AERONAUTICAL PROPELLER COMPRISING A BLADE WITH RELIEF DETAILS

Abstract

The invention relates to an aeronautical propulsion system comprising a blade, the blade comprising a streamlined portion (85) having an aerodynamic profile extending radially from the central piece (83), the streamlined portion (85) having a leading edge (86), a trailing edge (87), an intrados surface and an extrados surface (89), wherein the intrados surface (88) or the extrados surface (89) comprises a textured surface portion (90) having a series of protrusions (95) and / or depressions (96), the textured surface portion (90) being defined such that any point considered of the textured surface portion (90) is located at a radial distance from the principal axis of the aeronautical propulsion system equal to the sum of a minimum radius (Rmin) and between 20% and 95% of a span of the streamlined portion (85), and is located at an axial distance from the leading edge between 2% and 50% of the local rope length. Figure for the abbreviation: Figure 6

Description

Title of the invention: AERONAUTICAL PROPELLER COMPRISING A BLADE WITH RELIEF FIELD OF INVENTION

[0001] The invention relates to an aeronautical propulsion system, such as a gas turbine engine for example. STATE OF THE ART

[0002] Aeronautical propulsion systems include fixed parts and rotating parts, which are driven in rotation relative to the fixed parts when the aeronautical propulsion system is in operation.

[0003] The rotating parts and the fixed parts each have blades.

[0004] Thus, gas turbine engines generally include a blower module or propeller module, a compressor module, a combustion chamber and a turbine module.

[0005] The blower module (or propeller module), the compressor module and the turbine module each comprise rotating parts (or "rotor") and stationary parts (or "stator").

[0006] For example, the blower module comprises a blower stator and a blower rotor adapted to be driven in rotation relative to the blower stator. The blower rotor comprises one or more rows of movable blades. The rotation of the blower rotor compresses air, which is expelled to the rear in order to produce part of the engine's thrust.

[0007] In addition, the blower stator generally includes a set of fixed outlet vanes (also called "Outlet Guide Vanes" or "OGVs") arranged downstream of the blower rotor and acting as a straightener. This set of fixed vanes straightens and regulates the airflow downstream of the blower rotor to optimize engine thrust.

[0008] The airflow that passes through the set of fixed blades flows globally between the fixed blades in an upstream-downstream direction.

[0009] However, under certain engine operating conditions, particularly at partial power (i.e., during aircraft takeoff and landing phases), recirculation bubbles ("separation / recirculation bubble" in English) and / or areas with partially detached flow may appear on the blade surfaces, especially on the upper surfaces. These recirculation bubbles generate noise.

[0010] On the one hand, these recirculation bubbles produce high-frequency lines (tonal noise), which are linked to the presence of these bubbles and the associated vortex shedding. On the other hand, the recirculation bubbles increase the boundary layer thickness, which impacts the fan rotor wake and the interaction noise with the fan stator or the row of fixed blades located downstream. Indeed, an excessively large recirculation bubble increases the turbulent kinetic energy and the velocity deficit in the wake, which increases the interaction noise (broadband and tonal, respectively).

[0011] The same phenomenon can occur on the unfaired propellers of turboprop engines, of "Counter-Rotating Open Rotor" (CROR) type gas turbine engines and on the unfaired fans of "Unducted Single Fan" (USF) type gas turbine engines.

[0012] This same phenomenon can also occur on other fixed or moving blades of an aeronautical propulsion system. Summary of the invention

[0013] One object of the invention is to reduce the noise generated by a blade in the presence of a recirculation bubble and / or areas with partially detached flow which appear at certain regimes, without penalizing the operation of the aeronautical propulsion system at other regimes.

[0014] This objective is achieved within the framework of the present invention by means of an aeronautical propulsion system comprising a fixed part and a rotating part adapted to be driven in rotation relative to the fixed part around a principal axis of the aeronautical propulsion system, one of the fixed part and of the rotating part comprising a central piece and a blade, the blade comprising a profiled part having an aerodynamic profile extending radially from the central piece, the profiled part having a leading edge, a trailing edge, an intrados surface and an extrados surface, in which the intrados surface or the extrados surface comprises a portion of textured surface having a series of protrusions and / or hollows,the textured surface portion being defined such that any point considered of the textured surface portion is located at a radial distance from the principal axis of the aeronautical propulsion system equal to the sum of a minimum radius and between 20% and 95% of a span of the streamlined part, and is located at a distance from the leading edge of between 2% and 50% of the local chord length, in which the span of the streamlined part is defined as a difference between a maximum radius of the streamlined part and a minimum radius of the streamlined part, the , maximum radius being defined as a distance between a point on the airfoil furthest from the main axis of the aeronautical propulsion system, and the main axis of the aeronautical propulsion system, and the minimum radius being defined as a distance between a point on the leading edge of the airfoil closest to the main axis of the aeronautical propulsion system, and the main axis of the aeronautical propulsion system, or in the case where the blade has variable pitch, the minimum radius being defined as a distance between a point on the leading edge of the airfoil closest to the main axis of the aeronautical propulsion system, and the main axis of the aeronautical propulsion system, when the blade is positioned with a pitch angle in which the blade is feathered, and in which the local chord length is defined as a distance between a point on the leading edge and a point on the trailing edge,the leading edge point and the trailing edge point being located at the same radial distance from the principal axis as the point under consideration.

[0015] The portion of textured surface thus defined is located in an area in which a recirculation bubble and / or a partially detached flow is likely to appear.

[0016] The presence of a series of protrusions and / or depressions in this area allows for the recreation of turbulent flow at this location, instead of a recirculation bubble during partial operation (i.e., during the aircraft's takeoff and landing phases), and reduces wall friction and aerodynamic losses during nominal operation (i.e., during cruise). This thus limits the noise generated by the fan during takeoff and landing, without impacting the performance of the aircraft engine during cruise.

