Triple-flow aircraft turbine engine

Abstract

A triple-flow aircraft turbine engine (10), comprising: - two coaxial annular walls (12, 14), - rotor blading (30), - an annular separator (24) arranged downstream of the rotor blading (30) and between the two walls (12, 14), and having, upstream, an annular nose (24a), - stationary guide vanes (42) connected to the nose (24a), and - variable-pitch guide vanes (44) downstream of the stationary guide vanes (42).

Description

[0001]DESCRIPTION TITLE: TRIPLE-FLOW AIRCRAFT TURBOMACHINE Technical field of the invention The present invention relates to the general field of aeronautics. It relates more particularly to a triple-flow aircraft turbomachine. Technical background Conventionally, an aircraft turbomachine comprises a gas generator comprising along a longitudinal axis at least one compressor, a combustion chamber, and at least one turbine. An air flow enters the gas generator and is compressed in the compressor(s). This compressed air flow is mixed with fuel and burned in the combustion chamber, then the combustion gases are expanded in the turbine(s). This expansion causes the turbine rotor(s) to rotate, which causes the compressor rotor(s) to rotate. The combustion gases are ejected through a nozzle to provide thrust that can be added to the thrust provided by at least one propeller or fan, whether ducted or not.propulsion shroud of the turbomachine. The gas flows into the turbomachine through annular veins. As can be seen in Figure 1a, the turbomachine 10 thus comprises coaxial annular walls, respectively internal 12 and external 14, extending one around the other and defining between them a main annular flow path 16 for a main gas flow 18. In the case where the main gas flow 18 must be divided into two secondary gas flows, respectively internal 20 and external 22, an annular separator 24 is arranged between the two walls 12, 14 and defines respectively with these walls 12, 14 two secondary annular flow paths, respectively internal 26 and external 28, for the flow of the secondary gas flows 20, 22. This separator 24 comprises at an upstream end an annular nozzle 24a configured to separate the main gas flow 18 in two and form the secondary gas flows 20, 22. A rotor blade 30 can extendradially through the main vein 16, therefore upstream of the separator 24. As illustrated in Figure 1a, structural arms 32 may extend radially through the main vein 16 downstream of the rotor blade 30 and upstream of the separator 24. In the present application, the term arm 32 or structural arm means a stator element which has in section a generally aerodynamic shape such as that shown in Figure 1b, but which does not include a lower or upper surface. An arm 32 is therefore not comparable to a blade or vane which is profiled so as to include a lower and an upper surface. An arm 32 generally has a symmetry with respect to a plane P passing through the axis of the turbomachine. The number of arms 32 is generally less than 10 and may be 4. At least one of the arms 32 may be hollow and tubular in shape in the radial direction to be crossed by services and thus serve for the passage of these services in the engine through theveins. For certain types of turbomachine, such as those with multiflow or variable cycle, it would be useful to have a stator blading 34 directly downstream of the rotor blading 30 and integrated into the flow separation nozzle 24a instead of being positioned between the rotor 30 and the separator 24 (see Figure 2a), so as to reduce the length of the module between the concept illustrated in Figure 1a and that illustrated in Figure 2a. The stator blading 34 would comprise several blades distributed around the axis of the turbomachine. As mentioned above and illustrated in Figure 2b, each of these blades would have in section an aerodynamic profile comprising an intrados 34a and an extrados 34b (Figure 2b), therefore a non-symmetrical profile which is not the case of the arm 32 visible in Figure 1a. The stator blading 34 would extend radially across the main vein 16. In the case where the nozzle 24a would be connected to the vanes of the stator blading 34, these vanes would comprise edgesattack edges 36 located upstream of the nozzle 24a, in the main vein 16, and trailing edges, respectively internal 38a and external 38b, located in the internal 26 and external 28 veins. The stator blading 34 would impose a particular direction on the gas flows 16, 20, 22. However, in the case of a variable cycle turbomachine, it would be useful to provide a variable geometry downstream of the stator blading 34 to be able to adapt to the different operating regimes and variations in the bypass ratio of the turbomachine. However, for reasons of space, the addition of variable pitch blading downstream of the stator blading 34 can be complex. Indeed, this addition would require extending the axial dimension of the turbomachine, which would result in an increase in the mass of the turbomachine and a reduction in its performance. Furthermore, for reasons of noise pollution, it would not be possible to axially