Turbomachine stator assembly having tandem vane rows

Abstract

The invention relates to a stator assembly (32) of a turbomachine, characterized in that it comprises: internal and external supports; two successive rows of stator blades (33, 35) defining a tandem configuration, comprising an annular row of upstream blades (33) extending substantially radially between the internal and external supports, and an annular row of downstream blades (35), located downstream of the upstream blades (33), extending substantially radially between the internal and external supports.

Description

TECHNICAL FIELD OF THE INVENTION

[0001] The present invention relates to the general field of turbomachines, and more specifically to the field of turbomachine rectifiers or stators, in particular turbomachine compressor rectifiers or stators.

[0002] The invention applies to any type of aeronautical turbomachine, and in particular to aircraft turbomachines such as turbojets and turboprops. The invention can be applied to aircraft turbomachines comprising at least one unducted propeller, and also a pair of unducted co-rotating or contra-rotating propellers, this type of turbomachine also being called “with unducted fan(s)”, or also having the English names “open rotor” or “propfan”.

[0003] The invention thus provides a turbomachine stator assembly comprising successive rows of stator blades in a tandem configuration, a turbomachine compressor comprising such a stator assembly, as well as a turbomachine comprising such a stator assembly or such a compressor. STATE OF THE ART

[0004] Climate change is a major concern for many legislative and regulatory bodies around the world. Various states have, are, or will adopt various carbon emission restrictions. In particular, an ambitious standard applies to both new aircraft types and those already in operation, requiring the implementation of technological solutions to comply with current regulations. For several years now, civil aviation has been mobilizing to contribute to the fight against climate change.

[0005] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft.

[0006] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes and minimizing greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0007] This sustained research and development work focuses on new generations of aircraft engines, the weight reduction of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to ensure propulsion, and, as an essential complement to technological progress, aeronautical biofuels.

[0008] Thus, in order to reduce the size and weight of a turbomachine, it is possible to make the turbomachine modules more compact, and in particular the compressors, especially low pressure ones, by reducing the number of compression stages. However, in order to obtain an equivalent compression ratio with fewer stages, this requires driving the rotor faster in rotation.

[0009] It is possible to consider speeds close to the speed of sound. The compressor is then said to be transonic when at least one radially external part of the rotor moves at a speed greater than that of sound.

[0010] Transonic compressors are characterized by a significant load on the rotor and stator blades because they must accelerate and straighten an airflow over a shorter axial distance. This load is even higher when the compressor includes variable-orientation blades, which, depending on the engine speed, impose a large deflection on the flow and whose straightening therefore induces an even greater load. In this context, the camber of the blades must be significant, but too great a camber presents the risk of reaching boundary flow conditions, conditions in which the airflow separates from the blades under the effect of the adverse static pressure gradient. In the event of separation of the boundary layer, acceleration and / or straightening of the flow is no longer guaranteed, which affects the compressor's efficiency.

[0011] Also, too large a difference in angle of incidence, also called "swing" in English, between two extreme operating points of the turbomachine associated with a large deviation to be achieved can prove problematic at the compressor level, and even more so if, in addition to the angle of incidence to be tolerated, there is a slowdown that is too difficult to manage. In particular, a point with a strong negative incidence coupled with a large Mach number at the inlet can generate sonic blockages just as a point with a strong positive incidence coupled with a large slowdown to manage can present a high risk of separation. The operability of the compressor as well as the rectification function for the correct supply of the downstream parts can then be endangered.

[0012] A particularly critical point in this regard is the last compression stage of the compressor, which must imperatively restore a substantially axial flow.

[0013] To straighten a flow in such a compressor by limiting the aerodynamic load experienced by the blades, it is possible to provide a so-called "tandem" straightener. This is a succession of two rows or grids of stator blades, each of which partially participates in straightening the flow. Such examples are described in particular in patent applications BE 1 030 421 A1, EP 2 913 480 A1, EP 2 409 002 A1, US 2020 / 0240283 A1 and DE 10 2018 108 940 A1.

[0014] Thus, a first grid of blades, whose purpose is to accommodate a flow with a strong variation in the angle of incidence, is followed by a second grid of blades which is responsible for completing the remaining deviation with an operation always adapted thanks to the work of filtering the incidence of the first grid.

[0015] However, the implementation of such a tandem configuration requires fine optimization of several geometric parameters which contribute to guaranteeing the operability required of the rectifier and its performance in terms of aerodynamic losses. STATEMENT OF THE INVENTION

[0016] The invention aims to at least partially remedy the needs mentioned above and the drawbacks relating to the achievements of the prior art.

