Aeronautical thruster

Abstract

The invention relates to an aeronautical thruster comprising a blade, the blade comprising a profiled part (85) having an aerodynamic profile extending radially from the central component, the profiled part (85) having a leading edge (86), a trailing edge (87), a lower surface and an upper surface (89), wherein the lower surface (88) or the upper surface (89) comprise a textured surface portion (90) having a series of projections and / or recesses, the textured surface portion (90) being defined such that any given point of the textured surface portion (90) is located at a radial distance from the main axis of the aeronautical thruster equal to the sum of a minimum radius (Rmin) and between 20% and 95% of a span of the profiled part (85), and is located at an axial distance from the leading edge comprised between 2% and 50% of the local chord length.

Description

[0001] DESCRIPTION

[0002] TITLE: AERONAUTICAL PROPELLER

[0003] FIELD OF THE INVENTION

[0004] The invention relates to an aeronautical propulsion system, such as a gas turbine engine for example.

[0005] STATE OF THE ART

[0006] Aircraft thrusters comprise fixed parts and rotating parts, which are rotated relative to the fixed parts when the aircraft thruster is in operation.

[0007] The rotating parts and the fixed parts each have blades.

[0008] Thus, gas turbine engines typically consist of a fan module or propeller module, a compressor module, a combustor, and a turbine module.

[0009] The fan module (or propeller module), the compressor module and the turbine module each include rotating parts (or "rotor") and fixed parts (or "stator").

[0010] For example, the fan module includes a fan stator and a fan rotor adapted to be rotated relative to the fan stator. The fan rotor includes one or more rows of moving blades. Rotation of the fan rotor compresses air, which is expelled rearward to produce a portion of the engine's thrust.

[0011] In addition, the fan stator typically includes a set of outlet guide vanes (also called "Outlet Guide Vanes" or "OGVs") located downstream of the fan rotor and acting as a rectifier. This set of fixed vanes serves to straighten and regulate the airflow downstream of the fan rotor to optimize engine thrust.

[0012] The airflow passing through the set of fixed blades flows generally between the fixed blades in an upstream-downstream direction.

[0013] However, under certain engine operating conditions, particularly in partial engine speed (i.e. during aircraft takeoff and landing), recirculation bubbles (separation / recirculation bubbles) and / or areas with partially detached flow may appear on the blade surfaces, particularly on the extrados surfaces. These recirculation bubbles generate noise.

[0014] On the one hand, these recirculation bubbles produce high-frequency lines (tonal noise), which are related to the presence of these bubbles and the associated vortex shedding. On the other hand, recirculation bubbles increase the boundary layer thickness, which has an impact on the fan rotor wake and the interaction noise with the fan stator or the downstream blade row. Indeed, a recirculation bubble that is too large increases the turbulent kinetic energy and the velocity deficit in the wake, which increases the interaction noise (broadband and tonal, respectively).

[0015] The same phenomenon can occur on unducted propellers of turboprops, Counter-Rotating Open Rotor (CROR) gas turbine engines and on unducted fans of Unducted Single Fan (USF) gas turbine engines.

[0016] This same phenomenon can also occur on other fixed or moving blades of an aeronautical propeller.

[0017] Document EP 3 467 258 A1 discloses in particular a blade for a gas turbine engine, the blade comprising a pressure surface and a suction surface, wherein the pressure surface or the suction surface comprises a roughness zone which is configured to provide greater flow resistance in a direction along the blade than in a direction across the blade.

[0018] STATEMENT OF THE INVENTION

[0019] An aim of the invention is to reduce the noise generated by a blade in the presence of a recirculation bubble and / or zones with partially detached flow which appear at certain speeds, without penalizing the operation of the aeronautical propeller at other speeds.

[0020] This object is achieved within the scope of the present invention by means of an aeronautical propeller comprising a fixed part and a rotating part suitable for being driven in rotation relative to the fixed part around a main axis of the aeronautical propeller, one of the fixed part and the rotating part comprising a central part and a blade, the blade comprising a profiled part having an aerodynamic profile extending radially from the central part, the profiled part having a leading edge, a trailing edge, a pressure surface and an extrados surface, in which the pressure surface or the extrados surface comprises a textured surface portion having a series of projections and / or hollows,the textured surface portion being defined such that any considered point of the textured surface portion is located at a radial distance from the main axis of the aeronautical propeller equal to the sum of a minimum radius and between 20% and 95% of a span of the profiled portion, and is located at a distance from the leading edge of between 2% and 50% of the local chord length, wherein the span of the profiled portion is defined as a difference between a maximum radius of the profiled portion and a minimum radius of the profiled portion, the maximum radius being defined as a distance between a point of the profiled portion furthest from the main axis of the aeronautical propeller, and the main axis of the aeronautical propeller, and the minimum radius being defined as a distance between a point of the leading edge of the profiled portion closest to the main axis of the aeronautical propeller, and the main axis of the aeronautical propeller,or in the case where the blade is of variable pitch, the minimum radius being defined as a distance between a point on the leading edge of the profiled part closest to the main axis of the aeronautical propeller, and the main axis of the aeronautical propeller, 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 point on the leading edge and the point on the trailing edge being located at the same radial distance from the main axis as the point considered.,

[0021] 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.