bring the stator blade 34 closer together.towards the rotor blading 30. In the present application, a variable cycle turbomachine is understood to mean a turbomachine whose specific thrust can be modified at a given engine speed, by controlling variable geometries of the turbomachine. An example of variable geometry is a variable pitch stator blading. In the present application, blading is understood to mean an annular row of blades. The invention thus proposes to optimize a turbomachine as illustrated in Figure 2a so as to be able to use it in several configurations and in particular in the context of a turbomachine with several flows (at least two) and / or a variable cycle turbomachine. Summary of the invention The present invention proposes a triple flow aircraft turbomachine, comprising a gas generator comprising along a longitudinal axis at least one compressor, a combustion chamber and at least one turbine, the turbomachine further comprising: - two coaxial annular walls,respectively internal and external, extending one around the other and defining between them a main annular flow vein of a main air flow, - a rotor blade extending radially through said main flow vein and forming a shrouded propeller, - an annular separator arranged downstream of the rotor blade and between the two walls, the separator defining respectively with the internal and external walls two secondary annular flow veins, respectively internal and external, of secondary air flows, respectively internal and external, the separator comprising at an upstream end an annular nozzle configured to separate the main air flow in two and form the secondary air flows, - stator elements extending radially on the one hand through said main flow vein and on the other hand through said secondary flow veins, these stator elements being connected to said annular nozzle, and - an unshrouded propeller arranged upstream ofthe outer wall, characterized in that said stator elements comprise: - fixed stator vanes which are distributed around said axis and which each have a leading edge located upstream of said beak, and trailing edges, respectively internal and external, located respectively in the internal and external secondary veins, these fixed stator vanes being connected to said beak, and - variable-pitch stator vanes which are distributed around said axis and which extend radially through at least one of said secondary veins, each of the variable-pitch stator vanes having a leading edge and a trailing edge, and in that: - the leading edges of the variable-pitch stator vanes are located upstream of the internal and / or external trailing edges of the fixed stator vanes, or - the leading edges of the variable-pitch stator vanes are located directly downstream of the internal and / or external trailing edges of the fixed stator vanes,fixed rectifier, and are separated by predetermined axial clearances from these trailing edges. The present invention thus proposes to put both fixed rectifier vanes and variable-pitch rectifier vanes in place of the arms of Figure 1a or the stator vane of Figure 1b. The fixed and variable-pitch rectifier vanes are very close axially to each other or are axially embedded in each other so that they are considered as an assembly forming the stator elements within the meaning of the invention. Indeed, either the variable-pitch rectifier vanes have their leading edges located upstream of the trailing edges of the fixed rectifier vanes, or the variable-pitch rectifier vanes are separated by predetermined axial clearances, preferably as small as possible, from the trailing edges of the fixed rectifier vanes. Minimizing these axial clearances makes it possible to limit or even prevent the passage ofoperating gas between the trailing edges of the fixed stator vanes and the leading edges of the variable-pitch stator vanes. It is thus understood that the gases which flow on the intrados of the fixed stator vanes must then flow on the intrados of the variable-pitch stator vanes, and that the gases which flow on the extrados of the fixed stator vanes must then flow on the extrados of the variable-pitch stator vanes. This configuration is particularly advantageous because it makes it possible to optimize the operation of the turbomachine, by allowing multi-flow or variable-cycle applications, while limiting the impact on the length or axial dimension as well as the mass of the turbomachine. Indeed, reducing the axial clearance between the stator vanes and positioning them at the level of the nozzle makes it possible to limit the impact of these vanes on the axial dimension of the turbomachine. Upstream of the rotor blades locatedin the first vein, any configuration for the turbomachine can be seen there. In the present application, the term "annular" means a shape of revolution around an axis, this shape being able to be continuous or interrupted. In addition, in the present application, the term "variable pitch" means an element of which at least one part has a position which can be adjusted around an axis, which is called a pitch axis. The entirety of