Aircraft propulsion assembly

Abstract

The invention relates to an aircraft propulsion assembly (1) comprising a nacelle (3) surrounding a three-flow turbine (2), the turbine comprising a gas generator (4), a fan (16), and an annular element (18) through which a fan flow (FO) is accelerated, located between the generator (4) and the nacelle (3), defining a first duct (20) and a second duct (21), the annular element comprising a nose (19), the invention relates to an assembly (1) for separating a flow (F0) into an air flow (F1) through a first duct (20) and an air flow (F2) through a second duct (21), the assembly (1) comprising stator blades (22) mounted between the nose (19) and a fan (16), and a stator blade (24) between the generator (4) and the annular element (18), mounted between the nose (19) and a rotor blade of the compressor (8) of the generator (4), or mounted between the annular element (18) and the nacelle (3).

Description

Field of the Invention

[0001] The present invention relates to the field of general aviation. More specifically, the present invention relates to an aircraft propulsion assembly including a triple-flow turbine and a nacelle. The present invention also relates to an aircraft including such a propulsion assembly. Background Art

[0002] Traditionally, a propulsion assembly includes a nacelle surrounding a turbine that generates the thrust required to propel the aircraft. To this end, the turbine successively includes at least one compressor, a combustion chamber, and at least one turbine. The at least one compressor compresses the air flow entering the nacelle. The previously compressed air is mixed with fuel in the combustion chamber and then ignited to generate a hot propulsion gas flow. The at least one turbine is arranged to be rotated by this hot gas flow, and the turbine is connected to the compressor by a shaft. These elements form an engine, which is also referred to as a gas generator. Then, the hot gas flow is discharged through a nozzle at the outlet of the turbine. The rotor blades, also referred to as a fan, are typically mounted upstream of the gas generator to accelerate the primary air flow.

[0003] There are also dual-flow turbines in which an annular diverter is installed between the nacelle and the gas generator to divide the flow entering the nacelle into a primary air flow and a cold secondary air flow. The primary air flow flows into the gas generator, and the cold secondary air flow circulates in a duct formed by the space between the nacelle and the diverter. The main advantages of these turbines are less fuel consumption and quieter operation.

[0004] The propulsion of certain aircraft can also be ensured by a triple-flow turbine, such as that described in application No. FR-A1-3 074 476, in which non-ducted rotor blades forming a propeller are mounted upstream of the fan. The span of this additional rotor blade is generally greater than the span of the fan, such that the upstream edge of the nacelle divides the flow accelerated by the non-ducted rotor blades into a main flow and a tertiary air flow. The main flow enters the nacelle, and the tertiary air flow flows around the nacelle. Then, as in the case of dual-flow turbines, the main flow can be divided into a primary flow and a secondary flow.

[0005] The use of dual-flow turbines and triple-flow turbines is characterized by their bypass ratio, which corresponds to the ratio of the mass of the secondary / tertiary flow to the mass of the primary flow. This bypass ratio can also vary according to the flight phase of the aircraft, especially in variable cycle turbines. However, changes in the bypass ratio can lead to losses in the secondary / tertiary flow and thus a decrease in the efficiency and operability of the turbine. Summary of the Invention

[0006] The object of the present invention is to overcome this drawback by proposing an architecture that allows the straightening of the air flow entering the turbine and minimizes the impact of changes in the bypass ratio on the gas generator.

[0007] To this end, the present invention relates to an aircraft propulsion assembly, the propulsion assembly comprising a three-stream turbine and a nacelle surrounding the turbine, the turbine comprising:

[0008] - A gas generator, the gas generator comprising at least one compressor, a combustion chamber and a turbine, the gas generator being arranged along a longitudinal axis,

[0009] - A first propeller, the first propeller being mounted within the nacelle and around the longitudinal axis, and being configured to accelerate an incoming air flow entering the nacelle,

[0010] - At least one annular element, the at least one annular element being arranged radially between the gas generator and the nacelle and defining a first internal annular duct for supplying the gas generator and a second external annular duct for the nacelle, the annular element comprising upstream a first annular diverter nose configured to divide the incoming air flow into a first air flow flowing in the first duct and a second air flow flowing in the second external annular duct,

[0011] - A second propeller, the second propeller being mounted upstream of the nacelle and around the longitudinal axis, and being configured to accelerate a main air flow, the nacelle comprising upstream a second annular diverter nose configured to divide the main air flow into the incoming air flow flowing into the nacelle and a third air flow flowing around the nacelle.