[0022] The presence of a series of protrusions and / or hollows in this area makes it possible to recreate a turbulent flow at this location, instead of a recirculation bubble in partial regime (i.e. during the takeoff and landing phases of the aircraft) and to reduce wall friction and aerodynamic losses in nominal regime (i.e. in cruise regime). This thus makes it possible to limit the noise generated by the fan during takeoff and landing, without impacting the performance of the aeronautical propeller in cruise regime.

[0023] The aeronautical propellant may further have one or more of the following characteristics:

[0024] - the textured surface portion is a portion of the extrados surface and the series of projections and / or hollows consists of an alternation of projections and hollows,

[0025] - the textured surface portion is defined such that any considered point of the textured surface portion is located at a radial distance from the main axis of the aeronautical propeller equal to the sum of a minimum radius and between 50% and 90% of a span of the profiled part,

[0026] - the textured surface portion is defined such that any considered point of the textured surface portion is located at a radial distance from the main axis of the aeronautical propeller greater than a radial distance from a point on the leading edge of the profiled part located furthest upstream, considering that the main axis of the propeller extends from upstream to downstream in the direction of flow of the gases through the aeronautical propeller, when the aeronautical propeller is in normal operation,

[0027] - the textured surface portion is defined such that any considered point of 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,

[0028] - for any point considered on the textured surface portion, the thickness of the profiled part 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,

[0029] - the thickness of the profiled part at the point considered is greater than 0.1% of the local chord, preferably greater than 0.25% of the local chord, or even preferably greater than 1% of the local chord,

[0030] - for a given radial distance, the profiled part has a maximum thickness at a given point on the intrados surface or the 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,

[0031] - for a given radial distance, the profiled part has a maximum thickness at a given point on the intrados surface or the 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,

[0032] - the profiled part comprises a core of composite material and an insert fixed to the core and forming the leading edge of the profiled part, and the projections and recesses are formed only on the insert,

[0033] - the insert is a metal reinforcement piece fixed to the core or a heating mat fixed to the core,

[0034] - the projections and / or hollows are located at a radial distance from the main axis greater than a radial distance of the maximum local chord of the blade from the main axis, or greater than the radial distance of a leading edge belly from the main axis,

[0035] - the projections and / or hollows are formed by machining in the added part,

[0036] - alternatively, the projections and / or hollows are also formed on the composite material core,

[0037] - the projections and / or hollows are formed in a film applied to the composite material core,

[0038] - the intrados surface or the extrados surface comprises several textured portions having a series of projections and / or depressions, and in which a portion of textured surface is separated from another portion of textured surface closest to it by a distance of between 1% and 35% of the span, preferably between 2% and 15% of the span,

[0039] - the projections have a height between 0.04% of the local chord, 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,

[0040] - the projections and / or recesses form a repetitive pattern with a constant pitch between two consecutive projections or between two consecutive recesses, 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,

[0041] - the hollows comprise cavities, each cavity having the shape of a portion of a sphere, for example a hemispherical shape,

[0042] - the projections include ribs,

[0043] - each rib has a base and an edge having a height measured relative to the base which increases in a strictly monotonous manner from upstream to downstream,

[0044] - the edge has an upstream end and a downstream end, and each rib comprises 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°,

[0045] - 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,

[0046] - the first angle is between 20° and 70°,

[0047] - the second angle is between 20° and 70°,

[0048] - the downstream end face has a rounded joining portion extending from the downstream end of the edge and a flat portion extending from the rounded joining portion to the base,

[0049] - each rib has a triangular cross-sectional area, with a base and a top, the top having a height measured from the base greater than the width of the base,

[0050] - 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 propeller, 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,

[0051] - the first angle and the second angle are adjacent and the first angle is 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,

[0052] - the alternation of first zones and second zones has a spatial period between 5% and 20% of the local chord,

[0053] - each first rib converges towards a respective second rib in the upstream-downstream direction,

[0054] - the aeronautical propeller comprises a ducted fan or an unducted propeller, and the blade is a blade of the ducted fan or of the unducted propeller of the aeronautical propeller,

[0055] - the blade has variable pitch,

[0056] - the aeronautical propeller comprises a fan or propeller drive turbine, a fan shaft or propeller shaft connected to the fan or propeller, a turbine shaft connected to the turbine, and a reduction mechanism having an inlet connected to the turbine shaft and an outlet connected to the fan shaft or propeller shaft, such that in operation, the fan or propeller is rotated by the turbine at a rotational speed lower than a rotational speed of the turbine.

[0057] PRESENTATION OF THE DRAWINGS

[0058] Other characteristics and advantages will emerge from the following description, which is purely illustrative and not limiting, and must be read in conjunction with the appended figures, among which:

[0059] - figure 1 schematically represents a first example of an aeronautical propeller comprising an unducted propeller module,

[0060] - figure 2 schematically represents a second example of an aeronautical propeller comprising a ducted fan module,

[0061] - figure 3 schematically represents, in longitudinal section, a gas turbine engine with a ducted fan,

[0062] - figure 4 schematically represents a fan or propeller blade,

[0063] - figure 5 schematically represents, in cross-section, the fan or propeller blade of figure 4,

[0064] - figure 6 schematically represents a fan or propeller blade according to an embodiment of the invention,

[0065] - figure 7 schematically represents, in cross-section, the fan or propeller blade of figure 6,