[0017] To this end, the invention is the result of technological research aimed at significantly improving aircraft performance and, in this sense, contributes to reducing the environmental impact of aircraft.

[0018] In particular, the invention aims to propose a turbomachine rectifier or stator configuration making it possible to achieve the necessary flow deflection at all operating points of the turbomachine, and for a wide range of angles of incidence, to axially straighten the upstream flow so as to correctly supply a downstream zone. It thus aims to propose a rectifier or stator design making it possible to improve the compactness of the turbomachine without harming the efficiency, and this over the entire operating range, including transonic, of the turbomachine.

[0019] The invention thus relates, according to one of its aspects, to a turbomachine stator assembly, characterized in that it comprises: an inner support and an outer support, two successive rows of stator blades defining a tandem configuration, comprising: an annular row of upstream blades extending substantially radially between the inner support and the outer support, each upstream blade having a radial height, a chord connecting the leading edge to the trailing edge, an inlet angle and an outlet angle, an annular row of downstream blades, located downstream of the upstream blades, extending substantially radially between the inner support and the outer support, each downstream blade having a radial height, a chord connecting the leading edge to the trailing edge, an inlet angle and an outlet angle, an angular pitch being defined between two upstream blades, or two downstream blades, circumferentially adjacent, and a circumferential spacing length between a circumferentially adjacent upstream blade and a downstream blade being defined between the trailing edge of one of the upstream blade and the downstream blade and the leading edge of the other of the upstream blade and the downstream blade.

[0020] Thanks to the invention, it is possible to design a rectifier or stator principle in tandem configuration capable of tolerating a large range of flow incidence and of achieving a large flow deflection, by means of optimizing the design parameters of the rows of blades constituting the stator assembly.

[0021] Furthermore, the optimization proposed by the invention makes it possible to obtain better performances in terms of aerodynamic losses, residual gyration at the outlet and surge margin. In addition, the capacities of the tandem configuration in terms of resistance to incidence and aerodynamic load (high deviation) are improved.

[0022] The stator assembly according to the invention may further comprise one or more of the following characteristics taken in isolation or in any possible technical combinations.

[0023] The relative azimuthal offset of a downstream blade relative to an upstream blade can be defined as: Da = t / S, where: t is the circumferential spacing length between the upstream blade and the downstream blade, S is the angular pitch defined between two circumferentially adjacent upstream blades or two downstream blades.

[0024] The incidence protection of a downstream blade relative to an upstream blade can be defined as: Pi = β 1.2 - β 2.1 , where: β 1.2 is the inlet angle of the downstream blade, β 2.1 is the outlet angle of the upstream blade.

[0025] The deflection distribution between an upstream blade and a downstream blade can be defined as: Rd = Δβ 35 Δβ 33 , Or : Δβ 33 is the deflection of the upstream blade, equal to the difference between the exit angle and the entry angle of the upstream blade, Δβ 35 is the deflection of the downstream blade, equal to the difference between the exit angle and the entry angle of the downstream blade.

[0026] The chord ratio between an upstream blade and a downstream blade can be defined as: Rc = C 35 C 33 , Or : C 33 is the chord of the upstream blade, C 35 is the chord of the downstream blade.

[0027] In particular, between 10% and 90% of the radial height of an upstream blade and the radial height of a downstream blade, the azimuthal offset can be chosen such that 0 < Da < 0.50, in particular 0.15 < Da < 0.35. Such values ​​for the azimuthal offset make it possible in particular to ensure that there is a circumferential spacing length that is sufficiently small to obtain flow continuity while being sufficiently high to limit or avoid any aerodynamic blockage, in particular of the pumping type.

[0028] Between 10% and 90% of the radial height of an upstream blade and the radial height of a downstream blade, the incidence protection can be chosen such that 0 < Pi < 12°, in particular 3° < Pi < 9°. Such values ​​for the incidence protection make it possible in particular to obtain an inlet angle of the downstream blade which is close to the outlet angle of the upstream blade, while however being higher, so as to be able to recover the outlet flow and also recover part of the flow which would be incorrectly deflected by the upstream blade.

[0029] Between 10% and 90% of the radial height of an upstream blade and the radial height of a downstream blade, the deflection distribution can be chosen such that 1.1 < Rd < 3.9, especially 1.7 < Rd < 3.3. In order to obtain the desired aerodynamic deflection while maintaining the upstream incidence, the value of the deflection distribution must be high enough to be able to deflect more on the downstream blade while remaining low enough to maintain the aerodynamic constraints of maximum deflection of a blade. The upstream blade can allow the incidence and possible deflection, while the downstream blade can allow the necessary deflection.