[0012] The propulsion assembly is characterized in that the propulsion assembly further comprises:

[0013] - A first stator vane, the first stator vane extending radially between the housing of the gas generator and the nacelle, upstream of the first annular diverter nose and downstream of the first propeller, and

[0014] - A second stator vane, the second stator vane extending radially between the housing of the gas generator and the annular element, downstream of the first diverter nose and upstream of the first rotor vane of the at least one compressor of the gas generator, and / or between the annular element and the nacelle, downstream of the first annular diverter nose,

[0015] At least one of the first stator vane and the second stator vane is a variable pitch vane or comprises at least one variable pitch portion.

[0016] Thus, thanks to the present invention, the straightening of the air flow entering the nacelle is carried out upstream of the splitter, such that the vanes present in the duct only have the function of protecting the turbine against variations in the bypass ratio. Such an architecture makes it possible to simplify the construction and assembly of the various vanes present in a limited space such as a duct. The present invention also makes it possible to achieve a greater degree of freedom in the positioning of variable pitch vanes (the variable pitch vanes as a whole have a variable pitch, or only include a variable pitch part and thus include another fixed part), depending on the space available for the devices actuating the pitch. For example, at the level of the first annular splitter nose, the available space is generally very limited (since the thickness available in this area needs to be small), so it may be more attractive to place the first annular splitter nose in the nacelle.

[0017] The propulsion assembly may also have one or more of the following features taken individually or in combination with each other:

[0018] - The first stator vane has a variable pitch,

[0019] - The second outer annular duct does not have stator vanes from the first annular splitter nose to a plane that is perpendicular to the longitudinal axis and substantially passes through the first stator vane of the at least one compressor of the gas generator,

[0020] -- In particular, in the latter configuration, when the first stator vane has a variable pitch, the second stator vane as a whole may be fixed without having a variable pitch; in such a configuration, in fact, it is not necessary to have two consecutive variable pitch stator vanes,

[0021] - The second outer annular duct includes a third stator vane located downstream of the first annular splitter nose,

[0022] -- In particular, in the latter configuration, when the first stator vane has a variable pitch and the second stator vane is completely fixed, the third stator vane may include a variable pitch part and a fixed part; moreover, in such a configuration, it is not necessary to have two consecutive variable pitch stator vanes,

[0023] - The third stator vane is located downstream of the leading edge of the blade of the second stator vane or in line with the leading edge of the blade of the second stator vane, and is located upstream of the leading edge of the blade of the first stator vane of the at least one compressor of the gas generator or in line with the leading edge of the blade of the first stator vane of the at least one compressor of the gas generator,

[0024] - The second stator vane extends radially between the annular element and the nacelle, and the first inner annular duct does not have stator vanes upstream of the first rotor vane of the at least one compressor of the gas generator,

[0025] --Especially in the latter configuration, the second stator vane preferably has a variable pitch,

[0026] - The second stator vane is located downstream of the leading edge of the blade of the first rotor vane of the at least one compressor of the gas generator or in line with the leading edge of the blade of the first rotor vane of the at least one compressor of the gas generator, and is located upstream of the trailing edge of the blade of the first stator vane of the at least one compressor or in line with the trailing edge of the blade of the first stator vane of the at least one compressor,

[0027] - The first stator vane and / or the second stator vane includes a blade, the upstream portion of the blade includes a leading edge rotatable about a substantially radial axis, and the downstream portion of the blade includes a fixed trailing edge,

[0028] - The first stator vane and / or the second stator vane includes a blade, the downstream portion of the blade includes a trailing edge rotatable about a substantially radial axis, and the upstream portion of the blade includes a fixed leading edge,

[0029] The first propeller and the first rotor blade are connected to a single shaft, and

[0030] - The at least one first propeller and the first rotor blade are connected to a single shaft, preferably connected to the single shaft through a mechanical reduction gear.

