Propellant for an aircraft

Abstract

The invention relates to an aeronautical propeller comprising a central engine (3), an upstream series and a downstream series of blades (9, 6), the blades (90, 60) of at least one of these series being adapted to be able to be driven in rotation around the central longitudinal axis (X) by a turbine, a nacelle (5) which contains the central engine. At least some of the blades of the two series are variable-pitch. At least one of the blades (90) of the first series of blades (9) has a trailing edge having undulations (93), and / or at least one of the blades (60) of the second series of blades (6) has a leading edge having undulations (63). Figure to be published with the abstract: [Fig. 5]

Description

Title of the invention: Propellant for an aircraft Technical field of the invention

[0001] The invention relates to an aeronautical propeller, in particular for an airplane, along which a gas flow can circulate from upstream to downstream, the propeller having a central longitudinal axis (X), and comprising: - a first series of blades, - a second series of blades positioned downstream of the first series of blades, the first and second series of blades each defining an unducted propeller, including at least one unducted rotor ("open rotor" or "unducted fan" in English) which can be driven in rotation, - (at least) one motor (which can be called a central motor), for said rotational drive around the central longitudinal axis of the blades of at least one of the first series of blades and the second series of blades, and - a nacelle which can contain the engine.

[0002] The expression “non-ducted” therefore corresponds to the Anglo-Saxon expressions “open” (like open rotor) or “unducted” (like unducted fan).

[0003] Hereinafter the terms blade and vane designate the same thing.

[0004] The engine may be a thermal engine, in particular a turboshaft engine, a turbojet engine, a low bypass ratio turbofan engine, a high bypass ratio turbofan engine, a geared or speed reduction gearbox turbofan engine, a turbojet engine with counter-rotating turbines, an electric motor, a hydrogen engine, or a hybrid engine: thermal and / or electric and / or hydrogen.

[0005] The use of several engines is therefore of course not excluded. As energy source(s) for the engine(s), we can cite kerosene-based fuels, aviation gasoline, diesel, aviation biofuels, electricity, hydrogen.

[0006] The invention is therefore applicable in particular: - to turbomachines of the “open rotor(s)” type (for example, a “Contra-Rotating Open Rotor”, CROR: contra-rotating unducted propellers) and “open rotor-stator” type (for example, an “Unducted Single Fan”, USF: single unducted propeller or fan followed by a rectifier), hereinafter “unducted rotor(s)” or “unducted rotor and stator” forming a propeller system, - to turbomachines which may be a turboprop, but which may alternatively be a turbojet with a counter-rotating fan or with unducted rotor and stator pair.

[0007] It is recalled that in aeronautics a turbomachine is a gas turbine(s) propeller. State of the prior art

[0008] Particularly among gas turbine engines, some are known which use an architecture of the type with unducted rotor(s) or with unducted rotor and stator.

[0009] For example, a turbofan operates on the principle that a gas turbine engine drives a fan, the fan being located at a radial location between an engine nacelle and the engine hub.

[0010] An engine with unducted rotor(s) or unducted rotor and stator operates differently, with the fan located, radially to the axis of rotation of the central engine, outside the nacelle of the central engine. This allows the use of fan blades (or propeller) which can be larger and capable of acting on a greater volume of air than for a ducted turbofan. This can improve the bypass ratio (BPR) and the propulsive efficiency compared to conventional engines.

[0011] On a gas turbine engine(s), the invention detailed below applies here whether the said unducted rotor(s) or unducted rotor and stator are arranged upstream of the combustion chamber (configuration called "puller" or tractor) or downstream of it (configuration called "pusher" or pusher).

[0012] In “puller” configuration at least the first series of blades: - is rotated around the central longitudinal axis, and - is located at the engine compressor(s) and / or the speed reduction box (“gearbox”, in English), if there is one and if the engine is a gas turbine engine.

[0013] In “pusher” configuration at least the first series of blades: - is rotated around the central longitudinal axis, and - is located at the level of the engine turbine(s) if the engine is a gas turbine engine.

[0014] In each of these two cases, within the propeller system(s), it may be envisaged to place the power turbine of the central engine which drives the rotor(s), upstream, downstream or at the level of these counter-rotating propellers or of a rotor propeller and stator rectifier pair.

[0015] This is also applicable for the position of a speed reduction box (for example, an epicyclic differential reducer, as disclosed by EP2521851) in the case of a central engine with a gear system with respect to the rotor propeller(s).

[0016] Indeed, on a gas turbine engine(s), in particular for a CROR case, it may be very relevant to interpose a speed reduction box between the blades considered and the (or one of the) turbine(s), so that the blades of the upstream and / or downstream propeller in cause them to rotate at a lower speed compared to the (or one of the) turbine(s).

[0017] This is also applicable for the position of the epicyclic gear train in the case of a turboprop (with epicyclic gear train) with respect to the rotor propeller(s).

[0018] Thus, in the field, an aeronautical thruster is known along which a gas flow can circulate from upstream to downstream, the thruster having a central longitudinal axis (X), and comprising: - a first series of blades, - a second series of blades positioned downstream of the first series of blades, - drive means (sometimes referred to hereinafter as first means) for driving the blades of at least one of the first series of blades and the second series of blades in rotation around the central longitudinal axis, - a nacelle which has an aerodynamic external surface with respect to which the first series of blades and the second series of blades project, radially to the central longitudinal axis, each blade of the first series of blades and of the second series of blades having: — a free end opposite a connection end forming a blade root close to the nacelle, — an intrados face and an extrados face, — a chord C located at a defined radius centered on the central longitudinal axis, — a radius Ri between the central longitudinal axis and a location which is level with the external surface of the nacelle, this location being located: — on the blade or — on a blade setting arm, — a radius Re between the central longitudinal axis and a location on the free end of the blade furthest from the central longitudinal axis X, in a direction transverse to the central axis, and — a span (distance Re-Ri) defined, radially to the central longitudinal axis, between the free end and the connecting end, in said transverse direction, at least some of the blades of the first series of blades being of variable pitch (angle), so that each of them can pivot around a said pitch arm to which said blade is fixed, around a pitch axis which passes through the blade, and / or at least some of the blades of the second series of blades also being of variable pitch (angle), so that each of them can pivot around a said pitch arm to which said blade is fixed, around a pitch axis which passes through the blade.

[0019] The first series of blades and the second series of blades (or first propeller and second propeller) are therefore axially spaced from each other.

[0020] It has been understood that an aeronautical propeller is a device for producing energy which ensures, in the field of air navigation, the movement of a mobile and / or the operation of a motor.

[0021] In fact, we already know of aeronautical gas turbine engines (where the engine is often called a "core engine") which comprise drive means for pivoting the blade concerned via the pitch arm to which it is fixed, around its pitch axis.

[0022] As is known in other applications, this is a blade pitch change mechanism (PCM), connected to the blade pitch change arm at the blade root.

[0023] A problem encountered in any case still concerns the noise generated by the propeller. Mechanical strength, efficiency, aerodynamic performance may also be affected, as well as other aspects mentioned below.

[0024] In particular by improving efficiency and / or aerodynamic performance, an impact on combating global warming is also targeted. Presentation of the invention

[0025] The invention aims to respond to all or part of these problems, in a simple, reliable and inexpensive manner.

[0026] To this end, the invention therefore relates to an aeronautical propeller in accordance with that mentioned above, with first and second series of non-ducted blades (CROR / USF type for example) and which therefore comprises in particular: - a said first series of blades and a said second series of blades positioned downstream of the first series of blades, - first drive means for driving the first series of blades and / or the second series of blades in rotation around the central longitudinal axis, - second drive means for pivoting at least one of said blades via its wedging arm to which it is fixed, around its wedging axis, so that: — at least some of the blades of the first series of blades are variable pitch, and / or — at least some of the blades of the second series of blades are variable pitch.

[0027] In addition to the above, provision will be made on this aeronautical propeller: - that at least one of the blades of the first series of blades has a trailing edge (hereinafter sometimes referred to as BF) having undulations (serrations in English), and / or - that at least one of the blades of the second series of blades has a leading edge (hereinafter sometimes referred to as BA) having undulations.

[0028] The undulations comprise the alternating succession of at least two tooth peaks and two troughs (or bottoms).

