Variable pitch blade of an unfaired aeronautical propulsion system
The variable-pitch blade design with optimized thickness ratios and stator configurations addresses propulsive efficiency and noise reduction issues in unfaired turbomachines by enhancing robustness to airflow variations and reducing noise levels.
Patent Information
- Authority / Receiving Office
- FR · FR
- Patent Type
- Patents
- Current Assignee / Owner
- SAFRAN AIRCRAFT ENGINES SAS
- Filing Date
- 2023-04-28
- Publication Date
- 2026-05-15
AI Technical Summary
Unfaired turbomachines face challenges in maintaining propulsive efficiency and reducing noise levels, particularly during takeoff and landing phases, due to non-uniform airflows and high clipping rates that affect blade performance and generate transverse forces.
A variable-pitch blade design with specific thickness ratios and structural features, including a collection zone thickness relative to the leading edge thickness, optimized for different flight phases, and a stator with non-homogeneous blade lengths and clipping ratios, to enhance robustness to airflow variations and reduce noise.
The blade design improves aerodynamic performance and reduces noise generation across various flight conditions, maintaining efficient propulsion without significant efficiency loss.
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Abstract
Description
Title of the invention: Variable pitch blade for an unfaired aeronautical propulsion system. Technical field of the invention
[0001] The invention relates to a variable pitch blade of an unfaired aeronautical propulsion system, as well as a turbomachine comprising such blades. Prior art
[0002] The search for minimizing polluting emissions related to air transport involves, in particular, improving the efficiency of propulsion systems, and more specifically the propulsive efficiency which characterizes the efficiency with which the energy communicated to the air passing through the engine is converted into useful thrust effort for propulsion.
[0003] A known principle for improving propulsive efficiency involves modifying the elements of the low-pressure system of the thrusters, which contribute directly to thrust generation, in combination with other known turbomachine elements, such as the high-pressure casing and the combustion chamber. These elements typically include a low-pressure turbine, a low-pressure transmission system driving a fan, and a secondary flow straightener guiding the fan's flow. One solution aims to reduce the fan's compression ratio, thereby decreasing the flow velocity at the engine outlet and the associated kinetic energy losses.
[0004] One of the main consequences of this decrease in flow velocity at the engine outlet is that a greater mass flow rate of air must pass through the low-pressure section (or secondary flow) to ensure a given level of thrust. This therefore leads to an increase in the engine's bypass ratio (or BPR), defined as the ratio between the mass flow rate passing through the secondary flow (cold flow) and the mass flow rate passing through the primary flow (hot flow), which notably supplies the combustion chamber.
[0005] The very large fan diameters required for such dilution would lead to a very large increase in the dimensions of the casing and the nacelle, leading to the development of unfaired turbomachines to overcome this problem.
[0006] Figure 1 illustrates such an unducted turbomachine 1, of the type designated by the acronym USF (for Unducted Single Fan). The turbomachine 1 comprises an upstream propeller wheel 3 (or rotor 3) with variable pitch and a downstream stator wheel 5 (or stator 5) with fixed or variable pitch.
[0007] The terms "upstream" and "downstream" are understood in relation to a main axis X of the turbomachine 1, which coincides with the axis of rotation of the rotor 3, and in relation to a normal airflow direction during the operation of the turbomachine 1.
[0008] The turbomachine 1 is in a "puller" type configuration, i.e. with the rotor 3 and the stator 5 upstream of the turbine, and is generally mounted on the fuselage or under a wing of the aircraft by means of a mast 7. Alternatively, a turbomachine may be in a so-called "pusher" configuration with the propellers downstream of the turbine for mounting at the rear of the aircraft.
[0009] The rotor 3 and the stator 5 each comprise a plurality of blades 10 distributed circumferentially around the main axis X.
[0010] The rotor 3 is a movable propeller rotating around the X axis so as to drive the air and generate a primary flow Fp sent into the turbomachine and a secondary flow Fs flowing outside the turbomachine.
[0011] This rotor 3 is a variable-pitch, low-speed propeller, meaning that each blade 10 comprises a movable blade that rotates about a radial axis, so as to modify its pitch angle to maximize thrust depending on the flight state (takeoff, cruise, landing, etc.). The slow rotational speed maximizes propulsion energy efficiency. Such an unshrouded turbomachine does not include an external casing surrounding the secondary flow Fs. Only the primary flow Fp is guided within a central casing of the turbomachine.
