Turbomachine stator assembly, compressor and turbomachine

Abstract

The invention relates to a turbomachine stator assembly comprising two rows of successive tandem stator blades (121, 124), the blades extending radially between internal and external supports and comprising an equidistant line (20) between an intrados and an extrados, a chord (26) extending between a leading edge and a trailing edge, a skeleton angle (29) between the tangent (28) to the equidistant line (20) and the driving axis, a camber as the difference between the skeleton angle at 95% of the chord from the leading edge and the skeleton angle at 5% of the chord from the leading edge, the rows comprising an annular row of upstream blades, an annular row of downstream blades, the camber ratio between the downstream (124) and upstream (121) blades being greater than 2 and less than 4 on at least one cross-section along the height between the internal and external supports.

Description

technical field

[0001] The present invention relates to a turbomachine stator assembly, a compressor and a turbomachine. Previous art

[0002] With a view to developing aircraft engines that significantly improve their environmental performance, reducing the size and weight of turbomachinery is crucial. To this end, it is possible to make turbomachine modules more compact, particularly compressors, especially low-pressure compressors, by reducing the number of compression stages. However, achieving the same compression ratio with fewer stages requires a higher rotor speed.

[0003] It is possible to consider speeds close to the speed of sound. The compressor is then said to be transonic when at least a radially external part of the rotor moves at a speed greater than that of sound.

[0004] Transonic compressors are characterized by a significant load on the rotor and stator blades because these must accelerate and straighten an airflow over a shorter axial distance. This load is even greater when the compressor includes variable-pitch blades, which, depending on the engine speed, impose a large deflection on the flow, and whose straightening therefore induces an even greater load.

[0005] Therefore, there is a need for stator blades that can withstand a large deflection over a wide range of incidence. Description of the invention

[0006] To this end, the invention proposes a turbomachine stator assembly comprising an internal support and an external support, two successive rows of stator blades defining a tandem configuration, the blades extending radially between the internal and external supports and comprising: an intrados and an extrados, an equidistant line between the intrados and extrados, a leading edge and a trailing edge, a chord extending between the leading and trailing edges, a skeleton angle between the tangent to the equidistant line and the driving axis, a camber as the difference between the skeleton angle at 95% of the chord from the leading edge and the skeleton angle at 5% of the chord from the leading edge; the rows comprising an annular row of upstream blades and an annular row of downstream blades, located downstream of the upstream blades the camber ratio between the downstream blade and the upstream blade of a tandem is greater than 2 and less than 4 on at least one section along the height between the inner support and the outer support.

[0007] According to one variant, on at least one cross-section of the upstream blade of a tandem along the height between the inner support and the outer support, the angle between the tangent to the intrados and the chord has a minimum value between 5% of the leading edge and 95% of the chord, preferably the minimum value of the angle being positioned between 30 and 60% of the chord.

[0008] According to one variant, on at least one cross-section of the upstream blade of a tandem along the height between the inner support and the outer support, the difference between the minimum angle between the tangent to the lower surface and the chord and the lower surface angle at 5% of the leading edge is at least three times greater than the difference between the lower surface angle at 95% of the chord and the lower surface angle at 5% of the leading edge.

[0009] According to one variant, on at least one cross-section of the upstream blade of a tandem along the height between the inner support and the outer support, the evolution of the dimensionless skeleton angle β − β 5 % β 95 % − β 5 % with βx%: the skeleton angle at x% of chord, β5%: the skeleton angle at 5% of chord, and β95%: the skeleton angle at 95% of chord has a deviation of less than 10% over the interval from 20% of chord to 40% of chord from the leading edge.

[0010] According to one variant, on at least one cross-section of the upstream blade of a tandem along the height between the inner support and the outer support, the skeleton angle at 5% of its chord is greater than 50°.

[0011] According to one variant, the blades include a thickness between the intrados and extrados, the upstream blade has a thickness equivalent to at least 8% of the chord value reached at 15% of the chord from the leading edge, over at least one section along the height between the inner support and the outer support.

[0012] According to one variant, for an absolute skeleton deviation corresponding to the absolute value of the difference between the skeleton angle at 95% of chord and the skeleton angle at 5% of chord, the absolute skeleton deviation of the downstream blade of a tandem is greater than 35°, on at least one section along the height between the inner support and the outer support.

