AIRCRAFT TOWER INTERNAL ENGINE

Description

Domaine technique de l'invention

[0001] The present invention relates to the general field of aeronautics. It relates more particularly to an aircraft turbomachine. Arrière-plan technique

[0002] The technical background includes, in particular, documents US-A1-2019 / 078536, US-A1-2016 / 2014608 and US-A1-2011 / 171007.

[0003] Conventionally, an aircraft turbomachine comprises a gas generator including, along a longitudinal axis, at least one compressor, one combustion chamber, and at least one turbine.

[0004] An airflow enters the gas generator and is compressed in the compressor(s). This compressed airflow is mixed with fuel and burned in the combustion chamber. The combustion gases are then expanded in the turbine(s). This expansion causes the turbine rotor(s) to rotate, which in turn causes the compressor rotor(s) to rotate. The combustion gases are expelled through a nozzle to provide thrust, which can be added to the thrust provided by at least one propeller or propulsion fan of the turbomachine.

[0005] The gas flows through the turbomachine via annular channels. As can be seen in the figure 1a The turbomachine 10 thus comprises coaxial annular walls, respectively internal 12 and external 14, extending around each other and defining between them a main annular flow 16 of a main gas stream 18. In the case where the main gas stream 18 is to be divided into two secondary gas streams, respectively internal 20 and external 22, an annular separator 24 is disposed between the two walls 12, 14 and defines respectively with these walls 12, 14 two secondary annular flow ducts, respectively internal 26 and external 28, of the secondary gas streams 20, 22. This separator 24 has at an upstream end an annular nozzle 24a configured to divide the main gas stream 18 into two and form the secondary gas streams 20, 22.

[0006] A rotor blade 30 can extend radially through the main channel 16, therefore upstream of the separator 24.

[0007] As illustrated in the figure 1a , structural arms 32 can extend radially through the main channel 16 downstream of the rotor blade 30 and upstream of the separator 24.

[0008] In this application, the term arm 32 or structural arm means a stator element which has a general aerodynamic cross-sectional shape such as that shown in the figure 1b but which does not include an intrados or extrados. Therefore, an arm 32 is not comparable to a blade or fan blade, which is profiled to include an intrados and an extrados. An arm 32 generally exhibits 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 can be as low as 4. At least one of the arms 32 may be hollow and tubular in the radial direction to allow passage of auxiliary lines and thus facilitate the transmission of these lines into the engine through the intake manifolds.

[0009] For certain types of turbomachinery, such as those with multiflow or variable cycle, it would be useful to have a stator blade 34 directly downstream of the rotor blade 30 and integrated into the flow separation nozzle 24a instead of being positioned between the rotor 30 and the separator 24 (cf. figure 2a ), in order to reduce the length of the module between the concept illustrated on the figure 1a and the one illustrated on the figure 2a The stator blade 34 would comprise several blades distributed around the axis of the turbomachine. As mentioned above and illustrated in the figure 2b , each of these blades would have in cross-section an aerodynamic profile comprising an intrados 34a and an extrados 34b ( figure 2b ), therefore a non-symmetrical profile, which is not the case for arm 32 visible on the figure 1a The stator blade 34 would extend radially through the main vein 16. These blades would include leading 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 blade could be connected to the nozzle 24a.

[0010] The stator blade 34 would impose a specific direction on the gas flows 16, 20, 22. However, in the case of a variable-cycle turbomachine, it would be useful to incorporate a variable geometry downstream of the rotor blade 30 to adapt to the different operating regimes and variations in the turbomachine's bypass ratio. However, due to space constraints, adding a variable-pitch blade downstream of the stator blade 34 could be complex. Indeed, this addition would require lengthening the axial dimension of the turbomachine, resulting in an increase in the turbomachine's mass and a decrease in its performance.

[0011] Furthermore, for reasons of noise pollution, it would also not be feasible to bring the stator blade 34 closer axially to the rotor blade 30.

