FR3131601A1 - PROPULSION ASSEMBLY FOR AN AIRCRAFT - Google Patents

Abstract

The propulsion assembly (1) for an aircraft comprises a nacelle (3) surrounding a multi-flow turbomachine (2) which comprises a gas generator (4), a propeller (16) accelerating an air flow (FO) in the nacelle (3), an annular element (18) between the generator (4) and the nacelle (3) defining a first (20) and a second stream (21), comprising a nozzle (19) separating the flow (F0) into one air flow (F1) in the first stream (20) and in an air flow (F2) in the second stream (21), the assembly (1) comprising a stator blade (22) mounted between the blade (19) and the propeller (16), and a stator blade (24) between the generator (4) and the annular element (18), mounted between the blade (19) and a rotor blade of a compressor (8) of the generator (4), or between the annular element (18) and the nacelle (3). Figure for abstract: Figure 7

Description

PROPULSIVE ASSEMBLY FOR AN AIRCRAFT Technical field of the invention

The present invention relates to the general field of aeronautics. It is more particularly aimed at a propulsion assembly for an aircraft comprising a multi-flow turbomachine and a nacelle. The invention also relates to an aircraft comprising such a propulsion assembly.

Technical background

Conventionally, a propulsion assembly comprises a nacelle surrounding a turbomachine which makes it possible to generate the thrust necessary for the propulsion of an aircraft, whether it is an airliner or a fighter plane, etc. For this purpose, the turbomachine successively comprises at least one compressor which compresses a flow of air entering the nacelle, a combustion chamber in which the previously compressed air is mixed with fuel then ignited in order to generate a flow of hot gas propulsion, and at least one turbine which is rotated by this flow of hot gas, the turbine being connected by a shaft to the compressor. These elements form the engine also called gas generator. The hot gas flow then escapes through a nozzle at the outlet of the turbomachine. A rotor blade also called a fan is generally mounted upstream of the gas generator so as to accelerate the primary air flow.

There are also multi-flow turbomachines in which an annular separator is mounted between the nacelle and the gas generator so as to separate the flow entering the nacelle into a primary air flow flowing into the gas generator and a cold secondary air flow which circulates in the vein formed by the space between the nacelle and the separator. The main advantage of these turbomachines is that they consume less fuel and are less noisy.

The use of multi-flow turbomachines is characterized by their dilution rate which corresponds to the ratio of the mass of the secondary/tertiary flow to the mass of the primary flow. This dilution rate can also vary depending on the flight phases of the aircraft, particularly in variable cycle turbomachines. However, variations in the dilution rate can lead to secondary/tertiary flow losses and therefore to reductions in the efficiency and operability of the turbomachine.

The present invention aims to overcome this drawback by proposing an architecture allowing both the straightening of the air flows entering the turbomachine and the minimization of the impact of changes in dilution rate on the gas generator.

To this end, the invention relates to a propulsion assembly for an aircraft, this propulsion assembly comprising a multi-flow turbomachine and a nacelle which surrounds the turbomachine, said turbomachine comprising:
- a gas generator comprising at least one compressor, a combustion chamber and a turbine, said gas generator being arranged along a longitudinal axis,
- at least a first propeller mounted inside the nacelle and around the longitudinal axis and configured to accelerate an air inlet flow entering the nacelle,
- at least one annular element arranged radially between the gas generator and the nacelle and defining a first internal annular vein supplying the gas generator, and a second external annular vein with the nacelle, said annular element comprising upstream a first nozzle of seperation annular which is configured to separate said air inlet flow into a first air flow flowing in said first vein and into a second air flow flowing in said second external annular vein,

The propulsion assembly being characterized in that it further comprises:
- a first stator blade extending radially between a casing of the gas generator and the nacelle, upstream of the first annular separation nozzle and downstream of said first propeller, and
- a second stator blade extending radially between a casing of the gas generator and the annular element, downstream of said first separation nozzle and upstream of a first rotor blade of said at least one compressor of the gas generator, and/or between the annular element and the nacelle, downstream of said first annular separation nozzle,

and in that at least one of said first and second stator blades is variable pitched or comprises a variable pitch portion.