[0066] - figure 8 schematically represents a fan or propeller blade profile and the skeleton line in the section plane,

[0067] - Figure 9 is an enlarged view of a portion of textured surface of the fan or propeller blade, in accordance with a possible embodiment of the invention,

[0068] - figure 10A schematically represents the shape of a projection made in the extrados portion according to a first embodiment of the invention,

[0069] - figure 10B schematically represents the shape of a projection made in the extrados portion according to a second embodiment of the invention, - figure 11 schematically represents an alternation of projections and hollows made in the textured surface portion of the fan or propeller blade,

[0070] - figure 12 schematically represents the orientation angles of the projections, the projections having the shape of ribs,

[0071] - figure 13 schematically represents, in cross section, a blade comprising a projection,

[0072] - Figure 14 is an enlarged view of a portion of textured surface of the fan blade, in accordance with another possible embodiment of the invention,

[0073] - figure 15 schematically represents, in cross-section, the blade of figure 14,

[0074] - Figures 16A to 16F illustrate different configurations of the textured surface portion.

[0075] DETAILED DESCRIPTION OF AN EMBODIMENT

[0076] In Figure 1, the aeronautical propeller 1 shown is a gas turbine engine with unducted 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 Figure 1).

[0077] The gas turbine engine 1 comprises a nacelle 2 intended to be fixed to a fuselage of an aircraft, and an unducted fan 8. The fan 8 comprises two counter-rotating fan rotors 81 and 82. In other words, when the engine 1 is in operation, the rotors 81 and 82 are rotated relative to the nacelle 2 around the same axis of rotation X (which coincides with a main axis of the engine), in opposite directions.

[0078] In the example illustrated in Figure 1, the engine 1 is an unducted engine with counter-rotating fan rotors (also called a “contra-rotating open rotor” or “CROR” in English), in a “pusher” configuration. However, the invention is not limited to this configuration. The invention also applies to “Open Rotor” type engines, in a “puller” configuration (i.e. the fan is placed upstream of the power generator with an air inlet located before, between or just behind the two fan rotors). In addition, the invention also applies to engines having different architectures, such as an architecture comprising a fan rotor comprising moving blades (or an “Open Fan” in English) and a fan stator comprising blades (USF), or a single fan rotor.

[0079] The fan stator blades can be fixed or variable pitch. In this case, each of the blades is pivotally mounted relative to the nacelle 2 along a pitch axis.

[0080] The invention is applicable to turboprop-type architectures (comprising a single fan rotor).

[0081] In Figure 1, each fan rotor 81, 82 comprises a hub 83 rotatably mounted 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.

[0082] The blades 84 may be fixed or variable pitch. The fixed blades are mounted fixedly on the hub 83. In the case where the blades 84 are variable pitch, each blade 84 is pivotally mounted relative to the hub 83 around a respective pitch axis Y. The blades 84 are connected to a pitch change mechanism making it possible to adjust the pitch angle of the blades 84 relative to the hub 83, and thus their angle of incidence, depending on the flight phases.

[0083] In Figure 2, the aircraft propeller 1 shown is a gas turbine engine with a ducted fan. The gas turbine engine 1 shown comprises a nacelle 2 intended to be fixed to a fuselage of an aircraft, a fan 8 and a fairing surrounding the fan 8, the fairing being fixedly mounted on the nacelle 2. In the example illustrated in Figure 2, the fairing of the fan 8 is arranged inside the nacelle 2.

[0084] The fan 8 comprises a fan rotor 81 capable of being driven in rotation relative to the nacelle 2 around an axis of rotation X (which coincides with the main axis of the engine 1). The fan 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.

[0085] The fan 8 may be a variable pitch fan (called a “Variable pitch fan” or “VPF”), that is to say that the fan includes a mechanism for pivoting each blade 84 around a pitch axis so as to modify the pitch of the blades depending on the flight phases.

[0086] In Figure 3, the aircraft propeller 1 shown is a twin-spool, twin-flow gas turbine engine. The gas turbine engine may be a gas turbine engine having a high bypass ratio (“Ultra High Bypass Ratio” or “UHBR”), i.e. having a bypass ratio of between about 15 and about 40.

[0087] The gas turbine engine 1 has a main axis X (or longitudinal axis).

[0088] The gas turbine engine 1 comprises a nacelle 2, a fan module 3, a compressor module 4, a combustion chamber 5, and a turbine module 6.

[0089] In the example illustrated in Figure 3, the fan module 3 comprises a fan casing 7 mounted fixedly relative to the nacelle, a fan 8 capable of being driven in rotation relative to the fan casing 7. The fan casing 7 comprises fixed outlet vanes 9 (or “OGV”) having the function of straightening the secondary air flow which flows out of the fan 8.

[0090] In the example illustrated in Figure 3, the compressor module 4 comprises a low pressure compressor 10 and a high pressure compressor 11.

[0091] In addition, the turbine module 6 comprises a high pressure turbine 12 and a low pressure turbine 13.

[0092] The gas turbine engine 1 comprises a low pressure shaft 14 connecting the low pressure turbine 13 to the low pressure compressor 10 and to the fan 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 rotatably mounted relative to the nacelle 2, around the main axis X of the engine.