[0012] In this application, a variable-cycle turbomachine is defined as a turbomachine whose specific thrust can be modified at a given engine speed by controlling variable geometries of the turbomachine. An example of a variable geometry is a stator blade with variable pitch. In this application, a blade is defined as an annular row of blades.

[0013] The invention thus proposes to optimize a turbomachine as illustrated in the figure 2a so that it can be used in several configurations and in particular in the context of a multi-flow turbomachine (at least two) and / or a variable cycle turbomachine. Résumé de l'invention

[0014] The present invention proposes an aircraft turbomachine, comprising a gas generator including, along a longitudinal axis, at least one compressor, one 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 channel of a main airflow, a rotor blade extending radially through said main flow channel, an annular separator disposed downstream of the rotor blade and between the two walls, the separator defining with the internal and external walls respectively two secondary annular flow channels, respectively internal and external, of secondary airflows, respectively internal and external, the separator having at an upstream end an annular nozzle configured to split the main airflow into two and form the secondary airflows, stator elements extending radially on one side through said main flow channel and on the other side through said secondary flow channels,characterized in that said stator elements comprise first variable-pitch rectifier blades distributed around said axis and each comprising 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 stator blades which are at least partly fixed and which are distributed around said axis in the external secondary flow and downstream of the external trailing edges of said first variable pitch blades.

[0015] The present invention thus proposes to place variable-pitch rectifier blades at the separator nozzle. To allow angular displacement of these blades around their pitch axes, it is understood that the blades will be separated by small clearances between the nozzle and the separator, in order to limit gas leakage in these areas.

[0016] Fixed stator blades are combined with variable-pitch blades and are located in the outer secondary flow. This configuration optimizes turbomachine operation, enabling multiflow or variable-cycle applications while minimizing the impact on the turbomachine's axial length, dimensions, and mass. Positioning the variable-pitch blades at the nozzle reduces the axial distance between the rotor and the flow separation nozzle, allowing for adjustments to the gas flow in the inner and outer secondary flows.

[0017] Fixed rectifier blades can be entirely fixed or each comprise a fixed part and a moving part, particularly with variable pitch. Each of these blades includes, for example, an upstream part with variable pitch including a leading edge, and a fixed downstream part including a trailing edge.

[0018] In this application, "annular" means a shape of revolution around an axis, this shape being either continuous or interrupted. Furthermore, in this application, "variable-pitch element" means an element at least part of which has a position that can be adjusted around an axis, referred to as the pitch axis. The entire element or only a part of it may have variable pitch. In the case of a blade, for example, it may be a single piece with an adjustable position around a pitch axis. Alternatively, it could comprise only a part, such as a leading or trailing edge, whose position is adjustable around a pitch axis relative to the rest of the blade. In the case of a blade assembly with multiple blades, each blade has an adjustable position around its own pitch axis.For the same blade, there are therefore as many pitch axes as there are variable pitch blades. Each of these axes can have a radial or inclined orientation relative to the longitudinal axis of the turbomachine.

[0019] The turbomachine may include one or more of the following features, taken alone or in combination with each other: the stator elements further comprise second variable pitch rectifier blades which are located in said internal secondary vein; the second variable pitch blades 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 from 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 during 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 flowing over the lower surfaces of the first variable-pitch blades must then flow over the lower surfaces of the second variable-pitch blades, and that the gases flowing over the upper surfaces of the first variable-pitch blades must then flow over the upper surfaces of the second variable-pitch blades; when the clearances are larger, part of the gas flow over the upper surfaces of the first variable-pitch blades then flows towards the lower surfaces of the second blades and provides energy to the fluid flowing over the lower surfaces of the first blades. The stator blades, which are at least partially fixed, 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 of said first variable-pitch blades; the number of said blades, which are at least partially fixed, is equal to the number of said first variable-pitch blades; the number of said blades, which are at least partially fixed, is equal to a multiple 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 blades that are at least partly fixed have profiles different from these other blades, and thus form a multi-profile blade grid; the rotor blade is a propulsion fan or a compressor rotor blade; and the stator blades that are at least partly fixed are entirely fixed; the stator blades that are at least partly fixed each comprise a fixed part and a movable part, and in particular a variable-pitch part; each of these blades comprises, for example, an upstream variable-pitch part having a leading edge, and a downstream fixed part having a trailing edge; said stator blades that are at least partly fixed comprise an upper and lower surface, and said variable-pitch stator blades comprise an upper and lower surface.