Thus, thanks to the invention, the straightening of the air flows entering the nacelle is carried out upstream of the separator so that the blades present in the veins only have the function of protecting the turbomachine from changes in the dilution rate. Such an architecture makes it possible to simplify the construction and assembly of the different blades present in limited spaces such as veins. It also allows more freedom in the positioning of the variable stator blade (which is entirely variable-pitch, or which only includes a variable-pitch portion and therefore another fixed portion) depending on the space available. For example, the available space is often very limited at the level of the first separation nozzle (because the thickness available in this area is necessarily lower), so it may be more interesting to put it in the nacelle.

The propulsion assembly may also have one or more of the following characteristics, taken alone or in combination with each other:
- the first stator blade has variable pitch,
- said second external annular vein is devoid of stator blade from said first annular separation nozzle to a plane perpendicular to said longitudinal axis and passing substantially through a first stator blade of said at least one compressor of the gas generator,

-- particularly in the above configuration, when the first stator blade has variable pitch, the second stator blade can be entirely fixed and not have variable pitch; in this configuration in fact, it is not necessarily necessary to have two consecutive stator blades with variable pitch,
- said second external annular vein comprises a third stator blade downstream of said first annular separation nozzle,

-- particularly in the above configuration, when the first stator blade has variable pitch and the second stator blade is entirely fixed, the third stator blade may comprise a portion with variable pitch and therefore a fixed portion; in this configuration also, it is not necessarily necessary to have two consecutive stator blades with variable pitch,
- the third stator blade is located downstream or in line with the leading edges of the blades of said second stator blade, and upstream or in line with the leading edges of the blades of a first rotor blade of said at least one gas generator compressor,
- the second stator blade extends radially between the annular element and the nacelle, and said first internal annular vein is devoid of stator blade upstream of a first rotor blade of said at least one compressor of the gas generator ,

-- in particular in the preceding configuration, the second stator blade is preferably with variable pitch,
- the second stator blade is located downstream or to the right of the leading edges of the blades of said first rotor blade of said at least one compressor of the gas generator, and upstream or to the right of the trailing edges of the blades of a first stator blade of said at least one compressor,
- the first stator blade and/or said second stator blade comprises blades of which an upstream portion comprises a leading edge movable in rotation around a substantially radial axis, and a downstream portion of which comprises a fixed trailing edge,

-- the first stator blade and/or the second stator blade comprises blades of which a downstream portion has a trailing edge movable in rotation around a substantially radial axis, and an upstream portion has a fixed trailing edge, and
- said at least one first propeller and said first rotor blade are connected to the same shaft, preferably via a mechanical speed reducer.

The present invention also relates to an aircraft, in particular a transport aircraft, comprising a propulsion assembly such as that mentioned above.

Brief description of the figures

Other characteristics and advantages of the invention will appear during reading of the detailed description which follows, for the understanding of which we will refer to the appended drawings in which:

there shows a schematic longitudinal section of a propulsion assembly comprising a dual-flow turbomachine;

there represents a double longitudinal schematic section of a civilian propulsion assembly comprising a dual-flow turbomachine;
there represents a double longitudinal schematic section of a propulsion assembly, comprising a dual-flow turbomachine;

there represents a double longitudinal schematic section of a propulsion assembly comprising a turbomachine with several flows;

there represents a double longitudinal schematic section of a propulsion assembly according to a first embodiment of the invention;

there represents a double longitudinal schematic section of a propulsion assembly according to a variant of this first embodiment of the invention;

there represents a double longitudinal schematic section of a propulsion assembly according to a second embodiment of the invention;

there represents a double longitudinal schematic section of a propulsion assembly according to a variant of the second embodiment of the invention; And

there represents a schematic radial section of a stator blade with variable pitch according to the second embodiment of the invention.

Detailed description of the invention

The propulsion assembly 1 for an aircraft (hereinafter “assembly 1”), whether civil or other, is represented schematically in Figures 1 to 8. The assembly 1 comprises a turbomachine 2 which is arranged along 'a longitudinal axis XX. The turbomachine 2 can be with several flows in the context of a civil aircraft, as shown in Figures 1 and 3, or in the context of a different aircraft as shown in . Although not exclusively, the turbomachine 2 can be surrounded by a nacelle 3.

