressure compressor. The mechanical power of the low-pressure turbine and the high-pressure turbine is transmitted to the low-pressure compressor and the high-pressure compressor respectively via shafts, and to the fan via a reduction gearbox.

[0017] In this document, the use of the verb "including" and its variations and combinations thereof does not exclude the existence of elements other than those mentioned. The use of the indefinite article "a" or "one," or the definite article "the," to introduce an element does not exclude the existence of multiple such elements. The terms "first," "second," "third," etc., within the scope of this document are used only to distinguish similar elements and do not imply any order among these elements.

[0018] The preferred embodiments and advantages of the blades according to the invention are also applicable, with appropriate modifications, to aircraft compressors and turbine engines, and vice versa. Attached Figure Description

[0019] Other features and advantages of the invention will become apparent from the following detailed description, and with reference to the accompanying drawings for understanding that detailed description: - Figure 1 A schematic cross-section of an aircraft turbine engine is shown; - Figure 2 An example of the blade shape is shown; - Figure 3 The changes in the shape of the impeller were shown; - Figure 4 The variation in the shape of the blades is shown.

[0020] The figures in the accompanying drawings are not drawn to scale. Generally, similar elements are represented by similar reference numerals in the drawings. Within the scope of this document, identical or similar elements may have the same reference numerals. Furthermore, the presence of reference numerals or letters in the drawings is not to be considered limiting, even if such numerals or letters are shown in the claims. Detailed Implementation

[0021] This section describes preferred embodiments of the invention in detail. The accompanying drawings have been used and referenced, but the invention is not limited to the drawings. The figures and / or accompanying drawings described below are merely illustrative and not restrictive.

[0022] This invention relates to an aircraft turbine engine assembly comprising at least one row of variable-pitch blades. Downstream of the row of variable-pitch blades, the assembly further comprises a compressor having at least one blade, the blade comprising a pressure side and a suction side, a thickness between the pressure side and the suction side, a leading edge and a trailing edge, and a chord extending between the leading edge and the trailing edge. The blade extends into an airflow passage, and the blade has a thickness variation between the leading edge and the trailing edge. The blade such that the maximum thickness is located before 25% of the chord from the leading edge. Therefore, the compressor blade is particularly thick in the upstream portion in the direction of airflow, causing the blade to resist large changes in the angle of attack of the airflow.

[0023] In some of these figures, the reference frame is shown as an abstract geometric reference frame, primarily for quantifying and / or visualizing the characteristics of embodiments of the invention. In the context of this document, the terms "axial," "circumferential," and "radial" correspond to directions parallel to the engine axis, directions generally circular around the engine axis, and directions perpendicular to the engine axis, respectively. Figure 1 The reference numerals on the attached figures specifically indicate the direction of the engine axis, represented as direction X. The terms "inner" and "inward" naturally correspond to the direction toward the engine axis X in the radial direction, and the terms "outer" and "outward" correspond to the opposite directions in that direction. When referring to the location of an element in an aircraft turbine engine or compressor, the terms "at the inlet" (and correspondingly "at the outlet") preferably refer to the first (correspondingly, the last) such element approximately upstream (correspondingly, approximately downstream) of the aircraft turbine engine or compressor.

[0024] Figure 1 A cross-section of an aircraft turbine engine 100 is shown, on which the blades according to the invention are planned to be integrated. This can be a dual-flow axial turbine engine comprising multiple components. Along the engine axis X, 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 are sequentially present. These components are known to those skilled in the art. The fan 110 can be a variable-pitch fan, as shown in the diagram. Figure 1109 in the diagram. This is an example of an architecture, as the invention is applicable to other turbine engine architectures. During 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, and via shaft 101 to the fan 110. A reduction gear 111 can be inserted into shaft 101 such that the rotational speed of the fan 110 is proportional to the rotational speed of the low-pressure compressor 120. The rotors of these compressors rotate about the engine axis X, causing the compressors to draw in and compress air to bring it to the inlet of the combustion chamber 160 at a suitable speed, pressure, and temperature. The fan 110 generates a primary airflow 106 and a secondary airflow 107 upstream of the low-pressure compressor 120. The primary airflow 106 is primarily configured to pass axially through the aircraft turbine engine 100 to supply the combustion chamber 160, while the secondary airflow 107 is primarily configured to generate the thrust reaction force required for aircraft flight.