[0017] The aeronautical propulsion system may also have one or more of the following characteristics:

[0018] - the textured surface portion is a portion of the extrados surface and the series of protrusions and / or hollows consist of an alternation of protrusions and hollows,

[0019] - the textured surface portion is defined such that any point considered of the textured surface portion is located at a radial distance from the main axis of the aeronautical propulsion system equal to the sum of a minimum radius and between 50% and 90% of a span of the streamlined part,

[0020] - the textured surface portion is defined such that any point considered of the textured surface portion is located at a radial distance from the main axis of the aeronautical propulsion system greater than a radial distance from a point on the leading edge of the airfoil located furthest upstream, considering that the main axis of the propulsion system extends from upstream to downstream in the direction of gas flow through the aeronautical propulsion system, when the aeronautical propulsion system is in normal operation,

[0021] - the textured surface portion is defined such that any point considered of the the textured surface portion is located at an axial distance from the leading edge of between 5% and 40% of the local chord length, preferably between 10% and 30% of the local chord length,

[0022] - for any point considered on the textured surface portion, the thickness of the part the profile at the point considered is less than 12% of the local chord, preferably less than or equal to 10% of the local chord, or preferably less than or equal to 7% of the local chord,

[0023] - the thickness of the profiled part at the point considered is greater than 0.1% of the chord local, preferably greater than 0.25% of the local cord, or even preferably greater than 1% of the local cord,

[0024] - for a given radial distance, the profiled part has a maximum thickness at a given point on the intrados or extrados surface, and any point considered on the textured surface portion is located upstream of the point where the thickness of the profiled part is maximum,

[0025] - for a given radial distance, the profiled part has a maximum thickness at a given point on the intrados or extrados surface, the given point being located at a distance from the leading edge of between 10% and 45% of the local chord, preferably between 15% and 30% of the local chord,

[0026] - the profiled part comprises a core made of composite material and an added piece fixed to the core and forming the leading edge of the profiled part, and the projections and hollows are formed only on the added piece,

[0027] - the added part is a metal reinforcement piece fixed to the core or a mat heating element fixed to the core,

[0028] - the projections and / or hollows are located at a radial distance from the main axis greater than a radial distance from the maximum local chord of the blade relative to the main axis, or greater than the radial distance from a leading edge antinode relative to the main axis,

[0029] - the projections and / or hollows are formed by machining in the added part,

[0030] - alternatively, the projections and / or hollows are also formed on the core in composite material,

[0031] - the protrusions and / or hollows are formed in a film applied to the core of material composite,

[0032] - the intrados surface or the extrados surface comprises several textured portions featuring a series of protrusions and / or depressions, and in which a portion of textured surface is separated from the nearest other portion of textured surface by a distance of between 1% and 35% of the span, preferably between 2% and 15% of the span,

[0033] - the projections have a height between 0.04% of the local chord, of preferably 1%, and 3% of the local chord and / or the hollows have a depth between 0.04% of the local chord, preferably 1%, and 3% of the local chord,

[0034] - the projections and / or hollows form a repeating pattern with a constant pitch between two consecutive projections or between two consecutive hollows, the spacing being between 1 / 3 and 3 times the height of a projection or between 1 / 3 and 3 times the depth of a hollow,

[0035] - the hollows comprise cavities, each cavity having a portion-like shape sphere, for example a hemispherical shape,

[0036] - the projections include ribs,

[0037] - each rib has a base and an edge having a height measured by ratio to the base which increases in a strictly monotonic manner from upstream to downstream,

[0038] - the edge has an upstream end and a downstream end, and each rib includes an upstream end face extending from the base to the upstream end of the edge, forming a first non-zero angle with the base, less than or equal to 90°, and a downstream end face extending from a downstream end of the edge to the base, forming a second non-zero angle with the base, less than or equal to 90°,

[0039] - the upstream end face has a flat portion extending from the base and a rounded junction portion extending from the flat portion to the upstream end of the edge,

[0040] - the first angle is between 20° and 70°,

[0041] - the second angle is between 20° and 70°,

[0042] - the downstream end face has a rounded junction portion extending to starting from the downstream end of the edge and a flat portion extending from the rounded junction portion to the base,

[0043] - each rib has a triangular cross-section, with a base and a vertex, the vertex having a height measured relative to the base greater than the width of the base,

[0044] - the textured surface portion comprises first zones and second zones arranged alternately with the first zones along a radial direction relative to the main axis, and each first zone has a series of first ribs, oriented at a first angle relative to the main axis of the aeronautical propulsion system, and each second zone has a series of second ribs oriented at a second angle relative to the main axis, the second angle being different from the first angle,

[0045] - the first angle and the second angle are adjacent and the first angle is included between +15° and +45° relative to the main axis of the motor, and the second angle is between -15° and -45° relative to the main axis of the motor.

[0046] - the alternation of first zones and second zones presents a spatial period between 5% and 20% of the local rope,

[0047] - each first rib converges towards a respective second rib in the upstream-downstream direction,

[0048] - the aeronautical propulsion system comprises a ducted fan or a non- streamlined, and the blade is a blade of the streamlined fan or of the unstreamlined propeller of the aeronautical propulsion system,

[0049] - the blade has variable pitch,

[0050] - the aeronautical propulsion system includes a fan drive turbine or of the propeller, a blower shaft or a propeller shaft connected to the blower or propeller, a turbine shaft connected to the turbine, and a reduction mechanism having an input connected to the turbine shaft and an output connected to the blower shaft or propeller shaft, so that in operation the blower or propeller is driven into rotation by the turbine at a rotational speed lower than a rotational speed of the turbine. PRESENTATION OF THE DRAWINGS

[0051] Other features and advantages will become apparent from the following description, which is purely illustrative and not exhaustive, and should be read in conjunction with the accompanying figures, among which:

[0052] - Fig. 1 schematically represents a first example of a propulsion system aeronautical component comprising an unfaired propeller module,

[0053] - [Fig.2] schematically represents a second example of a propulsion system aeronautical component comprising a streamlined fan module,

[0054] - Figure 3 schematically represents, in longitudinal section, an engine with enclosed fan gas turbine,

[0055] - Figure 4 schematically represents a fan or propeller blade,

[0056] - Figure 5 schematically represents, in cross-section, the blade of blower or propeller of the [Fig.4],

[0057] - Figure 6 schematically represents a fan or propeller blade in accordance with an embodiment of the invention,

[0058] - Figure 7 schematically represents, in cross-section, the blade of blower or propeller of the [Fig.6],

[0059] - Figure 8 schematically represents a fan blade profile or of the helix and the skeleton line in the cutting plane,

[0060] - [Fig. 9] is an enlarged view of a portion of the textured surface of the dawn of blower or propeller, according to a possible embodiment of the invention,

[0061] - [Fig. 1OA] schematically represents the shape of a projection provided in the extrados portion according to a first embodiment of the invention,

[0062] - [Fig.1OB] schematically represents the shape of a projection provided in the extrados portion according to a second embodiment of the invention,

[0063] - [Fig. 1 1] schematically represents an alternation of projections and hollows formed in the textured surface portion of the fan or propeller blade,

[0064] - [Fig. 12] schematically represents the orientation angles of the projections, the projections having the shape of ribs,

[0065] - Fig. 13 schematically represents, in cross-section, a blade including a projection,

[0066] - [Fig. 14] is an enlarged view of a portion of the textured surface of the dawn of blower, conforming to another possible embodiment of the invention,

[0067] - Figure 15 schematically represents, in cross-section, the blade of the [Fig.14]

[0068] - Figures 16A to 16F illustrate different configurations of the surface portion textured. DETAILED DESCRIPTION OF A METHOD OF IMPLEMENTATION

[0069] In [Fig.1], the aeronautical propulsion unit 1 shown is a gas turbine engine with unfaired propellers. The gas turbine engine 1 is an "Open Rotor" type gas turbine engine, in a configuration commonly referred to as "pusher" (i.e. the fan is placed downstream of the power generator with an air inlet located upstream, on the left in [Fig.1]).