[0029] The terms axial, radial and circumferential are defined relative to the X axis of the propellant.

[0030] Each pitch arm is the arm which rotates about an axis (the pitch axis) extending (this may be radially) across the central longitudinal axis and about which the blade, fixed to this arm, pivots to change the angle of attack of the gas flow which passes through the rotor propeller or the stator in question.

[0031] Each wedging axis can pass through a blade and a wedging arm.

[0032] Furthermore, the terms upstream (AM) and downstream (AV) are defined in relation to the meaning of circulation of gases within the propellant.

[0033] The first drive means may comprise, arranged in the nacelle, a drive motor rotating around the central longitudinal axis of the first series of blades and / or the second series of blades. A gas turbine engine is particularly targeted. But, as already mentioned, the engine may be in particular thermal (such as a turboshaft engine, turbojet, turbofan), electric, hydrogen, hybrid (in particular thermal and / or electric and / or hydrogen).

[0034] Said rotary drive engine can therefore comprise at least one compressor, a combustion chamber and at least one gas turbine and thus be of the aeronautical turbomachine type.

[0035] Can be provided: - a speed reducer for the rotation of the first series of blades and / or the second series of blades around the central longitudinal axis; and / or - both a “pusher” and “puller” configuration, particularly in USF configuration.

[0036] Transmission members are typically interposed between this rotating drive motor and the blades.

[0037] Arranged in the nacelle, said rotational drive motor, and possibly the transmission members, will then be enclosed in the nacelle.

[0038] Other features that can complement the above basic solution are presented below.

[0039] Some are included in this chapter “Presentation of the invention”, others only in the chapter “Detailed description of the invention”, in order to avoid repetitions.

[0040] As additional characteristics, it can already be noted, concerning the (first) means for driving in rotation around the central longitudinal axis of the blades of at least one of the first series of blades and second series of blades: - that they may include a gas turbine (in other words a gas turbine engine(s)), and / or - that these (first) drive means, or this motor, may include a speed reducer engaged with the blades of at least one of the first series of blades and second series of blades, to adapt the speed of rotation of said blades around the central longitudinal axis (X).

[0041] It should be noted in this regard that, on such a turbine engine(s), the speed reducer would be placed between the (or one of the) turbine(s) driving the blades of the first series of blades or of the second series of blades (or of the rotating drive shaft of the turbine in question) and the blades of the propeller concerned, to reduce their rotation speed.

[0042] A double turbine, axially high pressure then low pressure, could in particular be used.

[0043] In the case of a gas turbine engine, one or more compressors, one or more combustion chambers, one or more turbines driving the compressor(s), via one or more axial drive shafts, and one or more gas exhaust nozzles will traditionally be found in succession along the X axis.

[0044] In connection with all of the above, and in particular the case where the blades of the first series of blades are driven in rotation by said drive turbine via a speed reducer, a particularly interesting case could be that where the upstream / downstream propeller pair is in “puller” configuration: propeller pair located towards the upstream end of the nacelle of the central engine, upstream of the combustion chamber, whether in rotor / stator propeller pair configuration (upstream rotor and downstream stator) or rotor / rotor (upstream rotor and downstream rotor).

[0045] One of the advantages of a speed reducer between the turbine(s) and the unducted rotor(s) is to improve performance and therefore optimize the operation of each module of the aeronautical propeller. Furthermore, a "puller" configuration is compatible with both an installation of the aeronautical propeller under the wing (as on most commercial aircraft) or at the rear of the aircraft using a mast or pylon.

[0046] An interest may be found:

[0047] - that the blades of the first series of blades and the second series of blades are arranged to be able to be driven in rotation around the central longitudinal axis (X) by the turbine(s) considered, and

[0048] - that the turbine concerned is adapted to drive, around the longitudinal axis central (X), the first series of blades at a rotation speed greater than or equal to the rotation speed of the second series of blades.

[0049] It is also possible to note, concerning the blades, the following characteristics a) and following, to be considered independently or, in whole or in part, in combination: a) the first set of blades and the second set of blades can have different numbers of blades; This helps reduce noise because it prevents all the wakes from the upstream propeller (rotor) simultaneously impact the downstream propeller (rotor or stator) and therefore allow the acoustic radiation from the blades to be decorrelated; b) we can provide: - that each blade of the first series of blades and / or of the second series of blades has a maximum thickness between the intrados face and the extrados face, and - on the blades, that the maximum thickness of the blades (or between the intrados and extrados profiles which define the blade in question) is located near the blade root, at a distance less than 0.1x(Re - Ri) from the connection end (often called blade root); this helps to promote the mechanical strength of the blades in nominal operation, as well as in the event of bird ingestion. The mechanical forces on blades with BA / BF undulations could be higher than the nominal case with smooth BA / BF.In this case, this characteristic would be necessary in order to improve the mechanical strength of blades with undulations; A minimum value of the thickness of the blade at the embedding will then be preferred, dictated by mechanical reasons rather than by aero-acoustic reasons; c) each blade of the first series of blades and / or of the second series of blades has a hub-to-free end Ri / Re ratio such that Ri / Re is between 0.10 and 0.50; A hub ratio that varies in this range makes it possible to guarantee a good compromise between aerodynamic needs (large blades, therefore low Ri / Re) and the needs for integrating systems into the hub (PCM, oil pipes, etc., hence large Ri / Re); . d) we can provide: - that the blades of the first series of blades and / or the second series of blades are located on a closed line centered on the central longitudinal axis and having a diameter D, - that the pitch axes or the pitch arms of the blades of the first series of blades and of the second series of blades are axially separated, in pairs, by a distance S, and - that the S / D ratio is between 0.005 and 1, and preferably between 0.15 and 0.70; Increasing the axial spacing between the upstream and downstream propellers (rotor or stator) makes it possible to reduce noise in most cases, thanks to the dissipation of the turbulent wake of the upstream propeller during its propagation downstream. Thus, the wakes which arrive at the leading edge of the downstream propeller (rotor or stator) are less energetic; but increasing this distance between the propellers too much could pose problems of mass (longer turbomachine in the axial direction) and of integration of the turbomachine in the more complex aircraft.Furthermore, the trailing edge of each of the blades of the first series of blades is located longitudinally upstream of a leading edge of each of the blades of the second series of blades. This avoids interference between the series of blades. e) clipping may be beneficial, viz. a situation where the span L (or Re - Ri) of the upstream propeller (rotor) blades is greater than that of the downstream propeller (rotor or stator) measured along the pitch axis. Clipping (as [Fig.4] below) reduces noise because by reducing the span of the downstream propeller (rotor or stator) the impact of the upstream propeller (rotor) blade tip vortex, which is very energetic, can be avoided. From an aerodynamic point of view, clipping can however have a negative impact, because part of the gyration of the flow produced by the upstream propeller is not recovered by the downstream propeller, especially in the upper part.

[0050] In other words, we can then provide: - that the blades of the first series of blades are arranged so as to be able to be driven in rotation around the central longitudinal axis by said drive turbine of the aforementioned series or two series of blades, and - that the distance Re - Ri, measured from the connection end, of the blades of the first series of blades is greater than that of at least one of the blades of the blades of the second series of blades. f) the maximum diameter (2xRe) of the propeller (upstream or downstream) is between 1m and 6m.

[0051] In other words, the diameter of the blades of the first series of blades and the blades of the second series of blades will then be respectively inscribed on a first circle and a second circle, each with a circumference between 1m and 6m.

[0052] As before, this is a compromise to be found, particularly in relation to strength, weight, and noise.

[0053] It should also be noted that, on a leading edge or trailing edge zone having undulations, a variation in skeleton angle (A[31 below) at the leading edge or (A[32 below) at the trailing edge, between a tooth tip and a tooth trough (or bottoms), adjacent to each other, along the span of a blade, or radially to the central longitudinal axis (X), may favorably be greater than 0° and less than 45°, or even in certain cases 30°. This avoids cases of isolated profiles.

[0054] Indeed, it turned out: - that the angle of attack of the air perceived by the blade profile(s) at the level of the troughs and at the level of the peaks of the BA undulations is then suitable, and / or - that there are no stall phenomena that penalize the performance of the blade(s).