[0031] The present invention also relates to an aircraft, especially a transport aircraft, including a propulsion assembly such as those mentioned above. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] Further features and advantages of the present invention will become apparent from the following detailed description. For understanding the following detailed description, reference is made to the drawings, in which:

[0033] Figure 1 Figure 1 A schematic longitudinal section of a propulsion assembly including a dual-flow turbine is shown;

[0034] Figure 2 Figure 2 A schematic folded longitudinal section of a civil-type propulsion assembly including a dual-flow turbine is shown;

[0035] Figure 3 Figure 3 A schematic longitudinal section of a propulsion assembly including a triple-flow turbine is shown;

[0036] Figure 4 Figure 4 A schematic folded longitudinal section of a civil-type propulsion assembly including a triple-flow turbine is shown; ​​​​​​​​

[0037] Figure 5 Figure 5 Shows a schematic folded longitudinal section of a propulsion assembly according to a first embodiment of the present invention;

[0038] Figure 6 Figure 6 Shows a schematic folded longitudinal section of a propulsion assembly according to a variant of the first embodiment of the present invention;

[0039] Figure 7 Figure 7 Shows a schematic folded longitudinal section of a propulsion assembly according to a second embodiment of the present invention;

[0040] Figure 8 Figure 8 Shows a schematic folded longitudinal section of a propulsion assembly according to a variant of the second embodiment of the present invention; and

[0041] Figure 9 Figure 9 Shows a schematic radial section of a stator blade with variable pitch according to a second embodiment of the present invention. Detailed Description of the Invention

[0042] In Figures 1 to 8 is schematically shown a propulsion assembly 1 for an aircraft (hereinafter referred to as "assembly 1"), whether the aircraft is a civil aircraft or not. Assembly 1 includes a turbine 2, which is arranged along a longitudinal axis X-X. In the context of a civil aircraft, the turbine 2 is a three-stream turbine, for example as Figure 3 shown. The turbine 2 may be surrounded by a nacelle 3, but other cases are not excluded.

[0043] The turbine 2 generally includes a gas generator 4, which includes at least one compressor 8, a combustion chamber, and a turbine 7. As Figure 1 and Figure 3 shown, the gas generator 4 forms a compartment 5, preferably in which a high-pressure body 6 and a low-pressure body are arranged. The high-pressure body 6 is formed by a high-pressure compressor, a high-pressure combustion chamber, and a high-pressure turbine (not detailed in the figure). The low-pressure body includes at least one low-pressure turbine 7 arranged downstream of the high-pressure body 6 and a low-pressure compressor 8 arranged upstream of the high-pressure body 6. The high-pressure compressor and the low-pressure compressor 8 include rotor blades 9 and stator blades 10 arranged alternately in sequence around the longitudinal axis X-X from upstream to downstream. In the present invention, a rotor blade is defined as a wheel attached with blades or vanes and rotating around the longitudinal axis X-X. In addition, a stator blade is defined as a wheel attached with blades or vanes and not rotating around the longitudinal axis X-X.

[0044] ​​​​​​​​​​In addition, by convention, in the present application, the terms "upstream" and "downstream" and "inner / below" and "outer / above" are used with reference to the positioning relative to the flow axis of the air flow along the longitudinal axis X-X of the turbine 2. Thus, the cylinder extending along the axis X-X includes an inner surface facing the axis X-X and an outer surface opposite thereto. "Longitudinal" or "longitudinally" means any direction parallel to the axis X-X, and "radial" or "radially" means any direction perpendicular to the axis X-X.

[0045] The low-pressure turbine 7 drives the shaft 11. In a civil aircraft, the reduction gear 12 located upstream of the gas generator 4 transmits the torque applied by the shaft 11 to at least one wheel 13. In the case of two wheels 13, the two wheels 13 can rotate in opposite directions around the longitudinal axis X-X. The shaft 11 and the wheels 13 are arranged in a housing or cover 15, which also houses the drive components for the wheels from the reduction gear 12. Each wheel 13 carries blades to define a propeller.

[0046] The three-flow turbine 2 according to the present invention includes a plurality of propellers 16, 30. Figure 1 A two-flow turbine 2 according to the prior art is shown. This turbine 2 includes a first rotor propeller 16 (hereinafter referred to as "propeller"), which is formed by a plurality of blades 17 distributed around the longitudinal axis X-X and extending radially from the cover 15. As Figure 1 shown, the propeller 16 (also called a fan) is rotatably mounted in the nacelle 3 such that each of its blades 17 is attached to the wheel 13 via the blade root 17A through the cover 15. Each blade 17 includes a free radial end 17B, which is opposite to the root 17A and faces the inner surface 3A of the nacelle 3. The angle of attack of each blade 17 varies from the root 17A to the free radial end 17B. In the present invention, the "angle of attack" means the angle formed between the plane in which the blade is arranged and the longitudinal axis X-X. The rotation of the propeller 16 can accelerate the incoming air flow F0 entering the interior of the nacelle 3.