this element or only a part of this element can be variable pitch. In the case of a blade for example, it can be in one piece and have an adjustable position around a pitch axis. Alternatively, it could comprise only one part, comprising for example a leading edge or a trailing edge, the position of which would be adjustable around a pitch axis relative to the rest of the blade. In the case of a blade comprising several blades, each of the blades has an adjustable position around a pitch axis whichis clean. For the same blading, there are therefore as many pitch axes as there are variable pitch blades. Each of these axes may have a radial or inclined orientation relative to the longitudinal axis of the turbomachine. The turbomachine may comprise one or more of the following characteristics, taken alone or in combination with each other: -- said clearances are less than 10 mm, and preferably less than or equal to 5 mm; -- said clearances are less than 10% of the chord of one of the fixed or variable pitch blades, and preferably less than or equal to 5% of this chord; -- said fixed stator blades comprise a lower surface and an upper surface, and said variable pitch stator blades comprise a lower surface and an upper surface; - the number of said variable pitch stator blades is greater than or equal to the number of said fixed stator blades; - said variable-pitch rectifier vanes are located in said internal secondary vein; -the trailing edges of said variable-pitch stator vanes are located downstream of the outer trailing edges of the fixed stator vanes; -- the turbomachine further comprises a system for controlling the angular pitch of the variable-pitch stator vanes, this system being mounted in said separator; - the turbomachine further comprises a system for controlling the angular pitch of the variable-pitch stator vanes, this system being mounted radially outside said outer wall; - said variable-pitch stator vanes are located in said outer secondary vein; - first variable-pitch stator vanes are located in said inner secondary vein, and second variable-pitch stator vanes are located in said outer secondary vein; - the turbomachine further comprises a common system for controlling the angular pitch of the first and second variable-pitch stator vanes, or systemsindependent control of the angular setting of the first and second variable-pitch stator vanes respectively; - the turbomachine further comprises structural arms distributed around said axis in said external secondary vein; - the number of structural arms is less than the number of fixed stator vanes; - the structural arms are connected to some of said fixed stator vanes; - the rotor blade is a fan or a compressor rotor blade; - the leading edges of the variable-pitch stator vanes are located at a distance from the internal and / or external trailing edges of the fixed stator vanes, which is greater than 10% of the chord of one of these vanes, and more preferably greater than or equal to 20% of this chord; -- at least some of the fixed stator vanes have profiles or cambers different from the other fixed stator vanes; -- at least some of the arms have profiles different from the other arms.The present invention also relates to an aircraft, in particular a transport aircraft, comprising a turbomachine as described in the foregoing. Brief description of the figures Other characteristics and advantages of the invention will appear during the reading of the detailed description which follows for the understanding of which reference will be made to the appended drawings in which: [Fig.1a] Figure 1a is a very schematic half view in axial section of an aircraft turbomachine, according to the art prior to the invention; [Fig.1b] Figure 1b is a very schematic cross-sectional view of an arm of the turbomachine of Figure 1a; [Fig.2a] Figure 2a is a very schematic half view in axial section of a part of an aircraft turbomachine; [Fig.2b] Figure 2b is a very schematic cross-sectional view of a stator blade of the turbomachine of Figure 2a; [Fig.3a] Figure 3a is a very schematic half-view in axial section of aaircraft turbomachine, according to a first embodiment of the invention; [Fig.3b] Figure 3b is a very schematic cross-sectional view of two fixed stator vanes followed by two variable-pitch stator vanes of the turbomachine of Figure 3a, and illustrate, respectively on the left and right of the figure, two distinct pitch positions of the variable-pitch stator vanes; [Fig.4a] Figure 4a is a very schematic half-view in axial section of an aircraft turbomachine, according to a second embodiment of the invention; [Fig.4b] Figure 4b is a very schematic cross-sectional view of two fixed stator vanes followed by three variable-pitch stator vanes of the turbomachine of Figure 4a, and illustrate, respectively on the left and right of the figure, two distinct pitch positions of the variable-pitch stator vanes; [Fig.5a] Figure 5a is a very schematic half-view in axial sectionof an aircraft turbomachine, according to a third embodiment of the invention; [Fig.5b] Figure 5b is a very schematic