[0012] The acoustic attenuation of the noise generated by the turbomachine is significantly reduced, particularly for low frequencies (below 500 Hz). Permissible noise levels are severely constrained, especially during takeoff and landing phases, which necessitates optimization of the blades to reduce noise levels, as conventional attenuation solutions for enclosed turbomachinery are not available.
[0013] Finally, the blades of variable-pitch turbines in an unfaired turbomachine are particularly sensitive to non-uniform airflows with non-axial incidence, i.e., airflows forming a non-zero angle with the direction of the turbomachine's principal X-axis, due to the absence of an external casing to guide the secondary flow. Such airflows appear notably during the aircraft's takeoff and landing phases and can generate transverse forces and moments in the propeller plane. These forces are transient and vary for each blade during its rotation, with differences depending on whether the blade is in the up or down position. Consequently, over one engine revolution, the same propeller blade is subjected to forces variables dependent on its azimuthal position. The stator blades located downstream of the propeller will also have a variable incidence depending on their azimuthal position.
[0014] A known solution for reducing the noise level generated by the blades is to uniformly decrease the radial length of the blades of the downstream wheel, i.e., the stator 5 in the case shown. In this way, the impact of the vortices formed at the radially external ends of the rotor 3 blades on the stator blades is limited because these vortices pass radially outside the stator blades. This solution is called "clipping," "cropping," "truncating," or "trimming" of the downstream wheel blades. A "clipping" ratio can be defined as the ratio of the difference in radius between the rotor and the stator to the rotor radius, generally expressed as a percentage.
[0015] However, this solution can still be improved. Indeed, a high clipping rate significantly reduces the turbomachine's propulsive efficiency. Moreover, noise reduction is primarily effective at zero angle of attack and does not necessarily provide satisfactory results at high angles of attack. Presentation of the invention
[0016] The invention aims to remedy these drawbacks, by proposing an unfaired turbomachine robust to variations in the incidence of airflows at different operating points corresponding to different phases of flight, and offering satisfactory aerodynamic and acoustic behavior for a wide range of rotational speeds.
[0017] To this end, the invention relates to a variable-pitch blade for an unfaired aeronautical propulsion system, comprising a blade extending along a blade axis from a root to a blade tip, the blade defining a leading edge, a trailing edge, and intrados and extrados surfaces extending from the leading edge to the trailing edge,
[0018] the blade comprising, in any cross-section plane orthogonal to the blade axis:
[0019] - a skeleton line extending from the leading edge to the trailing edge, equidistant of the intrados surface and the extrados surface, having a total length measured along the skeleton line from the leading edge to the trailing edge,
[0020] - a leading edge thickness, defined as the length of a first segment extending from the intrados edge to the extrados edge and intersecting the skeleton line perpendicularly at a first point on the skeleton line located at a distance from the leading edge, measured along the skeleton line, equal to 0.2% of the total length of the skeleton line,
[0021] - a collection zone thickness, defined as a length of a second segment extending from the intrados edge to the extrados edge and intersecting the skeleton line perpendicularly at a second point on the skeleton line located at a distance from the leading edge, measured along the skeleton line, equal to 5% of the total length of the skeleton line,
[0022] characterized in that, for each section plane orthogonal to the blade axis, the ratio between the thickness of the collection zone and the thickness of the leading edge is between 2.5 and 8.
[0023] Such a blade makes it possible to significantly limit aerodynamic separation at the leading edge of the propellers when they operate at low speed and with a high angle of attack to achieve the target thrust during the aircraft's takeoff phase. This then greatly reduces the formation of a downstream vortex, which would otherwise constitute a very energy-intensive and acoustically detrimental propeller wake.
[0024] For each section plane orthogonal to the blade axis, the ratio between the thickness of the collection zone and the thickness of the leading edge can be between 3.5 and 5.
[0025] Such a characteristic allows a better compromise between the robustness of the blade to variations in angle of attack and aerodynamic efficiency.
[0026] On a lower portion of the blade extending from the foot over a height between 0% and 35% of a total blade height measured between the foot and the top, for each section plane of said lower portion, the ratio between the thickness of the collection zone and the thickness of the leading edge can be between 3 and 8.
[0027] Such a characteristic makes it possible to improve robustness to variations in angle of attack in particular on the bottom of the blade, which generates the primary flow supplying the turbine in the case of the rotor, and therefore requires less performance from the blade regardless of the flight domain, or in the case of the stator, which takes on the most thrust.