[0055] Other considerations of a comparable nature may usefully be given priority, namely at least one of the following considerations, it being specified that the reference “y” defines a blade pitch angle. This is the angle between a parallel (XI) to the central longitudinal axis X passing through the BA of the blade, at a given radius, and a straight line passing through the BA and the BF of this blade. - a variation in pitch angle (Ay), between a profile containing a tooth tip and a profile containing a tooth trough, adjacent to each other, — depending on the span (L) of the blade, or — radially to the central longitudinal axis (X), is less than 45°, and preferably greater than 0°; and / or, - on a leading edge or trailing edge area of ​​a blade having undulations, a variation in skeleton angle (A[31) at the leading edge or (A[32) at the trailing edge, between two tooth tips and / or two tooth troughs, adjacent to each other, — depending on the span (L) of the blade, or — radially to the central longitudinal axis (X), is less than 45°, and preferably greater than 0°; and / or, - a variation in the setting angle (Ay), between two adjacent profiles containing tooth peaks and / or between two profiles containing tooth valleys, adjacent to each other, — depending on the span (L) of the blade, or — radially to the central longitudinal axis (X), is less than 45°, and preferably greater than 0°; and / or, - a variation in the setting angle (Ay), between two adjacent profiles containing tooth peaks and / or between two profiles containing tooth valleys, adjacent to each other, — depending on the span (L) of the blade, or — radially to the central longitudinal axis (X), is less than 45°, and preferably greater than 0; and / or, - at least one of the blades of one of the propellers has a greatest deflection at a radial position located over a radial length of 0.4x(Re-Ri) of said free end of the blade considered (this consideration is preferably for the blades of the upstream propeller, but it would also be valid for the blades of the downstream propeller); and / or, - at least one of the blades of the downstream series of blades has a radius Re greater at the leading edge than at the trailing edge, when the leading edge and trailing edge lines are not coincident, and / or at least one of the blades of the first series of blades has a radius Re smaller (preferably) or greater at the leading edge than at the trailing edge, when the leading edge and trailing edge lines are not coincident. Brief description of the figures

[0056] [Fig.l] illustrates an aircraft powered by propeller-type unducted fan engines which may utilize the invention, [Fig.2] illustrates a direct drive turbine system that can drive the propellers of [Fig.l] [Fig.3] illustrates a polygonal ring that surrounds a turbine stage and supports the propeller (or rotor) blades, in connection with [Fig.2], [Fig.4] is a side view of a part of a propellant according to the invention, according to one embodiment, [Fig.5] is a side view of a part of a thruster according to the invention, according to another embodiment, [Fig.6] is a side view of a part of a thruster according to the invention, according to another embodiment, but with the air inlet of the primary flow located between the upstream propeller and the downstream propeller, as in the embodiment of [Fig.5], [Fig.7] is a side view of a part of a thruster according to the invention, according to another embodiment, [Fig.8] is a schematic view of a part of a blade root and its surroundings, in accordance with a possible embodiment, in the case of a variable-pitch stator blade, seen in section parallel to the axis of the motor, mounted in the nacelle, precisely in the casing of the nacelle, [Fig.9] is a sectional view of a blade according to the invention, according to a possible embodiment, [Fig. 10] shows a diagram from the front of a possible configuration of certain stator blades, [Fig. 1 1] shows a diagram of a possible evolution of several noise sources on a sound spectrum SPL (“Sound Pressure Level” in dB) as a function of the frequency f (in Hz), for a CROR configuration, with therefore two propellers, respectively upstream and downstream, counter-rotating, [Fig. 12] shows a possible evolution of several noise sources on a sound spectrum SPL (“Sound Pressure Level” in dB) as a function of the frequency (in Hz), for a USF configuration, with therefore an upstream rotor propeller and a downstream stator propeller, [Fig.13] at [Fig.26] schematically shows a side view of a blade according to the invention, according to several possible embodiments, [Fig.27] and [Fig.28] schematically show in side view a local external surface of a blade (at BF or BA) in accordance with the invention, according to two possible forms of undulations and construction of the body of the blade (with the presence of zones in acoustically damping foam or porous surface), [Fig.29] is a side view of a part of a thruster according to the invention, according to another embodiment, [Fig.30] illustrates a schematic longitudinal cross-section through a counter-rotating double-flow propeller of a gas turbine engine employing the invention, [Fig.31] illustrates an enlarged view of a system of counter-rotating propellers equipping a turbomachine using the invention, [Fig.32] illustrates section AA' of [Fig.7] [Fig.33] illustrates section BB' of [Fig.7] [Fig.34] and [Fig.35] illustrate the characteristic angles of a profile (see [Fig.35] which corresponds to the section CC' or DD' of [Fig.34]), and, [Fig.36], and [Fig.37] illustrate other blade configuration variations. Detailed description of the invention

[0057] In particular, Figures 1 and 2 schematically illustrate an aeronautical propeller 1, in particular a dual-flow aircraft turbojet with unducted propellers, according to one embodiment of the invention. The propeller 1 extends along an axis X.

[0058] In the following, when the propulsion means of the propulsion unit 1 are mentioned, reference is made to a turbomachine, therefore to a gas turbine(s) propulsion unit. This should not, however, be considered as limiting, as already mentioned.

[0059] Thus, [Fig.l] illustrates an aircraft powered by at least one (for example two) aeronautical propeller 1, the or each aeronautical propeller 1 comprising: - an engine, which can be called central, 3 having a central longitudinal axis X (which can be parallel to the longitudinal axis of the aircraft), along which a gas flow can circulate from upstream (AM) to downstream (AV), - a first series of blades 9, which can be considered as defining a propeller, - a second series of blades 6 positioned downstream of the first series of blades and which can also be considered as defining a propeller.

[0060] Note that parallelism between the central longitudinal axis X and the longitudinal axis of the aircraft is not a necessary condition for implementing the invention. The central axis of the turbomachine may have a non-zero angle relative to the axis of the aircraft ("cant angle" in English) in order to minimize installation effects.

[0061] The central engine 3 is enclosed in the nacelle 5 which surrounds it circumferentially.

[0062] According to a CROR configuration, the propellers 6 and 9 are, in the example of [Fig.2], both rotors and are counter-rotating: They turn in opposite directions, around the common axis X.

[0063] They can be of the type with high-flow fan(s) and unducted rotor(s) also called unducted propellers (open rotor / ultra-high bypass ratio, unducted fan type). The directions of rotation are indicated by arrows 12 and 15.

[0064] The solution of [Fig.l] could also be suitable for a turbomachine with unducted upstream 9 and downstream 6 propellers of the USF (Unducted Single Fan) type, where the downstream propeller 6 does not rotate around the axis of the turbomachine X. In other words, the downstream propeller 6 is a rectifier or stator. This case would be obtained for example by removing the arm 19 mentioned below. [Fig.2] illustrates a type of turbine system which can be used to drive the rotation of at least one of the propellers 6 and / or 9 around the axis of the turbomachine X and the setting of at least one of these same propellers 6 and / or 9 around their respective axes 360, 390.

[0065] In [Fig.2], the forward or upstream propeller 9 (hatched) is attached to a first turbine 18 (also hatched) which rotates in direction 15 as shown in [Fig.l]. The aft or downstream propeller 6 is attached, by arm 19, to a second turbine 21 and rotates in direction 12 of [Fig.l]. A gas flow passes through the turbines 18 and 21. The air inlet of the air flow F intended for the turbine system is located at the upstream end of the central engine 3, and marked 31. The bearings 140 support the turbines and allow rotation. The hot and high energy gas flow F is supplied by a combustion chamber (not shown) and causes rotation of the turbines. The bearings 140 support a rotating frame 141 fixed with the blades of the turbine stage 23.

[0066] The propeller blades 60 and 90 (which are sometimes called fan blades, propeller fans or propeller blades because they have hybrid characteristics between propellers and fans) are of the variable pitch type. Variable pitch means that each blade, such as 60 and 90 respectively, can rotate about a respective pitch axis (or pitch axis) 360, 390, as indicated by the circular arrows 34. The main reason for changing the pitch is to give the blades the angle of attack that is appropriate under the flight conditions of the aircraft and the engine power setting. Furthermore, it is thus possible to adopt both a "puller" configuration (tractor: propellers upstream of the combustion chamber, like [Fig.4] or 5) and a "pusher" configuration (pusher: propellers downstream of the combustion chamber, like [Fig.l]), with therefore unducted propellers on a CROR type turbomachine (therefore two upstream propellers 9 and downstream 6 counter-rotating) and / or USF type (therefore with an upstream propeller 9 rotor and a downstream propeller 6 stator), [Fig.2] being able to relate a priori to a "pusher" case. We could apply this solution to a "puller" case, by shifting the propellers 9,6 upstream of the nacelle, therefore upstream of the combustion chamber); typically at the level of the compressor(s).