[0093] In one embodiment, the gas turbine engine 1 may comprise a fan shaft for rotating the fan 8 and a reduction mechanism having an inlet connected to the low-pressure shaft 14 and an outlet connected to the fan shaft. In this embodiment, the fan 8 is rotated at a speed lower than the rotational speed of the low-pressure turbine 13. The reduction mechanism thus makes it possible to independently optimize the rotational speed of the fan 8 and the rotational speed of the low-pressure turbine 13 and the low-pressure compressor 10. The fan module 3, the low-pressure compressor 10, the low-pressure turbine 13, and the low-pressure shaft 14 (and, if applicable, the reduction mechanism and the fan shaft) together form the low-pressure body of the engine 1.The low pressure turbine 13 is capable of driving the low pressure compressor 10 and the blower 8 in rotation via the low pressure shaft 14 (as well as, where appropriate, via the reduction mechanism and the blower shaft).

[0094] More specifically, the low-pressure compressor 10 comprises a low-pressure compressor casing 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 capable of being driven in rotation relative to the low-pressure compressor stator 18, around the main axis X of the engine 1. The low-pressure compressor rotor 17 comprises moving blades. The low-pressure compressor stator 18 comprises fixed blades (also called “guide blades” or “straightener blades”) which are fixedly mounted on the low-pressure compressor casing 16 by being interposed between the moving blades. These fixed blades have the function of guiding the primary air flow through the low-pressure compressor 10.

[0095] Likewise, the low pressure turbine 13 comprises a low pressure turbine casing 19, mounted fixedly relative to the nacelle 2, a low pressure turbine rotor

[0096] 20, and a low pressure turbine stator 21. The low pressure turbine rotor 20 is adapted to be rotated relative to the low pressure turbine stator

[0097] 21, around the main axis X of the engine 1. The low pressure turbine rotor 20 comprises moving blades. The low pressure turbine stator 21 comprises fixed blades which are fixedly mounted on the low pressure turbine casing 19 by being interposed between the moving blades. These fixed blades have the function of guiding the primary air flow through the low pressure turbine 13.

[0098] The low pressure turbine rotor 20 is connected to the low pressure compressor rotor 17 via the low pressure shaft 14. Thus, when the engine 1 is in operation, rotation of the low pressure turbine rotor 20 causes rotation of the low pressure compressor rotor 17.

[0099] The high-pressure compressor 11, the high-pressure turbine 12 and the high-pressure shaft 15 together form the high-pressure body of the engine 1. The high-pressure turbine 12 is capable of driving the high-pressure compressor 11 in rotation by means of the high-pressure shaft 15. More specifically, the high-pressure compressor 11 comprises a high-pressure compressor casing 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 capable of being driven in rotation relative to the high-pressure compressor stator 24, around the main axis X of the engine 1. The high-pressure compressor rotor 23 comprises moving blades.The high-pressure compressor stator 24 comprises fixed vanes (also called "guide vanes" or "rectifier vanes") which are fixedly mounted on the casing 22 of the high-pressure compressor by being interposed between the moving vanes. These fixed vanes have the function of guiding the primary air flow through the high-pressure compressor 11.

[0100] In one embodiment, the blades of the high pressure compressor stator 24 may be variable pitch, in order to ensure the operability of the high pressure compressor 11 and increase its pumping margin.

[0101] Similarly, the high-pressure turbine 12 comprises a high-pressure turbine casing 25, mounted fixedly 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 capable of being 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 moving blades. The high-pressure turbine stator 27 comprises fixed blades which are mounted fixedly on the casing 25 of the high-pressure turbine by being interposed between the moving blades. These fixed blades have the function of guiding the primary air flow through the high-pressure turbine 12.

[0102] The high pressure turbine rotor 26 is connected to the high pressure compressor rotor 23 via the high pressure shaft 15. Thus, when the engine 1 is in operation, rotation of the high pressure turbine rotor 26 causes rotation of the high pressure compressor rotor 23.

[0103] The fixed outlet vanes 9 of the fan module 3, the fixed stator vanes 18 of the low pressure compressor 10, the fixed stator vanes 24 of the high pressure compressor 11, the fixed stator vanes 27 of the high pressure turbine 12 and the fixed stator vanes 21 of the low pressure turbine 13 are examples of guide vanes.

[0104] When the engine 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.

[0105] Air is drawn in by the fan 8. The air drawn in by the fan 8 is divided into a primary air flow and a secondary air flow, which flow from upstream to downstream of the gas turbine engine 1.

[0106] The primary air flow flows from upstream to downstream of the gas turbine engine 1 in a primary vein, 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 air flow through the high-pressure turbine 12 and the low-pressure turbine 13 causes rotation of the rotors 26 and 20 of the turbines which in turn drive rotation of the rotors 23 and 17 of the high-pressure and low-pressure compressors, as well as the fan 8 via the high-pressure shaft 15 and the low-pressure shaft 14. The primary air flow escapes from the engine 1 through an exhaust casing 28, located downstream of the low-pressure turbine casing 19.

[0107] The secondary airflow (also called "bypass airflow" or "bypass flow" in English) flows from upstream to downstream of the gas turbine engine 1 in a secondary vein. This secondary airflow does not pass into 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 body and to generate the majority of the thrust provided by the gas turbine engine. The secondary airflow flows through the fixed blades 9 mounted on the fan casing 7, downstream of the fan 8.