[0025] Despite Figure 1 While not systematically mentioned, each compressor and each turbine comprises one or more axial stages arranged in series. Each stage includes a stator (or rectifier) ​​with fixed blades (or blade assemblies) and a rotor with movable blades (or blade assemblies) that can rotate about the engine axis. For the low-pressure compressor 120 and the high-pressure compressor 130, these fixed and movable blades are designated by 121, 122 and 131, 132, respectively. Within a stage, the rotor draws in and accelerates the main airflow by deflecting the main airflow 106 relative to the engine axis X, and the subsequent stator or rectifier rectifies the flow along the engine axis X and decelerates it by converting a portion of the flow's velocity into pressure. The fixed blade assembly includes blades that can be fixed in orientation and / or variable in orientation (or pitch) (the blades occupying angular pitch positions). These adjustable blades are also referred to as variable pitch blades or variable stator vanes (VSVs). Their special characteristic lies in the fact that the inclination of their chords can vary relative to the engine axis X, and particularly relative to the axes of compressors 120 and 130. The pressure side 16 and suction side 18 surfaces of the impellers can be more or less exposed to the main airflow 106. The pressure side 16 extends more or less relative to the main airflow 106 to regulate and thus rectify the direction of deflection applied to the main airflow 106. The compressor may include one or more stages with adjustable impellers.

[0026] The angular pitch (or in other words, orientation, angular pitch position, or angular position) of the blades within the stator of a fixed-blade assembly is controlled by a variable-pitch system. Figure 1In this configuration, the 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. The 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 description does not limit the number or location of the variable pitch systems that compressors 120 and 130 may include. This type of architecture is given as an example, and intermediate compressors may also be present.

[0027] The compressor extends from upstream to downstream along the engine axis X between an upstream air inlet end and a downstream air outlet end. The compressor includes an airflow passage 12 located between the upstream air inlet end and the downstream air outlet end. Passage 12 is a wall that guides the airflow. For example, passage 12 guides a first airflow 106. The first airflow 106 enters the compressor at the upstream air inlet end and exits the compressor at the downstream air outlet end.

[0028] Figure 2 An example of the shape of a blade 121 is shown. This blade is of a fixed blade type, may have an angular pitch, but preferably no angular pitch, and is used in a low-pressure compressor stator. The blade 121 may have a variable orientation relative to the stator element supporting it, or it may preferably have a fixed orientation relative to the stator supporting it. The blade 121 includes a pressure side 16 and a suction side 18. The blade 121 includes an equidistant line 20 between the pressure side 16 and the suction side 18. The blade 121 also includes a leading edge 22 located upstream in the airflow and a trailing edge 24 located downstream in the airflow. 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.

[0029] The blade extends into the airflow passage 12. The airflow passage 12 has an annular structure. The airflow passage 12 has a circular structure. Therefore, the passage 12 may include an inner wall in a radially inward direction. The passage 12 may include an outer wall in a radially outward direction. In cross-section, the passage 12 includes a height along which the blade 121 extends. The height of the passage is between the inner wall and the outer wall. This height extends in a radial direction transverse to the engine axis X. In the following, at different heights in the direction transverse to the engine axis X, the shape of the blade varies according to the blade cross-section along the engine axis X. The present invention relates to the shape of the blade, and particularly to the profile of the blade (corresponding to the cross-sections in the blade at different passage heights, formed by streamlines).

[0030] The angle formed at the leading edge 22 by the direction of the airflow relative to the tangent 28 of the equidistant line 20 is called the angle of attack (angle of attack setting angle or angle of attack setting). Figure 2In the diagram, this is the angle between any of the arrows 30 indicating the direction of airflow and the tangent 28 of the equidistant line 20 (the airflow can be "above" or "below" the tangent 28). Arrows 30 indicate different angles of attack (maximum values). The distance between two arrows 30 corresponds to the range of angles of attack. In the context of this invention, the blade is capable of providing large air deflection for operation over a wide range of air angles of attack (e.g., a range of 20° between the two arrows 30 (the distribution of this range can vary on both sides of the tangent 28)). The variation of angle 29 between the tangent 28 of the equidistant line 20 and the engine axis X along chord 26 is referred to as the skeleton angle rule. The variation of the blade thickness 50 along chord 26 is referred to as the thickness rule.

