Inhomogeneous pitch of the blades of a stator of an aircraft turbine engine

Abstract

The present invention relates to an assembly (1) for a compressor (120, 130) of an aircraft turbomachine (100) and comprising: a stator (10) comprising a row of blades (121), each of said blades being arranged along a pitch angle (65, 66, 67); and a rotor (20) comprising a row of blades (122) and located downstream of the stator (10) along an axial direction; characterized in that the pitch angle of the stator blades (10) is inhomogeneous along a circumferential direction and is a function of a circumferential distortion of a static pressure difference between the upstream and downstream of the rotor (20).

Description

technical field

[0001] The present invention relates to an assembly for a compressor of an aircraft turbomachine and comprising a stator and a rotor located downstream of the stator. State of the art

[0002] In the prior art, a rotor is generally designed assuming circumferential homogeneity of the static pressure difference between the upstream and downstream sides of the rotor, i.e., zero distortion of said static pressure difference. However, since it is known to a person skilled in the art that the distortion of the static pressure difference between the upstream and downstream sides of a rotor may not be homogeneous along a circumferential direction, for example due to an arm located upstream or downstream of the rotor, a stability margin is taken into account during the rotor design, this margin being quantified as a function of the maximum amplitude of the circumferential distortion of the static pressure difference between the upstream and downstream sides of the rotor.

[0003] In practice, the circumferential distortion of static pressure difference between the upstream and downstream sides of a rotor can reach 10% of the average value of said static pressure difference. In such a case, a rotor can be locally driven, opposite the peak of the static pressure difference, to a value up to 5% higher or lower than the average static pressure difference for which the rotor was designed, thus consuming a significant portion of its stability margin. It is then necessary to constrain the rotor design with a larger stability margin, which can degrade the rotor's efficiency.

[0004] Therefore, there is a need to be able to design a rotor with better efficiency when exposed to circumferential distortion from a static pressure difference between its upstream and downstream sides. Description of the invention

[0005] To this end, the invention proposes an assembly for a compressor of an aircraft turbomachine and comprising: a stator comprising a row of blades, each of said blades being arranged at a pitch angle; and a rotor comprising a row of blades and located downstream of the stator along an axial direction; characterized in that the pitch angle of the stator blades is inhomogeneous along a circumferential direction and is a function of a circumferential distortion of a static pressure difference between the upstream and downstream sides of the rotor, in that a stator blade has a more axial pitch (or in other words, an orientation) for (or in other words, facing) a local value of the static pressure difference higher than an average value of the static pressure difference.in that a stator blade has a more circumferential pitch (or orientation) for (or in other words, facing) a local value of the static pressure difference that is lower than an average value of the static pressure difference, in that the inhomogeneous pitch angle of the stator blades is fixed (or in other words, imposed, or determined) at the time of stator assembly in order to take account of the circumferential distortion of the static pressure difference, and in that the stator is intended to (or in other words, suitable for, or designed to) be the furthest upstream stator within the compressor along the axial direction, or in other words, along an airflow passing through the compressor during its normal operation.

[0006] Preferably, in the assembly according to the invention, the rotor is located directly downstream of the stator, i.e. directly follows the stator along the axial direction.

[0007] Generally, a compressor in an aircraft turbomachine can comprise a plurality of stator-rotor pairs and therefore a plurality of assemblies according to the invention. In particular, the stator included in the assembly according to the invention is intended to (or in other words, suitable for, or designed to) be the upstream stator within a compressor (in particular a low-pressure compressor) along an airflow passing through the compressor during its normal operation. Preferably, the stator included in the assembly according to the invention is the stator located furthest upstream along a primary airflow passing through a compressor (in particular a low-pressure compressor) in or included in an aircraft turbomachine. In other words, the stator included in the assembly according to the invention is preferably the IGV (Integrated Gas Ventilation) of the primary flow.The term "IGV" stands for "Inlet Guide Vanes," a term familiar to anyone skilled in the art. The term "primary flow" refers to the airflow passing through the components at the core of a turbomachine—such as a compressor, combustion chamber, or turbine—along the turbomachine's drive shaft; this term is also familiar to anyone skilled in the art.