[0030] Between 10% and 90% of the radial height of an upstream blade and the radial height of a downstream blade, the chord ratio can be chosen such that 0.5 < Rc < 1.5, in particular 0.7 < Rc < 1.2. Such values ​​for the chord ratio allow in particular the stator in tandem configuration to guarantee better performance, in particular in terms of resistance to incidence and aerodynamic load.

[0031] The number of upstream blades and the number of downstream blades can be the same. The internal support can be an internal shroud, and the external support can be an external shroud or a casing.

[0032] The chord of the upstream blades can be different from the chord of the downstream blades. In other words, the chord ratio can be chosen such that Rc = 1.

[0033] Furthermore, the invention also relates, according to another of its aspects, to a turbomachine compressor, in particular a high pressure compressor or a low pressure compressor, characterized in that it comprises a stator assembly as defined previously.

[0034] Preferably, the compressor is a low pressure compressor.

[0035] Preferably again, the stator assembly belongs to the last compression stage of the compressor.

[0036] Furthermore, the invention also relates, according to another of its aspects, to a turbomachine, characterized in that it comprises a stator assembly as defined previously or a compressor as defined previously.

[0037] The turbomachine may comprise a moving wheel and a separation nozzle, arranged downstream of the moving wheel and separating an annular air flow into a primary flow traveling through an internal vein and a secondary flow traveling through an external vein, the stator assembly being arranged in the internal vein upstream of a swan neck shape.

[0038] Furthermore, the turbomachine may comprise a high-pressure compressor arranged downstream of the stator assembly, the swan-neck shape being located in particular between the stator assembly and the high-pressure compressor. BRIEF DESCRIPTION OF THE FIGURES

[0039] Other advantages, aims and particular characteristics of the invention will emerge from the following non-limiting description of at least one embodiment of the present invention, with reference to the appended figures, in which: there Figure 1schematically represents, in an axial sectional view, an example of a turbomachine comprising a stator assembly in accordance with the invention; Figure 2 schematically illustrates, in a transverse view relative to the radial direction, an example of a stator assembly in accordance with the invention comprising two rows of stator blades, only one blade being represented per row; Figure 3 is a sectional view along AA' of the Figure 2 illustrating very schematically, perpendicular to the radial direction, the example of a stator assembly comprising two rows of stator blades, only two blades being represented per row; and the Figure 4 schematically represents, according to a partial axial sectional view, an example of installation of a stator assembly according to the invention in a turbomachine similar to that of the Figure 1 .

[0040] Throughout these figures, like references may designate identical or similar elements.

[0041] Furthermore, the different parts represented in the figures are not necessarily on a uniform scale, to make the figures more readable. DETAILED DESCRIPTION OF THE INVENTION

[0042] Throughout the description, given as a non-limiting example of embodiment, it is noted that the terms upstream and downstream are to be considered in relation to a main direction F of normal gas flow (from upstream to downstream) for a turbomachine 1. Furthermore, the axis X of the turbomachine 1 is called the axis of radial symmetry of the turbomachine 1. The axial direction of the turbomachine 1 corresponds to the axis of rotation X of the turbomachine 1. A radial direction of the turbomachine 1 is a direction perpendicular to the axis X of the turbomachine 1.

[0043] Furthermore, unless otherwise specified, the adjectives and adverbs axial, radial, axially and radially are used with reference to the aforementioned axial and radial directions. Furthermore, unless otherwise specified, the terms inner and outer are used with reference to a radial direction such that the inner part of an element is closer to the X axis of the turbomachine 1 than the outer part of the same element.

[0044] Furthermore, the term height refers to a dimension measured along the direction of the largest dimension of the blades, which can be substantially radial. The chord is the straight line segment connecting the leading edge to the trailing edge in a plane perpendicular to a radius. Camber refers to the median curve connecting the leading edge to the trailing edge equidistant from the intrados and the extrados.

[0045] On the Figure 1, an example of a turbomachine 1 according to the invention is shown in an axial sectional view.

[0046] An inner casing 2 guides a primary flow F1 successively passing through a low-pressure compressor 4, a high-pressure compressor 4', a combustion chamber 6 and high- and low-pressure turbines 8 before escaping through a nozzle 10. The energy of the combustion drives the turbines 8 in rotation around the longitudinal axis X of the turbomachine 1. The turbines 8 drive the compressors 4, 4', directly by means of transmission shafts or indirectly by means of a reduction gear 23. The turbines 8 also drive in rotation a rotor 12 with fan blades 14 which set in motion a secondary flow F2.