[0013] According to one variant, the azimuthal gap between the trailing edge of the upstream blade and the leading edge of the downstream blade is between 3mm and 8mm, over at least one section along the height between the inner support and the outer support.

[0014] According to one variant, the axial gap between the leading edge of the downstream blade and the trailing edge of the upstream blade is between 0.5mm and 3mm, over at least one section along the height between the internal support and the external support.

[0015] The invention also relates to an aircraft turbomachine compressor comprising the assembly as described above, the compressor being low pressure.

[0016] According to one variant, the stator assembly belongs to the last stator stage of the compressor.

[0017] The invention also relates to an aircraft turbomachine comprising the assembly as described above or a compressor as described above.

[0018] According to one variant, the turbomachine includes a variable-speed propulsion unit driven by a speed reducer, the whole assembly being downstream of the propulsion unit.

[0019] According to one variant, the propulsion unit is unfaired.

[0020] According to one variant, the assembly is upstream of a swan neck.

[0021] The stator assembly according to the invention may comprise one or more of the characteristics taken individually or according to any possible technical combinations.

[0022] 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.

[0023] The terms "first", "second", "third", etc., are used in this document exclusively to differentiate different elements, without implying any order between these elements.

[0024] All preferred embodiments and all advantages of the stator assembly according to the invention are applicable mutatis mutandis to the present compressor and turbomachine. The various embodiments may be considered individually or in combination. Brief description of the figures

[0025] Other features and advantages of the present invention will become apparent upon reading the detailed description that follows, for understanding of which reference should be made to the accompanying figures which show: there figure 1 , a schematic view of a turbomachine according to the invention; the figure 2 , a schematic view of an example stator assembly comprising a pair of stator blades; the figure 3 , a schematic view of a blade of the stator assembly; the figure 4 , a schematic view of an example of the stator assembly comprising a tandem of stator blades.

[0026] 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. Detailed description of embodiments of the invention

[0027] 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 "axial," "circumferential," and "radial" directions, corresponding respectively to directions parallel to the drive axis X, essentially circular around the drive axis, and perpendicular to the drive axis. A reference point on the figure 1This illustrates, in particular, the direction of the referenced engine axis X and the radial direction Z, with a specific direction. The terms "in" and "inward" naturally correspond to a direction toward the engine axis X along a radial direction, and the terms "outward" and "outward" to the opposite direction. The terms "in / inlet" (and respectively "in / outlet"), when referring to the position of an element in an aircraft turbomachine or compressor, preferentially refer to the first (respectively, last) such element essentially upstream (respectively, essentially downstream) of the aircraft turbomachine or compressor.

[0028] There figure 1Figure 100 illustrates a cross-section of an aircraft turbomachine into which the assembly 10 according to the invention is to be integrated. This is an example of an architecture, as the invention applies to other turbomachine architectures. It may be an axial multiflow turbomachine comprising a number of elements. In succession, along the engine axis X, there is a propulsion unit 110, a fan 210, a low-pressure compressor 120, a high-pressure compressor 130, a combustion chamber 160, and then high-pressure and low-pressure turbines. These elements are known to a person skilled in the art. The propulsion unit 110 may be shrouded (by a fairing 112, in which case it is called a fan) or unshrouded (in which case it is called a propeller). The propulsion unit 110 can have variable timing, the timing being schematically represented by reference 109.In operation, the mechanical power of the low pressure and high pressure turbines is transmitted to the low pressure compressors 120 and high pressure compressors 130 respectively, as well as to the propulsion unit 110. A reducer 111 can be interposed, so that the rotational speeds of the propulsion unit 110 and the low pressure compressor 120 are proportional. The rotors of these compressors rotate around the motor axis X (indicating the direction of air intake and the direction of the thrust reaction), allowing them to draw in and compress air to bring it to suitable speeds, pressures, and temperatures, up to the inlet of the combustion chamber 160. From an airflow 200, the propulsion unit 110 generates a flow that splits into an airflow 201 and an airflow 202. The airflow 201 is primarily intended to generate a thrust reaction for the aircraft's flight.The airflow 202 in turn splits into an airflow 203 and an airflow 204, by a separating nozzle 70. The airflow 203 is mainly intended to generate an additional thrust reaction for the flight of the aircraft and the airflow 204 is mainly intended to pass axially through the aircraft turbomachine 100.

