Triple-flow aircraft turbomachine

Abstract

The invention relates to a triple-flow aircraft turbomachine (10) comprising: • - two coaxial annular walls defining between them a main annular duct (16) for the flow of a primary airflow (18); • - a rotor blading (30) extending radially through the main duct (16) and forming a ducted fan (Hl); • - an annular separator (24) arranged downstream of the rotor blading (30) and configured to divide the primary airflow (18) into two to form the secondary air flows (20, 22); • - first variable guide vanes (40) which are distributed around the shaft and each comprise a leading edge (40a) which is located upstream of the separator (24) and trailing edges (40b, 40c) which are located in the secondary airflows (20, 22); and • - fixed guide vanes (42) which are distributed around the shaft in the external airflow and downstream of the first variable vanes (40). • - an unducted fan (H2) arranged upstream of the outer wall (14).

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 blades 34 would extend radially across the main vein 16. These blades would include leading edges 36 located upstream of the nozzle 24a, in the main vein 16, andtrailing edges, respectively internal 38a and external 38b, located in the internal 26 and external 28 veins. The stator blading could be connected to the nozzle 24a. 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 rotor blading 30 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 a 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 also not be possible to axially move the stator blading 34 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 blade. In the present application, a blade 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 around each 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 helix, - 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, and - an unshrouded helix arranged upstream of the external wall, characterized in that said stator elements comprise firstvariable-pitch stator vanes which are distributed around said axis and which each comprise a leading edge located upstream of said nozzle, and trailing edges, respectively internal and external, located respectively in the internal and external secondary veins, and in that the turbomachine further comprises fixed stator vanes which are distributed around said axis in the external secondary vein and downstream of the external trailing edges of said first variable-pitch vanes. The present invention thus proposes to place variable-pitch stator vanes at the separation nozzle. To allow the angular movement of these vanes around their pitch axes, it is understood that the vanes will be separated by small clearances from the nozzle and the separator, this in order to limit gas leaks in these zones. Fixed stator vanes are associated with the variable-pitch vanes and are located in the external secondary vein. This configuration allowsto 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, providing the variable-pitch blades at the nozzle makes it possible to axially reduce the distance between the rotor and the flow separation nozzle while allowing a modification of the gas flows flowing in the internal and external secondary veins. 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" element means an element of which at least one part has a position that can be adjusted around an axis, which is called the pitch axis. The entirety of this element or only a part of this element may be variable-pitch. In the case of a blade, for example, it may be a single piece and havean adjustable position around a setting axis. Alternatively, it could comprise only a part, comprising for example a leading edge or a trailing edge, the position of which would be adjustable around a setting axis relative to the rest of the blade. In the case of a blading comprising several blades, each of the blades has an adjustable position around a setting axis of its own. For the same blading, there are therefore as many setting 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 features, taken alone or in combination with each other: - the stator elements further comprise second variable-pitch stator vanes which are located in said internal secondary flow path; - the second variable-pitch vanes comprise leading edges and trailing edges,the leading edges of these second variable-pitch blades being located directly downstream of the internal trailing edges of the first variable-pitch blades, and being separated by predetermined axial clearances of these trailing edges; the first and second variable-pitch blades are therefore very close axially to each other so that they are considered as an assembly forming the stator elements within the meaning of the invention; the aforementioned axial clearances between these blades are preferably as small as possible. Minimizing these axial clearances makes it possible to limit or even prevent the passage of gas in operation between the trailing edges of the first variable-pitch blades and the leading edges of the second variable-pitch blades; it is thus understood that the gases which flow on the intrados of the first variable-pitch blades must then flow on the intrados of the second variable-pitch blades, and that the gases which flow on theextrados of the first variable-pitch blades must then flow onto the extrados of the second variable-pitch blades; when the clearances are greater, a portion of the gas flow flowing onto the extrados of the first variable-pitch blades then flows towards the intrados of the second blades and provides energy to the fluid flowing onto the intrados of the first blades.; - the fixed rectifier blades comprise leading edges separated by predetermined axial clearances from said trailing edges of the first variable-pitch blades; -- said clearances are preferably less than 10 mm, more preferably less than or equal to 5 mm; - the number of said second variable-pitch blades is equal to the number of said first variable-pitch blades; - the number of said second variable-pitch blades is equal to a multiple number of said first variable-pitch blades; - the number of said fixed blades is equal to the number of said firstvariable-pitch blades; - the number of said fixed blades is equal to and equal to a multiple number of said first variable-pitch blades; - the turbomachine further comprises at least one system for controlling the angular pitch of the variable-pitch blades; - said control system is mounted in said separator or radially outside said outer wall; - at least some of said fixed blades have profiles different from the other fixed blades and therefore form a multi-profile blade grid; - the rotor blade is a propulsion fan or a compressor rotor blade; and -- 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 present invention also relates to an aircraft, in particular a transport aircraft, comprising a turbomachine as described in the above. Brief description of the figures Other characteristics andAdvantages of the invention will become apparent upon reading 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 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 an aircraft turbomachine, according to a first embodiment of the invention; [Fig.3b] Figure 3b is a very schematic cross-sectional view of a variable pitch stator blade followed by a fixed stator bladeof the turbomachine of figure 3a, and illustrates, respectively on the left and right of the figure, two distinct positions of setting of the variable-pitch rectifier blade; [Fig.3c] figure 3c is a view similar to that of figure 3b on the left and shows an alternative embodiment of the invention; [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 a first variable-pitch rectifier blade followed by a second variable-pitch rectifier blade of the turbomachine of figure 4a, and illustrates, respectively on the left and right of the figure, two distinct positions of setting of these blades; [Fig.4c] figure 4c is a view similar to that of figure 4b on the left and shows an alternative embodiment of the invention; [Fig.5] Figure 5 is a very schematic half-sectional viewaxial flow of an aircraft turbomachine, according to a third embodiment of the invention in which the stator grid is composed of at least two different blade profiles; and [Fig.6] Figure 6 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 6, the turbomachine 10 is of the triple flow type 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 comprises a ducted propeller denoted H1 and an unducted propeller denoted H2. The propeller H1 is surrounded by a nacelle 4 which extends around the axis X downstream of the propeller H2. The airflow passing through the propeller H2 is separated by the nacelle 4 into a main flow F2 which enters the nacelle 4 and another flow F3which flows around the nacelle 4. The main flow F2 is then divided into two other flows F1, F2 as explained in the following. 