[0047] A fairing 16 and a nacelle 18 delimit a passage 19 which is traversed by the secondary flow F2. Structural arms 20 take up the forces between the nacelle 18 and the inner casing 2.

[0048] An annular row of stator vanes 22, also called OGV for “outlet guide vanes” in English, is arranged downstream of the rotor 12 to straighten the secondary flow F2.

[0049] The turbomachine 1 has a separation nozzle 48 for separating the annular flow F into two primary flows F1 and secondary F2. The invention can be applied downstream of any type of flow separation nozzle and is not limited to the separation nozzle 48.

[0050] Each compressor 4, 4' is formed from a succession of compression stages. Each stage comprises rotating or rotor blades providing a quantity of movement to the flow and fixed or stator blades straightening the direction of flow of the flow.

[0051] In this example, and in no way limiting, the low pressure compression 4 of the turbomachine 1 comprises a final compression stage, as detailed below with reference to the Figure 4, in which is located a stator assembly 32 according to the invention comprising rows of blades in tandem.

[0052] An example of a stator assembly 32 in accordance with the invention is described below with reference to figures 2 And 3 The stator assembly 32 thus has a tandem configuration with two rows or grids of stator blades forming a bi-grid.

[0053] Specifically, a first row or grid of upstream stator blades 33 is provided, moderately deflecting and intended to accommodate the flow of the primary flow F1 with a wide range of incidence angle and to provide the second row or grid of downstream stator blades 35 with a supply that is always uniform at any operating point of the turbomachine 1. The second row or grid of stator blades 35 operates in quasi-single incidence, and completes the deflection without having to manage the incidence variations.

[0054] THE figures 2And 3 allow to define design parameters helping to define the geometry of the stator assembly 32, which here forms, and in a non-limiting manner, the stator of the last compression stage of the low pressure compressor 4, also comprising a rotor upstream of the stator. The rotor comprises rotor blades which accelerate the flow of the fluid thanks to the energy transmitted by the transmission shaft while the stator transforms the kinetic energy into pressure thanks to the shape of the stator blades 33, 35.

[0055] The stator assembly 32 comprises a row of upstream blades 33 and a row of downstream blades 35, extending from a radially internal support 36 to a radially external support 38, as visible in the Figure 2The internal 36 and external 38 supports may be platforms for attachment to a supporting structure. The internal 36 and external 38 supports may also be ferrules describing 360° around the X axis or angular sectors of ferrules describing a few degrees or a few tens of degrees of angle around the X axis. The upstream 33 and downstream 35 blades are carried by a common internal support and by a common external support.

[0056] The camber of the upstream 33 and downstream 35 stator blades is noted A1 and A2 respectively and is shown in broken lines. The angular pitch between two circumferentially adjacent upstream 33 or downstream 35 blades is noted S as visible on the Figure 3 , and may be identical for both rows of blades 33, 35.

[0057] The geometry of the upstream stator blades 33 can be described in part by their inlet angle β 1,1 and their outlet angle β 2,1 . Similarly, the geometry of the downstream stator blades 35 can be described in part by their inlet angle β 1,2 and their outlet angle β 2,2 .

[0058] The "inlet angle" of a blade is the angle formed in a plane parallel to the X axis of the turbomachine 1 and perpendicular to a radius, between the tangent to the camber line and the X axis, at the leading edge of the blade. The "exit angle" of a blade is the angle formed in a plane parallel to the X axis of the turbomachine 1 and perpendicular to a radius, between the tangent to the camber line and the X axis, at the trailing edge of the blade.

[0059] The deflection of each upstream blade 33, denoted Δβ 33 , is the difference between the exit angle β 2.1 and the entry angle β 1.1 . Similarly, the deflection of each downstream blade 35, denoted Δβ 35 , is the difference between the exit angle β 2.2 and the entry angle β 1.2 . The radial height, along a radial axis perpendicular to the X axis, of each upstream blade 33 and downstream blade 35 is respectively denoted H33 and H35 on the Figure 2 .