[0070] The gas turbine engine 1 comprises a nacelle 2 for attachment to an aircraft fuselage, and an unfaired fan 8. The fan 8 comprises two counter-rotating fan rotors 81 and 82. In other words, when the engine 1 is running, the rotors 81 and 82 are driven in rotation relative to the nacelle 2 around the same axis of rotation X (which coincides with a main axis of the engine), in opposite directions.

[0071] In the example illustrated in [Fig. 1], the motor 1 is a non-ducted, counter-rotating open rotor (CROR) motor in a pusher configuration. However, the invention is not limited to this configuration. The invention also applies to open rotor motors in a puller configuration (i.e., the fan is positioned upstream of the power generator with an air inlet located before, between, or just behind the two fan rotors).

[0072] In addition, the invention also applies to motors with different architectures, such as an architecture comprising a blower rotor including movable blades (or an "Open Fan" in English) and a blower stator including blades (USF), or a single blower rotor.

[0073] The fan stator blades can be fixed or have variable pitch. In this case, each of the blades is mounted to pivot relative to the nacelle 2 along a pitch axis.

[0074] The invention is applicable to turboprop type architectures (comprising a single fan rotor).

[0075] In [Fig.1], each blower rotor 81, 82 comprises a hub 83 mounted rotatably relative to the nacelle 2 and a plurality of blades 84 connected to the hub 83. The blades 84 extend substantially radially relative to the axis of rotation X of the hub.

[0076] The blades 84 can be fixed or have variable pitch. Fixed blades are mounted fixedly on the hub 83. In the case where the blades 84 have variable pitch, each blade 84 is mounted to pivot relative to the hub 83 around a respective pitch axis Y. The blades 84 are connected to a pitch-change mechanism allowing adjustment of the pitch angle of the blades 84 relative to the hub 83, and thus their angle of incidence, according to the flight phases.

[0077] In [Fig. 2], the aircraft propulsion unit 1 shown is a ducted fan gas turbine engine. The ducted fan gas turbine engine 1 shown comprises a nacelle 2 for attachment to an aircraft fuselage, a fan 8, and a duct surrounding the fan 8, the duct being fixedly mounted on the nacelle 2. In the example shown in [Fig. 2], the duct of the fan 8 is located inside the nacelle 2.

[0078] The blower 8 comprises a blower rotor 81 adapted to be driven in rotation relative to the nacelle 2 about an axis of rotation X (which coincides with the main axis of the motor 1). The blower rotor 81 comprises a hub 83 and a plurality of blades 84 fixed to the hub 83 and extending in substantially radial directions from the hub 83. In the example illustrated in [Fig. 2], the blades 84 are all identical and arranged with a constant angular spacing between two successive blades.

[0079] The fan 8 can be a variable pitch fan (called "Variable pitch fan" or "VPF" in English), that is to say that the fan includes a mechanism allowing each blade 84 to be rotated around a pitch axis so as to modify the pitch of the blades according to the phases of flight.

[0080] In [Fig. 3], the aeronautical propulsion unit 1 shown is a twin-spool, twin-flow gas turbine engine. The gas turbine engine can be a turbine engine gas with a high dilution ratio ("Ultra High Bypass Ratio" or "UHBR" in English), that is to say with a dilution ratio ("Bypass ratio") between about 15 and about 40.

[0081] The gas turbine engine 1 has a main axis X (or longitudinal axis).

[0082] The gas turbine engine 1 comprises a nacelle 2, a blower module 3, a compressor module 4, a combustion chamber 5, and a turbine module 6.

[0083] In the example illustrated in [Fig.3], the blower module 3 comprises a blower housing 7 mounted fixed relative to the nacelle, a blower 8 adapted to be driven in rotation relative to the blower housing 7. The blower housing 7 comprises fixed outlet vanes 9 (or “OGV”) whose function is to straighten the secondary airflow which flows out of the blower 8.

[0084] In the example illustrated in [Fig.3], the compressor module 4 includes a low pressure compressor 10 and a high pressure compressor 11.

[0085] In addition, the turbine module 6 includes a high-pressure turbine 12 and a low-pressure turbine 13.

[0086] The gas turbine engine 1 includes a low-pressure shaft 14 connecting the low-pressure turbine 13 to the low-pressure compressor 10 and the blower 8, and a high-pressure shaft 15 connecting the high-pressure turbine 12 to the high-pressure compressor 11. The high-pressure shaft 15 is coaxial with the low-pressure shaft 14 and extends around the low-pressure shaft 14. The high-pressure shaft 15 and the low-pressure shaft 14 are mounted to rotate relative to the nacelle 2, around the main axis X of the engine.

[0087] In one embodiment, the gas turbine engine 1 may include a blower shaft for rotating the blower 8 and a reduction mechanism having an inlet connected to the low-pressure shaft 14 and an outlet connected to the blower shaft. In this embodiment, the blower 8 is rotated at a speed lower than the rotational speed of the low-pressure turbine 13. The reduction mechanism thus allows for independent optimization of the rotational speed of the blower 8 and the rotational speed of the low-pressure turbine 14 and the low-pressure compressor 10.

[0088] The blower module 3, the low-pressure compressor 10, the low-pressure turbine 13, and the low-pressure shaft 14 (as well as, where applicable, the reduction mechanism and the blower shaft) together form the low-pressure body of the motor 1. The low-pressure turbine 13 is suitable for driving the low-pressure compressor 10 and the blower 8 in rotation via the low-pressure shaft 14 (as well as, where applicable, via the reduction mechanism and the blower shaft).

[0089] More specifically, the low-pressure compressor 10 comprises a low-pressure compressor housing 16, fixedly mounted relative to the nacelle 2, a low-pressure compressor rotor 17, and a low-pressure compressor stator 18. The low-pressure compressor rotor 17 is adapted to be driven in rotation relative to the low-pressure compressor stator 18, around the main axis X of the motor 1. The low-pressure compressor rotor 17 includes movable vanes. The low-pressure compressor stator 18 includes fixed vanes (also called "guide vanes" or "straightener vanes") which are fixedly mounted on the low-pressure compressor housing 16, interposed between the movable vanes. These fixed vanes serve to guide the primary airflow through the low-pressure compressor 10.