[0029] Although not systematically referenced on the figure 1Each compressor and 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 that can be rotated around the drive shaft. Within a stage, the rotor draws in and accelerates an incoming airflow by deflecting it relative to the drive shaft axis X, and the following stator or rectifier straightens the flow along the drive shaft axis X and slows it down by converting some of its velocity into pressure. The fixed blades may have a fixed orientation and / or variable orientation (or pitch) (the blades occupying angular pitch positions). A variable pitch system controls the angular pitch. These movable blades are also known as variable-pitch vanes, or by the acronym "VSV" for "Variable Stator Vane".Their distinctive feature is that the inclination of their blades can vary relative to the X-axis of the motor (corresponding to the blade pitch), and particularly to the axes of the 120 and 130 compressors. The lower surfaces (intrados 16) and upper surfaces (extrados 18) of the blades can be more or less exposed to the airflow. The lower surfaces (intrados 16) extend to varying degrees in relation to the airflow in order to modulate the deflection, and therefore the straightening, imposed on the airflow. The compressors can have one or more stages of adjustable blades.

[0030] The low-pressure compressor 120, particularly of the transonic type, experiences a significant load on its rotor and stator blades because they must accelerate and straighten an airflow over a shorter axial distance. This load is even greater when the compressor includes variable-pitch blades, which, depending on the engine speed, impose a large deflection on the flow, and whose straightening therefore induces an even greater load. This is further accentuated with a variable-pitch propulsion unit 110. Thus, without limiting the scope of this designation, the low-pressure compressor 120 comprises a stator assembly 10 consisting of tandem rows of blades. This consists of a succession of two rows or grids of stator blades, each of which partially contributes to straightening the flow. The two rows are placed successively and directly one after the other.Along the engine axis, from upstream to downstream, the blade rows are without any other blade rows and / or arms interposed axially between them. More specifically, the arrangement of two blade rows is such that the leading edge of the blades in both rows is without any other blade rows and / or arms interposed between them, particularly axially. This does not preclude the blades of successive rows from partially overlapping. The same applies to the blade arrangement within the tandem. Within the tandem, the two blades are close together. The rows are annular, around the engine's X-axis. A particularly critical point for flow straightening is the final compression stage of the compressor, which must deliver a substantially axial flow. Thus, to straighten the flow in such a compressor while limiting the aerodynamic load experienced by the final stage blades, assembly 10 comprises the blade rows in tandem, for example, in the final stage.Thus, a first array of blades, designed to accommodate a flow with a significant variation in the angle of attack, is followed by a second array of blades responsible for completing the remaining deflection. This second array maintains its optimal performance thanks to the filtering action of the first array. Each blade in the tandem configuration has specific profile characteristics (as seen in a streamlined cross-section) to withstand high loads and large deflections over a wide range of angles of attack. With this assembly, it is possible to design a tandem stator or rectifier capable of tolerating a wide range of flow angles of attack (for example, 20°, potentially with both positive and negative angles) and achieving significant flow deflection (up to 60°).In addition, the optimization proposed by set 10 allows for better performance in terms of aerodynamic losses, residual gyration at the outlet and pumping margin.

[0031] An example of a stator assembly 10 is described below with reference to figures 2 to 4 The stator assembly 10 thus presents a tandem configuration with two rows or grids of stator blades 121, 124 forming a double grid. figures 3 And 4 are cuts according to AA' of the figure 2 .

[0032] The stator assembly 10 comprises an upstream row of blades 121 and a downstream row of blades 124. The blades extend from a radially internal support 52 to a radially external support 54, as can be seen in the figure 2The internal support 52 and external support 54 can be mounting platforms to a supporting structure. The internal support 52 and external support 54 can also be ferrules describing 360° around the X-axis or angular sectors of ferrules describing a few degrees or a few tens of degrees around the X-axis. The internal support 52 and the external support 54 define an airflow channel 12, particularly for channeling an airflow such as the flow 204. The upstream blades 121 and downstream blades 124 can be supported by a common internal support 52 and a common external support 54. The blades 121 and 124 can be fixed. The blades 121 and 124 can have a fixed pitch relative to the internal support 52 supporting them or a variable pitch relative to the internal support 52 supporting them. The blades 121, 124 extend radially between the two supports. The blades 121, 124 extend along a radial Z direction of the compressor, transversely to the motor axis X.The blades 121, 124 can also extend in another direction, inclined with respect to the X and Z directions.