[0008] The circumferential distortion of a static pressure difference between the upstream and downstream sides of the rotor can be caused by the presence of an obstacle, such as an arm, upstream or downstream of the rotor. This circumferential distortion of a static pressure difference between the upstream and downstream sides of the rotor can be determined by calculation during the design of the rotor included in the assembly according to the invention. Furthermore, an inhomogeneous pitch angle of the stator blades along a circumferential direction means that the pitch angle of the stator blades varies along the circumferential direction and is therefore not homogeneous along the circumferential direction. According to the invention, the inhomogeneity of the stator blade pitch angle is imposed during the stator assembly. In addition, according to the invention, the inhomogeneity of the stator blade pitch angle can concern some or all of the stator blades.

[0009] Preferably, the stator blades are distributed homogeneously (or uniformly) along the stator's circumference. In other words, the stator blades are preferably equally spaced along the stator's circumference. More precisely, the spacing between successive stator blades along a circumferential direction is preferably homogeneous (or uniform), or even identical. This does not preclude the stator blade pitch angle from being inhomogeneous along a circumferential direction.

[0010] In this document, the term "local" refers to a point along a direction. Thus, a local pressure difference along a circumferential direction refers to the pressure difference at a point along the circumferential direction. Similarly, a local rotor flow along a circumferential direction refers to the flow through the rotor at a point along the circumferential direction.

[0011] According to the invention, the local flow rate of a rotor is adapted to a local static pressure difference between the upstream and downstream sides of the rotor by varying the pitch angle of the stator blades located upstream of the rotor. A more axial pitch of a blade increases the local flow rate at the blade, while a more circumferential pitch decreases the local flow rate at the blade. In the case of a rotor exposed to a circumferential distortion of the static pressure difference between its upstream and downstream sides, i.e.Given an inhomogeneous static pressure difference between the upstream and downstream sides of the rotor along a circumferential direction, the invention proposes to balance the local rotor flow rate with the local value of the static pressure difference along the circumferential direction by varying the pitch angle of the stator blades located upstream of the rotor according to the local value of the static pressure difference at each stator blade, thus resulting in an inhomogeneous pitch angle of the stator blades along the circumferential direction. Indeed, a circumferential distortion of a static pressure difference causes the local value of the static pressure difference to vary around an average value of said difference, said local value being higher than, equal to, or lower than said average value.

[0012] For the purposes of this document, the pitch angle of a stator blade is defined as the angle between a blade chord and the stator's axis of revolution. A blade chord is a straight line segment connecting the leading and trailing edges of the blade and lying in a plane perpendicular to a longitudinal (or radial) axis of the blade. A stator blade may be oriented in a direction more parallel to the stator's axis of revolution, in which case, for the purposes of this document, it is referred to as a "more axial orientation" or "more axial pitch" of the blade, which corresponds to a reduced (or smaller) pitch angle.In addition, a stator blade can take an orientation more perpendicular to the axis of revolution of the stator, in which case, within the framework of this document, we will speak of a "more circumferential orientation" or a "more circumferential pitching" of the blade, which corresponds to a higher pitching angle.

[0013] According to the invention, the pitch angle of the stator blades is a function of a circumferential distortion of a static pressure difference between the upstream and downstream sides of the rotor, as follows: the pitch (or orientation) of a stator blade being more axial, i.e. the stator blade having a reduced pitch angle (or in other words, smaller), facing (or for) a higher local value of the static pressure difference between the upstream and downstream sides of the rotor, thus resulting in a higher local flow rate at the blade; the pitch (or orientation) of a stator blade being nominal, i.e. the stator blade having a nominal pitch angle, facing (or for) an average local value of the static pressure difference between the upstream and downstream sides of the rotor, thus resulting in a nominal local flow rate at the blade; the pitching (or orientation) of a stator blade being more circumferential, i.e. the stator blade having a higher pitching angle, facing (or for) a reduced (or in other words lower) local value of the static pressure difference between upstream and downstream of the rotor, thus resulting in a reduced (or in other words lower) local flow at the blade.

[0014] Thanks to the inhomogeneous positioning of the stator blades imposed (or in other words, fixed or determined) during stator assembly, the invention makes it possible to account, during the design of the rotor located downstream of the stator, for a circumferential distortion of a static pressure difference between the upstream and downstream sides of the rotor. This distortion can be caused by the presence of an obstacle, such as an arm, upstream or downstream of the rotor. Advantageously, the invention reduces the stability margin required during the design of such a rotor and thus allows for the design of such a rotor with improved efficiency.