[0090] Similarly, the low-pressure turbine 13 comprises a low-pressure turbine housing 19, fixedly mounted relative to the nacelle 2, a low-pressure turbine rotor 20, and a low-pressure turbine stator 21. The low-pressure turbine rotor 20 is adapted to be driven in rotation relative to the low-pressure turbine stator 21, around the main axis X of the engine 1. The low-pressure turbine rotor 20 comprises movable blades. The low-pressure turbine stator 21 comprises fixed blades that are fixedly mounted on the low-pressure turbine housing 19, interposed between the movable blades. These fixed blades serve to guide the primary airflow through the low-pressure turbine 13.

[0091] The low pressure turbine rotor 20 is connected to the low pressure compressor rotor 17 via the low pressure shaft 14. Thus, when the motor 1 is in operation, the rotation of the low pressure turbine rotor 20 causes a rotation of the low pressure compressor rotor 17.

[0092] The high-pressure compressor 11, the high-pressure turbine 12 and the high-pressure shaft 15 together form the high-pressure body of the motor 1. The high-pressure turbine 12 is adapted to drive the high-pressure compressor 11 in rotation via the high-pressure shaft 15.

[0093] More specifically, the high-pressure compressor 11 comprises a high-pressure compressor housing 22, fixedly mounted relative to the nacelle 2, a high-pressure compressor rotor 23, and a high-pressure compressor stator 24. The high-pressure compressor rotor 23 is adapted to be driven in rotation relative to the high-pressure compressor stator 24, about the main axis X of the motor 1. The high-pressure compressor rotor 23 comprises movable vanes. The high-pressure compressor stator 24 comprises fixed vanes (also called "guide vanes" or "straightener vanes") which are fixedly mounted on the housing 22 of the high-pressure compressor, interposed between the movable vanes. These vanes fixed are designed to guide the primary airflow through the high-pressure compressor 11.

[0094] In one embodiment, the high-pressure compressor stator blades 24 may have variable pitch, in order to ensure the operability of the high-pressure compressor 23 and increase its pumping margin.

[0095] Similarly, the high-pressure turbine 12 comprises a high-pressure turbine housing 25, fixedly mounted relative to the nacelle 2, a high-pressure turbine rotor 26, and a high-pressure turbine stator 27. The high-pressure turbine rotor 26 is adapted to be driven in rotation relative to the high-pressure turbine stator 27, around the main axis X of the engine 1. The high-pressure turbine rotor 26 comprises movable blades. The high-pressure turbine stator 27 comprises fixed blades that are fixedly mounted on the housing 25 of the high-pressure turbine, interposed between the movable blades. These fixed blades serve to guide the primary airflow through the high-pressure turbine 12.

[0096] The high-pressure turbine rotor 26 is connected to the high-pressure compressor rotor 11 via the high-pressure shaft 15. Thus, when the motor 1 is in operation, the rotation of the high-pressure turbine rotor 26 causes a rotation of the high-pressure compressor rotor 23.

[0097] The fixed outlet blades 9 of the blower module 3, the fixed stator blades 18 of the low pressure compressor 10, the fixed stator blades 24 of the high pressure compressor 11, the fixed stator blades 27 of the high pressure turbine 12 and the fixed stator blades 21 of the low pressure turbine 13 are examples of guide blades.

[0098] When the motor 1 is in operation, the blower 8 and the low pressure compressor 10 are driven in rotation by the low pressure turbine 13. Similarly, the high pressure compressor 11 is driven in rotation by the high pressure turbine 12.

[0099] Air is drawn in by the blower 8. The air drawn in by the blower 8 is divided between a primary airflow and a secondary airflow, which flow from upstream to downstream of the gas turbine engine 1.

[0100] The primary airflow flows from upstream to downstream of the gas turbine engine 1 in a primary channel, passing successively through the low-pressure compressor 10, the high-pressure compressor 11, the combustion chamber 5 where it is mixed with fuel to serve as an oxidizer, the high-pressure turbine 12 and the low-pressure turbine 13. The passage of the primary airflow through the high-pressure turbine 12 and the low-pressure turbine 13 causes rotation of the turbine rotors 26 and 20, which in turn drive the rotation of the rotors 23 and 17 of the high-pressure and low-pressure compressors, as well as the blower 8 via the high-pressure shaft 15 and the low-pressure shaft 14. The primary airflow escapes from the engine 1 through an exhaust casing 28, located downstream of the low-pressure turbine casing 19.

[0101] The secondary airflow (also called "bypass airflow") flows from upstream to downstream of the gas turbine engine 1 in a secondary channel. This secondary airflow does not pass through the combustion chamber 5 and does not drive the turbines 12 and 13. The secondary airflow serves both to cool the periphery of the engine casing and to generate most of the thrust supplied by the gas turbine engine. The secondary airflow flows through the fixed blades 9 mounted on the fan casing 2, downstream of the fan 8.

[0102] Figures 4 and 5 schematically represent a blower or propeller blade 84.

[0103] However, the invention also applies to other blades of the aircraft propulsion system, such as turbine or compressor blades, for example. The blade may be a blade of a rotating part of the aircraft propulsion system or a blade of a stationary part of the aircraft propulsion system.

[0104] In figures 4 and 5, the fan blade 84 includes a profiled part 85 having an aerodynamic profile extending radially from the fan hub 83.

[0105] The profiled part 85 has a leading edge 86, a trailing edge 87, an intrados surface 88 and an extrados surface 89.

[0106] Furthermore, in the example illustrated in Figures 4 and 5, the blower blade 84 is mounted to rotate relative to the hub 83 around a shimming axis Y, which allows the angle of incidence of the profiled part 85 to be changed relative to the main axis X of the motor, and consequently relative to a direction of the incoming airflow.

[0107] As illustrated in [Fig.4], a minimum radius Rmin of the profiled part 85 is defined as a distance between a point on the leading edge 86 of the profiled part 85 closest to the principal axis X of the aeronautical propulsion system, and the principal axis X of the aeronautical propulsion system.

[0108] In the case where the blade has variable pitch, the minimum radius Rmin of the profiled part 85 is defined as a distance between a point on the leading edge 86 of the profiled part 85 closest to the principal axis X of the aircraft propulsion system, and the principal axis X of the aircraft propulsion system, when the blade 84 is positioned with a pitch angle in which the blade 84 is feathered.