[0031] The impeller 121 can rectify the airflow at its trailing edge, ensuring that the rectified airflow is as close as possible to the engine axis X or at a suitable angle to the downstream rotor, but as constant as possible and independent of what happens upstream. The impeller 121 ensures a high degree of airflow deflection. Figure 2 In the diagram, the rectified airflow is indicated by arrow 25. Air deflection is the angle between arrow 25 and arrow 30. Therefore, air deflection varies with arrow 30. Depending on the angle of arrow 30, the angle of arrow 25 may vary less due to the stator's rectification efficiency at the angle of attack. Blade 121 can provide a large deflection of up to 60°. Furthermore, blade 121 is capable of withstanding significant changes in air angle of attack, as indicated by arrow 30.

[0032] Figure 3 The variation in the shape of the blade 121 (or the variation in curvature or the variation in camber) is shown. Figure 3 The dimensionless skeleton angle rule ("A") is shown on the ordinate, proportional to the position ("C") of the chord 26 along the contour on the abscissa (from the leading edge 22 - dimensionless position). This shows the variation of angle 29 from the leading edge 22 to the trailing edge 24. Therefore, the proportional value of angle 29 is 1 at the leading edge and 0 at the trailing edge. Figure 3 This shows how the variation of angle 29 is distributed, particularly how 50% of the variation of angle 29 is located. The special feature of the impeller is that its shape (or profile) varies according to the height of the channel (or in other words, according to the height of the impeller). Figure 3The diagram illustrates the variation in shape of the blade in its lower portion (between 0 and 50% of its height within the channel). Curve 36 shows the shape or profile of the blade in a cross-section at a height within the channel corresponding to the inner wall of the channel. This cross-section is located at the base of the blade in a radially outward direction. Curve 38 shows the shape or profile of the blade in a cross-section at a height within the channel corresponding to 1 / 5 of its height from the inner wall of the channel. This cross-section is located at 1 / 5 of the blade's height in a radially outward direction. Curve 40 shows the shape or profile of the blade in a cross-section at a height within the channel corresponding to the intermediate height from the inner wall of the channel. This cross-section is located at the intermediate height of the blade.

[0033] The average variation of angle 29 of blade 121 between the leading and trailing edges is less than the average variation between the 60% of the chord from the leading edge and the trailing edge. In other words, the average variation of angle 29 is greater in the last 40% of the chord from the leading edge. Specifically, the average variation of angle 29 between the leading edge 22 and the trailing edge 24 is less than the average variation between a first point (located at 60% of the chord 20 from the leading edge 22) and a second point located at the trailing edge 24 (and therefore in the last 40% of the chord). Blade comprises a certain angle 29 at the leading edge 22 and a certain angle 29 at the trailing edge. Angle 29 follows a variation (or "evolution" or "deflection") from the leading edge 22 to the trailing edge 24. This variation is more or less amplified from the leading edge to the trailing edge. The average distribution (i.e., average variation) from the leading edge 22 to the trailing edge 24 is less than the average distribution (i.e., average variation) from the first point to the second point. The slope of the dimensionless average change (or "evolution" or "deflection") of angle 29 from leading edge 22 to trailing edge 24 is less than the slope of the dimensionless average change (or "evolution" or "deflection") of angle 29 from the first point to the second point. The total change of angle 29 is the difference between angle 29 at leading edge 22 and angle 29 at trailing edge.

[0034] In the impeller according to the invention: - Angle 29 varies between 0% and 50% of the height of the impeller in the channel. - At the height of the blade in the channel corresponding to the inner wall of the channel, at 1 / 5 of the height of the blade in the channel, and at the middle height of the blade in the channel, 50% of the change in the angle 29 between the tangent of the equidistant line and the engine axis is between the first point and the second point.

[0035] This type of blade allows for an optimal trade-off between performance and stability. In fact, the greater the rule of the skeleton angle (the average change in the angle between the tangent of the equidistant line and the engine axis) downstream of the chord (more than halfway from the leading edge), the more stable the cross-section at the height in question, but it also generates a greater pressure drop. In this way, the average change in the chord as a function of height means that only the critical height (20%) incurs a significant pressure loss cost, the peripheral heights (0-50%) incur a moderate cost (at the peripheral heights, stability difficulties are less pronounced but real), and there is no cost at all at the stable heights (above 50% height).

[0036] Figure 3 The curves 36, 38, and 40 are steeper when the x-coordinate is above 0.6 than before 0.6. The curves are steeper between 60% of the chord from the leading edge and the trailing edge.