[0015] The presence of an obstacle, such as an arm, upstream or downstream of the rotor included in the assembly according to the invention can disrupt the airflow through said assembly, and can therefore induce a circumferential distortion of the static pressure difference between the upstream and downstream sides of the rotor, also referred to as "circumferential inhomogeneity of a static pressure difference" in the context of this document. More specifically, such an obstacle located upstream or downstream of the rotor can induce a circumferential distortion of the static pressure upstream or downstream of the rotor, which can then lead to a circumferential distortion of the static pressure difference between the upstream and downstream sides of the rotor.Furthermore, inhomogeneity in the airflow through the rotor due to dynamic factors occurring during rotor operation can also lead to circumferential distortion of the pressure difference between the upstream and downstream sides of the rotor. However, this would be a circumferential distortion of a dynamic pressure difference between the upstream and downstream sides of the rotor, as opposed to a circumferential distortion of a static pressure difference between the upstream and downstream sides of the rotor, which depends solely on factors known at the time of the rotor's design and / or assembly. In particular, a circumferential distortion of a static pressure difference between the upstream and downstream sides of a rotor depends on factors that are essentially geometric in nature and that can be determined during the rotor's design, without having to operate the rotor or simulate its operation.Among these factors, we can cite for example the presence of one or more obstacles upstream or downstream of the rotor, such as an arm, such an arm being for example a structural arm and / or a radial arm, or even a pylon, said arm being able to also allow the passage of servitudes.

[0016] The inventors propose several possible embodiments of the invention including optional features, some of which can be combined.

[0017] According to one embodiment of the invention, an assembly according to the invention further comprises an arm, the circumferential distortion of a static pressure difference being caused at least in part by the presence of said arm in the assembly. This arm may be, for example, a structural arm and / or a radial arm, or even a pylon, and may also allow the passage of service lines. This arm may constitute an obstruction in the flow of air passing through the assembly according to the invention, thus causing at least in part said circumferential distortion of a static pressure difference between the upstream and downstream sides of the rotor. The presence of said arm in the assembly according to the invention does not preclude the presence of other arms, or even other obstructions, in said assembly, and said other arms or obstructions may also cause at least in part or contribute to said circumferential distortion of a static pressure difference.

[0018] Preferably, the arm is located upstream of the stator, preferably directly upstream of the stator, or downstream of the rotor, preferably directly downstream of the rotor. If the arm is located upstream of the stator, the assembly according to the invention consists of the following successive elements along the axial direction: arm, stator, rotor. If the arm is located directly upstream of the stator, it means that the arm directly precedes the stator along the axial direction. If the arm is located downstream of the rotor, the assembly according to the invention consists of the following successive elements along the axial direction: stator, rotor, arm. If the arm is located directly downstream of the rotor, it means that the arm directly follows the rotor along the axial direction.

[0019] According to one embodiment of the invention, an assembly according to the invention further comprises a variable blade pitch system capable of uniformly modifying the pitch angle of the stator blades. This embodiment has the advantage of allowing the airflow to be adapted to the operating conditions of a device comprising an assembly according to the invention, such as, for example, a turbomachine compressor. In particular, controlling the orientation of the stator blades ensures that an airflow passing through the stator at a given velocity is directed onto the rotor located downstream of the stator at an acceptable angle, taking into account the airflow velocity.

[0020] According to this embodiment, the control of the variable pitch of the stator blades, i.e., the control of the stator blade orientation, is common to all the stator blades. In other words, although the pitch angle of the stator blades is inhomogeneous along a circumferential direction, the modification of this pitch by the variable pitch system is uniform for all the stator blades; i.e., the same increment or decrement of pitch angle is applied to each of the stator blades by the variable pitch system. Indeed, according to the invention, the inhomogeneity of the stator blade pitch is intended to balance a local flow rate of a rotor located downstream of the stator with a local static pressure difference between the upstream and downstream sides of the rotor, i.e., at a point along a circumferential direction.Thus, within the framework of the present invention, there is no need to individually adapt the pitch angle of a stator blade to the operating conditions of a device comprising the stator.