[0109] A variable-pitch fan blade is said to be "feathered" when the blade pitch angle is such that the aerodynamic drag generated by the blade in the airflow passing through the fan is minimal. The feathered position is the position taken by a fan blade when it is allowed to orient itself naturally in the airflow passing through the fan from upstream to downstream parallel to the main axis of the aircraft propulsion system, the rotating part not being driven in rotation.

[0110] A maximum radius Rmax is defined as a distance between a point on the profiled part 85 furthest from the main axis X of the aeronautical propulsion system, and the main axis of the aeronautical propulsion system.

[0111] A span L of the profiled part 85 is defined as a difference between the maximum radius Rmax of the profiled part 85 and the minimum radius Rmin of the profiled part 85.

[0112] On [Fig.4], the cutting plane AA is a plane parallel to the main axis of the X motor and which cuts the profiled part at a point on the leading edge 86 and a point on the trailing edge 87, the cutting plane AA being orthogonal to the alignment axis Y.

[0113] As illustrated in [Fig. 5], a local chord is defined as a segment connecting a point on the leading edge 86 and a point on the trailing edge 87, the leading edge point and the trailing edge point being located at the same radial distance from the principal axis X of the motor. Thus, the local chord length C is defined as a distance between a point on the leading edge 86 and a point on the trailing edge 87 located at the same radial distance from the principal axis X of the motor.

[0114] In [Fig.5], the angle y is the blade pitch angle. The pitch angle y is defined as the angle between the local chord C measured at a radial distance equal to 0.75 x Rmax from the main axis X of the motor, and the main axis of the motor.

[0115] In the case of a variable pitch blade, when the blade is in feathering, the pitch angle is generally equal to about 90° (within 15°).

[0116] As can be seen in Figures 6 and 7, the upper surface 89 of the profiled part 85 has a textured surface portion 90 with alternating protrusions and depressions. The remainder of the upper surface 89, extending beyond the textured surface portion 90, is smooth, i.e., it has no protrusions or depressions.

[0117] The textured surface portion 90 is delimited by a first boundary line 91, a second boundary line 92, a third boundary line 93 and a fourth boundary line 94.

[0118] The textured surface portion 90 extends radially between the first boundary line 91 and the second boundary line 92.

[0119] The first delimitation line 91 is located at a radial distance from the main axis X of the aeronautical propulsion greater than or equal to the sum of the minimum radius Rmin and 20% of the span L of the profiled part 85. In other words, the first delimitation line 91 is located at a distance from the main axis X greater than or equal to Rmin + 0.2 x L.

[0120] Preferably, the first boundary line 91 is located at a radial distance from the main axis X of the aeronautical propulsion unit greater than or equal to the sum of the minimum radius Rmin and 50% of the span L of the streamlined part 85. In other words, the first boundary line 91 is located at a distance from the main axis X greater than or equal to Rmin + 0.5 x L.

[0121] The second delimitation line 92 is located at a radial distance from the main axis X of the aeronautical propulsion less than or equal to the sum of the minimum radius Rmin and 95% of the span L of the profiled part 85. In other words, the second delimitation line 92 is located at a distance from the main axis X less than or equal to Rmin + 0.95 x L.

[0122] The second delimitation line 92 is located at a radial distance from the main axis X of the aeronautical propulsion less than or equal to the sum of the minimum radius Rmin and 90% of the span L of the profiled part 85. In other words, the second delimitation line 92 is located at a distance from the main axis X less than or equal to Rmin + 0.90 x L.

[0123] The textured surface portion 90 extends axially between the third boundary line 93 and the fourth boundary line 94.

[0124] The third delimitation line 93 is defined as the set of points on the extrados surface 89 located at an axial distance from the leading edge 86, greater than or equal to 2% of the local chord length C. In other words, the third delimitation line is located at an axial distance from the leading edge 86 greater than or equal to 0.02 x C.

[0125] Preferably, the third delimitation line 93 is defined as the set of points on the extrados surface 89 located at an axial distance from the leading edge 86, greater than or equal to 5% of the local chord length C, more preferably greater than or equal to 10% of the local chord C. In other words, the third delimitation line 93 is located at an axial distance, measured in the direction of the local chord at the same radial distance from the principal axis X of the motor, greater than or equal to 0.05 x C, preferably greater than or equal to 0.1 x C.

[0126] The fourth delimitation line 94 is defined as the set of points on the extrados surface 89 located at an axial distance from the leading edge 86, less than or equal to 50% of the local chord length C. In other words, the fourth delimitation line 94 is located at an axial distance, measured in the direction of the local chord at the same radial distance from the main axis X of the motor, less than or equal to 0.5 x C.

[0127] Preferably, the fourth boundary line 94 is located at an axial distance from the leading edge 86 less than or equal to 40% of the local chord length C, more preferably less than or equal to 30% of the local chord C. In other words, the fourth boundary line 94 is located at an axial distance, measured in the direction of the local string at the same radial distance from the main X axis of the motor, less than or equal to 0.4 x C, preferably less than or equal to 0.3 x C.

[0128] The portion of textured surface 90 thus defined is located in an area in which a recirculation bubble and / or a partially detached flow is likely to appear.

[0129] The presence of protrusions and / or depressions in this area allows for the recreation of turbulent flow at this location, instead of a recirculation bubble during partial operation (i.e., during the aircraft's takeoff and landing phases), and reduces wall friction and aerodynamic losses during nominal operation (i.e., during cruise). This thus limits the noise generated by the fan during takeoff and landing, without impacting the performance of the aircraft engine during cruise.

[0130] In the example illustrated in Figures 6 and 7, the alternation of protrusion and hollow extends over the entire portion of textured surface 90, that is to say from the first junction line 91 to the second junction line 92 and from the third junction line 93 to the fourth junction line 94.

[0131] Figure 7 schematically represents a profile of the profiled portion 85 of the blade 84, i.e., a cross-section of the profiled portion 85 in a transverse plane. The transverse plane is defined as a plane parallel to the principal axis X and containing a point on the leading edge and a point on the trailing edge located equidistant from the principal axis X.

[0132] In the case where the blade 84 is a variable pitch blade, the transverse cutting plane can be defined as a plane orthogonal to the pitch axis Y.

[0133] As illustrated in [Fig. 7], in each cross-sectional plane, the profiled portion 85 has a maximum thickness emax at a given point on the upper surface 89, located at an axial distance xemax from the leading edge 86, measured parallel to the local chord. The fourth boundary line 94 is located upstream of the point where the thickness of the profiled portion is maximum. In other words, each of the points on the fourth boundary line 94 is located at an axial distance from the leading edge 86 strictly less than xemax.