[0108] Figures 4 and 5 schematically represent a fan or propeller blade 84.

[0109] However, the invention also applies to other blades of the aeronautical propeller, such as turbine or compressor blades for example. The blade may be a blade of a rotating part of the aeronautical propeller or a blade of a fixed part of the aeronautical propeller.

[0110] In Figures 4 and 5, the fan blade 84 includes a profiled portion 85 having an aerodynamic profile extending radially from the hub 83.

[0111] The profiled portion 85 has a leading edge 86, a trailing edge 87, a pressure surface 88 and an extrados surface 89. Furthermore, in the example illustrated in FIGS. 4 and 5, the fan blade 84 is rotatably mounted relative to the hub 83 around a setting axis Y, which makes it possible to modify the angle of incidence of the profiled portion 85 relative to the main axis X of the engine, and consequently relative to a direction of the incoming air flow.

[0112] As illustrated in Figure 4, a minimum radius R m in of the profiled portion 85 is defined as a distance between a point of the leading edge 86 of the profiled portion 85 closest to the main axis X of the aeronautical propeller, and the main axis X of the aeronautical propeller.

[0113] In the case where the blade is variable-pitch, the minimum radius Rmin of the profiled portion 85 being defined as a distance between a point of the leading edge 86 of the profiled portion 85 closest to the main axis X of the aeronautical propeller, and the main axis X) of the aeronautical propeller, when the blade 84 is positioned with a pitch angle in which the blade 84 is in a feathered position.

[0114] A variable pitch 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 left free to orient itself naturally in the airflow passing through the fan from upstream to downstream parallel to the main axis of the aeronautical propeller, the rotating part not being driven in rotation.

[0115] A maximum radius R ma x is defined as a distance between a point of the profiled portion 85 furthest from the main axis X of the aeronautical propeller, and the main axis of the aeronautical propeller.

[0116] A span L of the profiled part 85 is defined as a difference between the maximum radius R ma x of the profiled part 85 and the minimum radius Rmin of the profiled part 85.

[0117] In Figure 4, the section plane A-A' is a plane parallel to the main axis of the engine X and which intersects the profiled part 85 at a point on the leading edge 86 and a point on the trailing edge 87, the section plane A-A' being orthogonal to the setting axis Y.

[0118] As illustrated in Figure 5 showing a section of the profiled portion 85 along the section plane A-A', 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 point on the leading edge and the point on the trailing edge being located at the same radial distance from the main axis X of the engine. 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 main axis X of the engine.

[0119] In Figure 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 engine, and the main axis of the engine.

[0120] In the case of a variable pitch blade, when the blade is feathered, the pitch angle is generally equal to approximately 90° (to within 15°).

[0121] As seen in Figures 6 and 7, the extrados surface 89 of the profiled portion 85 has a textured surface portion 90 having alternating projections and hollows. The remainder of the extrados surface 89, extending outside the textured surface portion 90, is smooth, that is to say it has no projections or hollows. The arrow D represents the direction of the main flow.

[0122] The textured surface portion 90 is delimited by a first delimiting line 91, a second delimiting line 92, a third delimiting line 93 and a fourth delimiting line 94.

[0123] The textured surface portion 90 extends radially between the first boundary line 91 and the second boundary line 92.

[0124] The first delimitation line 91 is located at a radial distance from the main axis X of the aeronautical propeller 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.

[0125] Preferably, the first delimitation line 91 is located at a radial distance from the main axis X of the aeronautical propeller greater than or equal to the sum of the minimum radius Rmin and 50% 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 + 0.5 x L.

[0126] The second delimitation line 92 is located at a radial distance from the main axis X of the aeronautical propeller 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.

[0127] The second delimitation line 92 is located at a radial distance from the main axis X of the aeronautical propeller 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 R m in + 0.90 x L.

[0128] The textured surface portion 90 extends axially between the third boundary line 93 and the fourth boundary line 94.

[0129] 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.

[0130] Preferably, the third delimitation line 93 is defined as the set of points of 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 main axis X of the engine, greater than or equal to 0.05 x C, preferably greater than or equal to 0.1 x C.

[0131] 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 engine, less than or equal to 0.5 x C.

[0132] Preferably, the fourth delimitation 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 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 engine, less than or equal to 0.4 x C, preferably less than or equal to 0.3 x C.

[0133] The textured surface portion 90 thus defined is located in an area in which a recirculation bubble and / or a partially detached flow is likely to appear.

[0134] The presence of protrusions and / or hollows in this area makes it possible to recreate a turbulent flow at this location, instead of a recirculation bubble in partial regime (i.e. during the takeoff and landing phases of the aircraft) and to reduce wall friction and aerodynamic losses in nominal regime (i.e. in cruise regime). This thus makes it possible to limit the noise generated by the fan during takeoff and landing, without impacting the performance of the aeronautical propeller in cruise regime.

[0135] In the example illustrated in Figures 6 and 7, the alternation of projection and hollow extends over the entire portion of textured surface 90, that is to say from the first delimitation line 91 to the second delimitation line 92 and from the third delimitation line 93 to the fourth delimitation line 94.

[0136] Figure 7 schematically represents a profile of the profiled portion 85 of the blade 84, that is to say a section of the profiled portion 85 in a transverse section plane. The transverse section plane is defined as a plane parallel to the main axis X and containing a point on the leading edge and a point on the trailing edge located at an equal distance from the main axis X.