[0021] Preferably, the variable blade pitch system comprises a control ring essentially coaxial with the stator, said control ring being mechanically coupled to each of the stator blades via respective levers. According to this embodiment, the length of each of said levers can be configured so that the pitch angle of each stator blade depends on the local value, at the blade, of the static pressure difference between the upstream and downstream sides of the rotor. This embodiment allows for easy implementation of the invention. Indeed, a simple variation in the length of the levers connecting the stator blades to the control ring makes it possible to achieve inhomogeneous pitching of the stator blades along a circumferential direction in order to balance the local flow of the rotor located downstream of the stator with the local value, at the blades, of the static pressure difference between the upstream and downstream sides of the rotor.

[0022] The invention further proposes a compressor for an aircraft turbomachine and comprising an assembly according to the invention.

[0023] Preferably, said compressor is a low pressure compressor, and the stator included in the assembly according to the invention is the stator located furthest upstream within said low pressure compressor.

[0024] The invention further proposes an aircraft turbomachine comprising a compressor according to the invention.

[0025] The invention further proposes an aircraft comprising a turbomachine according to the invention.

[0026] The invention further proposes a method for timing the blades of a stator assembly for a compressor of an aircraft turbomachine and comprising the following steps: a. provide an assembly according to the invention; b. determine a circumferential distortion of a static pressure difference between the upstream and downstream sides of the rotor of the assembly; c. impose on the stator blades of the assembly a pitch angle: inhomogeneous along a circumferential direction, and a function of the circumferential distortion of a static pressure difference determined in the previous step, a method in which a stator blade has a more axial pitch (or orientation) for (or facing) a local value of the static pressure difference higher than an average value of the static pressure difference, in which a stator blade has a more circumferential pitch (or orientation) for (or facing) a local value of the static pressure difference lower than an average value of the static pressure difference, in which the inhomogeneous pitch angle of the stator blades stator is fixed (or otherwise imposed, or determined) to the stator mounting in order to take account of the circumferential distortion of the static pressure gap, and in which the stator is intended to (or otherwise suitable to, or intended to) be the furthest upstream stator within the compressor along an axial direction, or otherwise along an airflow passing through the compressor during its normal operation.

[0027] All possible embodiments and all advantages of any one of the devices according to the invention apply mutatis mutandis to the process according to the invention. Brief description of the figures

[0028] Other features and advantages of the present invention will become apparent upon reading the detailed description that follows, and for understanding which reference should be made to the accompanying figures, among which: there figure 1is a schematic cross-sectional view of an aircraft turbomachine consisting of a turbofan engine; the figure 2 illustrates an operating field of a rotor exposed to circumferential distortion of a static pressure difference between its upstream and downstream sides according to the prior art; the figure 3 illustrates a circumferential distortion of a static pressure difference between the upstream and downstream sides of a rotor; the figure 4 resumes the representation of a circumferential distortion of a static pressure difference between the upstream and downstream sides of a rotor of the figure 3 , and also includes a schematic view along a circumferential direction of the blades included in a stator located upstream of the rotor according to the prior art; the figure 5 resumes the representation of the figure 4 , where the angle of the stator blades located upstream of the rotor has been modified according to an embodiment of the invention; the figure 6illustrates an operating field of a rotor exposed to a circumferential distortion of a static pressure difference between its upstream and downstream according to an embodiment of the invention.

[0029] The drawings in the figures are not to scale. Generally, similar features are 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. Detailed description of particular embodiments of the invention

[0030] The present invention is described with particular embodiments and references to figures, but the invention is not limited by them. The drawings or figures described are schematic only and are not limiting. Furthermore, the functions described can be performed by structures other than those described in this document.

[0031] The use of the verb "comprendre" (to understand), its variants, and its conjugations in this document does not in any way preclude the presence of elements other than those mentioned. The use of the indefinite article "un" (a / an) or the definite article "le" (the / it) to introduce an element does not preclude the presence of multiple such elements.

[0032] Furthermore, within the context of this document, the terms "first", "second", "third", etc. are used only to differentiate different elements and do not imply any order between these elements.