[0134] As illustrated in [Fig.8], in a given cross-sectional plane, the thickness e is defined as the distance between the intrados surface 88 and the extrados surface 89 of the profiled part 85 measured perpendicular to a skeleton line S.

[0135] The skeleton line S is defined as the set of points located midway between the intrados line and the extrados line in the transverse plane (the intrados line being defined as the line of intersection between the intrados surface 88 and the radial plane, and the extrados line being defined as the line of intersection between the extrados surface 89 and the radial plane). The skeleton line S can, for example, be obtained by positioning inscribed circles I inside the profile of the profiled part. The skeleton line is defined by the set of points that constitute the centers of the inscribed circles.

[0136] For any point considered of the textured surface portion 90, the thickness e of the profiled part 85 at the point considered is less than 12% of the local chord, preferably less than or equal to 10% of the local chord, or even more preferably less than or equal to 7% of the local chord.

[0137] The thicknesses of the profiled part 85 are small, especially in the upper part to reduce centrifugal forces and thus promote the mechanical strength of the blade 84.

[0138] Fig. 9 schematically represents an alternation of protrusions 95 and hollows 96 formed on the portion of textured surface 90.

[0139] In this example, the projections 95 have the shape of ribs, in relief with respect to the rest of the extrados surface 89.

[0140] The textured surface portion 90 comprises first zones 90A and second zones 90B, the second zones 90B being arranged alternately with the first zones 90A along a radial direction relative to the main axis X of the aeronautical propulsion system.

[0141] Each first zone 90A has a series of first ribs, oriented with a first angle with respect to the principal axis X of the aeronautical propulsion system, and each second zone 90B has a series of second ribs oriented with a second angle with respect to the principal axis X of the aeronautical propulsion system, the second angle being different from the first angle.

[0142] The angle α (α, α2) of a rib can be defined as the angle between the principal axis X of the aeronautical propulsion system and the projections or ribs (when these are projected onto a plane passing through the principal axis X and partially traversing the blade). If the blade has variable pitch, this definition is valid when the blade has the pitch corresponding to the Aerodynamic Design Point (ADP) or the operating point in cruise mode. For an unfaired propeller, this corresponds to a pitch angle γ, as defined in [Fig. 5], which varies between 60° and 70° on the cross-section at 0.75 x Rmax.

[0143] As illustrated in [Fig. 12], the first angle ai can be between +15° and 45°. The second angle can be symmetrical with respect to the first angle, i.e., the second angle a2 is between -45° and -15°. In a preferred embodiment, a2 = -aL

[0144] In this way, the first ribs and the second ribs together form a chevron pattern.

[0145] Fig. 1OA schematically represents a first example of rib 95.

[0146] Rib 95 has a base 97 facing the extrados surface and a summit 98 located at a distance from base 97.

[0147] A cross-section of the rib is defined as a section taken in a cutting plane orthogonal to a longitudinal direction of the rib 95.

[0148] In this first example, the rib 95 has a rectangular cross-section. That is to say, the rib 95 has a constant thickness t in cross-section from its base 97 to its apex 98.

[0149] In a variant illustrated to the right of [Fig.1OA], rib 95 may have a rounded apex 98.

[0150] Fig. 11 schematically represents a second example of rib 95.

[0151] Rib 95 has a base 97 facing the extrados surface and a summit 98 located at a distance from base 97.

[0152] In this second example, the rib 95 has a triangular cross-section. That is to say, the rib 95 has a cross-sectional thickness t which decreases continuously (or in a strictly monotonic manner) from its base 97 to its apex 98.

[0153] In a variant illustrated to the right of [Fig.1OB], rib 95 may have a rounded apex 98.

[0154] In these two examples, the apex 98 of the rib 95 has a height h measured relative to the base 97, greater than the width t of the base 97.

[0155] In addition, each rib 95 has an upstream end face 101 and a downstream end face 102.

[0156] Each rib 95 has an edge 103 extending from the upstream end face 101 to the downstream end face 102.

[0157] Edge 103 has an upstream end and a downstream end.

[0158] The upstream end face 101 has a flat portion extending from the base 97 and a rounded junction portion extending from the flat portion to the upstream end of edge 103. The flat portion forms a first non-zero angle [3 with the base 97. The first angle [3 is less than or equal to 90°.

[0159] Preferably, the first angle [3 is between 20° and 70°.

[0160] The downstream end face 102 has a rounded junction portion extending from the downstream end of edge 103 and a planar portion extending from the rounded junction portion to the base 97. The planar portion forms a second non-zero angle y with the base 97. The second angle y is less than or equal to 90°.

[0161] Preferably, the second angle rp is between 20° and 70°.

[0162] The edge 103 has a height h measured relative to the base 97 which increases in a strictly monotonic manner from the upstream end face 101 to the downstream end face 102.

[0163] For example, the edge has a first height measured at its upstream end greater than or equal to 0.04% of the local chord (i.e., h > 0.0004xC), preferably greater than or equal to 1% of the local chord (i.e., h > 0.01xC), and a second height measured at its downstream end less than or equal to 3% of the local chord (i.e., h < 0.03xC). This makes it possible to include ribs having a characteristic height corresponding to that of boundary layers and / or recirculation bubbles that may appear in partial flow on the upper surface of fan blades near the leading edge.

[0164] As illustrated in [Fig. 1 1], the projections and recesses form a repeating pattern with a constant pitch w between two consecutive projections 95 or between two consecutive recesses 96, the pitch being between 1 / 3 and 3 times the height of a projection or between 1 / 3 and 3 times the depth of a recess. In other words, 1 / 3 < w / h < 3.

[0165] The step size w is preferably between 1% and 3% of the local chord, i.e. 0.01 < w / C < 0.03.

[0166] As illustrated in [Fig. 12], the radial width of a rib pattern is between 5% and 20% of the chord. This rib pattern or motif can be repeated identically or homothetically in one or more directions (radial, axial, etc.) on the textured surface portion 90.

[0167] As illustrated in [Fig. 13], the base 97 of the ribs 95 can be located below the level of the smooth part of the extrados surface 89 which extends out of the portion of the extrados surface 90, while the top of the ribs 95 can be located above the level of this smooth part.

[0168] In the embodiment illustrated in figures 14 and 15, the textured surface portion 90 has hollows 96. The rest of the extrados surface 89, extending outside the textured surface portion 90, is smooth, that is to say, it does not have any protrusions or hollows.