[0033] Between the internal support 52 and the external support 54, the assembly 10 is distinguished by the profile of the blades 121, 124 according to at least one cross-section along the height H121, H124 of the blades. The blades 121, 124 comprise an intrados 16 and an extrados 18. The blades 121, 124 comprise an equidistant line 20 between the intrados 16 and the extrados 18. The blades 121, 124 also comprise a leading edge 22 upstream in the airflow and a trailing edge 24 downstream in the airflow. The blades 121, 124 also comprise a chord 26 extending between the leading edge 22 and the trailing edge 24. A tangent 28 to the equidistant line 20 is shown.

[0034] 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. 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 tandem blades being capable of righting the airflow, which can be on either side of the tangent 28, i.e., with a positive or negative angle of attack). The arrows 30 show different (maximum) angles of attack. The distance between the two arrows 30 corresponds to the range of angles of attack.

[0035] The evolution, along the chord 26, of the skeleton angle 29 between the tangent 28 to the equidistant line 20 and the motor axis X is called the skeleton angle law. We position ourselves at X% of the chord to position the skeleton angle law.

[0036] The thickness law describes the evolution, along chord 26, of the thickness 50 of the blade between the intrados 16 and the extrados 18. The thickness 50 is measured perpendicular to the skeleton line 20.

[0037] The blade pair 121, 124 is capable of providing a large air deflection for operation over a wide range of air incidence angles on the upstream blade 121, 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.). The blade pair 121, 124 is also capable of straightening the airflow at the trailing edge of the downstream blade 124 so that the straightened airflow is as close as possible to the drive axis X – particularly when the pair is the last stage of the compressor. Thus, the blade pair 121, 124 is capable of providing a large airflow deflection. On the figure 3The straightened airflow is represented schematically by arrow 25 (of the downstream blade 124). The total air deflection by the blade pair is the angle between arrow 25 (of the downstream blade 124) and arrow 30 (of the upstream blade 121). The blade pair 121, 124 is capable of providing a large deflection (greater than 40° to 50°, even up to 60°). Thus, the pair can withstand a significant skeleton angle 29 for the upstream blade and a significant variation in angle of attack (corresponding to the variation of arrow 30 relative to the tangent 28).

[0038] The blades 121, 124 each have a certain camber defined as the difference between the skeleton angle 29 at 95% of the chord 26 from the leading edge 22 and the skeleton angle 29 at 5% of the chord 26 from the leading edge 22. The skeleton angle 29 at X% of the chord 26 from the leading edge 22 means the angle 29 (between the tangent 28 to the equidistant line 20 and the driving axis X) taken at a position at the intersection between a perpendicular to the chord 26 and the equidistant line 20, the perpendicular being at a distance 27 from the leading edge 22 corresponding to X% of the length of the chord 26 from the leading edge 22 (approximately 25% on the figure 3The camber of the blades 121 and 124 allows for the straightening of the airflow. The camber of each blade 121 and 124 is determined to avoid reaching boundary flow conditions, conditions in which the airflow separates from the blades due to the adverse static pressure gradient. In the event of boundary layer separation, the acceleration and / or straightening of the flow is no longer guaranteed, which impairs compressor efficiency. There is thus a link between deflection (air angle in operation) and camber (geometric angle of the blade).

[0039] The camber ratio between the downstream blade 124 and the upstream blade 121 of a tandem is greater than 2 and less than 4 over at least one section along the height between the inner and outer supports (at the same height for both blades). In this way, the profile shape of the forward blade 121 is designed to deviate only slightly compared to the aft blade 124. Thus, the first row or grid of upstream stator blades 121 is designed with moderate deviance and is intended to accommodate the flow upstream of the assembly 10 over a wide range of angles of incidence and to provide the second row or grid of downstream stator blades 124 with a substantially uniform supply at every operating point of the turbomachine 100. The second row or grid of stator blades 124 operates at a near-single angle of incidence and completes the deflection without having to manage variations in angle of incidence.In absolute terms, the camber of each upstream blade 121 is between 10° and 25°, preferably between 12° and 22°. Assembly 10 prevents a single blade from having to withstand an excessively large difference in angle of incidence between two extreme operating points of the turbomachine, which would be associated with a significant deflection. According to the tandem configuration of assembly 10, the upstream blade 121 supports the airflow angle of incidence over a wide range, and the downstream blade 124 ensures the airflow deflection over a wide range. This tandem configuration meets the need for blades that are more tolerant and more resistant to angle of incidence and aerodynamic deflections with respect to flow separation on the blade face. This is particularly important within the technical context described.