[0033] Reference numerals are shown in some of the figures as abstract geometric coordinates primarily to quantify and / or visualize properties of embodiments of the invention. Reference numeral Z, for example, generally designates a "drive shaft" of an aircraft turbomachine. This shaft is directed from "upstream" to "downstream." The compressor and turbine stages of the aircraft turbomachine are stacked essentially along this drive shaft. The terms "inlet" and "outlet" of a compressor refer to the upstream and downstream ends of the compressor, respectively.Within the framework of this document, reference is made to the directions "axial", "circumferential", and "radial", corresponding preferentially and respectively to a direction essentially parallel to the drive axis, a direction essentially circular around the drive axis and contained in a plane essentially perpendicular to the drive axis, and a direction essentially perpendicular to the drive axis. A reference frame on the... figure 1This illustrates these axial, circumferential, and radial directions, respectively denoted Z, C, and R, each direction being assigned a sense. The direction and sense of the driving axis and the Z vector correspond. The terms "axially" and "radially" are derived from the terms "axial" and "radial," respectively, with a similar preferred meaning. The terms "circumferential" and "radial" preferably refer to a polar coordinate system known to a person skilled in the art in each plane essentially perpendicular to the driving axis.

[0034] There figure 1This illustrates an aircraft turbomachine 100 comprising an assembly 1 according to the invention, said assembly 1 comprising a stator 10 and a rotor 20 located downstream of the stator 10. The stator 10 comprises a row of blades 121. The rotor 20 comprises a row of blades 122. In the context of this document, a stator may also be referred to as a "stator assembly". The aircraft turbomachine 100 is, for example, an axial twin-flow turbomachine comprising successively along the engine axis Z, a fan 110, a first compressor 120 (or low-pressure compressor), a second compressor 130 (or high-pressure compressor), 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.In operation, the mechanical power of the low pressure turbines 150 and high pressure turbines 140 is transmitted respectively via the low pressure shafts 101 and high pressure shafts 102 to the low pressure compressors 120 and high pressure compressors 130, as well as to the fan 110 via a reduction gear 111 interposed at the shaft 101. The fan 110 generates a primary flow 106 through the aircraft turbomachine 100 in a primary aerodynamic channel and a secondary flow 107 externally around the compressors 120, 130 and the turbines 140, 150.

[0035] Although not systematically referenced on the figure 1Each compressor 120, 130 and each turbine 140, 150 comprises at least one stage, each stage including a fixed blade and a movable blade capable of being rotated about the drive shaft. For the low-pressure compressor 120, the fixed blades are designated 121, 123, 125, and the movable blades are designated 122, 124. The fixed blades of compressors 120, 130 are generally called "straighteners." The movable and fixed blades of the stages of such a compressor 120, 130 alternate along the drive shaft. Parameters such as the dimensions and surface geometry of the blades are determined so that the operating conditions of each stage are adapted to those of the stages upstream and / or downstream along the drive shaft.In particular, the moving blades provide energy by increasing the relative velocity of the airflow passing through the compressor 120, 130, while the straighteners redirect the flow parallel to the drive axis, increasing the pressure and decreasing the absolute flow velocity. Each fixed blade has a specific aerodynamic profile and a pitch angle relative to the drive axis to dictate the flow direction. The low-pressure 120 and high-pressure 130 compressors thus synergistically draw in and compress air to bring it to the appropriate velocity, pressure, and temperature for entry into the combustion chamber 160. These concepts are familiar to anyone skilled in the art.

[0036] One or more compressor stages 120, 130 are equipped with a variable-pitch system for their blades to optimize airflow between stages of these compressors 120, 130 according to the operating conditions of the aircraft turbomachine 100. A variable-pitch blade can pivot about an essentially radial axis of rotation. For the purposes of this document, the pitch angle of a variable-pitch blade is defined as the angle between a chord of the blade and the engine axis, a chord of a blade being a straight line segment connecting the leading and trailing edges of the blade and lying in a plane perpendicular to the radial axis of rotation of the blade. It should be noted that in the case of a stator blade, the blade pitch angle can be defined equivalently to the above definition as the angle between a chord of the blade and the axis of revolution of the stator.Indeed, the axis of revolution of a stator within a turbomachine is essentially parallel to, or even coincides with, the turbomachine's drive axis. A variable-pitch stator blade can be oriented in a direction more parallel to the stator's axis of revolution; in this case, for the purposes of this document, we will refer to it as "more axial pitching," which corresponds to a reduced (or in other words, smaller) pitch angle. Furthermore, by pivoting around its radial axis of rotation, a variable-pitch stator blade can assume an orientation more perpendicular to the stator's axis of revolution; in this case, for the purposes of this document, we will refer to it as "more circumferential pitching," which corresponds to a higher pitch angle.As is well known to anyone skilled in the art, each of the aforementioned variable-pitch systems for the stator blades may include a control ring mechanically coupled to the blades by respective levers (or connecting rods) and circumferentially and externally adjusted around a housing of the associated compressor 120, 130, and a displacement unit, typically at least one cylinder, for moving this ring and thereby synchronously altering the pitch angle of each of the stator blades. In particular, each of said levers may have, on the one hand, a first end rigidly mounted on a pivot integral with a blade and, on the other hand, a second end connected to the control ring. According to the... figure 1The low-pressure compressor 120 includes at least one rectifier 121 equipped with such a variable-timing system, referenced by 126. This rectifier 121 is preferably the rectifier 121 located furthest upstream within the low-pressure compressor 120. The representation of the figure 1 is not limiting insofar as one or more of the other rectifiers 123, 125 could also be equipped with such a variable timing system.