[0047] Advantageously, the rotor blade 9 and the propeller 16 are connected together by a single body or shaft S as shown in the drawings.

[0048] The turbine 2 also includes one or more annular elements 18, 18'. In particular, as Figure 2 shown, the annular element 18 is arranged radially between the gas generator 4 and the nacelle 3. As Figure 4As shown, the annular element 18' is arranged upstream of a plurality of compressor stages. The length of the annular element 18 extending along the longitudinal axis X-X is substantially similar to the length of the gas generator 4. The annular element 18 is provided upstream with an annular diverter nose 19. The arrangement of the annular element 18 relative to the gas generator 4 defines a first internal annular duct 20, which is delimited by the casing 5B of the compartment 5 of the gas generator 4 and the inner surface 18A of the annular element 18. The arrangement of the annular element 18 relative to the nacelle 3 also defines a second external annular duct 21, which is delimited by the outer surface 18B of the annular element 18 and the inner surface 3A of the nacelle 3. The annular diverter nose 19 divides the incoming air flow F0 entering the nacelle 3 into a first air flow F1 flowing into the internal annular duct 20 and a second air flow F2 flowing into the external annular duct 21. The air flow F1 flowing in the internal annular duct 20 is typically compressed by the multi-stage low-pressure compressor 8 and the high-pressure compressor and then enters the high-pressure combustion chamber. The low-pressure compressor 8 and the high-pressure compressor are formed by a series of rotor blades 9 and stator blades 10. The combustion energy is recovered by the multi-stage high-pressure turbine and the subsequent low-pressure turbine 7 that drive the multi-stage compressor 8 and the upstream propeller 16. The air flow F2 flowing into the external annular duct 21, for its part, participates in providing the thrust of the turbine 2.

[0049] The ratio of the air flow F2 flowing into the external duct 21 to the air flow F1 flowing into the internal duct 20 is generally referred to as the bypass ratio. In a non-limiting manner, the propulsion assembly 1 according to the invention has a variable cycle, i.e., the bypass ratio of the assembly 1 can be modified according to the flight phase. As an example, during the takeoff or landing phase of the aircraft AC, the bypass ratio of the assembly 1 is higher in order to reduce noise and the specific fuel consumption rate.

[0050] In a first preferred embodiment, the assembly 1 further comprises a first stator vane 22 which is arranged upstream of the splitter nose 19 and downstream of the propeller 16. The blades 23 of the stator vane 22 are circumferentially distributed around the longitudinal axis X-X and radially extend over the entire distance D0 between the gas generator 4 and the nacelle 3, such that each vane 22 is attached to the cover 15 by a first inner end 23A and to the inner surface 3A of the nacelle 3 by an outer end 23B opposite the inner end 23A. Alternatively, the blade 23 can be fixed only by one of its radial ends, for example, it can be fixed to the nacelle 3 by its radially outer end and be suspended. Since the blade 23 extends over the entire distance D0 between the cover 15 and the nacelle 3, the blade 23 does not affect the flow velocity of the incoming air flow F0 entering the nacelle 3 and thus does not affect the efficiency and operability of the assembly 1. In addition, the presence of the stator vane 22 greatly reduces the turbulence in the incoming air flow F0 located upstream of the splitter nose 19, such that the angle of attack of the rotor blades 9 and the blades of the stator vanes 10 of the compressor 8 does not change. Therefore, the gas generator 4 is not adversely affected by the change in the bypass ratio of the variable cycle assembly 1. In addition, the reduction of the turbulence on the incoming air flow F0 can reduce the aerodynamic losses of the flow F0, as well as reduce the aerodynamic losses of the air flows F1 and F2 flowing on the edges of the annular element 18, the cover 15, the nacelle 3, etc. in the internal flow passage 20 and the external flow passage 21 respectively by reducing the friction surface.

[0051] In this first embodiment, as Figure 5 shown, the stator vane 22 is fixed such that the angle of attack of each blade 23 does not change. Within the scope of the present invention, the term "fixed vane" defines a set of blades radially mounted around the longitudinal axis X-X, each blade not pivoting around the radial axis along which each blade is arranged.