[0067] The source of motive power that causes the pitch change is generally located within the annular path of the turbines 18 and 21, such as in region 35 of [Fig. 2]. Therefore, a mechanism is required to transport the mechanical torque from region 35, through the flow path of the flux (flux Fs in the example), towards the blades 60 and 90. Such means (called second means) of drive 40 for pivoting, around its setting axis 360, 390, each blade concerned via the setting arm 36 or 39 to which it is fixed can be organized as follows: The blades of the propeller considered can be carried by an annular support 24, such as a ferrule or polygonal ring surrounding the turbine stage considered (23 on the section of [Fig.2]) which is shown in schematic section in [Fig.2], the turbine stage also being visible in [Fig.3]. The annular support 24 therefore supports the blades 90 of the propeller 9, on the section. For each open rotor blade, a radial shaft 87 passes through a turbine blade (radial shaft 87 and stage 23 figures 2 and 3). The shaft 87 is extended radially by a blade wedging arm 39 fixed to its foot.The change in pitch indicated by arrows 34 may be caused by rotation between gears, via sub-planets of a compound planetary gear attached to a drive shaft. If the transmission ratio between the sub-planets and a fixed crown 54 is different from the transmission ratio between said sub-planets and a mobile crown 52 (i.e. the system is a differential planetary system), then the rotation of the compound planetary can cause a relative rotation of the crowns 52 and 54, and thus cause the change of pitch 34. The same solution can be used for the timing of the downstream propeller 6: A second frame 142, rotating around the axis X and fixed with the blades of a downstream turbine stage located at the downstream rotor 6 can be coupled to a frame 141 also rotating around the axis X and fixed with the blades of the upstream turbine stage 23. The two frames 141, 142 are supported by bearings 140.Furthermore, the principle of [Fig.3] explained above can be applied: For each open rotor blade 6, a radial shaft 87b identical to that 87 can pass through a polygonal ring surrounding a blade of said downstream turbine stage.

[0068] In all the cases referred to here, if the central engine 3 is a turbojet, it therefore comprises successively, along the axis X, one or more compressor(s), one or more combustion chamber(s), one or more turbine(s) for driving the compressor(s), via one or more axial drive shafts, and at least one hot gas outlet or nozzle downstream, 33.

[0069] In a "puller" configuration, such as [Fig.4], in addition to the case of upstream 9 and downstream 6 propellers forming two rotors, the upstream propeller 9 can be a rotor alone, and the downstream propeller 6 a stator.

[0070] The stator then comprises rectifier blades 60, for straightening a part of the flow Fs, precisely a secondary air flow Fs which passes through the two propellers, in the case of a central engine 3 of the turbojet type having a flow zone of such a secondary air flow Fs, surrounding a vein of primary gas flow Fp.

[0071] In the "puller" configuration, the air inlet 35 intended for the primary gas flow Fp passing through the central engine 3 can be downstream of the two upstream 9 and downstream 6 propellers, as [Fig.4].

[0072] A relevant solution is however to locate the air inlet 35 axially between the two upstream 9 and downstream 6 propellers, as [Fig.5] or 6. This location can in particular be between the respective setting axes 390, 360 of the blades of the first series of blades 9 and of the second series of blades 6.

[0073] The nacelle 5 which contains the central engine 3 has an aerodynamic external surface 50 relative to which the first series of blades 9 and the second series of blades 6 project radially from the central longitudinal axis X.

[0074] Each blade 90,60 of the first series of blades 9 and of the second series of blades 6 having: - a free end 51 opposite a connection end 53 forming a blade root close to the nacelle 5, an intrados face 55 and an extrados face 57 (see in particular figures 1-2), - a chord C located at a defined radius centered on the central longitudinal axis X (see for example [Fig.5]), - a radius Ri between the central longitudinal axis X and a location which is level with the external surface 50 of the nacelle, this location being located: — on the blade (at the base of the blade) or — on a blade setting arm 39, - a radius Re between the central longitudinal axis X and a location on the free end 51 of the blade furthest from the central longitudinal axis X, in a direction transverse to the central longitudinal axis X, and - a span L (see for example [Fig.7]) defined as Re - Ri, which corresponds to a distance radially to the central longitudinal axis X, between the free end 51 and the connection end 53, in said transverse direction.

[0075] Note that this location on the free end 51 of the blade may be located on the axis 390 or 360 of the said blade, or not.

[0076] The chord C may vary in the direction from the radial blade to the central longitudinal axis X.

[0077] Before detailing this and other aspects below, it is important to recall that an element at the heart of the invention is, in combination, on the proposed propellant 1 and as existing in all the figures (even if not marked): - that at least one of the blades of the first series of blades 9 has a trailing edge 91 having undulations 93, and, - that at least one of the blades of the second series of blades 60 has a leading edge (BA) 61 having undulations 63, and - that at least some of the blades 90 of the first series of blades 9 are pitched variable, so that each of them can pivot around a wedging arm 39 to which the blade 90 concerned is fixed, around the wedging axis 390 which at least partially passes through said blade 90, and / or - that at least some of the blades 60 of the second series of blades 6 are variable pitch, so that each of them can pivot around a pitch arm 36 to which the blade 60 concerned is fixed, around the pitch axis 360 which passes through the blade 60.

[0078] An alternating succession of at least two tooth peaks and two troughs, such as for example marked respectively 630,631 and 632,634 in [Fig. 18], defines an undulating zone along the BA or the BF.

[0079] Preferably, the number of peaks and / or troughs will preferably vary between 2 and 100 along the span L.

[0080] The blade pitch angle 9 or 6 has been symbolized by "y" in Figures 9 and 35. The pitch of the profiles that define a blade can therefore vary depending on the radius of the blade. However, it is possible to define a reference pitch angle y for a blade at a given radius, for example at 75% of the radius Re.

[0081] Concerning the characteristic angles of a profile (for example in section AA' or BB' in [Fig.7] or section CC' or DD' in [Fig.34]), we will also note (see [Fig.35]) the interest in applying certain specificities to the blades 9,6 and / or in connection with all or part of the following: - the pitch angle (y) of the blade in question, which angle is therefore defined between the line or straight line A which connects the leading edge and the trailing edge of the blade in question, and the line XI parallel to the central longitudinal axis X, - the angle of the skeleton at the BA ( [31 ) considered, which angle is defined between: — a line tangent to the skeleton line, measured at a distance “a” close to the BA, i.e. a / C < 0.2 and preferably a / C <0.1 (“a” being the length, on the line A, between the BA and a point along the chord C), and — a line parallel to the central longitudinal axis, X - the angle of the skeleton at the BF ([32) considered, which angle is defined between: — a line tangent to the skeleton line, measured at a distance close to the BF, such that a / C > 0.8 and preferably a / C > 0.9, and — a line parallel to the central longitudinal axis, X; - the skeleton line, which is defined as the average camber line, dotted, midway between the intrados and the extrados of the blade profile 9 or 6.

[0082] It should be noted that the angles, y, [31, [32 vary in the radial direction to the central longitudinal axis X, i.e., these angles depend on the blade profiles at a given radial position (r) or along a streamline / air friction line near the blade surface 90,60.

[0083] The angles [31,

[32] make it possible to characterize the angle of incidence and exit of the flow upstream and downstream of the blade profile. Thus, it is necessary to ensure that the incidence (or angle of attack) perceived by the blade profile(s) at the troughs (such as 632,634 [Fig. 18] or 634 [Fig.27]) and at the peaks (such as 630,631 [Fig. 18] or 635 [Fig.27]) of the BA undulations are acceptable, and that there are no stall phenomena penalizing the blade performances.