[0060] The chord of the upstream 33 and downstream 35 blades, which connects the leading edge to the trailing edge, is respectively denoted C 33 and C 35 . The chord C 33 of the upstream 33 blades may be different from the chord C 35 of the downstream 35 blades. The length A0 formed axially between the two rows of upstream 33 and downstream 35 blades is also shown on the figures 2 And 3. This length A0 can be similar to an axial overlap length when the position of the blades 33, 35 is such that they overlap at least partially, thus providing axial overlap over at least part of their radial height. In the example shown in figures 2 And 3 , the rows of blades 33, 35 do not overlap.

[0061] The spacing between the trailing edge of an upstream blade 33 and the leading edge of an adjacent downstream blade 35 is quantified by a circumferential spacing length between blades of the tandem, denoted t on the Figure 3 , measured perpendicular to the X axis in a similar way to the measurement of the pitch S between two adjacent blades in the same row.

[0062] Design parameters are advantageously provided for the stator assembly 32 according to the invention. These parameters are determined in particular between 10% and 90% of the radial height H33 of an upstream blade 33 and between 10% and 90% of the radial height H35 of a downstream blade 35.

[0063] The relative azimuthal offset Da of a downstream blade 35 with respect to an upstream blade 33 is defined as: Da = t / S. In particular, the azimuthal offset Da is chosen such that 0 < Da < 0.50, in particular 0.15 < Da < 0.35.

[0064] The incidence protection Pi of a downstream blade 35 relative to an upstream blade 33 is defined as: Pi = β 1.2 - β 2.1 . In particular, the incidence protection Pi is chosen such that 0 < Pi < 12°, in particular 3° < Pi < 9°.

[0065] The deflection distribution Rd between an upstream blade 33 and a downstream blade 35 is defined as: Rd = Δβ 35 Δβ 33 . In particular, the deviation distribution Rd is chosen such that 1.1 < Rd < 3.9, notably 1.7 < Rd < 3.3.

[0066] The chord ratio Rc between an upstream blade 33 and a downstream blade 35 is defined as: Rc = C 35 C 33 . In particular, the chord ratio Rc is chosen such that 0.5 < Rc < 1.5, notably 0.7 < Rc < 1.2.

[0067] There Figure 4 illustrates an example of possible installation of the stator assembly 32 according to the invention in a turbomachine 1 such as that shown in the Figure 1 , preferably downstream of a flow separator.

[0068] The annular flow F is split into two flows F1 and F2. The annular flow F flows in an annular vein 42, and the flows F1, F2 flow respectively in a primary annular vein 44 and a secondary annular vein 46. The separation of the flows is carried out by the separation nozzle 48.

[0069] Directly or not upstream of the separation nozzle 48 is located a rotating or rotor assembly in the form of a mobile wheel 50, in particular a fan, the blades 52 of which extend radially upstream of the primary 44 and secondary 46 annular veins.

[0070] The stator assembly 32, consisting of an annular row of stator blades 33 and a row of stator blades 35 forming the tandem, is preferably arranged in the primary annular flow path 44, in the low-pressure compressor 4 also comprising rotor blades 30, and precedes a swan-neck shape 54 which is arranged upstream of the high-pressure compressor 4'. Thus, the stator assembly 32 constitutes the last blades 33, 35 of the low-pressure compressor 4 and makes it possible to axially straighten the primary flow F1 coming from the upstream stages in order to correctly supply the swan-neck 54 located downstream.

[0071] The low pressure compressor 4 may comprise variable stator vanes, or VSV for "Variable Stator Vanes" in English, and the stator assembly 32 may comprise the only stator vanes of the low pressure compressor 4 which are not variable. The low pressure compressor 4 may comprise between 1 and 4 compression stages, each formed of at least one row or annular grid of rotor vanes directly followed by at least one row or grid of stator vanes.

[0072] Of course, the invention is not limited to the exemplary embodiments which have just been described. Various modifications may be made thereto by those skilled in the art.

Claims

1. Stator assembly (32) of turbomachine (1), characterized in that it comprises: - an internal support (36) and an external support (38), - two successive rows of stator blades (33, 35) defining a tandem configuration, comprising: - an annular row of upstream blades (33) extending substantially radially between the internal support (36) and the external support (38), each upstream blade (33) having a radial height (H33), a chord (C 33 ) connecting the leading edge to the trailing edge, an entry angle (β 1,1 ) and an exit angle (β 2,1 ), - an annular row of downstream blades (35), located downstream of the upstream blades (33), extending substantially radially between the internal support (36) and the external support (38), each downstream blade (35) having a radial height (H35), a chord (C 35 ) connecting the leading edge to the trailing edge, an entry angle (β 1,2 ) and an exit angle (β 2,2), an angular pitch (S) being defined between two upstream blades (33), or two downstream blades (35), circumferentially adjacent, and a circumferential spacing length (t) between an upstream blade (33) and a downstream blade (35) circumferentially adjacent being defined between the trailing edge of one of the upstream blade (33) and the downstream blade (35) and the leading edge of the other of the upstream blade (33) and the downstream blade (35), wherein, between 10% and 90% of the radial height (H33) of an upstream blade (33) and the radial height (H35) of a downstream blade (35), the incidence protection (Pi) of a downstream blade (35) relative to an upstream blade (33) is defined as: Pi = β 1,2 − β 2,1 where: β 1,2 is the inlet angle of the downstream blade (35), β 2,1 is the exit angle of the upstream blade (33), and in which the incidence protection (Pi) is chosen such that 0 < Pi < 12°.