[0052] In this first preferred embodiment, the assembly 1 further comprises a second stator vane 24 with variable pitch. The blades of the variable pitch vane can rotate around a radial axis (or an axis slightly inclined with respect to the radial axis), and each blade extends along this radial axis. In fact, the blade can rotate around an axis extending from the root to the tip of the blade. The housing is not necessarily straight, so depending on the position of the blade, the blade does not necessarily rotate around a completely radial axis. The introduction of the variable pitch vane can in particular improve the operability of the turbine 2 under a series of flight conditions and reduce the acoustic impact of the turbine 2.

[0053] In Figure 5 , the propulsion assembly 1 further comprises another stator vane 22 located downstream of the propeller 16 and a second propeller 30 located upstream of this another stator vane 22, as Figure 3 shown. The reference numeral 30 refers to another propeller mounted upstream of the first propeller 16.

[0054] As Figure 5 and Figure 6 shown, the stator vanes 24 are arranged radially in the internal duct 20, and an air flow F1 supplied to the gas generator 4 flows in the internal duct 20. The blades 25 of the vanes 24 are distributed radially around the longitudinal axis X-X and extend over a distance D1, which corresponds to the distance between the housing 5A of the gas generator 4 and the annular element 18. Each blade 25 is fixed to the inner surface 18A of the annular element 18 by a radial end 25B and to the housing 5A of the gas generator 4 by a root 25A. Each blade 25 has an angle of attack capable of axially straightening the air flow F1 entering the internal duct 20. As mentioned above, as an alternative, the blade 25 can be fixed only by one of its radial ends. Each blade 25 is pivotable about a radial axis R1 (or an axis slightly inclined with respect to the radial axis), and each blade 25 is arranged along the radial axis R1. As Figure 5 and Figure 6 shown, the stator vanes 24 are arranged at the inlet of the internal duct 20, i.e., downstream of the splitter nose 19 and upstream of the first rotor blade 9 of the low-pressure compressor 8. For example, the variable stator vanes 24 can be inlet guide vanes (IGVs), which have a low curvature and low losses compared to conventional vanes. Selecting variable pitch guide vanes can ensure the operability of the variable cycle assembly 1. On the one hand, the air flow F1 enters the internal duct 20, and most of it is axially straightened by the stator vanes 22; thus, there is no need to have conventional straightening vanes at the inlet of the internal duct 20. On the other hand, the rotor blade 9 located at the most upstream of the low-pressure compressor 8 requires a certain level of co-turbulence, so this co-turbulence does not have to be eliminated by conventional straightening vanes at the inlet of the internal duct 20 that supplies the gas generator 4. Since not all of the turbulence has to be removed, it also facilitates the design of the stator vanes 22.

[0055] In this first specific embodiment, as Figure 5 shown, the external annular duct 21 does not have stator vanes from the annular splitter nose 19 to the radial plane P. This radial plane P is perpendicular to the longitudinal axis X-X and substantially passes through the stator vane 10 located at the most upstream of the low-pressure compressor 8 of the gas generator 4. The air flow F2 entering the external duct 21 does not encounter any vanes during its flow.

[0056] In a variant of this first particular embodiment, the external annular duct 21 includes a third stator vane 26 mounted downstream of the annular splitter nose 19. The stator vane 26 is provided with a plurality of vanes 27 which are circumferentially arranged around the longitudinal axis X-X, and each vane 27 extends in the radial direction over a distance D2 which corresponds to the distance which radially separates the annular element 18 and the nacelle 3. Each vane 27 is attached to the outer surface 18B of the annular element 18 by a vane root 27A and to the inner surface 3A of the nacelle by a radially outer end 27B. The presence of such vanes 26 can straighten the air flow F2 if the upper part of the incoming air flow F0 (the air flow F2 comes from the upper part of the incoming air flow F0) is deflected more axially than the lower part of the incoming air flow F0. In addition, each vane 27 of the vane 26 is preferably fixed and can be internally traversed by cables, in particular for powering the gas generator 4. For example, the stator vane 26 is arranged downstream of the leading edge 25C of the vanes of the stator vane 24 or in line with the leading edge 25C of the vanes of the stator vane 24, the vanes of the stator vane 24 not forming part of the low-pressure compressor 8, and at least a part of the stator vane 26 is also arranged upstream of the leading edge 10A of the vanes of the stator vane 10 of the low-pressure compressor 8 of the gas generator 4 or in line with the leading edge 10A of the vanes of the stator vane 10 of the low-pressure compressor 8 of the gas generator 4, so that the vane 26 is close to the inlet of the external duct 21 to be able to straighten the air flow F2. More specifically, the leading edge of the vanes of the vane 26 can be located downstream of the trailing edge 25D of the vane 24 or at the level of the trailing edge 25D of the vane 24 (or at right angles to the trailing edge 25D of the vane 24), and upstream of the leading edge 10A of the vanes of the vane 10 or at the level of the leading edge 10A of the vanes of the vane 10 (or at right angles to the leading edge 10A of the vanes of the vane 10). For example, the vane 26 can be a vane of the outer guide vane (OGV) type.