[0084] To this end, it is proposed that, on the BA (or BF) zones of the blades 90, 60 having undulations 93 and / or 63, the variation of the angle of the skeleton at the BA, A[31, (or at the BF, A[32) between a tooth tip 635 and a tooth trough 634 adjacent to each other, according to the span L of the blade considered 90 or 60, or in the radial direction to the axis X, is less than 45°, and preferably less than 25°, in an alternative / preferred embodiment

[0085] In order to reduce noise, it is further proposed: - that the first series of blades 9 and the second series of blades 6 may have different numbers of blades, as for example shown [Fig.l], preferably each between 5 and 20; this may make it possible to decorrelate the acoustic radiation of the upstream and downstream blades, and / or - that the blades of the first series of blades 9 and / or of the second series of blades 6 are located on a closed line centered on the central longitudinal axis X and having a diameter D, with - the alignment axes of the blades 390, 360 respectively of the first series of blades 9 and of the second series of blades 6 which will be axially separated, in pairs, by a distance S, and - an S / D ratio which will be between 0.005 and 1, and preferably between 0.15 and 0.70, or even between 0.20 and 0.50.

[0086] Thus, the turbulent wake of the upstream propeller 9 can be effectively dissipated as it propagates downstream. The wakes which arrive at the leading edge 61 of the downstream propeller 6 will be relatively low in energy.

[0087] D may thus, in one case, correspond (as conventionally) to the maximum diameter of the upstream propeller 9.

[0088] Furthermore, the trailing edge 91 of each of the blades 90 of the first series of blades 9 is located longitudinally upstream of a leading edge of each of the blades 60 of the second series of blades 6. Thus, interference between the series of blades is avoided.

[0089] In connection with the noise, it is also proposed that each blade has, on the first series of blades 9 and / or on the second series of blades 6 and at a defined radius, common: - a segment length of circumference E between two successive blades of the circumference, - a C / E ratio such that C / E < 2.5 at any radial position along the blade span, and such that C / E < 0.8 at the free end.

[0090] A low solidity, C / E, of the blades 90, 60 will increase the size of the inter-blade channels 59 ([Fig.5], i.e. the spacing between two consecutive blades in the azimuthal / circumferential direction). This can prevent the formation of shock waves in the inter-blade channel at high rotation speed at certain operating points, which is linked to the contraction of the flow in each channel 59. These shock waves are the origin of the "shock noise" or "buzz-saw noise", in English.

[0091] Various studies have also demonstrated that noise reduction can be promoted if, the two propellers 90, 60 being assumed to be counter-rotating rotors, the rotation speed around the X axis of the upstream propeller 90 is greater than that of the downstream propeller 60.

[0092] To adapt a propeller rotation speed, such as the upstream propeller 9, a speed reducer may be used, which may be epicyclic, such as the speed reducer 104 presented below.

[0093] In this case, the blades of the first series of blades 9 and of the second series of blades 6 will therefore be arranged so as to be able to be driven in rotation around the central longitudinal axis X by a free power turbine, in a counter-rotating manner. An example is disclosed in EP2368030.

[0094] It should be noted that increasing the rotational speed of the downstream propeller 6 can increase the relative speed at the tip of the blade 51 of the downstream propeller, which could become transonic (hence generation of shocks, and others).

[0095] In the example of [Fig. 30], the core engine 3 operates in a conventional manner, so that the air entering the intake 35 is accelerated and compressed by the low pressure compressor 920 and directed to the high pressure compressor 92 where further compression takes place. The compressed air discharged from the high pressure compressor 92 is directed to the combustion chamber 930 where it is mixed with the fuel and the mixture is burned. The resulting hot combustion products expand and thus drive the high pressure 94, low pressure 98 and free power 96 turbines before being discharged at 33 through the nozzle 97 to provide some propulsive thrust. The high pressure, low pressure and power turbines, free, respectively 94, 98, 96, drive respectively the high pressure compressors 92 and low pressure 920 and the unshrouded rotors 9, 6, by suitable interconnecting shafts.The two unducted propellers 9.6 are counter-rotating, fixed to the free turbine 96 and driven by it, via the sets of counter-rotating blades 90, 60 99 (contra-rotating blade arrays).

[0096] The example of [Fig.31] presents a situation where the aircraft turbomachine comprises: - a free power turbine 102 comprising a rotor 103; - a downstream propeller 6 and an upstream propeller 9, with the upstream propeller intended to be rotated around a central longitudinal axis X of the propeller system, relative to a stator of this system which can be the nacelle 5; this stator can also be a blade or fixed vane of another propeller, for example, blade 6 of the second series of blades on a USF type turbomachine.and - a mechanical transmission device comprising a speed reduction box or speed reducer 104 (or “gearbox” in English) between an input shaft driven by the turbine, such as the shaft 121 below, and the blades of the upstream propeller 9, in order to reduce the rotation speed of the propeller blades 90 and therefore the noise generated by this propeller.

[0097] This can be extrapolated to a USF, for example by removing the downstream part of the planet carrier shaft 117 fixed to the downstream propeller 6.

[0098] In this example, however, it is a CROR where the downstream propeller 6 and the upstream propeller 9 are therefore counter-rotating, intended to be set in rotation together around a longitudinal axis X of the propeller system, relative to the nacelle 5.

[0099] The rotor 103 is a first rotor 103.

[0100] The speed reducer 104 comprises an epicyclic gear train 105 provided with a sun gear 107 centered on said longitudinal axis X and driven by the first rotor 103 of the free power turbine 102, at least one satellite 106 meshing with the sun gear 107, a planet carrier(s) 108 driving the downstream propeller 6, as well as a ring gear 109 meshing with each satellite 106 and driving the upstream propeller 9.

[0101] The free power turbine 102 also comprises a second rotor 111 counter-rotating with respect to said first rotor 103, and driving the crown 109 in rotation.

[0102] The sun gear 107 centered on the longitudinal axis X is carried by a planetary shaft 113 of the same axis, securely connected upstream to the first rotor 103, by a flange 115. Thus, the first rotor 103 directly drives the sun gear 107 in rotation, which can take the form of an externally toothed wheel.

[0103] The planet carrier 108 is carried by a planet carrier shaft 117 of the same axis, integral with the downstream propeller 6. The crown 109, centered on the axis X, is carried by a crown shaft 119 of the same axis. The crown 109 meshes with each satellite 106. The shaft 119 extends downstream while being integral with the upstream propeller 9, so as to be able to drive it directly in rotation. The shaft 119 can be located around the planet carrier shaft 117 with which it is then concentric. The crown 109, taking the form of an internally toothed wheel, is also carried by another crown shaft 121, of the same axis, and extending upstream. The crown shaft 121, located around the planetary shaft 113 with which it is concentric, is securely connected to the second rotor 111, by means of a flange 123. The aircraft advances in the direction 101.

[0104] Further details can be found in WO2010070066.

[0105] In the solutions of both [Fig. 30] and [Fig. 31], (second) drive means 40 are of course further provided for pivoting, around its setting axis 390 or 360, each blade 90, 60 concerned via the setting arm 39 or 36 to which it is fixedly connected. For this purpose, the solution of FIGS. 2, 3 can be applied, for example.

[0106] Also to combat noise, it will be considered that the leading edge undulations 63 and / or those 93 of the trailing edge are located, preferably only, or essentially: - beyond 0.4x(Re-Ri), starting from the side of the connection end 53, and / or - at (near) the radial position where the chord C is maximum, over a length of 0.2x(Re-Ri) to 0.8x(Re-Ri) around the maximum chord position.

[0107] A length of the order of 0.3x(Re-Ri) around the maximum chord position may be suitable, in particular.

[0108] Near the free ends 51 of the blade, the undulations make it possible to reduce the noise linked to the blade tip vortex. If the undulations are at the trailing edge 91, towards the free end of the upstream propeller 9, the undulations 93 make it possible to improve the mixing of the wake and therefore to reduce the intensity of the blade tip vortex.

[0109] The trailing edge undulations 91 also make it possible to reduce the inherent noise linked to the passage of the turbulent boundary layer which develops on the intrados and the extrados of the blade 90 and which radiates noise at the trailing edge 91.

[0110] If the undulations 63 are located at the leading edge 61 of the downstream propeller 6, this makes it possible to reduce the interaction noise and to better decorrelate the noise sources along the span L.