[0020] The present invention also relates to an aircraft, in particular a transport aircraft, comprising a turbomachine as described above. Brève description des figures

[0021] Other features and advantages of the invention will become apparent upon reading the detailed description that follows, for an understanding of which reference should be made to the attached drawings in which: [ Fig.1a-1b ] there figure 1a is a very schematic half-view in axial cross-section of an aircraft turbomachine, according to the prior art; the figure 1b is a very schematic cross-sectional view of an arm of the turbomachine of the figure 1a ; Fig.2a-2b ] there figure 2a is a very schematic half-view in axial cross-section of an aircraft turbomachine; the figure 2b is a very schematic cross-sectional view of a stator blade of the turbomachine of the figure 2a ; Fig.3a-3c ] there figure 3a is a very schematic half-view in axial cross-section of an aircraft turbomachine, according to a first embodiment of the invention; the figure 3b is a very schematic cross-sectional view of a variable-pitch stator blade followed by a fixed-pitch stator blade of the turbomachine of the figure 3a and illustrates, respectively on the left and right of the figure, two distinct positions for the variable-pitch stator blade; the figure 3c is a view similar to that of the figure 3b from the left and shows a variant embodiment of the invention; [ Fig.4a-4c ] there figure 4a is a very schematic half-view in axial cross-section of an aircraft turbomachine, according to a second embodiment of the invention; the figure 4b is a very schematic cross-sectional view of a first variable-pitch stator blade followed by a second variable-pitch stator blade of the turbomachine of the figure 4a and illustrates, respectively on the left and right of the figure, two distinct positions for setting these blades; the figure 4c is a view similar to that of the figure 4b from the left and shows a variant embodiment of the invention; and [ Fig.5 ] there figure 5 is a very schematic half view in axial section 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. Description détaillée de l'invention

[0022] THE figures 1a, 1b , 2a et 2b have been described above.

[0023] THE figures 3a et 3b illustrate a first embodiment of the invention.

[0024] In a conventional manner, the turbomachine 10 comprises a gas generator (not shown) including along a longitudinal axis at least one compressor, one combustion chamber and at least one turbine, the turbomachine.

[0025] 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 16 of a main gas flow 18.

[0026] The main gas flow 18 is divided into two secondary gas flows, internal 20 and external 22 respectively, by an annular separator 24 which is disposed between the two walls 12, 14. This separator 24 has at an upstream end an annular spout 24a configured to separate the main gas flow 18 into two and form the secondary gas flows 20, 22.

[0027] A rotor blade 30 extends radially through the main channel 16, upstream of the separator 24.

[0028] Stator elements are located downstream of the rotor blade 30 and at the level of the separation nozzle 24a.

[0029] According to the invention, these stator elements include first variable-pitch rectifier blades 40.

[0030] Furthermore, "fixed" stator blades 42 are located in the external secondary vein 28 downstream of the first variable pitch stator blades 40.

[0031] The first variable-pitch blades 40 are distributed around the axis and each has a leading edge 40a located upstream of the nozzle 24a, and trailing edges, respectively internal 40b and external 40c, located in the internal 26 and external 28 secondary channels. It is thus understood that the first variable-pitch blades 40 are located at the nozzle 24a, as shown in the drawing. Concealed clearances are provided between the nozzle 24a and the first variable-pitch blades 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 nozzle 24a. As also visible, the leading edges 42a can be inclined and extend outwards from upstream to downstream. This inclination is determined, for example, by a compromise between the size of the engine and the optimization of the noise it generates.To minimize noise, it is best to increase the height at the top of the blade, which results in a greater blade tilt.