[0057] In Figure 7In the second embodiment of the present invention shown, the stator vanes 22 have a variable pitch such that the angle of attack of each blade 23 can be angularly varied. The variable pitch stator vanes 22 include blades 23 that are pivotable about a radial axis R2 (or an axis slightly inclined with respect to the radial axis). The outer annular duct 21 also includes stator vanes 26 mounted downstream of the annular splitter nose 19, and the stator vanes 26 are preferably fixed. The stator vanes 26 are provided with a plurality of blades 27 that are circumferentially arranged about the longitudinal axis X-X, and each blade 27 extends in the radial direction over a distance D2 that corresponds to the distance that radially separates the annular element 18 and the nacelle 3. Only the multi-stage arrangement of the low-pressure compressor 8 is in the inner duct 20, and the multi-stage of the low-pressure compressor 8 particularly includes a first rotor blade 9 and a following first stator blade 10. Thus, the first blade that encounters the air flow F1 after the stator vanes 22 is the rotor blade 9. This avoids introducing variable pitch blades and associated control mechanisms into the element 18 that has very limited space. Instead, the mechanism is moved to the nacelle 3 that has more available space. Thus, a single control mechanism can be used to affect both flows F1, F2.

[0058] The stator vanes 26 can be arranged downstream of the trailing edge 9B of the blades of the rotor blade 9 or at the level of the trailing edge 9B of the blades of the rotor blade 9 (or in a straight line with the trailing edge 9B of the blades of the rotor blade 9). More specifically, the trailing edge of the blades of the vanes 26 can be located downstream of the trailing edge 9B of the blade 9 or at the level of the trailing edge 9B of the blade 9 (or in a straight line with the trailing edge 9B of the blade 9). The stator vanes 26 can be arranged upstream of the leading edge 10A of the blades of the stator vanes 10 of the low-pressure compressor 8 of the gas generator 4 or at the level of the leading edge 10A of the blades of the stator vanes 10 of the low-pressure compressor 8 of the gas generator 4 (or in a straight line with the leading edge 10A of the blades of the stator vanes 10 of the low-pressure compressor 8 of the gas generator 4), such that the vanes 26 are positioned close to the inlet of the outer duct 21 to be able to straighten the secondary flow F2.

[0059] In Figure 6In a variant of this second embodiment as shown, the turbine 2 further includes variable pitch stator blades 24. The stator blades 24 are arranged radially between the annular element 18 and the nacelle 3, within the outer duct 21, through which the air flow F2 flows. The blades 24 are mounted downstream of the annular diverter nose 19. As described above, the variable pitch stator blades 24 can be arranged upstream of the fixed stator blades 26. Alternatively, the stator blades 26 can be the sole elements of the outer duct 21. The blades 25 are distributed radially about the longitudinal axis X-X, and each blade 25 is fixed to the outer surface 18B of the annular element 18 by its root 25A and to the inner surface 3A of the nacelle 3 by its outer end 25B. Each blade 25 is arranged along a radial axis R3 (or an axis slightly inclined with respect to the radial axis), and each blade 25 is pivotable about the radial axis R3. Further, in this second embodiment, the inner annular duct 20 does not have stator blades arranged upstream of the rotor blades 9 of the low-pressure compressor 8. There are no specific blades upstream of the first rotor blade 9 and the stator blade 10 of the compressor 8, enabling the length of the gas generator 4 to be reduced. Then, each blade 25 of the blades 24 is downstream of the leading edge 9A of the blade of the rotor blade 9 of the compressor 8 and upstream of the trailing edge 10B of the blade of the stator blade 10 within the inner duct 20. Each blade 25 of the blades 24 can also be positioned in line with the leading edge 9A of the blade of the rotor blade 9 and upstream of the trailing edge 10B of the blade of the stator blade 10. The blades 25 can also be downstream of the leading edge 9A of the blade of the rotor blade 9 and upstream of the trailing edge 10B of the blade of the stator blade 10. The blades 25 can also be arranged within the inner duct 20 downstream of the leading edge 9A of the blade of the rotor blade 9 of the compressor 8 and in line with the trailing edge 10B of the blade of the stator blade 10.