[0169] In this embodiment, the hollows 96 have a shape of a portion of a sphere, for example a hemispherical shape.

[0170] Each hollow 96 has a depth h and a diameter t, the depth h preferably being equal to 0.5 x t.

[0171] The hollows 96 form a repeating pattern with a constant step w between two consecutive or adjacent hollows 96, the step being between 1 / 3 and 3 times the depth h of a hollow.

[0172] Moreover, the step w is preferably between 1% and 3% of the local chord.

[0173] The hollows 96 are arranged to form first rows parallel to each other and oriented with a first angle with respect to the main axis X of the aeronautical propulsion system, and second rows parallel to each other and oriented with a second angle with respect to the main axis X of the aeronautical propulsion system, different from the first angle.

[0174] As illustrated in [Fig.14], the first angle ai can be between +15° and +45°. The second angle can be symmetrical with respect to the first angle, that is, the second angle a2=-ai is between -45° and -15°.

[0175] Thus, the second rows form an angle equal to 2ai with the first rows.

[0176] Figures 16A to 16F illustrate different configurations of the textured surface portion 90.

[0177] In a first configuration illustrated in [Fig. 16A], the profiled part 85 comprises a core 111 made of composite material and an added piece 112 fixed to the core 111. The added piece 112 forms the leading edge 86 of the profiled part 85.

[0178] The added part 112 can be a metallic part, such as a metallic reinforcing part attached to the web 111, or a heating mat attached to the web 111 (in particular, in the case of a propeller blade, the heating mat provides a de-icing function for the leading edge area most exposed to frost).

[0179] The protrusions 95 and / or hollows 96 are formed only on the added part 112, all along the leading edge 86 from the first delimitation line 91 to the second delimitation line 92 delimiting the portion of textured surface 90. This makes it easy to machine the patterns of the protrusions and / or hollows on the surface of the blade, as well as to limit the wear of the patterns which may appear in service.

[0180] In a second configuration illustrated in [Fig. 10B], the upper surface 89 comprises several textured surface portions 90a, 90b, 90c having protrusions and / or recesses. The textured surface portions 90a, 90b, 90c are arranged one after the other along the leading edge 86. The protrusions 95 and / or recesses 96 are formed only on the added part 112. Each textured surface portion 90a, 90b, 90c is separated from the nearest textured surface portion by a distance (in the radial direction) bb b2 of between 1% and 35% of the span L of the profiled part 85, preferably between 2% and 15% of the span of the profiled part 85.

[0181] In a third configuration illustrated in [Fig.10C], the projections 95 and / or the hollows 96 are formed only on the added part 112, in a part of the added part 112 located near an end furthest radially from the main axis X of the aeronautical propulsion unit.

[0182] In a fourth configuration illustrated in [Fig.10D], the leading edge 86 is formed partly by the metal reinforcement 112 and partly by the composite material core 111.

[0183] More specifically, a portion of the leading edge 86 further from the main axis of the aeronautical propulsion system is formed by the composite material core 111, and a portion of the leading edge 86 closer to the main axis of the aeronautical propulsion system is formed by the metal reinforcement 112.

[0184] The projections 95 and / or the hollows 96 are formed only on the metal insert 112, in a part of the insert 112 located near one end furthest radially from the main axis X of the aeronautical propulsion unit.

[0185] In a fifth configuration illustrated in [Fig.10E], the leading edge 86 is formed partly by the metal reinforcement 112 and partly by the composite material core 111.

[0186] More specifically, a portion of the leading edge 86 closer to the main axis of the aeronautical propulsion system is formed by the composite material core 111, and a portion of the leading edge 86 further from the main axis of the aeronautical propulsion system is formed by the metallic reinforcement 112.

[0187] The projections 95 and / or the hollows 96 are formed only on the added part 112, in a part of the added part 112 located near an end furthest radially from the main axis X of the aeronautical propulsion.

[0188] When the projections 95 and / or the hollows 96 are formed on the added part 112, these projections and hollows can be formed by machining directly in the metal reinforcement 112.

[0189] In a sixth configuration illustrated in [Fig. 16F], the projections 95 and / or the hollows 96 are formed only on the composite material core 111. This makes it possible to limit the weight of the blade, in particular when a metal reinforcement piece is not required for reasons of mechanical resistance to bird ingestion.

[0190] In figures 15C, 15D, 15E and 15F, the protrusions 95 and / or the hollows 96 are located in a radial position either above the radial position of the local chord maximum C, or above the position of the leading edge belly 86, in an area where partial regime separations are likely to occur.

[0191] When the protrusions 95 and / or hollows 96 are formed on the core of composite material 111, these protrusions and / or hollows can be formed directly in the composite material or in a film which is applied to the extrados surface 89, for example by bonding.

Claims

1. Demands Aeronautical propulsion system (1) comprising a fixed part (7) and a rotating part (8) adapted to be driven in rotation relative to the fixed part (7) about a principal axis (X) of the aeronautical propulsion system, one of the fixed part (7) and the rotating part (8) comprising a central piece (83) and a blade (84), the blade (84) comprising a profiled part (85) having an aerodynamic profile extending radially from the central piece (83), the profiled part (85) having a leading edge (86), a trailing edge (87), an intrados surface (88) and an extrados surface (89), in which the intrados surface (88) or the extrados surface (89) comprises a portion of textured surface (90) having a series of projections (95) and / or depressions (96),the textured surface portion (90) being defined such that any point considered of the textured surface portion (90) is located at a radial distance from the principal axis (X) of the aeronautical propulsion system equal to the sum of a minimum radius (Rmin) and between 20% and 95% of a span (L) of the streamlined part (85), and is located at a distance from the leading edge of between 2% and 50% of the local chord length (C), in which the span (L) of the streamlined part (85) is defined as a difference between a maximum radius (Rmax) of the streamlined part (85) and a minimum radius (Rmin) of the streamlined part (85), the maximum radius (Rmax) being defined as a distance between a point of the streamlined part (85) furthest from the principal axis (X) of the aeronautical propulsion system, and the principal axis (X) of the aeronautical propulsion system,and the minimum radius (Rmin) being defined as a distance between a point on the leading edge (86) of the airfoil (85) closest to the principal axis (X) of the aeronautical propulsion system, and the principal axis (X) of the aeronautical propulsion system, or in the case where the blade has variable pitch, the minimum radius (Rmin) being defined as a distance between a point on the leading edge (86) of the airfoil (85) closest to the principal axis (X) of the aeronautical propulsion system, and the principal axis (X) of the aeronautical propulsion system, when the blade (84) is positioned with a pitch angle in which the blade (84) is feathered, and in which the chord length, local (C) is defined as a distance between a point on the leading edge (86) and a point on the trailing edge (87), the point on the leading edge and the point on the trailing edge being located at the same radial distance from the principal axis (X) as the point considered, characterized in that the profiled part (85) comprises a core (111) of composite material and an added piece (112) fixed to the core (111) and forming the leading edge (86) of the profiled part (85), and in which the protrusions (95) and / or hollows (96) are formed only on the added piece (112).