cross-sectional view of two fixed rectifier vanes interposed with two variable-pitch rectifier vanes of the turbomachine of Figure 5a, and illustrate, respectively on the left and right of the figure, two distinct pitch positions of the variable-pitch rectifier vanes; [Fig.6] Figure 6 is a very schematic half-view in axial section of an aircraft turbomachine, according to a fourth embodiment of the invention; [Fig.7] Figure 7 is a very schematic half-view in axial section of an aircraft turbomachine, according to a fifth embodiment of the invention; [Fig.8] Figure 8 is a very schematic half-view in axial section of an aircraft turbomachine, according to a sixth embodiment of the invention; [Fig.9] Figure 9 is a very schematic half-view in axial section of a turbomachineof aircraft, according to a seventh embodiment of the invention; [Fig.10] Figure 10 is a very schematic half view in axial section of an aircraft turbomachine, according to an eighth embodiment of the invention; [Fig.11] Figure 11 is a very schematic half view in axial section of an aircraft turbomachine, according to a ninth embodiment of the invention; [Fig.12] Figure 12 is a very schematic half view in axial section of an aircraft turbomachine, according to a tenth embodiment of the invention; [Fig.13] Figure 13 is a very schematic half view in axial section of an aircraft turbomachine, according to an eleventh embodiment of the invention; ; and [Fig.14] Figure 14 is a schematic view of a triple-flow turbomachine within the scope of which the present invention is situated. Detailed Description of the Invention Figures 1a, 1b, 2a and 2b have been described in the above. With reference to Figure 14, the turbomachine 10 is of the type withtriple flow and conventionally comprises a gas generator 2 comprising along a longitudinal axis X at least one compressor, a combustion chamber and at least one turbine. The turbomachine also comprises a ducted propeller or fan denoted H1 and an unducted propeller or fan denoted H2. The propeller H1 is surrounded by a nacelle 4 which extends around the axis X downstream of the propeller H2. The air flow which passes through the propeller H2 is separated by the nacelle 4 into a main flow F2 which enters the nacelle 4 and into another flow F3 which flows around the nacelle 4. The main flow F2 is then divided into two other flows F1, F2 as explained below. In the context of the present invention illustrated in figures 3a and 3b, the turbomachine 10 comprises two coaxial annular walls, respectively internal 12 and external 14, extending around each other and defining between them a main annular flow vein 16 for a main gas flow.18. The main gas flow 18 is divided into two secondary gas flows, respectively internal 20 and external 22, by an annular separator 24 which is arranged between the two walls 12, 14. This separator 24 comprises at an upstream end an annular nozzle 24a configured to separate the main gas flow 18 in two and form the secondary gas flows 20, 22. A rotor blade 30 extends radially through the main vein 16, upstream of the separator 24. In the context of the turbomachine of FIG. 14, this rotor blade 30 forms the shrouded propeller H1. Stator elements 40 are located downstream of the rotor blade 30 and at the separation nozzle 24a. According to the invention, these stator elements 40 comprise fixed rectifier vanes 42 and variable-pitch rectifier vanes 44. The fixed vanes 42 are distributed around the axis and each comprise a leading edge 42a located upstream of the nozzle 24a, and trailing edges, respectively internal 42b and external 42b.42c, located respectively in the internal 26 and external 28 secondary veins. It is thus understood that the fixed blades 42 are connected to the nozzle 24a, as can be seen in the drawing. As can also be seen, the leading edges 42a can be inclined and extend from upstream to downstream towards the outside. This inclination is for example determined according to a compromise between the size of the engine and the optimization of the noise it generates. To minimize noise, it is preferable to increase the height at the top of the blade, which results in a greater inclination of the blade. Figure 3b shows that each of the fixed blades 42 has an aerodynamic profile and comprises a lower surface 46 (of concave curved shape) and an upper surface 48 (of convex curved shape). Furthermore, each of the fixed blades 42 has a certain curvature along its chord. C denotes the zone of greatest curvature of a fixed blade 42. This zone is preferably located upstream of the nozzle 24a. The bladesfixed blades 42 are preferably all identical. Their leading edges 42a are preferably crossed by the same transverse plane. The number of fixed blades 42 is for example between 10 and 200. The variable-pitch blades 44 are distributed around the axis in the internal secondary vein 26 only. The variable-pitch blades 44 each have a leading edge 44a located downstream of the nozzle 24a, and a trailing edge 44b located in the internal secondary vein 26. Figure 3b shows that each of the variable-pitch blades 44 has an aerodynamic