[0028] On an upper portion of the blade extending to the top, over a height between 35% and 100% of a total blade height measured between the base and the top, for each section plane of said upper portion, the ratio between the thickness of the collection zone and the thickness of the leading edge can be between 2.5 and 5.
[0029] Such a feature makes it possible to optimize the aerodynamic performance of the blade on the upper part while maintaining sufficient robustness to variations in angle of attack and satisfactory acoustic performance.
[0030] The ratio between the thickness of the collection zone and the thickness of the leading edge for any cross-sectional plane of a lower portion of the blade extending from the root over a height between 0% and 35% of the total blade height measured from the root to the tip may be greater than or equal to the ratio between the thickness of the collection zone and the thickness of the leading edge for any cross-sectional plane of a upper portion of the blade extending to the top over a height between 35% and 100% of the total height of the blade.
[0031] Such a feature makes it possible to distribute the aerodynamic performance of the blade over the most critical parts and to improve the robustness at incidence of the lower parts of the blade.
[0032] The blade may include, in each cross-section plane orthogonal to the blade axis, a maximum thickness, defined as the length of a third segment extending from the intrados edge to the extrados edge and intersecting the skeleton line perpendicularly at a third point, for which the length of the third segment is maximum over a portion of the skeleton line,
[0033] wherein said third point is located at a distance from the leading edge, measured along the skeleton line, greater than or equal to 15%, and advantageously between 15% and 40% of the total length of the skeleton line.
[0034] Such a feature makes it possible to move the maximum thickness away from the leading edge and thus improve the performance of the blade.
[0035] The ratio can be strictly increasing from the leading edge to said third point and strictly decreasing from said third point to the trailing edge.
[0036] The blade can be a rotor blade mounted on a rotating disc about a main axis.
[0037] The invention also relates to an unfaired aircraft propulsion system, comprising at least one rotor and one stator, spaced along a main axis of the propulsion system, at least one of the rotor and the stator comprising a plurality of blades as above, distributed circumferentially around the main axis, in particular between 3 and 25 blades, advantageously between 8 and 16 blades.
[0038] Such a number of blades constitutes an advantageous compromise between propulsion energy performance and the noise generated.
[0039] Among the rotor and the stator, the one disposed upstream relative to the main axis may comprise at least two more blades than the one disposed downstream.
[0040] Such a feature makes it possible to reduce the noise of the turbomachine. Indeed, in the case where the number of rotor and stator blades are equal, the entire rotor wake interacts with the stator blades simultaneously, which increases the noise levels.
[0041] The stator can be disposed downstream and have a clipping ratio between 5% and 15% and in particular between 7% and 12%.
[0042] Such a feature allows for an additional reduction of the noise generated at the stator level, without significantly reducing the propulsive efficiency of the turbomachine.
[0043] The lengths of the stator blades may be non-homogeneous, with blade lengths in the lower part of the stator shorter than blade lengths in the upper part of the stator.
[0044] Such a feature allows for a higher clipping ratio below the stator, where the generated noise is greatest and therefore where noise reduction is most necessary, and a lower clipping ratio at the top, where it is less required. Thus, the trade-off between noise reduction and propulsive efficiency is improved.
[0045] Each blade can have a chord length defined as the maximum over a span of the blade of a distance between the leading edge and the trailing edge in a plane of cross-section transverse to the blade axis,
[0046] the rotor and stator having gaps separating adjacent blades, measured along a circumferential direction,
[0047] in which a strength of the rotor and stator, defined as the ratio of the chord length to the spacing separating the neighboring blades, can be less than or equal to 3, and in particular less than or equal to 1 for the one located furthest upstream of the rotor and stator, relative to the main axis.
[0048] A ratio between an axial distance separating the rotor and the stator and an external diameter of the rotor can be between 0.01 and 0.5, preferably between 0.15 and 0.35.
[0049] Such a feature allows for efficient rectification of the flow at the rotor outlet by the stator and improves the aerodynamic performance of the propulsion system. Brief description of the figures
[0050] [Fig-1] is a side view of an unfaired turbomachine according to the invention,
[0051] [Fig.2] is a side view of a rotor blade of the turbomachine of [Fig. 1],
[0052] [Fig.3] is a cross-sectional view of the blade of the [Fig.2],
[0053] [Fig.4] is a cross-sectional profile of the blade of figures 2 and 3. Detailed description of the invention
[0054] An unfaired turbomachine 1 is shown in [Fig.1], defining a central axis X and comprising a rotor 3 and a stator 5 spaced along the main axis X. The stator 5 is positioned downstream of the rotor 3.