[0137] In the case where the blade 84 is a variable pitch blade, the cross-sectional plane can be defined as a plane orthogonal to the pitch axis Y.

[0138] As illustrated in Figure 7, in each cross-sectional plane, the profiled portion 85 has a thickness e ma x maximum at a given point on the extrados surface 89, located at an axial distance x em ax of the leading edge 86, measured parallel to the local chord. The fourth delimiting line 94 is located upstream of the point where the thickness of the profiled part is maximum. In other words, each of the points of the fourth delimiting line 94 is located at an axial distance from the leading edge 86 strictly less than x em ax.

[0139] As illustrated in Figure 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 portion 85 measured perpendicular to a skeleton line S.

[0140] 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 which constitute the centers of the inscribed circles. 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 preferably less than or equal to 7% of the local chord.

[0141] The thicknesses of the profiled part 85 are low, particularly in the upper part to reduce centrifugal forces and therefore promote the mechanical strength of the blade 84.

[0142] Figure 9 schematically represents an alternation of projections 95 and depressions 96 formed on the textured surface portion 90.

[0143] In this example, the projections 95 have the shape of ribs, in relief relative to the rest of the extrados surface 89.

[0144] 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 propeller.

[0145] Each first zone 90A has a series of first ribs, oriented at a first angle relative to the main axis X of the aeronautical propeller, and each second zone 90B has a series of second ribs oriented at a second angle relative to the main axis X of the aeronautical propeller, the second angle being different from the first angle.

[0146] The angle a (eu, 02, ... ) of a rib can be defined as the angle between the main axis X of the aeronautical propeller and the projections or ribs (when these are projected in a plane passing through the main axis X and crossing -partially- the blade). If the blade is of 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 unducted propeller, this corresponds to a pitch angle y, as defined in Figure 5, which varies between 60° and 70° on the cross section at 0.75 x R ma x.

[0147] As illustrated in Figure 12, the first angle eu can be between +15° and 45°. The second angle can be symmetrical with respect to the first angle, i.e. the second angle 02 is between -45° and -15°. In a preferred embodiment, 02 = -eu. The arrow D represents the direction of the main flow.

[0148] In this way, the first ribs and the second ribs together form a herringbone pattern. Figure 10A schematically shows a first example of rib 95.

[0149] The rib 95 has a base 97 facing the extrados surface and a top 98 located at a distance from the base 97.

[0150] 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.

[0151] 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 top 98.

[0152] In a variant illustrated on the right of Figure 10A, the rib 95 may have a rounded apex 98.

[0153] Figure 11 schematically represents a second example of rib 95. Arrow D represents the direction of the main flow.

[0154] The rib 95 has a base 97 facing the extrados surface and a top 98 located at a distance from the base 97.

[0155] In this second example, the rib 95 has a triangular cross-section. That is to say, the rib 95 has in cross-section a thickness t which decreases continuously (or in a strictly monotonic manner) from its base 97 to its apex 98.

[0156] In a variant illustrated on the right of Figure 10B, the rib 95 may have a rounded apex 98.

[0157] In these two examples, the top 98 of the rib 95 has a height h measured relative to the base 97, greater than the width t of the base 97.

[0158] In addition, each rib 95 has an upstream end face 101 and a downstream end face 102.

[0159] Each rib 95 has an edge 103 extending from the upstream end face 101 to the downstream end face 102.

[0160] Edge 103 has an upstream end and a downstream end.

[0161] The upstream end face 101 has a planar portion extending from the base 97 and a rounded junction portion extending from the planar portion to the upstream end of the edge 103. The planar portion forms a first non-zero angle P with the base 97. The first angle P is less than or equal to 90°.

[0162] Preferably, the first angle P is between 20° and 70°.

[0163] The downstream end face 102 has a rounded joining portion extending from the downstream end of the edge 103 and a planar portion extending from the rounded joining 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°.

[0164] Preferably, the second angle i is between 20° and 70°.

[0165] 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.

[0166] 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 > 0.01 xC), 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 the boundary layers and / or recirculation bubbles which may appear in partial regime on the extrados of the fan blades near the leading edge.

[0167] As illustrated in Figure 11, the projections and recesses form a repeating pattern having a constant pitch w between two consecutive projections 95 or between two consecutive recesses 96, the pitch being between 1 / 3 and 3 times a height of a projection or between 1 / 3 and 3 times a depth of a recess. In other words, 1 / 3 < w / h < 3.

[0168] The pitch w is preferably between 1% and 3% of the local chord, i.e. 0.01 < w / C < 0.03.

[0169] As illustrated in Figure 12, the radial width of a rib pattern is between 5% and 20% of the chord. This rib pattern or pattern may be repeated identically or homothetically in one or more directions (radial, axial, etc.) on the textured surface portion 90.

[0170] As illustrated in Figure 13, the base 97 of the ribs 95 may be located below the level of the smooth portion of the extrados surface 89 which extends out of the portion of the textured surface 90, while the top of the ribs 95 may be located above the level of this smooth portion.

[0171] In the embodiment illustrated in Figures 14 and 15, the textured surface portion 90 has depressions 96. The remainder of the extrados surface 89, extending outside the textured surface portion 90, is smooth, i.e., it does not have any projections or depressions.