The turbomachine 2 conventionally comprises a gas generator 4 comprising at least one compressor 8, a combustion chamber and a turbine 7. As shown in the , the gas generator 4 forms a compartment 5 in which are preferably arranged a high pressure body 6 formed of a high pressure compressor, a high pressure combustion chamber and a high pressure turbine, not detailed in the figures, and a low pressure body comprising 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 and low pressure compressors 8 are formed of an alternation rotor blades 9 and stator 10 arranged successively from upstream to downstream around the longitudinal axis XX. In the present invention, rotor blade means a wheel on which vanes or blades are fixed and which rotates around the longitudinal axis XX. By stator blade is also meant a wheel on which vanes or blades are fixed which do not rotate around the longitudinal axis XX.

Furthermore, by convention, in the present application, the terms "upstream" and "downstream", and "internal/below" and "external/above" are used with reference to positioning relative to a flow axis of flows. of air along the longitudinal axis X-X of the turbomachine 2. Thus, a cylinder extending along the axis “Longitudinal” or “longitudinal” means any direction parallel to the X-X axis, and “radially” or “radial” any direction perpendicular to the X-X axis.

The low pressure turbine 7 drives a shaft 11. In a civil aircraft, a reduction gear 12 with an epicyclic gear train, located upstream of the gas generator 4, transmits the torque exerted by the shaft 11 to at least one disc 13. In the case of two discs 13, they can rotate in opposite directions around the longitudinal axis XX. The shaft 11 and the disc(s) 13 are arranged in a casing or cover 15 which also houses the drive members of the disc(s) from the reducer 12. Each of the discs 13 carries blades to form a propeller.

In the case of the , the turbomachine 2 comprises a first rotor propeller 16 (hereinafter “propeller”) which is formed of a plurality of blades 17 distributed around the longitudinal axis XX and extending in radial directions from the cover 15.

Advantageously,

The rotor blade 9 and the propeller 16 are connected to the same drum or shaft 16, as illustrated in the drawings.

The turbomachine 2 also includes one or more annular elements 18, 18'.

As shown in Figures 2 to 4 in particular, the annular element 18 can be arranged radially between the gas generator 4 and the nacelle 3. The annular element 18 extends along the longitudinal axis X-X over a length substantially similar to the length of the gas generator 4. The annular element 18 is provided, upstream, with an annular separation nozzle 19. The arrangement of this annular element 18 relative to the gas generator 4 defines a first internal annular vein 20 delimited by a casing 5B of the compartment 5 of the gas generator 4 and an internal 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 vein 21 which is delimited by an external surface 18B of the annular element 18 and the internal surface 3A of the nacelle 3. The annular separation nozzle 19 separates the air inlet flow F0 entering the nacelle 3 into a first air flow F1 which flows into the internal annular vein 20 and into a second air flow F2 which flows into the annular vein external 21. The air flow F1 circulating in the internal annular vein 20 is conventionally compressed by stages of the low pressure 8 and high pressure compressors formed by the succession of rotor blades 9 and stator 10 before entering the high pressure combustion chamber. The combustion energy is recovered by the high pressure then low pressure turbine stages 7 which drive the compressor stages 8 and the rotation of the propeller 16 upstream. The air flow F2 which flows in the external annular vein 21 participates, for its part, in providing the thrust of the turbomachine 2.

The ratio between the air flow F2 flowing in the external vein 21 and the air flow F1 flowing in the internal vein 20 is generally called dilution ratio. In a non-limiting manner, the propulsion assembly 1 according to the invention has a variable cycle, that is to say that depending on the flight phases, the dilution rate of the assembly 1 can be modified. For example, the dilution rate of assembly 1 during a take-off or landing phase of the AC aircraft is high so as to reduce noise and specific fuel consumption.