[0111] The interest of the trailing edge undulations on the blade profiles 90, 60 with a large chord C (C > Cmoy, where Cmoy is the average chord of the blade) is to reduce the inherent noise linked to the boundary layer, which becomes thicker on profiles with a large chord (see figures 32, 33: sections AA' and BB'). The undulations 93 at the trailing edge 91 of the upstream propeller 9 also lead to a reduction in the interaction noise (broadband and tonal) between the wake and the downstream propeller, because the mixing of the turbulence and the average speed deficit in the wake are accelerated.

[0112] The disadvantage from an acoustic point of view is however to limit the areas of the leading edge and the trailing edge which can contribute to noise reduction by the presence of undulations. However, limiting the areas of the BA and the BF with undulations makes it possible to limit possible aerodynamic losses, which makes it possible to find an advantageous aero-acoustic compromise. Yet another noise reduction factor can be identified, when, as [Fig.4] or 5:

[0113] - the first series of blades 9 and the second series of blades 6 are located towards a upstream end of central engine 3 ("puller" configuration), - the central engine 3 has an air inlet 35, to bring the primary air flow Fp of gas flow into the central engine 3, and - the air inlet 35 of the primary flow is located axially between the first series of blades 9 and the second series of blades 6, and even more precisely, and preferably, between the respective axes 390, 360 of the blades of the first series of blades 9 and of the second series of blades 6.

[0114] This makes it possible to reduce the span L (and therefore the radius Re) of the downstream propeller 6, because it must have a "clipping". The size of a 90 or 60 blade, and in particular its span, is a contributor to the radiated noise. Thus, such a configuration will make it possible to reduce the noise of the turbomachine.

[0115] The air inlet 35 can be placed on an angular sector or on 360° (crown). FR3083207 refers to it.

[0116] To reduce the interaction of the wake of the upstream propeller 9 with the nozzle 37, and therefore a reduction in noise, it will be possible to provide, as [Fig.6]: - that the air inlet nozzle 37 is located at a radius Rb from the central longitudinal axis X, and - that the trailing edge 91 of the first series of blades 9 has undulations 93 located at the radius Rb.

[0117] In this case, we find as in the embodiment of [Fig.5] the interest linked to the fact that the doublet of propellers 9,6 is positioned in “puller” configuration, with the air inlet of the primary flow between the upstream propeller 9 and the downstream propeller 6.

[0118] However, undulations 93 at the trailing edge of the blades of the upstream propeller could disturb the supply (the air flow) which arrives at the inlet 35 of the primary flow.

[0119] Providing, according to a USF type configuration, that the downstream propeller 6 is a stator / flow rectifier which does not rotate around the central longitudinal axis X, unlike the upstream propeller 9, but whose blades 60 can each rotate around its pitch axis 360, could also be advantageous, from an aerodynamic and acoustic point of view.

[0120] In this case, the turbine of the central engine will be connected to the first series of blades 9 so as to only drive in rotation the blades of the first series of blades 9, the blades of the second series of blades 6 defining air rectifier blades. Only the blades 90 of the first series of blades 9 will therefore be arranged to be able to be driven in rotation around the central longitudinal axis X by the blade drive turbine 90 which may be a free power turbine. The blades of the second series of blades 6 define air rectifier blades (“outlet guide vanes” or “stator swirl recovery vanes”).

[0121] It will then be possible to usefully interpose a speed reducer (such as the one 104 mentioned above) between the blades 90 of the upstream propeller 9 and the shaft of the drive turbine of these 90 blades, in order to reduce the rotation speed of this propeller, which optimizes the aerodynamic performance of the propeller and reduces the noise it generates.

[0122] The advantage of a USF configuration compared to a CROR is to present less tonal noise, with fewer "lines / peaks" on the spectrum. Indeed, on a CROR, we can distinguish the blade passing frequencies f (or "Blade Passing Frequency", BPF, in English) of each propeller (upstream / downstream) and its harmonics, as [Fig.11]. In addition, there are "combination / interaction lines" on the SPL noise spectrum of a CROR. However, on an SPL noise spectrum of a USF there are only the BPFs linked to the upstream propeller (as [Fig. 12]), which should make it possible to reduce the acoustic levels. In the case of a USF, these BPFs are mainly related to the aerodynamic load noise on the rotor blades (upstream propeller) and the interaction of the rotor wakes - characterized by a deficit in average speed - with the stator.

[0123] From a mechanical / integration point of view, a USF type architecture is nevertheless simpler to implement (fewer rotating parts in particular) than a CROR architecture.

[0124] [Fig. 8] illustrates, by way of non-limiting example, a solution of (second) drive means 40 for pivoting, around its setting axis 390 or 360, each rectifier blade concerned via the setting arm 39 or 36 to which it is fixedly connected. This is an articulated system for setting each rectifier blade ("outline guide vane" or "swirl recovery vane") 60 if the downstream propeller 6 is of the stator type. This solution is therefore suitable for a USF configuration and can be coupled to an upstream propeller 9 which would be controlled, for example, as in FIGS. 2 and 3, after the arm 142 has been removed and the lever 613a or 613b cited below has been connected to suitable actuating members, known per se.

[0125] The variable-pitch rectifier blade 60 is in the example radially rotatable through the external surface 50, in a casing belonging to the nacelle 5. The blade comprises a blade part 601, a plate 603 and a rod or pivot defining the pitch arm 360. The pivot 36 is housed in a radial orifice formed in the casing 5. A bearing of the pivot 39 consists of a bushing 607 in sliding contact with the pivot 36. The bushing 607 secured to the casing is in contact with the plate 603 by means of an annular boss 609. The opposite face of the plate 603 relative to the bushing is swept by the air which passes through the two propellers 9, 6. A washer 611 holds the blade in its housing. A lever controls the rotation of the blade 60 considered around the 360 ​​setting axis of the pivot to put it in the required position (see double arrow [Fig.8]) relative to the air flow sweeping the blade.The relative movements result from the sliding of the surfaces in contact, here the pivot 39 and the plate 603 with the bushing 607. The lever extends either downstream (lines 613a, in particular in the case of a . pusher / pusher thruster), or upstream (dotted lines 613b, in particular in the case of a tractor / puller thruster). The members which actuate the lever 613a or 613b are not shown; they comprise an actuator and a control which can, together, also actuate the blades 90 to wedge each of them around their axis 390, as already explained in connection with figures 2, 3. For further details, please refer to EP17174450.

[0126] Concerning the noise sources characteristic of an engine with unducted propellers, the acoustic radiation of an aerodynamic profile is similar to that of a dipole (two main radiation lobes), whose preferred propagation direction is normal to the chord of the profile (see [Fig. 10]). The rectifier blades 60 located at 6 o'clock and 12 o'clock on USF architecture will thus contribute to the lateral noise levels (on the side) of the aircraft. The stators located at 3 o'clock and 9 o'clock on USF architecture will contribute to the noise levels above and below the aircraft. As an effect on the trace radiated noise during the takeoff and landing phases, one embodiment proposes to apply undulations only to the rectifier blades which contribute mainly to the trace radiated noise. Even if an optical effect may suggest this in [Fig.10], the blade span (L=Re-Ri) can be constant over the entire periphery of the propeller considered.

[0127] It is also proposed that, as [Fig. 10], circumferentially around the central longitudinal axis X, only the air straightener blades 60 located in a first angular range of + / - 60° relative to 3 o'clock, and in a second angular range of + / - 60° relative to 9 o'clock, have undulations, marked 63 elsewhere.

[0128] This is a compromise between noise reduction efficiency / cost and aerodynamic performance.

[0129] Yet another aspect can usefully contribute to combating noise: that undulations 93 and 63 are located respectively on the trailing edge 91 of the upstream propeller 9 and on the leading edge 61 of the downstream propeller 6.

[0130] In fact, undulations 93 at the trailing edge of the upstream propeller have a double advantage:

[0131] (i) to reduce the blade's own noise (linked to the acoustic radiation produced by the boundary layer on the blade surfaces as it passes over the trailing edge) and,

[0132] (ii) to improve the mixing of the wake, which reduces the intensity of the vortex at the blade tip and the average speed deficit when it impacts the downstream propeller 6.