[0032] There figure 3b This shows that each of the first variable-pitch blades 40 has an aerodynamic profile and comprises a concave curved lower surface 46 and a convex curved upper surface 48. Furthermore, each of the first variable-pitch blades exhibits a certain curvature along its chord. The area of ​​greatest curvature of a variable-pitch blade 40 is denoted by C. This area is preferably located upstream of the nozzle 24a. The first variable-pitch blades 40 are preferably all identical. Their leading edges 40a are preferably crossed by the same transverse plane.

[0033] The number of the first variable pitch blades 40 is, for example, between 10 and 200.

[0034] Each of the first variable-pitch blades 40 rotates about a pitch axis Y, which has a substantially radial orientation. The rotation of each of the first variable-pitch blades 40 is achieved by a control system 50, which is located radially outside the outer wall 14. This is advantageous because it allows the system to be located in a relatively cool environment compared to the high temperatures that can occur in the gas generator. Furthermore, this environment is less constrained and contains free spaces to accommodate this type of system.

[0035] The fixed blades 42 are distributed around the axis in the external secondary vein 28. Each has a leading edge 42a located downstream of the nozzle 24a, and a trailing edge 42b located in the external secondary vein 28. The figure 3b This shows that each of the fixed blades 42 has an aerodynamic profile and comprises an intrados 46 (with a concave curved shape) and an extrados 48 (with a convex curved shape). Furthermore, each of the fixed blades 42 exhibits a certain curvature along its chord.

[0036] The fixed blades 42 may be entirely fixed or each may comprise a fixed part and a movable part, particularly with variable pitch. Each of these blades 42 comprises, for example, an upstream part with variable pitch, including the leading edge 42a, and a fixed downstream part, including the trailing edge 42b. These parts are schematically delimited by dashed lines in the figures.

[0037] The number of fixed blades 42 is equal to the number of the first variable-pitch blades 40, or a multiple thereof, and the fixed blades 42 are located directly downstream of the first variable-pitch blades 40 and in their axial extension. 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 10 mm and more preferably less than or equal to 5 mm. Preferably, these clearances I are less than 10% of the chord of a blade 40 or a blade 42, and more preferably less than or equal 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 positioning of the blades 40 relative to the blades 42.

[0038] 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.

[0039] The number of fixed blades 42, for example, is between 10 and 200.

[0040] There figure 3b The diagram on the left shows 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 approximately 60° around their Y axes.

[0041] There figure 4c illustrates an alternative embodiment in which the number of fixed blades 424 is equal to a multiple of the number of the first variable pitch blades 40. This multiple is for example 2, 3, 4, etc.

[0042] THE figures 4a et 4b illustrate a second embodiment of the invention which differs from the previous embodiment essentially in that the turbomachine further comprises second variable pitch stator blades 44 located in the internal secondary channel 26 downstream of the trailing edges 40b of the first variable pitch blades 40.

[0043] The second 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.

[0044] Each of the second variable-pitch blades 44 has an aerodynamic profile and comprises an intrados and an extrados. Furthermore, each of the variable-pitch blades 44 exhibits a certain curvature along its chord.

[0045] The number of second variable-pitch blades 44 can be equal to the number of first variable-pitch blades 40.

[0046] The second variable-pitch blades 44 are located directly downstream of the fixed blades 42 and in their axial extension. 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 10 mm and more preferably less than or equal to 5 mm. 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 relevant edges 40b, 44a and therefore of the internal duct 26. Naturally, these clearances J are likely to vary during operation depending on the pitch positions of the blades 40, 44.

[0047] 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.

[0048] The number of variable-pitch secondary blades 44 is, for example, between 10 and 200. Each variable-pitch secondary blade 44 rotates about a pitch axis Z, which has a substantially radial orientation. The rotation of each variable-pitch secondary blade 44 is achieved by a control system 50', which is located in the separator 24.