[0037] The first compressor 120 is equipped with at least one row of rotor blades 122, 124, preceded directly upstream by a row of stator blades 121, 123, 125, each row of stator blades 121, 123, 125 forming a stator assembly. The invention can be applied to any of the stator assemblies of a compressor 120, 130, and in particular to the upstream stator assembly of the first compressor 120, i.e., the low-pressure compressor 120. In other words, the stator 10 included in assembly 1 according to the invention is preferably the upstream stator 10 within the low-pressure compressor 120.

[0038] The aircraft turbomachine 100 includes an inlet support housing 181 that extends around the inlet of the primary flow, through which the primary flow 106 passes, downstream of the fan 110. The inlet support housing 181 may be provided with one or more structural arms 183 extending radially through the primary flow. The aircraft turbomachine 100 also includes an intermediate support housing 182 that extends circumferentially between the first 120 and second 130 compressors. The intermediate support housing 182 includes an annular sleeve delimiting the primary aerodynamic flow between the first 120 and second 130 compressors. The intermediate support housing 182 may be provided with one or more structural arms 184 extending radially through the primary flow.Such a structural arm 183, 184 can constitute an obstacle in the flow of the airflow, such an obstacle being able to cause a circumferential distortion of a static pressure difference between the upstream and downstream of a rotor located upstream or downstream of the obstacle.

[0039] The invention can be applied to turbomachines other than the aircraft turbomachine 100 shown in the figure 1 , and in particular to aircraft turbomachinery having more than two flow paths and / or unfaired rotors.

[0040] There figure 2 Figure 99 illustrates an operating field of a rotor exposed to circumferential distortion due to a static pressure difference between its upstream and downstream sides. This rotor can, for example, be part of a low-pressure compressor 120. figure 2is equipped with axis markers 97 and 98 corresponding respectively to the airflow through the rotor and the static pressure downstream of the rotor. A nominal average operating point 80 of the rotor is located on a characteristic rotor compression curve 90 for a given rotor operating speed. The pumping line 71, also referred to as the "rotor aerodynamic stability line" in this document, defines and / or estimates a stability limit within the rotor's operating range 99. To ensure proper rotor operation, a margin, referred to as the "stability margin" in this document, must be maintained between the pumping line 71 and the rotor operating point, particularly considering uncertainties in the measurement of the rotor's operating parameters.The rotor operating point can shift along the characteristic curve 90 in response to local fluctuations in static pressure downstream of the rotor, along a circumferential direction. Thus, when the local static pressure downstream of the rotor increases, the rotor operating point moves up along the characteristic curve 90 and closer to the pumping line 71. When the local static pressure downstream of the rotor reaches its maximum, the rotor operating point corresponds to the operating point 81 along the rotor compression characteristic curve 90. When the local static pressure downstream of the rotor decreases, the rotor operating point moves down along the characteristic curve 90 and away from the pumping line 71.Faced with a minimum local value of static pressure downstream of the rotor, the local operating point of the rotor corresponds to the operating point 82 along the characteristic curve 90 of rotor compression.

[0041] There figure 3 illustrates a circumferential distortion of a static pressure difference between the upstream and downstream sides of a rotor. figure 3is equipped with a coordinate system with axes 51 and 52 corresponding respectively to a position along a circumferential direction and to a static pressure difference between the upstream and downstream sides of the rotor. An average static pressure difference between the upstream and downstream sides of the rotor is represented by the dashed line 53. A distortion of the static pressure difference along a circumferential direction is represented by the solid line 54. This is a simplified representation of a static pressure difference distortion, in the case of a single obstacle, where the distortion can take the form of a sinusoid as shown in the diagram. figure 3 In the case of a plurality of obstacles, the distortion of the static pressure difference can be a combination of several sinusoids of different amplitudes and phases.