[0055] The turbomachine 1 also includes at least one engine disposed in its internal space, said engine being able to be a thermal engine, in particular of the turboshaft, turbojet, turbofan type, and / or an electric motor, and / or a hydrogen engine, and / or a hybrid engine combining several of these technologies.
[0056] The rotor 3 and the stator 5 both comprise a plurality of blades 10 extending substantially radially from the main axis X and distributed regularly circumferentially around the axis X. The blades 10 of the rotor 3 may have different dimensions of the stator 5 blades 10, including different blade lengths.
[0057] Advantageously, the rotor 3 comprises at least two more blades 10 than the stator 5.
[0058] The rotor 3 has a radius RI, measured from the central axis to a vertex of each blade 10 of the rotor 3.
[0059] Similarly, the stator 5 has a radius R2 measured from the central axis to a vertex of each blade 10 of the stator 5.
[0060] According to one embodiment, the radius R2 of the stator 5 is smaller than the radius RI of the rotor 3, specifically smaller than the radius RI by between 5% and 15% of the value of RI, and in particular between 7% and 12% of the value of RI. In other words, the stator 5 has a clipping ratio, as defined above, of between 5% and 15%, and in particular between 7% and 12%. This clipping ratio allows for a further reduction in the noise generated at the stator, but remains sufficiently moderate so as not to significantly reduce the propulsive efficiency of the turbomachine.
[0061] A diameter D of the turbomachine 1 is defined as twice the larger of the two radii RI, R2.
[0062] The blades 10 are variable pitch blades, that is to say, whose blades are mobile in rotation around a radial axis in order to vary the pitch angle of each blade of the rotor 3 or of the stator 5 in a controlled manner.
[0063] The axes of rotation of the blades, or blade axes, of the rotor blades 3 are included in a plane PI perpendicular to the main axis X, and the axes of the stator blades 5 are included in a plane P2 perpendicular to the main axis X and offset from the plane PI by a distance S measured along the main axis X.
[0064] In the case where the stator blades 5 are fixed pitch blades, the plane P2 is defined at the center of gravity of the blades 10.
[0065] Advantageously, an S / D ratio, between the axial distance S separating the planes PI and P2, and thus separating the rotor 3 from the stator 5, and the external diameter of the turbomachine 1 is between 0.01 and 0.5, preferably between 0.15 and 0.35.
[0066] A blade 10 is shown in more detail in [Fig.2]. The blade 10 comprises a foot 12, a blade 14 and a tip 16. The blade 14 extends along a blade axis Z perpendicular to the principal axis X and included in the plane PI described above.
[0067] The blade 14 defines a leading edge 18 formed by the upstream line of the blade 14 and a trailing edge 20 formed by the downstream line.
[0068] The blade 14 comprises an intrados surface 22 and an extrados surface 24 extending from the leading edge to the trailing edge on either side of the blade 14.
[0069] A blade height H is defined as the distance measured along the blade axis Z separating the foot 12 from the tip 16, and a chord length C as the distance separating the leading edge 18 from the trailing edge 20, measured in a cross-section plane P.
[0070] A so-called solidity ratio ir = C / E of the rotor 3 or of the stator 5 is defined as the ratio between the chord length C measured at the apex 16 and a spacing E between the apexes 16 of two blades 10 close to the rotor 3 or the stator 5.
[0071] Advantageously, the strength is less than 3 for the rotor and the stator, and preferably less than 1 for the one placed furthest upstream, i.e. rotor 3 in the case represented.
[0072] A height h of a cross-section plane P is defined as the distance measured along the blade axis Z between the foot 12 and the cross-section plane P.
[0073] Figures 3 and 4 represent cross-sections of the blade 14 in a plane of section P located at a height h.
[0074] In plane P, a skeleton line LS is defined as the curved line extending from the leading edge 18 to the trailing edge 20 and equidistant from the intrados surface 22 and the extrados surface 24.
[0075] The length of the skeleton line LS is strictly greater than the chord length C in the plane P, which separates the leading edge 18 from the trailing edge 20, measured in a straight line.
[0076] The pitch angle y of the blade 14 is represented as the angle between the chord C and the transverse plane PL. The pitch angle y can be modified by rotating the blade 14 around the blade axis Z.