[0049] There figure 4c illustrates an alternative embodiment in which the number of variable pitch stator blades 44 is equal to a multiple of the number of first variable pitch blades 40. This multiple is for example 2, 3, 4, etc.

[0050] There figure 5This illustrates a third embodiment of the invention, which differs from the previous embodiment essentially in that the fixed blades 42 are not all identical. The fixed blades 42 are of at least two types, differing from one another in their dimensions and / or geometries and / or cambers, etc. The different types of fixed blades 42 are regularly distributed around the axis so as to obtain a cyclic distribution of these blades 42 around the axis.

[0051] 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

1. An 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 duct (16) for a main air flow (18), - a rotor blading (30) extending radially across said first duct (16), - an annular separator (24) arranged downstream of the rotor blading (30) and between the two walls (12, 14), the separator (24) defining, with the internal and external walls (12, 14) respectively, two secondary annular flow ducts, respectively internal (26) and external (28), of the secondary air flow, respectively internal (20) and external (22), the separator (24) comprising at an upstream end an annular splitter nose (24a) configured to split the main air flow (18) into two and to form the secondary air flows (20, 22), - stator elements extending radially on the one hand through said main duct (16) and on the other hand through said secondary ducts (26, 28), characterised in that said stator elements comprise first variable pitch straightener vanes (40) which are distributed around said axis and which each comprise a leading edge (40a) located upstream of said splitter nose (24a), and trailing edges, respectively internal (40b) and external (40c), located respectively in the internal (26) and external (28) secondary ducts, and in that the turbomachine (10) further comprises straightener vanes (42) that are at least partly stationary and that are distributed about said axis in the external secondary duct (28) and downstream of the external trailing edges (40c) of said first variable pitch vanes (40).

2. The turbomachine (10) of claim 1, wherein the stator elements further comprise second variable pitch straightener vanes (44) that are located in said internal secondary duct (26).

3. The turbomachine (10) of 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) from these trailing edges (40b).

4. The turbomachine (10) according to any of the preceding claims, wherein the straightener vanes (42) which are at least partly stationary comprise leading edges (42a) separated by predetermined axial clearances (I) from said trailing edges (40c) of the first variable pitch vanes (40).

5. The turbomachine (10) according to claim 3 or 4, wherein said clearances are less than 10mm, and preferably less than or equal to 5mm.

6. The turbomachine (10) according to any of claims 2 to 4, wherein the number of said second variable pitch vanes (44) is equal to the number of said first variable pitch vanes (40) or a multiple of said first variable pitch vanes (40).

7. The turbomachine (10) according to any of the preceding claims, wherein the number of said vanes (42) that are at least partly stationary is equal to the number of said first variable pitch vanes (40) or a multiple of said first variable pitch vanes (40).

8. The 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 vanes (40).

9. The turbomachine (10) according to claim 7, wherein said control system (50, 50') is mounted within said separator (24) or radially outside said external wall (14).

10. The turbomachine (10) according to any of the preceding claims, wherein at least some of said vanes (42) which are at least partly stationary, have different profiles from those other vanes (42).

11. The turbomachine (10) according to any of the preceding claims, wherein the rotor blading (30) is a fan or a compressor rotor blading.

12. The turbomachine (10) according to any of the preceding claims, wherein the straightener vanes which are at least partly stationary are fully stationary.

13. The turbomachine (10) according to any of claims 1 to 11, wherein the straightener vanes which are at least partly stationary comprise each a stationary portion and a movable portion and particularly with variable pitch.

14. The turbomachine (10) according to claim 13, wherein each of the straightener vanes comprises for example an upstream portion with variable pitch including a leading edge, and a stationary downstream portion comprising a trailing edge.

15. The turbomachine (10) according to any of the preceding claims, wherein said straightener vanes which are at least partly stationary comprise an intrados and an extrados, and said variable pitch straightener vanes comprise an intrados and an extrados.