profile and comprises a lower surface 46 (of concave curved shape) and an upper surface 48 (of convex curved shape). Furthermore, each of the variable-pitch blades 44 has a certain curvature along its chord. In this embodiment, the number of variable-pitch blades 44 is equal to the number of fixed blades 42 and the variable-pitch blades 44 are located directly downstream of the fixed blades 42 and in the axial extension ofthese. The leading edges 44a of the variable-pitch blades 44 are separated by predetermined axial clearances J from the trailing edges 42b of the fixed blades 42. Preferably, these clearances J are less than 10mm and more preferably less than or equal to 5mm. Preferably, these clearances J are less than 10% of the chord of a blade 42 or a blade 44, and more preferably less than or equal to 5% of this chord. Each of these clearances J is preferably constant over the entire radial extent of the edges 42b, 44a concerned and therefore of the internal vein 26. Naturally, these clearances J are likely to vary in operation depending on the pitch positions of the blades 44 relative to the blades 42. The variable-pitch blades 44 are preferably all identical. Their leading edges 44a are preferably in the same transverse plane or crossed by the same transverse plane. The number of variable-pitch blades 44 is for example between 10 and 200. Each of the variable-pitch bladesvariable 44 is rotatable about a setting axis Y which has a substantially radial orientation. The rotation of each of the variable-pitch blades 44 is obtained by means of a control system 50 which is here located in the separator 24. Figure 3b shows on the left a first angular or setting position of the variable-pitch blades 44 and on the right a second angular or setting position of these blades. The variable-pitch blades 44 can, for example, be moved over angular ranges of the order of 60° around their Y axes. Figures 4a and 4b illustrate a second embodiment of the invention which differs from the previous embodiment essentially in that the number of variable-pitch blades 44 is different from the number of fixed blades 42 and is here greater than the number of fixed blades 42. In this variant, there are twice as many variable-pitch blades 44 as fixed blades 42. It is therefore understood that the circumferential pitch between the fixed blades 42is twice as large as the circumferential pitch between the variable-pitch vanes 44. Alternatively, the number of fixed vanes 42 is equal to a multiple of the number of variable-pitch vanes 44, which is different from 2 and which is for example 3, 4, etc. Half of the variable-pitch vanes 44 extend downstream and in the axial extension of the fixed vanes 42, as is the case in the embodiment of Figures 3a and 3b. The other half of the variable-pitch vanes 44 are interposed between the fixed vanes 42 and therefore do not extend in the extension of the fixed vanes 42. The variable-pitch vanes 44 are preferably all identical. Their leading edges 44a are preferably located in the same transverse plane or crossed by the same transverse plane, as is the case with the fixed blades 42. Figures 5a and 5b illustrate a third embodiment of the invention which differs from the first embodiment essentially by the positioning of the blades with pitchvariable 44 relative to the fixed blades 42. The variable-pitch blades 44 are interposed axially between the fixed blades 42 and are arranged between these blades 42. The variable-pitch blades 44 are not located in the axial extension of the fixed blades 42 but are on the contrary angularly offset by half a pitch relative to the axis of the turbomachine and are therefore each located midway between two fixed blades 42. The leading edges 44a of the variable-pitch blades 44 are located upstream of the trailing edges 42b of the fixed blades 42. The trailing edges 44b of the variable-pitch blades 44 are located downstream of the trailing edges 42b of the fixed blades. The nesting distance of the variable-pitch blades 44 between the fixed blades 42 is noted W and can be estimated as a percentage of the chord of one of the blades 42 or one of the blades 44. Preferably, this distance W is greater than 10% of the chord of a blade 42 or a blade 44, and more preferably greater than or equal to 20% ofthis rope. Figure 6 illustrates a fourth embodiment of the invention which differs from the first embodiment essentially in that the control system 50 is here located radially outside the external wall 14. This is advantageous because it makes it possible to locate this system in a relatively cold environment compared to the high temperatures which can prevail in the gas generator. Furthermore, this environment is not very constrained and contains free spaces to accommodate this type of system. This system 50 is connected to the variable-pitch blades 44 and for this purpose pass through the fixed blades 42. These blades 42 can thus be extended in the axial direction and comprise an internal passage extending in the radial direction through the external vein 28 to allow the mounting of the system 50 and