In a first preferred embodiment, the assembly 1 further comprises a first stator blade 22 which is arranged upstream of the separation nozzle 19 and downstream of the propeller 16. The blades 23 of the blade 22 stator are distributed circumferentially around the longitudinal axis cover 15 and by an external end 23B which is opposite the internal end 23A to the internal surface 3A of the nacelle 3. Alternatively, the blades 23 could be fixed only via one of their radial ends, and could for example be suspended by being fixed by their radially external ends to the nacelle 3. As the blades 23 extend over the entire distance D0 between the cover 15 and the nacelle 3, they do not influence the flow rate air inlet F0 which enters the nacelle 3 therefore on the efficiency and operability of the assembly 1. Furthermore, the presence of the stator vane 22 makes it possible to very significantly reduce the turbulence of the flow air inlet F0 upstream of the separation nozzle 19 so that the incidence of the blades of the rotor 9 and stator 10 blades of compressor 8 is not modified. The gas generator 4 therefore does not suffer any undesirable effects due to changes in the dilution rate of the variable cycle assembly 1. In addition, the reduction in turbulence on the air inlet flow F0 makes it possible to reduce the aerodynamic losses of this flow F0 as well as of the air flows F1 and F2 circulating respectively in the internal 20 and external 21 veins on the edges of the annular element 18, of the hood 15, of the nacelle 3, etc. by reducing friction surfaces.

In this first embodiment, the stator blade 22 is fixed so that the incidence of each of the blades 23 does not vary, as shown in the figure. . In the context of the present invention, fixed blade means all the blades mounted radially around the longitudinal axis XX and each of the blades does not pivot around the radial axis along which it is arranged.

In this first preferred embodiment, assembly 1 also includes a second stator blade 24 which has variable pitch. The blades of a variable-pitch blade can rotate around a radial axis R1 or in an inclined manner relative to this axis, according to which each of the blades is arranged. In practice, the blades can rotate along an axis which extends from the root to the head of the blade. We see that the casing is not necessarily rectilinear so, depending on the position of the blade, it might not necessarily rotate along a perfectly radial axis. The introduction of variable-pitch blades makes it possible in particular to improve the operability of turbomachine 2 for a range of flight conditions and to reduce its acoustic impact.

In the , the propulsion assembly 1 further comprises another stator blade 22 upstream of the propeller 16, and a second propeller 30 upstream of this other stator blade 22. The reference 30 designates another propeller mounted upstream of the first propeller 16.

As shown in Figures 5 and 6, the stator blade 24 is arranged radially in the internal vein 20 in which flows the air flow F1 which supplies the gas generator 4. The blades 25 of the blade 24 are distributed radially around the longitudinal axis XX and extend over a distance D1 which corresponds to the distance between the casing 5A of the gas generator 4 and the annular element 18. Each of the blades 25 is fixed to the internal surface 18A of the annular element 18 by a radial end 25B and to the casing 5A of the gas generator 4 by a foot 25A. Each of the blades 25 has an incidence making it possible to axially straighten the air flow F1 entering the internal vein 20. As mentioned in the above, as a variant, the blades 25 could be fixed only by means of a single one of their radial ends. Each of the blades 25 can pivot around a radial axis R1 (or in a slightly inclined manner relative to such an axis, as mentioned in the above) according to which it is arranged. As shown in Figures 5 and 6, the stator blade 24 is arranged at the entrance to the internal vein 20, that is to say downstream of the separation nozzle 19 and upstream of the first rotor blade 9 of the low pressure compressor 8. For example, the stator vane 24 with variable pitch can be an inlet guide vane with variable pitch (IGV or “ Inlet Guide Vane ” in English) which has a low curvature and a low loss compared to conventional blading. The choice of a variable-pitch guide vane ensures the operability of assembly 1 with a variable cycle. On the one hand, the air flow F1 enters the internal vein 20 while being mostly straightened axially by the stator blade 22; it is therefore not necessary to have a conventional straightening blade at the inlet of the internal vein 20. On the other hand, the rotor blade 9 most upstream of the low pressure compressor 8 requires a certain level of co-turbulences which must therefore not be eliminated by a conventional straightening vane at the inlet of the internal vein 20 supplying the gas generator 4. All turbulence must not be removed, the design of the vane 22 of stator is also made easier.