[0077] As shown in [Fig.4], a thickness of the blade 14 is defined in the plane of section P at a height h as the length of a segment perpendicular to the skeleton line LS and extending from the edge of the intrados 22 to the edge of the extrados 24.
[0078] A leading edge thickness Ep0,2 is thus defined as the length of a first segment S1 which intersects the skeleton line LS perpendicularly at a first point Al located at a distance, measured along the skeleton line LS, equal to 0.2% of the total length of the skeleton line LS.
[0079] Similarly, a capture zone thickness Ep5 is defined as the length of a second segment S2 which intersects the skeleton line LS perpendicularly at a second point A2 located at a distance, measured along the skeleton line LS, equal to 5% of the total length of the skeleton line LS.
[0080] The ratio R = Ep5 / Ep0.2 between the thickness of the collection zone Ep5 and the leading edge thickness Ep0.2 is characteristic of the blade's performance and its robustness to variations in angle of attack. In particular, low R values correspond to blades with high aerodynamic performance, while high R values correspond to blades that are robust to changes in angle of attack.
[0081] Thus, R values between 2.5 and 8 provide an advantageous compromise allowing good performance for the blade while reducing the risks of delamination and having satisfactory acoustic performance.
[0082] These minimum and maximum values of the ratio R are valid for any section plane P of the blade 14, that is to say over the entire height H of the blade 14.
[0083] The behavior of the blade 14 is even more satisfactory for R values between 3.5 and 5 over the entire height H.
[0084] According to an advantageous embodiment, a lower part of the blade 14 is distinguished, for heights h between 0 and 35% of the total height H, and an upper part of the blade, for heights h between 35% and 100% of the total height H of the blade 14.
[0085] Over the entire lower portion of the blade 14, the R-ratio is advantageously between 3 and 8, emphasizing the blade's resistance to variations in angle of attack. Indeed, aerodynamic performance is less critical in the lower portion, which concentrates the primary flow Fp towards the turbine.
[0086] Over the entire upper part of the blade 14, the ratio R is advantageously between 2.5 and 5 in order to emphasize the aerodynamic performance of the blade 14 on the upper part generating the secondary flow Fs.
[0087] Moreover, the R ratio over the entire lower part is advantageously greater than the R ratio over the entire upper part of the blade 14.
[0088] The application of these criteria is valid regardless of the shape of the leading edge area, whether for a substantially circular leading edge, or for a substantially asymmetrical leading edge.
[0089] As before, a maximum thickness Epmax is defined as the length of a third segment S3 which intersects the skeleton line LS perpendicularly at a third point A3 and for which the measured length is maximum over the whole of the skeleton line.
[0090] Advantageously, the blade thickness varies monotonically from the leading edge to the maximum Epmax, and then again monotonically from the maximum Epmax to the trailing edge.
[0091] Preferably, the third point A3 is located at a distance from the leading edge, measured along the skeleton line LS, greater than or equal to 15% of the total length of said skeleton line LS. This allows the maximum thickness to be sufficiently removed from the leading edge to achieve satisfactory performance. Even more preferably, point A3 is located at a distance from the leading edge of between 15% and 40% of the total length of said skeleton line LS.
Claims
1. Demands Variable pitch aeronautical propulsion (1) blade (10), comprising a blade (14) extending along a blade axis (Z), from a root (12) to a tip (16) of the blade (14), the blade (14) defining a leading edge (18), a trailing edge (20), and intrados (22) and extrados (24) surfaces extending from the leading edge (18) to the trailing edge (20), the blade (14) comprising, in any cross-section plane (P) orthogonal to the blade axis (Z): - a skeleton line (SL) extending from the leading edge (18) to the trailing edge (20), equidistant from the intrados surface (22) and the extrados surface (24), having a total length measured along the skeleton line (SL) from the leading edge (18) to the trailing edge (20), - a leading edge thickness (Ep0,2), defined as the length of a first segment (SI) extending from the lower surface edge (22) to the upper surface edge (24) and intersecting perpendicularly the skeleton line (LS) at a first point (Al) of the skeleton line (LS) located at a distance from the leading edge (18), measured along the skeleton line (LS), equal to 0.2% of the total length of the skeleton line (LS), - a capture zone thickness (Ep5), defined as a length of a second segment (S2) extending from the lower surface edge (22) to the upper surface edge (24) and intersecting perpendicularly the skeleton line (LS) at a second point (A2) of the skeleton line (LS) located at a distance from the leading edge (18), measured along the skeleton line (LS), equal to 5% of the total length of the skeleton line (LS), characterized in that, for each section plane (P) orthogonal to the blade axis (Z), the ratio (R) between the capture zone thickness (Ep5) and the leading edge thickness (Ep0.2) is between 2.5 and 8, wherein the ratio (R) between the thickness of the collection zone (Ep5) and the thickness of the leading edge (Ep0.2) for any cross-sectional plane (P) of a lower portion of the blade (14) extending from the root (12) over a height (h) between 0% and 35% of a total height (H) of the blade (14) measured between the root (12) and the apex (16) is greater than or equal to the ratio between the thickness of the capture zone (Ep5) and the thickness of the leading edge (Ep0,2) for any section plane (P) of an upper portion of the blade (14) extending to the apex (16) over a height (h) between 35% and 100% of the total height of the blade (14).