[0042] There figure 4resumes the representation of a circumferential distortion 54 of a static pressure difference between the upstream and downstream sides of a rotor of the figure 3 , and also includes a schematic view along a circumferential direction of the blades 61 contained within a stator 60 located upstream of the rotor. On the figure 4 , the stator 60 blades 61 are in an axisymmetric configuration along a circumferential direction, i.e. the pitch angle of the stator 60 blades 61 is identical for all the blades 61.

[0043] There figure 5 resumes the representation of the figure 4where the pitch angle 65, 66, 67 of the blades 61' of a stator 60' located upstream of a rotor exposed to a circumferential distortion 54 of a static pressure difference between the upstream and downstream sides of the rotor has been modified according to an embodiment of the invention. Thus, according to the invention, a local flow rate of an air stream passing through a rotor exposed to a circumferential distortion 54 of a static pressure difference between its upstream and downstream sides is adapted to a local value 55, 56, 57 of the static pressure difference by varying a pitch angle 65, 66, 67 of a blade 62, 63, 64 of a stator 60' located upstream of the rotor as illustrated in the following cases: In the first case, the blade 62 has a more axial pitch, i.e. the blade 62 has a reduced pitch angle 65 (or in other words, a smaller pitch angle), facing a higher local value 55 of the static pressure difference, resulting in a higher local flow rate at the blade 62; in the second case, the blade 63 has a nominal pitch, i.e. the blade 63 has a nominal pitch angle 66, facing an average local value 56 of the static pressure difference, resulting in a nominal local flow rate at the blade 63; in the third case, the blade 64 has a more circumferential pitch, i.e. the blade 64 has a higher pitch angle 67, facing a reduced local value 57 (or in other words, a smaller static pressure difference), resulting in a reduced local flow rate (or in other words, a smaller flow rate) at the blade 64. In other words, the pitch angle 65, 66, 67 of the blades 62, 63, 64 of a stator 60' according to the invention depends on the local value 55, 56, 57 at the blade of the static pressure difference between the upstream and downstream of a rotor located downstream of the stator 60', a higher value of the static pressure difference corresponding to more axial pitching, i.e. a reduced (or in other words lower) pitch angle, and a lower value of the static pressure difference corresponding to more circumferential pitching, i.e. a higher pitch angle.

[0044] There figure 5This also illustrates how the pitch angles 65, 66, 67 of the stator blades 62, 63, 64 according to the invention are defined. The vertical dashed lines 41, 43, and 45 are parallel to the axis of revolution of the stator 60' and therefore also parallel to the drive axis. The oblique dashed lines 42, 44, and 46 are each a chord of the blade 62, 63, and 64, respectively. The pitch angle 65 of the blade 62 is the angle between the axis 41 and the chord 42. The pitch angle 66 of the blade 63 is the angle between the axis 43 and the chord 44. The pitch angle 67 of the blade 64 is the angle between the axis 45 and the chord 46.

[0045] There figure 6 partly reiterates the representation of the figure 2and illustrates an operating field 99' of a rotor exposed to circumferential distortion of a static pressure difference between its upstream and downstream sides according to an embodiment of the invention. This rotor can, for example, be included in the low-pressure compressor 120. The figure 6 illustrates, for a given rotor operating regime, the evolution of the characteristic curve 90 of rotor compression and of the operating point 80 of the rotor in the operating field 99' for the different cases illustrated in the figure 5 and detailed below: In the first case, corresponding to a more axial pitch of a blade facing a higher local value of the static pressure difference between the upstream and downstream of the rotor, resulting in a higher local flow rate at the blade, the characteristic compression curve of the rotor shifts towards a higher flow rate and a higher downstream static pressure, and then corresponds to curve 91, the local operating point of the rotor shifting in the same way and then corresponding to point 83 located on curve 91; in the second case, corresponding to a nominal pitch of a blade facing an average local value of the static pressure difference between the upstream and downstream of the rotor, resulting in a nominal local flow rate at the blade, the characteristic compression curve of the rotor corresponds to curve 90, the local operating point of the rotor then corresponding to point 80 located on curve 90;in the third case, corresponding to a more circumferential setting of a blade facing a reduced (or in other words lower) local value of the static pressure difference between the upstream and downstream of the rotor, resulting in a reduced (or in other words lower) local flow at the blade, the characteristic compression curve of the rotor moves towards a reduced flow and a reduced downstream pressure, and then corresponds to curve 92, the local operating point of the rotor moving in the same way and then corresponding to point 84 located on curve 92. ; Thus, thanks to the invention, the operating point 80 of the rotor evolves along a line 72 more parallel to the line 71 of aerodynamic stability of the rotor, thus limiting the impact on the margin in stability of the rotor of the circumferential distortion of the static pressure difference between the upstream and downstream of the rotor, which makes it possible to design such a rotor with better efficiency.