In this first particular embodiment, the external annular vein 21 is devoid of stator blade from the annular separation nozzle 19 to a radial plane P as shown in the . This radial plane P is perpendicular to the longitudinal axis XX and passes substantially through the stator blade 10 most upstream of the low pressure compressor 8 of the gas generator 4. The air flow F2 entering the external vein 21 does not therefore encounters no blades during its flow.

In a variant of this first particular embodiment illustrated in , the external annular vein 21 comprises a third stator vane 26 mounted downstream of the annular separation nozzle 19. The stator blade 26 is provided with a plurality of blades 27 arranged circumferentially around the longitudinal axis X-X and each extending in a radial direction over a distance D2 which corresponds to the distance radially separating the annular element 18 and the nacelle 3. Each of the blades 27 is fixed to the external surface 18B of the annular element 18 by a blade root 27A and to the internal surface 3A of the nacelle by an external radial end 27B. The presence of such a vane 26 makes it possible to straighten the air flow F2 if the upper part of the air inlet flow F0, from which the air flow F2 comes, is more deviated axially than its lower part. Furthermore, each of the blades 27 of the blade 26 is preferably fixed and can be crossed internally by cables used in particular for the electrical supply of the gas generator 4. For example, the stator blade 26 is arranged downstream or to the right of the leading edges 25C of the blades of the stator blade 24 which is not here part of the low pressure compressor 8. At least part of the stator blade 26 is also arranged upstream or at right of the leading edges 10A of the blades of the stator blade 10 of the low pressure compressor 8 of the gas generator 4 so that the blade 26 is located close to the inlet of the external vein 21 to allow straightening of the air flow F2. More precisely, the leading edges of the blades of the blade 26 can be located downstream or to the right of the leading edges 25C of the blade 24, and upstream or to the right of the leading edges 10A of the blades of the blade 10. The blade 26 can be of the OGV type (Outer Guide Vane) For example.

In a second embodiment of the invention shown in , the stator blade 22 has variable pitch so that the incidence of each of the blades 23 can be angularly modified. The variable-pitch stator vane 22 comprises blades 23 capable of pivoting around the radial axis R2 or in a slightly inclined manner relative to this axis. The external annular vein 21 also comprises a stator vane 26, preferably fixed, mounted downstream of the annular separation nozzle 19. The stator blade 26 is provided with a plurality of blades 27 arranged circumferentially around the longitudinal axis X-X and each extending in a radial direction over a distance D2 which corresponds to the distance radially separating the annular element 18 and the nacelle 3. In the internal vein 20 are arranged only stages of the low pressure compressor 7 comprising in particular the first rotor blade 9 followed by the first stator blade 10. The first blade encountered by the air flow F1 after the The stator blade 22 is thus a rotor blade 9. This makes it possible to avoid introducing a variable-pitch blade and the associated control mechanism in the element 18, which has a very restricted space, and to instead move it in nacelle 3 where there is more space. A single control mechanism can thus be used to influence two flows. The stator blade 26 can be arranged downstream or to the right of the leading edges 9A of the blades of the rotor blade 8. More precisely, the leading edges of the blades of the blade 26 can be located downstream of or to the right of the leading edges 9A of the blade 9. The stator blade 26 can also be arranged upstream or to the right of the leading edges 10A of the blades of the stator blade 10 of the low pressure compressor 8 of the gas generator 4 so that the The blade 26 is located close to the inlet of the external vein 21 to allow straightening of the air flow F2.