[0040] The evolution of the angle between a tangent to the lower surface 16 and the chord 26 is called the intrados angle law. Within a tandem wing, the upstream blade 121 can be such that the angle between the tangent to the lower surface 16 and the chord 26 has a minimum value between 5% of the leading edge 22 and 95% of the chord 26 (i.e., 5% of the trailing edge 24) over at least one section along the height between the inner and outer supports. Thus, the upstream blades 121 have an intrados curvature along the chord that limits air separation, even when the air incidence range is large. Preferably, the minimum value of the angle between the tangent to the lower surface 16 and the chord 26 is positioned between 30% and 60% of the chord. This allows the airflow to be diverted, while allowing the blades to withstand a heavy load over a wide range of angles of attack.

[0041] For the leading edge blade 121 of a tandem wing, the difference between the minimum angle between the tangent to the lower surface and the chord 26 and the lower surface angle (between the tangent to the lower surface and the chord 26) at 5% of the leading edge can be at least three times greater than the difference between the lower surface angle at 95% of the chord and the lower surface angle at 5% of the leading edge, over at least one cross-section along the height between the inner and outer supports. Thus, the leading edges of the blades 121 have a lower surface close to the leading edge 22, which allows, in particular, for proper airflow guidance along the lower surface to limit air separation, even when the airflow angle of attack is wide.

[0042] Over at least one section along the height between the internal support and the external support, the evolution of the dimensionless skeleton angle of the upstream blade 121 defined by the formula βx % − β 5 % β 95 % − β 5 % (with βx%: the skeletal angle 29 at x% of the chord, x being within the interval [20%, 40%], β5%: the skeletal angle 29 at 5% of the chord, and β95%: the skeletal angle 29 at 95% of the chord) may exhibit a deviation of less than 10% over the interval from 20% of the chord to 40% of the chord (from the leading edge 22). For example, β 40 % − β 5 % β 95 % − β 5 % − β 20 % − β 5 % β 95 % − β 5 % is less than 10% (but also for all values ​​of x in the above range). This allows for good stability of the upstream blade's angle of attack while providing some deflection, the downstream blade providing the necessary deflection expected from the tandem. Within a tandem, the skeleton angle 29 of the upstream blade at 5% of its chord is greater than 50°. This ensures stability with respect to angles of attack that can vary significantly, and even be negative.

[0043] The blades 121 and 124 have a thickness of 50 between the lower surface 16 and the upper surface 18. The leading edge blade 121 of a tandem wing can have an equivalent thickness of between 8% and 25%, preferably 8% to 15%, of the chord length (i.e., at least 8% of the chord length from the leading edge) reached at 15% of the chord length from the leading edge (over at least one section along the height between the inner and outer supports). This allows the leading edge blades 121 to have high tolerance and resistance to high angles of attack and their variations.

[0044] The absolute skeleton deflection is defined as the absolute value of the difference between the skeleton angle 29 at 95% of chord 26 and the skeleton angle 29 at 5% of chord 26. Within a tandem turbine, over at least one section along the height between the inner and outer supports, the absolute skeleton deflection of the downstream blade 124 can exceed 35°. This allows the turbine to withstand a high load on a large deflection by the downstream blade 124 while maintaining optimal operation thanks to the filtering action of the upstream blade 121.