its connection to the variable-pitch blades 44. It is therefore understood that the trailing edges 42c of the fixed blades 42 can be located downstream of the trailing edges 42bof these blades. Figure 7 illustrates a fifth embodiment of the invention which differs from the first embodiment by the position of the variable-pitch blades 44. The variable-pitch blades 44 are distributed around the axis in the external secondary vein 28 only. The variable-pitch blades 44 each have a leading edge 44a located downstream of the nozzle 24a, and a trailing edge 44b located in the external secondary vein 28. Each of the variable-pitch blades 44 has an aerodynamic profile and comprises a lower surface and an upper surface. Furthermore, each of the variable-pitch blades 44 has a certain curvature along its chord. The number of variable-pitch blades 44 may be equal to the number of fixed blades 42 or greater than this number, as mentioned above in relation to Figures 3a to 4b. The variable-pitch blades 44 are located directly downstream of the fixed blades 42 and in the axial extension thereof. The leading edges 44a of the variable-pitch bladesvariable pitch 44 are separated by predetermined axial clearances J of the trailing edges 42c of the fixed blades 42. Preferably, these clearances J are less than 10mm and more preferably less than or equal to 5mm. Preferably, these clearances J are less than 10% of the chord of a blade 42 or a blade 44, and more preferably less than or equal to 5% of this chord. Each of these clearances J is preferably constant over the entire radial extent of the edges 42c, 44a concerned and therefore of the external vein 28. Naturally, these clearances J are likely to vary in operation depending on the pitch positions of the blades 44 relative to the blades 42. The variable pitch blades 44 are preferably all identical. Their leading edges 44a are preferably located in the same transverse plane or crossed by the same transverse plane. The number of variable-pitch blades 44 is for example between 10 and 200. Each of the variable-pitch blades 44 is movable in rotation aroundof a pitch axis Y which has a substantially radial orientation. The rotation of each of the variable pitch blades 44 is obtained by means of a control system 50 which is here located radially outside the external wall 14. Figure 8 illustrates a sixth embodiment of the invention which differs from the first embodiment in that variable pitch blades 44 are further distributed around the axis in the external secondary vein 28. The variable pitch blades 44 of the internal vein 26 may be similar to those described above in relation to Figures 3a and 3b, or 4a and 4b, and the variable pitch blades 44 of the external vein 28 may be similar to those described above in relation to Figures 7a and 7b. The angular pitch of the variable pitch blades 44 located in the two veins is here controlled by independent systems 50. A first control system 50 is located in the separator 24 andcontrols the pitch of the variable-pitch vanes 44 in the inner duct 26, and a second control system 50 is located radially outside the wall 14 and controls the pitch of the variable-pitch vanes 44 in the outer duct 28. In the seventh embodiment of FIG. 9, a single control system 50 is used to control the angular pitch of the variable-pitch vanes 44 located in both ducts 26, 28. This control system 50 is located radially outside the wall 14. FIG. 10 illustrates an eighth embodiment of the invention which differs from the first embodiment essentially in that the fixed vanes 42 are not all identical. The fixed vanes 42 are at least of two types which differ from each other in their dimensions and / or their geometries and / or their cambers, etc. The different types of fixed blades 42 are regularly distributed around the axis so as to obtain a cyclic distribution of theseblades 42 around the axis. In the ninth embodiment of the invention illustrated in Figure 11, structural arms 32 are located in the external vein 14 downstream of the trailing edges 42c of the fixed blades 42. The number of arms 32 is less than the number of fixed blades 42 and the arms 32 may extend in the axial extension of some of the fixed blades 42. The arms 32 may all be identical. In the tenth embodiment of the invention illustrated in Figure 12, the structural arms 32 are brought together axially upstream and are connected to some fixed blades 42. The arms 32 are therefore integrated with the fixed blades 42. The fixed blades 42 which are not connected to arms 32 have their trailing edges 42c which are located upstream of the trailing edges 32a of the arms. The arms 32 are all identical in Figure 10 and are different and have a cyclic distribution in the eleventh embodiment of Figure 11. In the embodiments of Figures 11 to 13,some arms 32 may be solid for example and others may be tubular for the passage of services from the external wall 14 to the separator 24. In general, the present invention applies to any turbomachine in which a main flow is separated into two secondary flows downstream of a shrouded rotor blade.