[0037] Therefore, regarding the blade curvature rule, due to the increased angle of attack which deflects the airflow near the leading edge, blade 121 can limit the deflection of the airflow in the upstream portion of the blade, while controlling the remaining deflection of the airflow in the downstream portion of the blade to remain compatible with its stability. Beyond 60% of the chord from the leading edge (i.e., beyond this first point), the change in the angle between the engine axis X and the tangent 28 depends on the position of the blade cross-section according to the channel height. In this document, the change in angle 29 refers to the average change in angle 29.

[0038] Curve 36 at the height of the blade in the channel (which corresponds to the inner wall of the channel) shows that the average change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 is less than the average change between the chord from the leading edge (60%) and the trailing edge. Beyond the 60% of the chord from the leading edge 22, the change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 is greater. The slope of curve 36 increases. According to the profile of the blade 121 along curve 36, 50% of the change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 occurs in the last 40% of the chord of the blade from the leading edge 22. The angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 at the leading edge 22 is greater than 50°. The total change of the angle 29 between the engine axis X and the tangent 28 of 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 cross-section at the base shows a total variation of more than 55°, with 50% of the total variation distributed over the last 40% of chord 26 from the leading edge.

[0039] Curve 38 at the height of the blade in the channel (which corresponds to 1 / 5 of the height in the channel) shows that the average change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 is less than the average change between the chord from the leading edge to the trailing edge (60%). Beyond the chord from the leading edge 22, the change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 is greater. The slope of curve 38 increases. According to the profile of the blade 121 along curve 38, 50% of the change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 occurs in the last 40% of the chord of the blade from the leading edge 22. The angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 at the leading edge 22 is 40°. The total change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 is greater than 50°. The cross-section at 1 / 5 of the height shows a total variation of more than 50°, with 50% of the total variation distributed over the last 40% of chord 26.

[0040] Curve 40 at the height of the blade in the channel (which corresponds to the middle height of the channel) shows that the average change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 is less than the average change between 60% of the chord from the leading edge and the trailing edge. Beyond 60% of the chord from the leading edge 22, the change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 is greater. The slope of curve 40 increases. According to the profile of the blade 121 along curve 40, 50% of the change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 occurs at the last, at most 50%, preferably 40%, of the chord of the blade from the leading edge 22. Thus, the middle height cross-section causes this change to occur earlier. The angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 at the leading edge 22 is greater than 35° and less than 50°. At the middle height of the channel, the total change of the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 is greater than 40°. At the intermediate height, the point where the change reaches 50% is 10% ahead of its position on the inner wall of the channel. At the intermediate height of the channel, the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 at the leading edge is at least 10° smaller than the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 at the leading edge at the height corresponding to the cross-section of the inner wall of the channel. The change in the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 at the leading edge is at least 10° smaller than the change in the angle 29 between the engine axis X and the tangent 28 of the equidistant line 20 at the leading edge at the height corresponding to the inner wall of the channel.

[0041] Figure 4 The shape of the blade 121 is shown to be different. Figure 4The thickness rule ("E") on the ordinate is shown based on the position ("C") of the chord 26 along the profile on the abscissa (from the leading edge 22 - dimensionless position). Figure 4 The variation in thickness between the pressure side 16 and the suction side 18 from the leading edge 22 to the trailing edge 24 is shown. Figure 4 This illustrates how the thickness variation of the impeller 121 is distributed. The impeller is unique in that its shape (or profile) varies according to the height of the channel (or in other words, according to the height of the impeller). Figure 4 The shape of the blade is shown to vary in the lower part of the blade (between 0 and 50% of the blade height in the channel). Figure 4 The diagram illustrates the variation in thickness of blade 121 in cross-section at its height within the channel. Curve 42 shows the shape or profile of the blade according to a cross-section at a height corresponding to the inner wall of the channel. This cross-section is located at the base of the blade in a radially outward direction. Curve 44 shows the shape or profile of the blade according to a cross-section at a height corresponding to 1 / 5 of the height from the inner wall of the channel. This cross-section is located at 1 / 5 of the blade's height in a radially outward direction. Curve 46 shows the shape or profile of the blade according to a cross-section at a height corresponding to the intermediate height of the channel from the inner wall of the channel. This cross-section is located at the intermediate height of the blade's height.

[0042] The blade 121 shows the variation in thickness between the leading and trailing edges. The blade is designed such that the maximum thickness occurs before 25% of the chord from the leading edge. In other words, the variation in blade thickness is greatest at the first 25% of the chord from the leading edge. Figure 2 The curves 42, 44, and 46 are steeper before 0.25 on the x-axis than after 0.25. The curves are steeper between the leading edge and 25% of the chord from the leading edge.