[0172] In this embodiment, the hollows 96 have a shape of a portion of a sphere, for example a hemispherical shape. Each hollow 96 has a depth h and a diameter t, the depth h preferably being equal to 0.5 x t.

[0173] The hollows 96 form a repeating pattern having a constant pitch w between two consecutive or adjacent hollows 96, the pitch being between 1 / 3 and 3 times the depth h of a hollow.

[0174] Furthermore, the pitch w is preferably between 1% and 3% of the local chord.

[0175] The hollows 96 are arranged to form first rows parallel to each other and oriented at a first angle relative to the main axis X of the aeronautical propeller, and second rows parallel to each other and oriented at a second angle relative to the main axis X of the aeronautical propeller, different from the first angle.

[0176] As illustrated in Figure 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°.

[0177] Thus, the second rows form an angle equal to 2ai with the first rows.

[0178] Figures 16A-16F illustrate different configurations of the textured surface portion 90.

[0179] In a first configuration illustrated in Figure 16A, the profiled portion 85 comprises a core 111 made of composite material and an insert 112 attached to the core 111. The insert 112 forms the leading edge 86 of the profiled portion 85.

[0180] The insert 112 may be a metal part, such as a metal reinforcement part fixed to the core 111, or a heating mat fixed to the core 111 (in particular, in the case of a propeller blade, the heating mat provides a de-icing function for the area of ​​the leading edge most exposed to frost).

[0181] The projections 95 and / or the recesses 96 are formed only on the insert 112, all along the leading edge 86 from the first delimiting line 91 to the second delimiting line 92 delimiting the textured surface portion 90. This makes it possible to easily machine the patterns of the projections and / or recesses on the surface of the blade, as well as to limit the wear of the patterns which may appear in service.

[0182] In a second configuration illustrated in Figure 16B, the extrados surface 89 comprises several textured surface portions 90a, 90b, 90c having projections and / or depressions. The textured surface portions 90a, 90b, 90c are arranged, one at the pick of the other, along the leading edge 86. The projections 95 and / or the depressions 96 are formed only on the insert 112. Each textured surface portion 90a, 90b, 90c is separated from a nearest textured surface portion by a distance (in the radial direction) bi, b2 of between 1% and 35% of the span L of the profiled portion 85, preferably of between 2% and 15% of the span of the profiled portion 85.

[0183] In a third configuration illustrated in Figure 16C, the projections 95 and / or the depressions 96 are formed only on the insert 112, in a portion of the insert 112 located near an end radially farthest from the main axis X of the aeronautical thruster.

[0184] In a fourth configuration illustrated in Figure 16D, the leading edge 86 is formed partly by the metal reinforcement 112 and partly by the composite material core 111.

[0185] More specifically, a portion of the leading edge 86 further from the main axis of the aeronautical propeller is formed by the composite material core 111, and a portion of the leading edge 86 closer to the main axis of the aeronautical propeller is formed by the metal reinforcement 112.

[0186] The projections 95 and / or the depressions 96 are formed only on the metal insert 112, in a portion of the insert 112 located near an end furthest radially from the main axis X of the aeronautical propeller.

[0187] In a fifth configuration illustrated in Figure 16E, the leading edge 86 is formed partly by the metal reinforcement 112 and partly by the composite material core 111.

[0188] More specifically, a portion of the leading edge 86 closer to the main axis of the aeronautical propeller is formed by the composite material core 111, and a portion of the leading edge 86 further from the main axis of the aeronautical propeller is formed by the metal reinforcement 112.

[0189] The projections 95 and / or the depressions 96 are formed only on the insert 112, in a portion of the insert 112 located near an end furthest radially from the main axis X of the aeronautical propeller.

[0190] When the projections 95 and / or the recesses 96 are formed on the insert 112, these projections and recesses may be formed by machining directly into the metal reinforcement 112.

[0191] In a sixth configuration illustrated in Figure 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 part is not necessary for reasons of mechanical resistance to bird ingestion. In Figures 15C, 15D, 15E and 15F, the projections 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 belly of the leading edge 86, in an area where separations in partial regime are likely to occur. When the projections 95 and / or the recesses 96 are formed on the composite material core 111, these projections and / or these recesses can be formed directly in the composite material or in a film which is applied to the extrados surface 89, for example by gluing.

Claims

CLAIMS 1. Aeronautical thruster (1) comprising a fixed part (7) and a rotating part (8) capable of being driven in rotation relative to the fixed part (7) around a main axis (X) of the aeronautical thruster (1), one of the fixed part (7) and the rotating part (8) comprising a central part (83) and a blade (84), the blade (84) comprising a profiled portion (85) having an aerodynamic profile extending radially from the central part (83), the profiled portion (85) having a leading edge (86), a trailing edge (87), a pressure surface (88) and an extrados surface (89), wherein the pressure surface (88) or the extrados surface (89) comprises a textured surface portion (90) having a series of projections (95) and / or depressions (96), the textured surface portion (90) being defined such that any considered point of the textured surface portion (90) is located at a radial distance from the main axis (X) equal to the sum of a minimum radius (Rmin) and between 20% and 95% of a span (L) of the profiled portion (85), and is located at a distance from the leading edge of between 2% and 50% of the local chord length (C), wherein the span (L) of the profiled portion (85) is defined as a difference between a maximum radius (Rma x) of the profiled part (85) and a minimum radius (Rmin) of the profiled part (85), the maximum radius (R ma x) being defined as a distance between a point of the profiled part (85) furthest from the main axis (X), and the main axis (X), and the minimum radius (Rmin) being defined as a distance between a point of the leading edge (86) of the profiled part (85) closest to the main axis (X), and the main axis (X), or in the case where the blade is variable pitch, the minimum radius (Rmin) being defined as a distance between a point of the leading edge (86) of the profiled part (85) closest to the main axis (X), and the main axis (X), when the blade (84) is positioned with a pitch angle in which the blade (84) is feathered, and in which 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), the point on the leading edge and the point on the trailing edge being located at the same radial distance from the main axis (X) as the point considered.