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 one around the other and defining between them a main annular flow vein 16 of 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 flow vein 16, upstream of the separator 24. In the context of the turbomachine of Figure 6, this rotor blade 30 formsthe shrouded H1 propeller. Stator elements are located downstream of the rotor blades 30 and at the separation nozzle 24a. According to the invention, these stator elements comprise first variable-pitch stator vanes 40. Furthermore, fixed stator vanes 42 are located in the external secondary vein 28 downstream of the first variable-pitch stator vanes 40. The first variable-pitch vanes 40 are distributed around the axis and each comprise a leading edge 40a located upstream of the nozzle 24a, and trailing edges, respectively internal 40b and external 40c, located respectively in the internal 26 and external 28 secondary veins. It is thus understood that the first variable-pitch vanes 40 are located at the nozzle 24a, as can be seen in the drawing. Invisible clearances are provided between the nozzle 24a and the first variable-pitch vanes 40 to allow their movement. These clearances are preferably as small as possible.to limit or prevent the passage of gas between these blades 40 and the nose 24a. As can also be seen, the leading edges 42a may be inclined and extend from upstream to downstream outwards. 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 first variable-pitch blades 40 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 first variable-pitch blades has a certain curvature along its chord. C denotes the zone of greatest curvature of a variable-pitch blade 40. This zone is preferably located upstream of the nose 24a. The first variable pitch blades 40 are preferably allidentical. Their leading edges 40a are preferably crossed by the same transverse plane. The number of first variable-pitch vanes 40 is for example between 10 and 200. Each of the first variable-pitch vanes 40 is rotatable about a pitch axis Y which has a substantially radial orientation. The rotation of each of the first variable-pitch vanes 40 is obtained by means of a control system 50 which 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. The fixed vanes 42 are distributed around the axis in the external secondary vein 28. They each comprise a leading edge 42a located downstream of the nozzle 24a, and a trailing edge 42blocated in the external secondary vein 28. 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. The number of fixed blades 42 is equal to the number of the first variable-pitch blades 40 or a multiple number of the first variable-pitch blades 40, and the fixed blades 42 are located directly downstream of the first variable-pitch blades 40 and in the axial extension thereof. The leading edges 42a of the fixed blades 42 are separated by predetermined axial clearances I from the trailing edges 40c of the first variable-pitch blades 40. Preferably, these clearances I are less than 10mm and more preferably less than or equal to 5mm. Preferably, these clearances I are less than 10% of the chord of a blade 40 or a blade 42, and more preferably less than orequal to 5% of this chord. Each of these clearances I is preferably constant over the entire radial extent of the edges 40c, 42a concerned and therefore of the external vein 28. Naturally, these clearances I are likely to vary in operation depending on the pitch positions of the blades 40 relative to the blades 42. The fixed blades 42 are preferably all identical. Their leading edges 42a are preferably in the same transverse plane or crossed by the same transverse plane. The number of fixed blades 42 is for example between 10 and 200. Figure 3b shows on the left a first angular or pitch position of the first variable-pitch blades 40 and on the right a second angular or pitch position of these blades. The first variable-pitch blades 40 can, for example, be moved over angular ranges of the order of 60° around their Y axes. Figure 4c illustrates an alternative embodiment in which the number of fixed blades 424 is equal to a multiple of thenumber of first variable-pitch vanes 40. This multiple is for example 2, 3, 4, etc. Figures 4a and 4b illustrate a second embodiment of the invention which differs from the previous embodiment essentially in that the turbomachine further comprises second variable-pitch rectifier vanes 44 located in the internal secondary vein 26 downstream of the trailing edges 40b of the first variable-pitch vanes 40. The second variable-pitch vanes 44 each comprise a leading edge 44a located downstream of the nozzle 24a, and a trailing edge 44b located in the internal secondary vein 26. Each of the second variable-pitch vanes 44 has an aerodynamic profile and comprises a lower surface and an upper surface. Furthermore, each of the variable-pitch vanes 44 has a certain curvature along its chord. The number of second variable-pitch blades 44 may be equal to the number of first variable-pitch blades 40. The second variable-pitch blades 44are located directly downstream of the fixed blades 42 and in the axial extension thereof. The leading edges 44a of the second variable-pitch blades 44 are separated by predetermined axial clearances J from 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 40 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 40b, 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 40, 44. The second 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 ofsecond variable-pitch vanes 44 is for example between 10 and 200. Each of the second variable-pitch vanes 44 is rotatable about a pitch axis Z which has a substantially radial orientation. The rotation of each of the second variable-pitch vanes 44 is obtained by means of a control system 50' which is here located in the separator 24. Figure 4c illustrates an alternative embodiment in which the number of variable-pitch rectifier vanes 44 is equal to a multiple of the number of first variable-pitch vanes 40. This multiple is for example 2, 3, 4, etc. Figure 5 illustrates a third embodiment of the invention which differs from the previous embodiment essentially by the fact that the fixed vanes 42 are not all identical. The fixed vanes 42 are at least of two types which differ from each other by their dimensions and / or their geometries and / or their cambers, etc. The different types of fixed blades 42 areregularly distributed around the axis so as to obtain a cyclic distribution of these blades 42 around the axis. 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 air flow (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 air flow 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), and - an unducted propeller (H2) arranged upstream of the external wall (14), characterized in that said stator elements comprise first variable-pitch rectifier vanes (40) which are distributed around said axis and which each comprise a leading edge (40a) located upstream of said nozzle (24a), and trailing edges, respectively internal (40b) and external (40c), located respectively in the internal (26) and external (28) secondary veins,and in that the turbomachine (10) further comprises fixed rectifier vanes (42) which are distributed around said axis in the external secondary vein (28) and downstream of the external trailing edges (40c) of said first variable-pitch vanes (40)., 2. Turbomachine (10) according to claim 1, wherein the stator elements further comprise second variable-pitch stator vanes (44) which are located in said internal secondary vein (26).