2. Blade (10) according to claim 1, wherein, for each section plane (P) orthogonal to the blade axis (Z), the ratio (R) between the thickness of the collection zone (Ep5) and the thickness of the leading edge (Ep0.2) is between 3.5 and 5.
3. Blade (10) according to claim 1 or 2, wherein, on a lower portion of the blade (14) extending from the foot (12) over a height (h) between 0% and 35% of a total height (H) of the blade (14) measured between the foot (12) and the top (16), for each section plane (P) of said lower portion, the ratio (R) between the thickness of the collection zone (Ep5) and the thickness of the leading edge (Ep0.2) is between 3 and 8.
4. Blade (10) according to any one of the preceding claims, wherein, on an upper portion of the blade (14) extending to the top (16), over a height (h) between 35% and 100% of a total height (H) of the blade (14) measured between the foot (12) and the top (16), for each section plane (P) of said upper portion, the ratio (R) between the thickness of the collection zone (Ep5) and the thickness of the leading edge (Ep0.2) is between 2.5 and 5.
5. Blade (10) according to any one of the preceding claims, wherein the blade (14) comprises, in each section plane (P) orthogonal to the blade axis (Z), a maximum thickness (Epmax), defined as a length of a third segment (S3) extending from the lower edge (22) to the upper edge (24) and intersecting perpendicularly the skeleton line (LS) at a third point (A3), for which the length of the third segment (S3) is maximum over a portion of the skeleton line (LS), wherein said third point (A3) is located at a distance from the leading edge (18), measured along the skeleton line (LS), greater than or equal to 15% of the total length of the skeleton line (LS), and advantageously between 15% and 40% of the total length of the skeleton line (LS).
6. Blade (10) according to any one of the preceding claims, wherein the blade (10) is a rotor blade (3) mounted on a rotating disc about a main axis (X).
7. Unfaired aircraft propulsion (1), comprising at least one rotor (3) and one stator (5) spaced about a principal axis (X) of the propulsion (1), at least one of the rotor (3) and the stator (5) comprising a plurality of blades (10) according to any one of the preceding claims, distributed circumferentially around the principal axis (X), in particular between 3 and 25 blades (10), advantageously between 8 and 16 blades (10).
8. Propeller (1) according to the preceding claim, wherein, among the rotor (3) and the stator (3), the one disposed upstream relative to the main axis (X) comprises at least two more blades (10) than the one disposed downstream.
9. Propellant (1) according to claim 7 or 8, wherein the stator (5) is disposed downstream and has a clipping rate of between 5% and 15% and in particular between 7% and 12%.
10. Propeller (1) according to any one of claims 7 to 9, wherein each blade (14) has a chord length (C) defined as the maximum over a span of the blade (14) of a distance between the leading edge (18) and the trailing edge (20) in a section plane (P) transverse to the blade axis (Z), the rotor (3) and the stator (5) having spacings (E) separating neighboring blades (14), measured along a circumferential direction, wherein a strength (ji) of the rotor (3) and the stator (5), defined as the ratio of the chord length (C) to the spacing (E) separating neighboring blades, is less than or equal to 3, and in particular less than or equal to 1 for the one located furthest upstream of the rotor (3) and the stator (5), relative to the main axis (Z).
11. Propeller (1) according to any one of claims 7 to 10, wherein a ratio (S / D) between an axial distance (S) separating the rotor (3) and the stator (5) and an external diameter (D) of the propeller (1) is between 0.01 and 0.5, preferably between 0.15 and 0.35.