[0043] For example, regarding the blade thickness rule, because the blade experiences a significant change in air angle of attack (especially in the upstream portion of the blade), a thicker blade upstream of the chord allows the blade to better resist changes in the upstream portion of the blade. Regarding the thickness rule, as air flows around the leading edge, the curvature of the leading edge produces a significant acceleration, which destabilizes the airflow. Shifting the maximum thickness of the profile towards the leading edge minimizes its curvature.

[0044] Thickness is considered between 0 and 50% of the blade height in the channel. Thickness near the leading edge stabilizes blade behavior but increases pressure drop. Therefore, the present invention aims to stabilize the blade profile at a height between 0 and 50%, where the blade is most unstable (pressure drop between 50-100% of the height is less critical because the blade does not exhibit stability difficulties). The maximum thickness, located before 25% of the chord from the leading edge, is considered in the cross-section of the blade in the channel at the following heights: corresponding to the inner wall of the channel (curve 42) and / or at 1 / 5 of the blade height in the channel (curve 44) and / or at 50% of the blade height in the channel (curve 46). Preferably, 80%, preferably 90%, of the maximum thickness is reached before 10% of the chord from the leading edge. This further strengthens the blade and enables it to withstand significant changes in the angle of attack at the leading edge. Relatively speaking, the maximum thickness is greater than 6% of the chord. Furthermore, thickening towards the leading edge improves the stability of the flow in the presence of an angle of attack by reducing its radius of curvature until a certain point. Too much thickness too close to the leading edge results in a very small local radius (in the extreme case, with 100% of the maximum thickness at 0% of the leading edge, we get a rectangle with two very unstable right angles at the point of angle of attack, although the radius of curvature at the leading edge is infinite because we have a leading edge that becomes a line segment). Achieving 90% of the maximum thickness at 10% of the chord is the most robust to ensure the maximum radius of curvature at the leading edge without excessively reducing the radius of curvature between 10% and 25% of the chord.

[0045] All these characteristics of the blade 121 can be considered individually or in combination. In particular, the average variation of angle 29 relative to the chord and the variation of thickness relative to the chord can be considered individually or in combination. The blade ensures uniform flow along the blade. The blade prevents air separation before reaching the trailing edge. The blade provides a combination of high deflection and high angle-of-attack setting of the airflow. The blade can withstand large angle-of-attack variations in the airflow. This improves the performance of a compressor implementing such a blade. This also enables improved performance of a turbine engine assembly according to the invention, which includes at least one row of variable-pitch blades and includes a compressor having at least one blade 121 as described above downstream of this row of variable-pitch blades. This is because the blade 121 located downstream of the variable-pitch blades in the airflow direction is subjected to significant angle-of-attack variations in the airflow due to the presence of the variable-pitch blades upstream. The described blade 121 is capable of withstanding large variations in the angle of attack of the airflow.

[0046] The invention also relates to an aircraft turbine engine compressor 120, which includes at least one blade 121, preferably at least one row of blades 121, the blades including all or some of the features described. The invention also relates to a turbine engine assembly including at least one row of variable-pitch blades (rotor or stator), and a compressor 120 downstream of this row of variable-pitch blades. The invention further relates to a turbine engine including a compressor or an assembly having a compressor. Specifically, the compressor is an axial compressor for a turbine engine. More specifically, the compressor is a variable geometry transonic compressor (having movable blades that can move according to the pitch on the stator). The compressor may be located downstream of a fan 110 of the turbine engine, the fan itself having a variable pitch. The row of blades 121 may be carried in the direction of airflow by the final stator 124 at the outlet of the compressor (referred to as an outlet guide vane (OGV)) (the final stator must be rectified toward pure axial velocity convection) and / or by the first stator of the compressor having a fixed pitch. In the direction of airflow, the blade row 121 is located downstream (directly downstream, i.e. immediately following) or indirectly downstream (i.e. separated by one or more other blade rows) of a variable pitch blade row (rotor, stator, or fan). The blade 121 preferably provides a combination of high airflow deflection and high angle-of-attack operation through a combination of thickness rules and skeleton curvature rules.

[0047] The aforementioned blade 121 with variations in angle 29 and / or thickness can be described as a stator blade, or as a shaped blade, or as a shaped stator blade.

[0048] The invention has been described in conjunction with specific embodiments, which are exemplary and should not be considered limiting. Generally, it will be apparent to those skilled in the art that the invention is by no means limited to the examples shown and / or described above. For example, the invention can be applied, in combination with or independently of variations in angle 29 between 0 and 50% of the height, to the blade tip.