2. Aeronautical propulsion according to claim 1, wherein the textured surface portion (90) is a portion of the upper surface and the series of protrusions (95) and / or hollows (96) consists of an alternation of protrusions (95) and hollows (96),

3. Aeronautical propulsion according to claim 1 or 2, wherein for any point considered of the textured surface portion (90), the thickness (e) of the profiled part (85) at the point considered is less than 12% of the local chord (C), preferably less than or equal to 10% of the local chord (C), or even more preferably less than or equal to 7% of the local chord (C).

4. Aeronautical propulsion according to any one of claims 1 to 3, wherein for a given radial distance, the profiled part (85) has a maximum thickness (e) at a given point on the intrados surface (88) or the extrados surface (89), and any point considered on the textured surface portion (90) is located upstream of the point where the thickness of the profiled part is maximum, in the direction of gas flow through the aeronautical propulsion.

5. Aeronautical propulsion according to any one of claims 1 to 4, wherein for a given radial distance, the profiled part (85) has a maximum thickness (e) at a given point on the intrados surface (88) or the extrados surface (89), the given point being located at a distance from the leading edge of between 10% and 45% of the local chord (C), preferably between 15% and 30% of the local chord (C).

6. Aeronautical propulsion according to any one of claims 1 to 5, wherein the added part (112) is a metal reinforcement part fixed to the core (111) or a heating mat fixed to the core (111).

7. Aeronautical propulsion according to any one of claims 1 to 6, wherein the protrusions (95) and / or the hollows (96) are located at a radial distance from the main axis (X) greater than a radial distance from the maximum local chord (C) of the blade (84) with respect to the main axis (X), or greater than the radial distance from a belly of the leading edge (86) with respect to the main axis (X).

8. Aeronautical propulsion according to any one of claims 1 to 7, wherein the lower surface (88) or the upper surface (89) comprises several textured surface portions (90a, 90b, 90c) having a series of protrusions (95) and / or hollows (96), and wherein a textured surface portion (90a) is separated from another textured surface portion (90b) nearest by a distance of between 1% and 35% of the wingspan (L), preferably between 2% and 15% of the wingspan of the streamlined portion (85).

9. Aeronautical propulsion according to any one of claims 1 to 8, wherein the projections (95) have a height (h) between 0.04% of the local chord (C), preferably 1%, and 3% of the local chord (C) and / or the hollows (96) have a depth (h) between 0.04% of the local chord (C), preferably 1%, and 3% of the local chord (C).

10. Aeronautical propulsion according to any one of claims 1 to 9, wherein the protrusions (95) and / or the hollows (96) form a repeating pattern having a constant pitch (w) between two consecutive protrusions (95) or between two consecutive hollows (96), the pitch being between 1 / 3 and 3 times a height (h) of a protrusion (95) or between 1 / 3 and 3 times a depth (h) of a hollow (96).

11. Aeronautical propulsion according to any one of claims 1 to 10, wherein the projections (95) comprise ribs.

12. Aeronautical propulsion according to claim 11, in which each rib has a base (97) and an edge (103) having a height (h) measured with respect to the base (97) which increases in a strictly monotonic manner from upstream to downstream in the direction of gas flow through the aeronautical propulsion.

13. An aeronautical propulsion system according to claim 12, wherein the edge (103) has an upstream end and a downstream end, and each rib comprises an upstream end face (101) extending from the base to the upstream end of the edge, forming a first non-zero angle (|3) with the base (97), lower or equal to 90°, preferably between 20° and 70°, and a downstream end face (102) extending from a downstream end of the edge to the base, forming a second non-zero angle (rp) with the base (97), less than or equal to 90°, preferably between 20° and 70°.

14. Aeronautical propulsion according to claim 13, wherein the upstream end face (101) has a flat portion extending from the base (97) and a rounded junction portion extending from the flat portion to the upstream end of the edge (103).

15. Aeronautical propulsion according to any one of claims 13 and 14, wherein the downstream end face (101) has a rounded junction portion extending from the downstream end of the edge (103) and a flat portion extending from the rounded junction portion to the base (97).

16. Aeronautical propulsion according to any one of claims 11 and 12, wherein each rib has a cross-section of triangular shape, with a base (97) and a vertex (98), the vertex (98) having a height (h) measured relative to the base (97) greater than the width (t) of the base (97).

17. Aeronautical propulsion according to any one of claims 11 to 16, wherein the textured surface portion (89) comprises first zones (90A) and second zones (90B) arranged alternately with the first zones (90A) along a radial direction with respect to the principal axis (X), and wherein each first zone (90A) has a series of first ribs, oriented with a first angle (ai) with respect to the principal axis (X) of the aeronautical propulsion, and each second zone (90B) has a series of second ribs oriented with a second angle (a2) with respect to the principal axis (X), the second angle being different from the first angle.

18. Aeronautical propulsion according to claim 17, wherein the first angle (aj) and the second angle (a2) are adjacent and the first angle is between +15° and +45° with respect to the principal axis (X) and the second angle is between -15° and -45° with respect to the principal axis (X).

19. Aeronautical propulsion according to any one of claims 17 and 18, wherein the alternation of first zones (90A) and second zones (90B) has a spatial period (dr) between 5% and 20% of the local chord (C).

20. Aeronautical propulsion according to any one of claims 17 and 18, wherein each first rib converges towards a respective second rib in the upstream-downstream direction.

21. Aeronautical propulsion according to any one of claims 1 to 10, wherein the hollows (96) comprise cavities each having a portion-sphere shape.

22. Aeronautical propulsion according to any one of claims 1 to 21, comprising a fan (8) or a propeller, and wherein the blade (84) is a blade of the fan (8) or of the propeller of the aeronautical propulsion (1).

23. Aeronautical propulsion according to claim 22, comprising a turbine (13) for driving the fan (8) or propeller, a fan shaft or propeller shaft connected to the fan or propeller, a turbine shaft connected to the turbine (13), and a reduction mechanism having an inlet connected to the turbine shaft and an output connected to the fan shaft or propeller shaft, such that in operation, the fan (8) or propeller is driven in rotation by the turbine (13) at a rotational speed lower than a rotational speed of the turbine (13).

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