[0046] In summary, the present invention relates to an assembly for a compressor of an aircraft turbomachine and comprising: a stator comprising a row of blades, each of said blades being arranged along a pitch angle; and a rotor comprising a row of blades and located downstream of the stator along an axial direction; characterized in that the pitch angle of the stator blades is inhomogeneous along a circumferential direction and is a function of a circumferential distortion of a static pressure difference between upstream and downstream of the rotor.

[0047] 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

1. Assembly (1) for a compressor (120, 130) of an aircraft turbomachine (100) and comprising: a stator (10) comprising a row of blades (121), each of said blades being arranged at a pitch angle (65, 66, 67); and a rotor (20) comprising a row of blades (122) and located downstream of the stator (10) along an axial direction; characterized in that the pitch angle of the stator blades (10) is inhomogeneous along a circumferential direction and is a function of a circumferential distortion of a static pressure difference between the upstream and downstream sides of the rotor (20), in that a stator blade (10) has a more axial pitch for a local value of the static pressure difference higher than an average value of the static pressure difference, in that a stator blade (10) has a more circumferential pitch for a local value of the static pressure difference lower than an average value of the static pressure difference, in thatThe inhomogeneous pitching angle of the stator blades (10) is fixed during stator mounting (10) to account for the circumferential distortion of the static pressure difference, and in that the stator (10) is intended to be the most upstream stator (10) within the compressor (120, 130) along the axial direction.

2. Assembly (1) according to the preceding claim and further comprising an arm (183, 184), said circumferential distortion of a static pressure deviation being caused at least in part by the presence of said arm (183, 184) in the assembly (1).

3. Assembly (1) according to the preceding claim, wherein said arm (183, 184) is located upstream of the stator (10), preferably directly upstream of the stator (10), or downstream of the rotor (20), preferably directly downstream of the rotor (20).

4. Assembly (1) according to any one of the preceding claims and further comprising a variable blade pitching system (126) capable of uniformly changing the pitching angle of the stator blades (10).

5. Assembly (1) according to the preceding claim, wherein the variable blade pitch system (126) comprises a control ring essentially coaxial with the stator (10), said control ring being mechanically coupled to each of the stator blades (10) via respective levers.

6. Compressor (120, 130) for an aircraft turbomachine (100) and comprising an assembly (1) according to any one of the preceding claims.

7. Compressor (120) according to the preceding claim, characterized in that said compressor (120) is a low-pressure compressor (120), and in that the stator (10) included in the assembly (1) is the stator (10) located furthest upstream within said low pressure compressor (120).

8. Aircraft turbomachine (100) comprising a compressor (120, 130) according to any one of the two preceding claims.

9. Aircraft comprising a turbomachine (100) according to the preceding claim.

10. Method for setting the blades of a stator (10) of an assembly (1) for a compressor (120, 130) of an aircraft turbomachine (100) and comprising the following steps: a. providing an assembly (1) according to any one of claims 1 to 5; b. determining a circumferential distortion of a static pressure difference between the upstream and downstream ends of the rotor (20) of the assembly (1); c.imposing on the stator blades (10) of the assembly (1) a pitch angle: - inhomogeneous along a circumferential direction, and - a function of the circumferential distortion of a static pressure difference determined in the previous step, a method in which a stator blade (10) has a more axial pitch for a local value of the static pressure difference higher than an average value of the static pressure difference, in which a stator blade (10) has a more circumferential pitch for a local value of the static pressure difference lower than an average value of the static pressure difference, in which the inhomogeneous pitch angle of the stator blades (10) is fixed at the stator (10) mounting to take into account the circumferential distortion of the static pressure difference, and in which the stator (10) is intended to be the furthest upstream stator (10) within the compressor (120, 130) along an axial direction.

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