2. Aeronautical thruster according to claim 1, in which the textured surface portion (90) is a portion of the extrados surface and the series of projections (95) and / or hollows (96) consists of an alternation of projections (95) and hollows (96), 3. Aeronautical thruster according to one of claims 1 and 2, in which for any point considered on 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 preferably less than or equal to 7% of the local chord (C).

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

5. Aeronautical propeller according to one of claims 1 to 4, in which 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 propeller according to one of claims 1 to 5, in which the profiled part (85) comprises a core (111) made of composite material and an insert (112) fixed to the core (111) and forming the leading edge (86) of the profiled part (85), and in which the projections (95) and / or the hollows (96) are formed only on the insert (112).

7. Aeronautical propeller according to claim 6, in which the added part (112) is a metal reinforcement part fixed to the core (111) or a heating mat fixed to the core (111).

8. Aeronautical propeller according to one of claims 6 and 7, in which the projections (95) and / or the hollows (96) are located at a radial distance from the main axis (X) greater than a radial distance of the local chord maximum (C) of the blade (84) from the main axis (X), or greater than the radial distance of a leading edge antinode (86) from the main axis (X).

9. Aeronautical thruster according to one of claims 1 to 8, in which the intrados surface (88) or the extrados surface (89) comprises several textured surface portions (90a, 90b, 90c) having a series of projections (95) and / or hollows (96), and in which a textured surface portion (90a) is separated from another textured surface portion (90b) closest to it by a distance of between 1% and 35% of the span (L), preferably of between 2% and 15% of the span of the profiled part (85).

10. Aeronautical thruster according to one of claims 1 to 9, in which the projections (95) have a height (h) of 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) of between 0.04% of the local chord (C), preferably 1%, and 3% of the local chord (C).

11. Aeronautical propeller according to one of claims 1 to 10, in which the projections (95) and / or the hollows (96) form a repetitive pattern having a constant pitch (w) between two consecutive projections (95) or between two consecutive hollows (96), the pitch being between 1 / 3 and 3 times a height (h) of a projection (95) or between 1 / 3 and 3 times a depth (h) of a hollow (96).

12. Aeronautical propeller according to one of claims 1 to 11, in which the projections (95) comprise ribs.

13. An aeronautical propeller according to claim 12, wherein each rib has a base (97) and an edge (103) having a height (h) measured relative to the base (97) which increases strictly monotonically from upstream to downstream in the direction of flow of the gases through the aeronautical propeller.

14. An aeronautical propeller according to claim 13, 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 by forming a first non-zero angle (P) with the base (97), less than 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, by forming a second non-zero angle (i) with the base (97), less than or equal to 90°, preferably between 20° and 70°.

15. Aeronautical thruster according to claim 14, in which 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).

16. Aeronautical thruster according to one of claims 14 and 15, in which 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).

17. Aeronautical propeller according to one of claims 12 and 13, in which each rib has a triangular cross-sectional section, with a base (97) and a top (98), the top (98) having a height (h) measured relative to the base (97) greater than the width (t) of the base (97).

18. Aeronautical thruster according to one of claims 12 to 17, 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 relative to the main axis (X), and wherein each first zone (90A) has a series of first ribs, oriented at a first angle (ai) relative to the main axis (X), and each second zone (90B) has a series of second ribs oriented at a second angle (02) relative to the main axis (X), the second angle being different from the first angle.

19. Aeronautical thruster according to claim 18, wherein the first angle (ai) and the second angle (02) are adjacent and the first angle is between +15° and +45° relative to the main axis (X) and the second angle is between -15° and -45° relative to the main axis (X).

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

21. Aeronautical propeller according to one of claims 18 to 19, in which each first rib converges towards a respective second rib in the upstream-downstream direction.

22. Aeronautical propeller according to one of claims 1 to 11, in which the hollows (96) comprise cavities each having the shape of a portion of a sphere.

23. Aeronautical propeller according to one of claims 1 to 22, comprising a fan (8) or a propeller, and in which the blade (84) is a blade of the fan (8) or of the propeller of the aeronautical propeller (1).

24. An aeronautical propeller according to claim 23, comprising a turbine (13) for driving the fan (8) or the propeller, a fan shaft or a propeller shaft connected to the fan or the propeller, a turbine shaft connected to the turbine (13), and a reduction mechanism having an inlet connected to the turbine shaft and an outlet connected to the fan shaft or the propeller shaft, such that in operation, the fan (8) or the propeller is rotated by the turbine (13) at a rotational speed lower than a rotational speed of the turbine (13).

Images