In a variant of this second embodiment shown in , the turbomachine 2 also includes the stator blade 24 with variable timing. The stator blade 24 is arranged radially between the annular element 18 and the nacelle 3, inside the external vein 21 in which the air flow F2 flows. The vane 24 is mounted downstream of the annular separation nozzle 19. The variable-pitch stator blade 24 could be arranged upstream of the fixed stator blade 26 described in the above. It can, as a variant, be the only element of the external vein 21. The blades 25 are distributed radially around the longitudinal axis XX and each is fixed, by its foot 25A, to the external surface 18B of the annular element 18 and, through its external end 25A, to the internal surface 3A of the nacelle 3. Each of the blades 25 is arranged along a radial axis R3 or in a slightly inclined manner relative to this axis, around which it can pivot. Furthermore, in this second embodiment, the internal annular vein 20 is devoid of stator blade arranged upstream of the rotor blade 9 of the low pressure compressor 8. The absence of particular blade upstream of the first blades rotor 9 and stator 10 of the compressor 8 makes it possible to reduce the length of the gas generator 4. Each of the blades 25 of the blade 24 is then located downstream of the leading edges 9A of the blades of the rotor blade 9 of the compressor 8 in the internal vein 20 and upstream of the trailing edges 10B of the blades of the stator blade 10 of the compressor 8. Each of the blades 25 of the blade 24 can also be located to the right of the leading edges 9A blades of the rotor blade 9 and upstream of the trailing edges 10B of the blades of the stator blade 10. The blades 25 can also be located downstream of the leading edges 9A of the blades of the rotor blade 9 and upstream of the trailing edges 10B of the blades of the stator blade 10. They can, in addition, be arranged downstream of the leading edges 9A of the blades of the rotor blade 9 of the compressor 8 in the vein internal 20, and to the right of the trailing edges 10B of the blades of the stator blade 10.

As shown in Figures 7 and 8, the blade 22 which is arranged between the propeller 16 and the annular separation nozzle 19 has variable pitch. The incidence of each of the blades 23 is adjusted so as to straighten the flow F0 flowing into the nacelle 3 before being separated into two flows F1 and F2. The axial straightening of the flow F0 by the blade 22 with variable pitch allows the flow F1 in the internal vein 20 to flow facing the leading edge 9A of each blade of the first rotor blade 9 which makes the presence of another useless straightening blade at the entrance to the internal vein 20.

The installation of the vane 24 with variable pitch in the external vein 21 has the advantage of not only axially straightening the flow F2 but also of eliminating the turbulence of the flow F0 which can be generated by variations in the incidence of the blades. 23 of the blade 22.

As shown on the , this dynamic adjustment can also be carried out by blades 25 comprising a downstream portion 29 and an upstream portion 28. The downstream portion 29 can be, for example, a structural element which includes the trailing edge 25D of the blade 25 This structural element is fixed, by its ends (not shown), to the nacelle 3 and to the annular element 18. This upstream part 28 can be hollow so that easements, such as cables, can pass through it in the radial direction at the end of supply of the gas generator 4. The upstream part 28 is movable in rotation around a substantially radial axis common with the axis along which the downstream part 29 extends. The upstream part 28 comprises the leading edge 25C of the blade 25 and which can pivot according to the incidence of the blade 22 upstream so as to ensure axial straightening of the flow F2 flowing in the external vein 21. To do this, the The blade 22 and the blade 24 are connected to each other mechanically (not shown). Alternatively, the changes in incidence of the blades 23 of the blade 22 are between 15 degrees and 20 degrees so that the blade 24 can be fixed. Such an arrangement offers the advantage of optimizing the integration of elements in the space available in the veins, this space being all the more reduced as the annular element may require a defrosting device.

The blade 22 also comprises blades 22 provided with a downstream portion 29 comprising a trailing edge 23D and an upstream portion 28 comprising a leading edge 23C.
In a particular embodiment illustrated in , assembly 1 comprises a turbomachine 2 with several flows. Assembly 1 then comprises a second rotor propeller 30 (hereinafter “propeller 30”). The propeller 30 comprises a plurality of blades extending radially around the longitudinal axis XX in radial directions. This propeller 30 can be a rotor blade arranged upstream of a plurality of rotor and stator blades forming the compressor stages present in the turbomachine 2. As shown in the , the rotation of the propeller 30 generates an acceleration of a main air flow FP. The nacelle 3 also includes, on an upstream end, a second annular separation nozzle 31. This separation nozzle 31 makes it possible to separate the main air flow FP accelerated by the propeller 30 into the air inlet flow F0 which flows into the space between the nacelle 3 and the cover 15 and which is accelerated by the rotation of the propeller 16 and in a third air flow F3 which flows above the nacelle 3.

Advantageously, the turbomachine 2 does not include arms directly downstream of the stator blade 22. This means that the first nozzle 19 is preferably not connected to arms and is not located downstream of edges of attack of these arms and upstream of the trailing edges of these arms. In general, this type of arm extends through the air inlet flow F0 and also through the air flows F1 and F2.