[0060] As Figure 7 and Figure 8 shown, the blade 22 arranged between the propeller 16 and the annular diverter nose 19 has variable pitch. The incidence of each blade 23 is adjusted to straighten the flow F0 flowing into the nacelle 3, which then divides into two flows F1 and F2. The axial straightening of the flow F0 by the variable pitch blade 22 enables the flow F1 within the inner duct 20 to flow relative to the leading edge 9A of each blade of the first rotor blade 9, eliminating the need for another straightening blade to be present at the inlet of the inner duct 20.

[0061] The installation of the variable pitch blade 24 in the outer duct 21 not only has the advantage of axially straightening the flow F2, but also has the advantage of eliminating the turbulence of the flow F0, which can be generated by changes in the angle of attack of the blades 23 of the blade 22.

[0062] As Figure 9As shown, this dynamic adjustment can also be achieved by means of a blade 25 including a downstream portion 29 and an upstream portion 28. For example, the downstream portion 29 can be a structural element including the trailing edge 25D of the blade 25. This structural element is attached to the nacelle 3 and the annular element 18 by its ends (not shown). The upstream portion 28 can be hollow so that auxiliary devices (such as cables) can pass through the upstream portion 28 along the radial direction to supply the gas generator 4. The upstream portion 28 can rotate about a substantially radial axis, which is collinear with the axis along which the downstream portion 29 extends. The upstream portion 28 includes the leading edge 25C of the blade 25, and the leading edge 25C can pivot upstream according to the angle of attack of the blade 22 to ensure that the flow F2 flowing in the outer duct 21 is rectified axially. For this purpose, the blade 22 and the blade 24 are mechanically connected to each other (not shown). Alternatively, the angle of attack of the blade 23 of the blade 22 varies between 15 degrees and 20 degrees so that the blade 24 can be fixed. Such an arrangement provides the advantage of optimizing the integration of the components in the available space in the duct, which is further reduced due to the annular element possibly requiring a defrosting device.

[0063] The blade 22 also includes a vane 22, the vane 22 being provided with a downstream portion 29 and an upstream portion 28, the downstream portion 29 including a trailing edge 23D and the upstream portion 28 including a leading edge 23C.

[0064] In Figure 3 In the specific embodiment shown, the assembly 1 includes a three-flow turbine 2. Thus, the assembly 1 includes a second rotor propeller 30 (hereinafter referred to as "propeller 30"). The propeller 30 includes a plurality of blades radially extending along the radial direction around the longitudinal axis X-X. The propeller 30 can be a rotor blade, arranged upstream of the plurality of rotor and stator blades forming a compressor stage in the turbine. The propeller is a ducted propeller. As Figure 3 shown, the rotation of the propeller 30 generates an acceleration of the main air flow FP. The nacelle 3 also includes a second annular diverter nose 31 at the upstream end. The diverter nose 31 divides the main air flow FP accelerated by the propeller 30 into an inlet air flow F0 and a third air flow F3. The inlet air flow F0 flows into the space between the nacelle 3 and the cover 15 and is accelerated by the rotation of the propeller 16, and the third air flow F3 flows above the nacelle 3.

[0065] In a variant, as Figure 4 shown, the propeller 30 is a rotor blade, arranged upstream of the plurality of rotor and stator blades forming a compressor stage (present in the three-flow turbine 2).

[0066] Advantageously, the turbine 2 does not include an arm that is located directly downstream of the stator vane 22. This means that the first diverter nose 19 is preferably not connected to an arm and is not located downstream of the leading edge of such an arm and upstream of the trailing edge of these arms. Typically, such an arm extends into the air flow F0 and into the air flows F1, F2.