Claims

CLAIMS 1. A triple-flow aircraft turbomachine (10), comprising a gas generator comprising along a longitudinal axis at least one compressor, a combustion chamber and at least one turbine, the turbomachine further comprising: - two coaxial annular walls, respectively internal (12) and external (14), extending around each other and defining between them a main annular flow path (16) for a main airflow (18), - a rotor blade (30) extending radially through said main flow path (16) and forming a shrouded propeller (H1), - an annular separator (24) arranged downstream of the rotor blade (30) and between the two walls (12, 14), the separator (24) defining respectively with the internal and external walls (12, 14) two secondary annular flow paths, respectively internal (26) and external (28), for the flow of airflow secondary, respectively internal (20) and external (22),the separator (24) comprising at an upstream end an annular nozzle (24a) configured to separate the main air flow (18) in two and form the secondary air flows (20, 22), - stator elements extending radially on the one hand through said main vein (16) and on the other hand through said secondary veins (26, 28), these stator elements being connected to said annular nozzle (24a), and - an unducted propeller (H2) arranged upstream of the external wall (14), characterized in that said stator elements comprise: - fixed rectifier vanes (42) which are distributed around said axis and which each comprise a leading edge (42a) located upstream of said nozzle (24a), and trailing edges, respectively internal (42b) and external (42c), located respectively in the internal (26) and external (28) secondary veins, these fixed rectifier vanes (42) being connected to said nozzle (24a),and - variable pitch rectifier vanes (44) which are distributed around said axis and which extend radially through at least one of said veins, secondary (26, 28), each of the variable-pitch stator vanes (44) having a leading edge (44a) and a trailing edge (44b), and in that: - the leading edges (44a) of the variable-pitch stator vanes (44) are located upstream of the inner (42b) and / or outer (42c) trailing edges of the fixed stator vanes (42), or - the leading edges (44a) of the variable-pitch stator vanes (44) are located directly downstream of the inner (42b) and / or outer (42c) trailing edges of the fixed stator vanes (42), and are separated by predetermined axial clearances (J) of these trailing edges (42b, 42c).

2. Turbomachine (10) according to claim 1, wherein the number of said variable-pitch stator vanes (44) is greater than or equal to the number of said fixed stator vanes (42). 3.Turbomachine (10) according to one of the preceding claims, wherein said variable-pitch stator vanes (44) are located in said internal secondary flow path (26).

4. Turbomachine (10) according to one of the preceding claims, wherein the trailing edges (44b) of said variable-pitch stator vanes (44) are located downstream of the external trailing edges (42c) of the fixed stator vanes (42).

5. Turbomachine (10) according to claim 4, wherein it further comprises a system (50) for controlling the angular pitch of the variable-pitch stator vanes (42), this system being mounted radially outside said external wall (14).

6. Turbomachine (10) according to one of claims 1 to 3, wherein said variable-pitch stator vanes (44) are located in said external secondary flow path (28). 7.Turbomachine (10) according to one of claims 1 to 3, in which first variable-pitch stator vanes (44) are located in said internal secondary vein (26), and second variable-pitch stator vanes (44) are located in said external secondary vein (28).

8. Turbomachine (10) according to claim 7, wherein it further comprises a common system (50) for controlling the angular pitch of the first and second variable-pitch stator vanes (44), or independent systems (50) for controlling the angular pitch of the first and second variable-pitch stator vanes (44) respectively.

9. Turbomachine (10) according to one of the preceding claims, wherein it further comprises structural arms (32) distributed around said axis in said external secondary vein (28).

10. Turbomachine (10) according to claim 9, wherein the number of structural arms (32) is less than the number of fixed stator vanes (42).

11. Turbomachine (10) according to claim 9 or 10, wherein the structural arms (32) are connected to some of said fixed stator vanes (42). 12.Turbomachine (10) according to one of the preceding claims, in which the rotor blade (30) is a fan or a compressor rotor blade.

13. Turbomachine (10) according to one of the preceding claims, in which the leading edges (44a) of the variable-pitch stator vanes (44) are located at a distance (W) from the inner (42b) and / or outer (42c) trailing edges of the fixed stator vanes (42), which is greater than 10% of the chord of one of these vanes (42, 44), and more preferably greater than or equal to 20% of this chord.

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