[0133] The advantage of undulations 63 at the leading edge of the downstream propeller 6 is to reduce the noise resulting from the interaction of the wake and the vortices at the blade tip on the upstream propeller 9 which can interact with the downstream propeller. An advantage is therefore to reduce (or avoid) “clipping” at the blade tip on the downstream propeller. Indeed, reducing the clipping on the blades 60 of the downstream propeller 6 makes it possible to improve the aerodynamic performances. because an additional part of the flow gyration at high radii, near the external limit of the radius Re, is recovered at the end 51 of the blade on the downstream propeller 6. This therefore makes it possible not to increase the chord of the downstream propeller 6 (in order to ensure the same bearing surface as that of a downstream propeller without "clipping"), particularly in the case of a CROR.

[0134] Combining the corrugations 93 in BF and 63 in BA in this way, however, increases the complexity of the system and the implementation costs, compared to a separate solution, as proposed further on in the description. Indeed, the more corrugations the blades have, the more complex and expensive the parts are to produce (possibly more deviations / non-conformities to deal with, more manufacturing defects, etc.).The mechanical impacts related to increasing the chord C of a propeller are further as follows: Increasing the chord C by decreasing the span L makes the blade less resistant to bird ingestion at BA / BF and it especially becomes more difficult to integrate because of the variable pitch: the blade can protrude from its platform which causes aerodynamic leaks radially inward at the blade root, the radially inward part of the blade root tends to be very mechanically loaded, the moment of inertia increases, which generates a torque around the 390 or 360 pitch axis and therefore a pitch change system and a feathering system to be adapted. The integration of such a blade can negatively impact the hub ratio, Ri / Re.

[0135] The leading and trailing edges of the blade 9 may in fact extend beyond the (circular) platform at the blade root. The circular platform is integral / rotates with the blade 9, but the annular support 24 and the nacelle wall 5 are fixed. This means that unwanted air leaks may exist between the blades and the nacelle wall at the BA and BF, at the blade root.

[0136] Several geometries or patterns of teeth of the corrugations can individually or in combination be used at the leading edge and / or trailing edge of the blades 9, 6: sinusoidal, square, rectangular shaped teeth, slots / grooves, etc. The advantage of using several patterns is to reduce the correlation of the noise sources along the BA and / or BF, which should promote noise reduction; see the information provided elsewhere on the angles y, [31, [32 and the shapes of the corrugations, such as the rounded 81 or elliptical 79 shapes presented below.

[0137] Depending on at least one of the following characteristics: pattern (or geometry), amplitude, spacing, radial positioning of the undulations along the span, undulations could vary (be different) between the different blades 90, 60 of the same propeller (upstream and / or downstream) and / or between at least some of the blades of the first series of blades 9 and at least some of the blades 6 of the second series of blades. This would make it possible to define several families of blades with, for example, different geometric patterns. We could therefore vary the geometry for one or more blades, the chord, the camber, the maximum thickness, etc. in order to define a 9.6 propeller with heterogeneous 90.60 blades. One advantage of using several patterns of different geometries at the leading edge of each 90 and / or 60 blade is to reduce the correlation of noise sources. This can be particularly interesting if the solidity of the 9 or 6 propeller is high or if the size of the vortex structures is large.

[0138] Furthermore, the advantage of using several of the different geometric patterns at the trailing edge of the upstream propeller is to improve the mixing of the wakes.

[0139] The search for noise reduction may also be associated with other considerations. The following order of these considerations is not related to their importance.

[0140] First consideration: It will be possible to seek to ensure improved mechanical strength of the blades 90 and / or 60, in nominal operation of the turbomachine, and in the event of bird ingestion. The mechanical forces on the blades having BA / BF undulations could be greater than a reference case having smooth BA / BF. To then improve the mechanical strength of the blades having undulations, it is proposed - that each undulating blade 90, 60 of the first series of blades 9 and / or of the second series of blades 6 has a maximum thickness (marked e[Fig.2]) between the intrados face 55 and the extrados face 57, and - that, on the blades, the maximum thickness of the blades is located near the blade root 53, at a distance less than 0.1x(Re - Ri) from the connection end 53 (end located at radius Ri). See section AA' and BB' figures 7, 32 and 33.

[0141] Second consideration: We may seek to ensure a hub ratio that varies within a range that guarantees a good compromise between aerodynamic needs (large blades —> low Ri / Re) and the needs for integration of additional systems in the hub, i.e. under surface 50 (device allowing variation in the blade timing, oil and air pipes, etc. —> large Ri / Re).

[0142] It is also proposed that each blade 90 and / or 60 with undulating BA and / or BF of the first series of blades and / or of the second series of blades may have a hub-to-free end Ri / Re ratio such that Ri / Re is between 0.10 and 0.50.

[0143] Third consideration: We could seek to improve the mechanical strength of the blades, particularly in the event of bird ingestion.

[0144] It is also proposed that the undulations, such as 93 or 63, have, per blade or on at least one of said blades 9 or 6, an amplitude h(r) (which may be maximum) between an adjacent peak 635 and trough 634, such that: 0.0005xCmax < h(r) < 0.5xCmax, where: - Cmax is the maximum chord of the blade, and - h(r) corresponds here to the difference (if necessary maximum) of chord C between a profile at the level of a peak 635 and a profile at the level of an adjacent trough 634, in the direction of the span (L) of the blade, or radially to the central longitudinal axis (X).

[0145] Figures 13 to 26 are consistent with this.

[0146] Amplitude (hr, such as hl,h2..) otherwise also has the meaning: depth of the undulations between the top and the bottom of the tooth considered, see [Fig.27] for an illustration whose principle will be applied to [Fig. 13] to 26. Undulations, such as 93,63, which are too deep could have a negative impact on the aerodynamic performance of the blades and present difficulties for the mechanical resistance of the blades in the event of bird ingestion.

[0147] Fourth consideration: We can seek to adapt the geometry of the undulations according to the local characteristics of the gas flow.

[0148] Thus, it will be possible to usefully plan: - that the undulations 93, 63 at the leading edge 61 and at the trailing edge 91 have different geometries or patterns between them, and / or - that the undulations 93 of some of the blades of the first series of blades 9 are different from the undulations 63 of some of the blades of the second series of blades 6.

[0149] In the case, for example, of trailing edge undulations, their amplitude and spacing could usefully be dimensioned by the turbulent quantities of the boundary layer at the trailing edge of the blade, where, for example, the undulations could be defined as a function of the spatial correlation length of the wall pressure fluctuations along the trailing edge or as a function of the boundary layer thickness. In the case of leading edge undulations, their amplitude and spacing would be dimensioned by the aerodynamic quantities of the incident flow (atmospheric turbulence, upstream blade wakes, etc.). In this case, one of the favorable dimensioning parameters could be the integral scale of the turbulence.

[0150] It is also proposed that, as illustrated in example figures 14, 19, 20 (where the amplitude and spacing, h(r) and X(r), can be understood from the illustration of [Fig.27]), on at least some of the blades 90 and / or 60, the undulations have an amplitude, h(r), and a spacing between two successive undulation peaks, X(r), which vary(s). It should be noted that h(r) and X(r) can be functions defined piecewise along the span L of the blade 90, 60 as a function of the radial position, r.

[0151] Fifth consideration: In addition to improving the mechanical strength of the blades and the geometry of the teeth, particularly in the event of bird ingestion, we may seek to simplify the manufacture of the blades, because it is a priori simpler to machine complex geometries on metal parts than on composite parts (typical fibrous material for manufacturing blades 90,60,). In addition, a cover or foil 73 to BA and / or BF corrugations could be sold or marketed as an option for aircraft manufacturers and / or airlines wishing to have improved acoustic performance. Providing corrugations only on a mechanical cover (for example, metal) can also simplify maintenance operations and improve resistance to erosion.

[0152] It is also proposed, as shown diagrammatically in Figures 13 and following, that at least some of the blades 90, 60 comprise a composite material 71 and, on the intrados face 55 and / or the extrados face 57, that at least one metal mechanical reinforcement cover 73 is fixed with the composite material and extending along at least a portion of the leading edge and / or the trailing edge. The metal cover(s) may not only be along the leading edges or trailing edges.

[0153] It may also be a question of reinforcement plates on the surface of the intrados 55 and / or the extrados 57. The weight is to be considered; the removable nature with respect to the composite material 71 of the blade, therefore replaceable, of the cover 73 also.