[0045] There figure 4This is a schematic view of an example stator assembly with two rows of stator blades, with only two blades 121 and 124 shown per row (corresponding to a tandem). The axial gap 60 formed along the X-axis between the two rows of blades, upstream 121 and downstream 124, can be likened to a gap along the X-axis between the trailing edge 24 of the upstream blade 121 and the leading edge 22 of the downstream blade 124 of a tandem. In the example shown in the figures, the rows of blades 121 and 124 do not overlap, but they could. With (XBA, YBA) the coordinates of the leading edge 22 of the downstream blade 124 and (XBF, YBF) the coordinates of the trailing edge 24 of the upstream blade 121, we can note the axial deviation 60 as ΔX = XBA - XBF.The azimuthal gap 62 (or circumferential spacing, or spacing in the circumferential direction transverse to the driving axis X) formed along the Y axis between the two rows of upstream blades 121 and downstream blades 124 can be likened to a gap along the Y axis between the trailing edge 24 of the upstream blade 121 and the leading edge 22 of the downstream blade 124 of a tandem turbine. The azimuthal gap 62 can be denoted as ΔY = YBF - YBA.

[0046] Regarding the relative positioning of each blade in the tandem assembly, the azimuthal gap 62 between the trailing edge 24 of the upstream blade 121 and the leading edge 22 of the downstream blade 124 can be between 3 mm and 8 mm, over at least one cross-section along the height between the inner and outer supports. This ensures a circumferential spacing that is sufficiently small to maintain flow continuity while being large enough to limit or prevent any aerodynamic blockage, particularly pumping. The axial gap 60 between the leading edge 22 of the downstream blade 124 and the trailing edge 24 of the upstream blade 121 can be between 0.5 mm and 3 mm, over at least one cross-section along the height between the inner and outer supports. This ensures that there is an axial spacing length that allows for continuous flow and that the downstream blade 124 operates in a quasi-single incidence.

[0047] The stator assembly's ability to withstand deflection and angle of attack is achieved through the thickness and / or angle laws of the upstream and downstream blades of the tandem, and / or through the relative positioning of the blades. Over at least one cross-section along the height between the inner and outer supports, the shape of the upstream blade's profile provides greater robustness and better resistance to angle of attack variations, while the shape of the downstream blade's profile allows for the majority of the deflection to be achieved.

[0048] The invention also relates to an (axial) compressor for a turbomachine, comprising the stator assembly 10, and also to an aircraft turbomachine with the compressor and the assembly 10. This can be any type of aeronautical turbomachine, and in particular aircraft turbomachines such as turbojets and turboprops. It can be a turbomachine comprising the propulsion unit 110 upstream (or inlet) of the turbomachine. The propulsion unit 110 can be of the variable-pitch type or of the variable-pitch type driven by the reduction gear 111. The propulsion unit 110 can be of the ducted (fan) type, or of the unducted type (propeller and also a pair of unducted co-rotating or counter-rotating propellers). This latter type is also called "open-fan" or "open-fan." It could be the 120 low pressure compressor.The compressor can be a transonic variable-geometry compressor, in which the stator vanes have variable pitch. The compressor (and assembly 10) can be located downstream of the propulsion unit 110 (downstream meaning after the propulsion unit, separated from it by other elements in the airflow). Assembly 10 can be any compressor rectifier. Primarily, these are fixed compressor rectifiers with variable pitch. The stator assembly 10 can be part of the last stator stage (or rectifier) ​​at the compressor outlet (called the OGV, or "outlet guide vane"), which must rectify the flow to a purely axial velocity, and / or it can be part of the first fixed rectifier (in the direction of the flow).The assembly 10 can be upstream of a gooseneck 56, and more specifically, directly upstream of the gooseneck 56 (directly upstream meaning that the gooseneck is at the outlet of the blade pair 121, 124, without any other elements in the airflow). The invention is particularly well-suited to propulsion units 110 with variable pitch, due to the significant variations in angle of attack encountered (which are greater than in the case of propulsion units 110 without variable pitch). Within the scope of the invention, for any interval relating to a ratio and including the value 1, this value is excluded.

[0049] It will be obvious to a person skilled in the art that the invention is not limited to the achievements and examples illustrated and / or described above, but that its scope is more broadly defined by the claims introduced below.