Claims

1. An aircraft propulsion assembly, the propulsion assembly (1) comprising a three - flow turbine (2) and a nacelle (3) surrounding the turbine (2), the turbine (2) comprising: - A gas generator (4), the gas generator comprising at least one compressor (8), a combustion chamber, and a turbine (7), the gas generator (4) being arranged along a longitudinal axis (X - X), - A first propeller (16), the first propeller being mounted within the nacelle (3) and around the longitudinal axis (X - X), and being configured to accelerate an incoming air flow (F0) entering the nacelle (3), - At least one annular element (18), the at least one annular element being arranged radially between the gas generator (4) and the nacelle (3) and defining a first internal annular duct (20) for supplying the gas generator (4) and a second external annular duct (21) for the nacelle (3), the annular element (18) comprising upstream a first annular splitter nose (19), the first annular splitter nose being configured to divide the incoming air flow (F0) into a first air flow (F1) flowing in the first duct (20) and a second air flow (F2) flowing in the second external annular duct (21), - A second propeller (30), the second propeller being mounted upstream of the nacelle (3) and around the longitudinal axis, and being configured to accelerate a main air flow (FP), the nacelle (3) comprising upstream a second annular splitter nose (31), the second annular splitter nose being configured to divide the main air flow (FP) into the incoming air flow (F0) flowing into the nacelle and a third air flow (F3) flowing around the nacelle (3), characterized in that the propulsion assembly further comprises: - A first stator vane (22), the first stator vane extending radially between the housing (5B) of the gas generator (4) and the nacelle (3), upstream of the first annular splitter nose (19) and downstream of the first propeller (16), and - A second stator vane (24), the second stator vane extending radially between the housing (5B) of the gas generator (4) and the annular element (18), downstream of the first splitter nose (19) and upstream of the first rotor vane (9) of the at least one compressor (8) of the gas generator (4), and / or extending radially between the annular element (18) and the nacelle (3), downstream of the first annular splitter nose (19), At least one of the first stator vane (22) and the second stator vane (24) is a variable pitch vane or comprises at least one variable pitch portion.

2. The propulsion assembly (1) according to claim 1, wherein, The second stator vane (24) extends radially between the gas generator (4) and the annular element (18), and the second outer annular duct (21) has no stator vanes from the first annular diverter nose (19) to a plane (P) that is perpendicular to the longitudinal axis (X-X) and substantially passes through the first stator vane (22) of the at least one compressor (8) of the gas generator (4).

3. The propulsion assembly (1) according to claim 1, wherein, The second stator vane (24) extends radially between the gas generator (4) and the annular element (18), and the second outer annular duct (21) includes a third stator vane (26) located downstream of the first annular diverter nose (19).

4. The propulsion assembly (1) according to the previous claim, wherein, The third stator vane (26) is located downstream of or in line with the leading edge (25C) of the blade (25) of the second stator vane (24), and is located upstream of or in line with the leading edge (9A) of the blade of the first stator vane (9) of the at least one compressor (8) of the gas generator (4).

5. The propulsion assembly (1) according to claim 1, wherein, The second stator vane (24) extends radially between the annular element (18) and the nacelle (3), and the first inner annular duct (20) has no stator vanes upstream of the first rotor vane (9) of the at least one compressor (8) of the gas generator (4).

6. The propulsion assembly (1) according to the preceding claim, wherein, The second stator vane (24) is located downstream of or in line with the leading edge (9A) of the blade of the first rotor vane (9) of the at least one compressor (8) of the gas generator (4), and is located upstream of or in line with the trailing edge (10B) of the blade of the first stator vane (10) of the at least one compressor (8).

7. The propulsion assembly (1) according to claim 5 or 6, wherein, The first stator vane (22) and / or the second stator vane (24) includes blades (23, 25), the upstream portion (28) of the blades includes leading edges (23C, 25C) that can rotate about a substantially radial axis, and the downstream portion (29) of the blades includes fixed trailing edges (23D, 25D).

8. The propulsion assembly (1) according to any one of the preceding claims, wherein, The at least one first propeller (16) and the first rotor vane (9) are connected to a single shaft (S), preferably connected to the single shaft (S) through a mechanical reduction gear.

9. An aircraft comprising at least one propulsion assembly (1) according to any one of the preceding claims.

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