Claims (10)

Propulsion assembly for an aircraft, this propulsion assembly (1) comprising a turbomachine (2) with several flows and a nacelle (3) which surrounds the turbomachine (2), said turbomachine (2) comprising:
- a gas generator (4) comprising at least one compressor (8), a combustion chamber and a turbine (7), said gas generator (4) being arranged along a longitudinal axis (X-X),
- at least a first propeller (16) mounted inside the nacelle (3) and around the longitudinal axis (X-X) and configured to accelerate an air inlet flow (FO) entering the nacelle ( 3),
- at least one annular element (18) arranged radially between the gas generator (4) and the nacelle (3) and defining a first internal annular vein (20) supplying the gas generator (4), and a second vein (21) external annular with the nacelle (3), said annular element (18) comprising upstream a first separation nozzle (19) annular which is configured to separate said air inlet flow (F0) into a first air flow (F1) flowing in said first vein (20) and into a second air flow (F2) flowing in said second external annular vein (21),
characterized in that it further comprises:
- a first stator blade (22) extending radially between a casing (5B) of the gas generator (4) and the nacelle (3), upstream of the first annular separation nozzle (19) and downstream of said first propeller (16), and
- a second stator blade (24) extending radially between the annular element (18) and the nacelle (3), downstream of said first annular separation nozzle (19),
and in that said first stator blade (22) has variable pitch or comprises a variable pitch portion, this first stator blade (22) being located directly upstream of said first annular separation nozzle (19). Propulsion assembly (1) according to claim 1, wherein said second stator vane (24) extends radially between the gas generator (4) and the annular element (18), and said second outer annular vein (21). is devoid of stator blade from said first annular separation nozzle (19) to a plane perpendicular (P) to said longitudinal axis (X-X) and passing substantially through a first stator blade (10) of said at least one compressor ( 8) of the gas generator (4). Propulsion assembly (1) according to claim 1, wherein said second stator vane (24) extends radially between the gas generator (4) and the annular element (18), and said second outer annular vein (21). comprises a third stator vane (26) downstream of said first annular separation nozzle (19). Propulsion assembly (1) according to the preceding claim, in which said third stator vane (26) is located downstream or to the right of the leading edges (25C) of the blades (25) of said second stator vane (24), and upstream or to the right of the leading edges (9A) of the blades of a first rotor blade (9) of said at least one compressor (8) of the gas generator (4). Propulsion assembly (1) according to claim 1, in which said second stator vane (24) extends radially between the annular element (18) and the nacelle (3), and said first internal annular vein (20) is devoid stator blade upstream of a first rotor blade (9) of said at least one compressor (8) of the gas generator (4). Propulsion assembly (1) according to the preceding claim, in which said second stator blade (24) is located downstream or to the right of the leading edges (9A) of the blades of said first rotor blade (9) of said at least one compressor (8) of the gas generator (4), and upstream or to the right of the trailing edges (10B) of the blades of a first stator blade (10) of said at least one compressor (8). Propulsion assembly (1) according to claim 5 or 6, in which said first stator blade (22) and/or said second stator blade (24) comprises blades (23, 25) including an upstream portion (28) comprising a leading edge (23C, 25C) is movable in rotation around a substantially radial axis, and of which a downstream portion (29) comprising a trailing edge (23D, 25D) is fixed. Propulsion assembly (1) according to any one of the preceding claims, in which said at least one first propeller (16) and said first rotor blade (9) are connected to the same shaft (S), preferably via a mechanical speed reducer. Propulsion assembly (1) according to any one of the preceding claims, in which it comprises a second propeller (30) mounted upstream of the nacelle (3) and around the longitudinal axis (X-X) and configured to accelerate a main flow of air (FP), said nacelle (3) comprising upstream a second annular separation nozzle (31) which is configured to separate said main air flow (FP) into said air inlet flow (F0) flowing into the nacelle (3), and in a third air flow (F3) flowing around the nacelle (3). Aircraft comprising at least one propulsion assembly (1) according to any one of the preceding claims.