3. Turbomachine (10) according to claim 2, wherein the second variable-pitch vanes (44) comprise leading edges (44a) and trailing edges (44b), the leading edges (44a) of these second variable-pitch vanes (44) being located directly downstream of the internal trailing edges (40b) of the first variable-pitch vanes (40), and being separated by predetermined axial clearances (J) of these trailing edges (40b).

4. Turbomachine (10) according to one of the preceding claims, in which the fixed rectifier vanes (42) comprise leading edges (42a) separated by predetermined axial clearances (I) from said trailing edges (40c) of the first variable-pitch vanes (40). 5.Turbomachine (10) according to one of claims 2 to 4, wherein the number of said second variable-pitch blades (44) is equal to the number of said first variable-pitch blades (40) or to a multiple number of said first variable-pitch blades (40).

6. Turbomachine (10) according to one of the preceding claims, wherein the number of said fixed blades (42) is equal to the number of said first variable-pitch blades (40) or to a multiple number of said first variable-pitch blades (40).

7. Turbomachine (10) according to one of the preceding claims, wherein it further comprises at least one system (50, 50') for controlling the angular pitch of the variable-pitch blades (42).

8. Turbomachine (10) according to claim 7, wherein said control system (50, 50') is mounted in said separator (24) or radially outside said outer wall (14). 9.Turbomachine (10) according to one of the preceding claims, in which at least some of said fixed blades (42) have profiles different from the other fixed blades (42).

10. Turbomachine (10) according to one of the preceding claims, in which the rotor blade (30) is a fan or a compressor rotor blade.

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