2. Stator assembly (32) according to claim 1, characterized in thatthe incidence protection (Pi) is chosen such that 3° < Pi < 9°.

3. Stator assembly (32) according to claim 1 or 2, characterized in that , between 10% and 90% of the radial height (H33) of an upstream blade (33) and the radial height (H35) of a downstream blade (35), the relative azimuthal offset (Da) of a downstream blade (35) with respect to an upstream blade (33) is defined as: Da = t / S where: t is the circumferential spacing length between the upstream blade (33) and the downstream blade (35), S is the angular pitch defined between two circumferentially adjacent upstream blades (33) or two downstream blades (35), and in that the azimuthal shift (Da) is chosen such that 0 < Da < 0.

50.

4. Stator assembly (32) according to claim 3, characterized in that the azimuthal shift (Da) is chosen such that 0.15 < Da < 0.

35.

5. Stator assembly (32) according to one of the preceding claims, characterized in that, between 10% and 90% of the radial height (H33) of an upstream blade (33) and the radial height (H35) of a downstream blade (35), the deflection distribution (Rd) between an upstream blade (33) and a downstream blade (35) is defined as: Rd = Δβ 35 Δβ 33 where: Δβ 33 is the deflection of the upstream blade (33), equal to the difference between the exit angle (β 2,1 ) and the entry angle (β 1,1 ) of the upstream dawn (33), Δβ 35 is the deflection of the downstream blade (35), equal to the difference between the exit angle (β 2,2 ) and the entry angle (β 1,2 ) of the downstream dawn (35), and in that the deviation distribution (Rd) is chosen such that 1.1 < Rd < 3.

9.

6. Stator assembly (32) according to claim 5, characterized in that the deviation distribution (Rd) is chosen such that 1.7 < Rd < 3.

3.

7. Stator assembly (32) according to any one of the preceding claims, characterized in that, between 10% and 90% of the radial height (H33) of an upstream blade (33) and the radial height (H35) of a downstream blade (35), the chord ratio (Rc) between an upstream blade (33) and a downstream blade (35) is defined as: Rc = C 35 C 33 where: C 33 is the chord of the upstream dawn (33), C 35 is the chord of the downstream blade (35), and in that the chord ratio (Rc) is chosen such that 0.5 < Rc < 1.

5.

8. Stator assembly (32) according to claim 7, characterized in that the chord ratio (Rc) is chosen such that 0.7 < Rc < 1.

2.

9. Compressor (4) of turbomachine (1), characterized in that it comprises a stator assembly (32) according to any one of the preceding claims.

10. Compressor (4) according to claim 9, characterized in that the compressor (4) is a low pressure compressor.

11. Compressor (4) according to claim 9 or 10, characterized in thatthe stator assembly (32) belongs to the last compression stage of the compressor (4).

12. Turbomachine (1), characterized in that it comprises a stator assembly (32) according to any one of claims 1 to 8 or a compressor (4) according to any one of claims 9 to 11.

13. Turbomachine (1) according to claim 12, characterized in that it comprises a moving wheel (50) and a separation nozzle (48), arranged downstream of the moving wheel (50) and separating an annular air flow (F) into a primary flow (F1) traveling through an internal vein (44) and a secondary flow (F2) traveling through an external vein (46), the stator assembly (32) being arranged in the internal vein (44) upstream of a swan neck shape (58).

14. Turbomachine (1) according to claim 12 or 13, characterized in thatit comprises a high-pressure compressor (4') arranged downstream of the stator assembly (32), the swan-neck shape (58) being in particular located between the stator assembly (32) and the high-pressure compressor (4').

15. Turbomachine (1) according to claims 13 and 14, characterized in that the swan neck shape (58) is located between the stator assembly (32) and the high pressure compressor (4').

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