Claims

1. Turbomachine stator assembly (10) comprising: - an inner support (52) and an outer support (54), - two successive rows of stator blades defining a tandem configuration, the blades (121, 124) extending radially between the inner support (52) and the outer support (54) and comprising: ∘ an intrados (16) and an extrados (18), ∘ an equidistant line (20) between the intrados and the extrados, ∘ a leading edge (22) and a trailing edge (24), ∘ a chord (26) extending between the leading edge and the trailing edge, ∘ a skeleton angle (29) between the tangent (28) to the equidistant line (20) and the driving axis (X), ∘ a camber as the difference between the skeleton angle (29) and 95% of the chord (26) from the leading edge and the skeleton angle (29) at 5% of the chord (26) from the leading edge the rows comprising • an annular row of upstream blades (121), • an annular row of downstream blades (124),located downstream of the upstream blades, the camber ratio between the downstream blade (124) and the upstream blade (121) of a tandem turbine is greater than 2 and less than 4 on at least one section along the height between the inner support (52) and the outer support (54).

2. Assembly (10) according to the preceding claim, wherein, on at least one cross-section of the upstream blade (121) of a tandem along the height between the inner support (52) and the outer support (54), the angle between the tangent to the intrados (16) and the chord (26) has a minimum value between 5% of the leading edge and 95% of the chord, preferably the minimum value of the angle being positioned between 30 and 60% of the chord.

3. Assembly (10) according to any one of the preceding claims, wherein, on at least one cross-section of the upstream blade (121) of a tandem along the height between the inner support (52) and the outer support (54), the difference between the minimum angle between the tangent to the lower surface (16) and the chord (26) and the lower surface angle at 5% of the leading edge is at least three times greater than the difference between the lower surface angle at 95% of the chord and the lower surface angle at 5% of the leading edge.

4. Assembly (10) according to any one of the preceding claims, wherein, on at least one cross-section of the upstream blade (121) of a tandem along the height between the inner support (52) and the outer support (54), the evolution of the dimensionless skeleton angle β − β 5 % β 95 % − β 5 % with βx%: the skeleton angle (29) at x% of chord, β5%: the skeleton angle (29) at 5% of chord, and β95%: the skeleton angle (29) at 95% of chord has a deviation of less than 10% over the interval from 20% of chord to 40% of chord from the leading edge (22).

5. Assembly (10) according to any one of the preceding claims, wherein, on at least one cross-section of the upstream blade (121) of a tandem along the height between the inner support (52) and the outer support (54), the skeleton angle (29) at 5% of its chord is greater than 50°.

6. Assembly (10) according to any one of the preceding claims, wherein, the blades comprise a thickness between the intrados (16) and the extrados (18), the upstream blade (121) has a thickness equivalent to at least 8% of the chord value reached at 15% of the chord from the leading edge (22), over at least one section along the height between the inner support (52) and the outer support (54).

7. Assembly (10) according to any one of the preceding claims, wherein, for an absolute skeleton deflection corresponding to the absolute value of the difference between the skeleton angle (29) at 95% of chord and the skeleton angle (29) at 5% of chord, the absolute skeleton deflection of the downstream blade (124) of a tandem is greater than 35°, over at least one section along the height between the inner support (52) and the outer support (54).

8. Assembly (10) according to any one of the preceding claims, wherein the azimuthal gap (62) between the trailing edge (24) of the upstream blade (121) and the leading edge (22) of the downstream blade (124) is between 3mm and 8mm, on at least one cross-section along the height between the inner support (52) and the outer support (54).

9. Assembly (10) according to any one of the preceding claims, wherein the axial gap (60) between the leading edge (22) of the downstream blade (124) and the trailing edge (24) of the upstream blade (121) is between 0.5mm and 3mm, on at least one section along the height between the inner support (52) and the outer support (54).

10. Aircraft turbomachine compressor (120) comprising the assembly (10) according to any one of the preceding claims, the compressor being low pressure.

11. The compressor according to the preceding claim, in which the stator assembly (10) belongs to the last stator stage of the compressor.

12. Aircraft turbomachine (100) comprising the assembly (10) according to any one of the preceding claims or a compressor according to any one of the preceding claims.

13. Turbomachine according to any one of the preceding claims, comprising a propulsion unit (110) with variable pitch and driven by a speed reducer, the assembly (10) being downstream of the propulsion unit (110).

14. Turbomachine according to the preceding claim, in which the propulsion unit (110) is unfaired.

15. Turbomachine according to any one of the preceding claims, wherein the assembly (10) is upstream of a gooseneck (56).

Images