Claims

1. An aircraft turbine engine assembly, comprising: -At least one row of variable pitch blades - A compressor having at least one blade (121) downstream of a row of variable pitch blades, the blade comprising a pressure side (16) and a suction side (18), a thickness between the pressure side (16) and the suction side (18), a leading edge (22) and a trailing edge (24), a chord (26) extending between the leading edge and the trailing edge, the blade being capable of extending into an airflow passage (12), the thickness of the blade varying between the leading edge and the trailing edge, the blade such that the maximum thickness is positioned before 25% of the chord from the leading edge, and 90% of the maximum thickness is reached before 10% of the chord from the leading edge, the thickness being considered at a height between 0 and 50% of the blade in the passage.

2. The component according to any one of the preceding claims, wherein, The thickness is taken into account at the following heights of the blades in the channel, which correspond to the inner wall of the channel and / or 1 / 5 and / or 50% of the height of the blades in the channel.

3. The component according to any one of the preceding claims further includes a rotor, said rotor comprising at least one row of variable pitch blades.

4. The component according to any one of the preceding claims, wherein, The at least one blade (121) further includes an equidistant line (20) between the pressure side and the suction side, the blade (121) having an angle (29) between the tangent (28) of the equidistant line (20) and the engine axis (X), the angle (29) having an average variation between the leading edge and the trailing edge less than the average variation between a first point located at 60% of the chord from the leading edge and a second point located at the trailing edge.

5. The component according to the preceding claim, wherein, At the height corresponding to the inner wall of the channel, at least one blade (121) in the channel, - 50% of the change in the angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) is in the last 40% of the chord of the blade from the leading edge.

6. The component according to any one of the preceding two claims, wherein, At the height corresponding to the inner wall of the channel, at least one blade (121) in the channel, - The angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) at the leading edge is greater than 55°.

7. The component according to any one of the preceding two claims, wherein, At the height corresponding to the inner wall of the channel, at least one blade (121) in the channel, - The total variation of the angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) is greater than 55°.

8. The component according to any one of the preceding four claims, wherein, At 1 / 5 of the height of at least one blade (121) in the channel, - 50% of the change in the angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) is in the last 40% of the chord of the blade from the leading edge.

9. The component according to any one of the preceding five claims, wherein, At 1 / 5 of the height of at least one blade (121) in the channel, - The angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) at the leading edge is greater than 40°.

10. The component according to any one of the preceding six claims, wherein, At 1 / 5 of the height of at least one blade (121) in the channel, - The total variation of the angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) is greater than 50°.

11. The component according to any one of the preceding seven claims, wherein, At the middle height of at least one blade (121) in the channel, - 50% of the change in the angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) is at most 50% of the last chord of the blade from the leading edge, preferably 40%.

12. The component according to any one of the preceding eight claims, wherein, At the middle height of at least one blade (121) in the channel, - The angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) at the leading edge is greater than 35° and less than 50°.

13. The component according to any one of the preceding nine claims, wherein, At the middle height of at least one blade (121) in the channel, - The total variation of the angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) is greater than 40°.

14. The component according to any one of the preceding ten claims, wherein, At the middle height of at least one blade (121) in the channel, - The angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) at the leading edge is at least 10° smaller than the angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) at the leading edge at the height corresponding to the inner wall of the channel.

15. The component according to any one of the preceding eleven claims, wherein, At the middle height of at least one blade (121) in the channel, The change in the angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) at the leading edge is at least 10° smaller than the change in the angle (29) between the engine axis (X) and the tangent (28) of the equidistant line (20) at the leading edge at the height corresponding to the inner wall of the channel.

16. An aircraft turbine engine comprising the components according to any one of the preceding claims.

17. The turbine engine according to the preceding claim, comprising a variable pitch fan (110), wherein the at least one blade (121) or the at least one row of blades (121) is located downstream of the variable pitch fan, or downstream of a rotor carrying a row of variable pitch blades, or downstream of a stator carrying a row of variable pitch blades.

18. The turbine engine according to any one of the preceding two claims further includes a low-pressure turbine (150) and a high-pressure turbine (140) and a high-pressure compressor (130), wherein the mechanical power of the low-pressure turbine (150) and the high-pressure turbine (140) is transmitted to the low-pressure compressor (120) and the high-pressure compressor (130) respectively via shafts (101, 102) and to the fan (110) via a reduction gearbox (111).

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