[0154] Sixth consideration: Reducing the weight of the blades will also make it possible to reduce the fuel consumption of the aeronautical propeller on which the double set of propellers 9.6 is installed. It is therefore proposed that the blade (or its composite part, such as part 71) be an assembly of organic materials (thermosetting resins, thermoplastic resins) reinforced or not by carbon fibers according to one of the multiple possible manufacturing processes (weaving, braiding, lamination, winding, etc.). The reinforcing fibers can also be glass, Kevlar or aramid fibers.

[0155] Seventh consideration: We can seek to limit the number of parts to be manufactured and optimize the mechanical strength of the blade.

[0156] It is also proposed that, for each blade 90 or 60, the reinforcement or cover 73 connects the BA to the BF by the free end 51 of the blade. This will be particularly interesting in the event of the presence of undulations on the reinforcement or cover 73. Examples of this can be seen in particular in figures 16, 17, 20, 22.

[0157] Eighth consideration: We may seek to reduce the intensity of the free end vortex 51 of the blade, in order to reduce the interaction of the blade tip vortex 51 of the upstream propeller 9 with the downstream propeller 6.

[0158] It is also proposed that, as illustrated in some examples figures 23, 24, the free end 51 of at least one of the blades 90 or 60 has a “proplet” or an abrupt change in the tilt angle, marked 77.

[0159] Ninth consideration: We may seek to reduce the intensity of the free end vortex 51 of the blade, and therefore the interaction of the blade tip vortex 51 of the upstream propeller 9 with the downstream propeller 6.

[0160] It is also proposed, as illustrated in some examples in figures 25,26: - that the free end 51 of at least one of the blades 90 or 60 has an elliptical shape 79, or - that this blade head 51 has a rounded shape 81 and therefore that the lines of BA and BF converge (or are merged) at the free end 51 of the blade.

[0161] Thus, the leading edge lines, such as 61, and the trailing edge lines, such as 91, will tend to be confused.

[0162] Tenth consideration: We could seek to facilitate overall manufacturing and maintenance, and thus limit intervention times.

[0163] It is also proposed that, as illustrated, the corrugations are formed only on the metal reinforcement cover 73, not on the composite material 71.

[0164] Eleventh consideration: We may seek to avoid the accumulation of ice on the leading / trailing edges, which can significantly degrade performance (reduction in thrust, reduction in efficiency, boundary layer detachment, etc.).

[0165] It is also proposed that the aeronautical propeller includes a heating device, for de-icing or anti-icing functions, and which could be connected to the metal reinforcement cover 73.

[0166] As illustrated in example [Fig.9], a heating device 75, which may comprise electrical resistors or channels for the passage of a hot fluid, in the metal reinforcement cover 73, can therefore be added and connected to this cover from the inside of the nacelle 5. It will then be necessary to control a possible increase in the thickness of the blade at the BA and / or BF in order to integrate the de-icing system in a relatively small footprint.

[0167] Twelfth consideration: We could seek to limit the weight of the blades 90 and / or 60.

[0168] It is also proposed, as illustrated as an example in several of the figures 13 and following, that the reinforcing cover 73 of the blade occupies less than 50% of the total volume of the blade, and preferably, less than 25% of the total volume of the blade.

[0169] Thirteenth consideration:

[0170] It may be possible to seek to increase the mechanical strength of the leading edge and / or trailing edge presenting undulations. Indeed, the teeth / undulations may have particularly elongated ([Fig.27]) or pointed ([Fig.28]) peaks.

[0171] It is also proposed, as illustrated as an example in several of the figures 13 and following, that the protection of the leading edge and / or trailing edge and / or blade head 90 and / or 60 (free end 51) be made of metal, with reinforcement cover 73 integrated or attached by removable or non-removable fixing, titanium or titanium alloy, steel, nickel or nickel alloy, stainless steel.

[0172] Fourteenth consideration: We may seek to increase the resistance to cracking of a blade reinforced by a metal reinforcement 73 comprising the corrugations.

[0173] It is also proposed that the removable metal reinforcement 73 comprising the on- The assemblies are attached to the composite material blade 71 by bonding using an epoxy adhesive. The adhesive can be reinforced with thermoplastic or elastomeric nodules. This assembly method is relevant for a metal / composite assembly.

[0174] It is also possible to use, in addition or as an alternative, a method of assembly by geometry, for example: metal insert 73 with a local dovetail shape and cavity of complementary shape in the composite 71, or vice versa. The advantage then lies in the possibility of changing only the metal reinforcement 73 (removable) in the event of damage caused by a bird strike or by erosion.

[0175] Whether there is a metal reinforcement 73 or a blade 90 or 60 made solely of composite material, the undulations can also be produced by a mechanical machining process, forging or foundry (if metallic), additive manufacturing or even chemical machining.

[0176] Fifteenth consideration: We can seek to limit the areas with undulations to just what is needed to maximize the acoustic gain while minimizing the impact on aerodynamics (efficiency) and mechanics (resistance in the event of bird ingestion, erosion, etc.).

[0177] It is also proposed that the undulations, at the BA and / or BF, are not in a single part / length but occupy several regions of the BA and / or BF separated by zones with a smooth BA and / or BF, as illustrated as an example in figures 14, 15, 21, 22 in particular.

[0178] Sixteenth consideration: We could also seek to define a connection zone between the smooth part(s) and the undulating part(s), which is interesting from an aerodynamic and mechanical point of view.

[0179] It is also proposed, on at least some of the blades 90,60: - that a first part of the leading edge and / or the trailing edge, such as 61 and / or 91, is smooth, without undulation, as along the zones 95 and / or 65 of the blades 90, 60, and in particular near the blade root, therefore at a distance less than 0.1x(Re - Ri) from the connection end 53,

[0180] - that a second part 97 and / or 67 of the leading edge and / or the trailing edge has said undulations, and - that the amplitude of the undulations at (near) the radial position where the second part connects to the first part, so that the connection is progressive, evolves monotonically and even strictly monotonically (in particular by decreasing): connection zones 99 and / or 69; see for example figures 14, 17, 20, 22, the marks 69 and 99 existing only on some of these figures.

[0181] Seventeenth consideration: We may also seek to limit possible aerodynamic losses and difficulties / costs linked to the manufacture of undulations at the leading edge and / or trailing edge.

[0182] It is also proposed that, along the leading edge and / or the trailing edge of certain at least 90.60 corrugated blades, following the radial direction or the span L, the corrugations extend along a length H which is limited to: H / (Re-Ri) <0.8, as in the example of [Fig.5] or 7.

[0183] If necessary, the length H will be accumulated over the length of the corrugated zones of the same BF or BA, if these zones are multiple (Hb H2, H3, ...), as in certain example figures (see in particular figures 9, 19, 22).

[0184] Eighteenth consideration: We may also seek to increase the amplitude of the undulations where the length scale of the turbulence is larger. Indeed, several studies show that the integral scale of the turbulence in the wake of a rotating part increases towards the tip of the blade.

[0185] It is also proposed that the amplitude (h(r)) and / or the spacing between two successive peaks of undulations, X(r), of at least some of the undulations in BF and / or BA varies monotonically, or even preferably or strictly monotonically, in the direction of (or even close to) the free end 51 of the blade considered 90 and / or 60.

[0186] Nineteenth consideration: To optimize the locations of the corrugations in terms of the noise reduction / cost / weight / resistance gains ratio, it may be necessary for the corrugations 93, 63 to be located at least on at least a portion of the trailing edge 91 of the first series of blades 9 and at least a portion of the leading edge 61 of the second series of blades 6. There may thus be leading edge and trailing edge corrugations on all the blades at the same time. Providing them only on these trailing edges 91 of the first series of blades 9 and leading edge 61 of the second series of blades 6 would, however, be a relevant compromise.

[0187] Twentieth consideration: The leading and / or trailing edge of the blades 90 and / or 60 could be locally porous. It can thus ultimately be provided that the leading and / or trailing edge of at least some of the blades 90, 60 is formed locally by a porous material 85, at the location of at least some of the undulations. As porous material 85, a metal foam can for example be provided, in particular on the regions with undulations 93 and / or 63, as shown diagrammatically in figures 27, 28, where X (or XI, X2) corresponds to X(r) and h (or h1, h2, h3) corresponds to h(r). In addition to allowing acoustic attenuation, the porous material 85 has the effect of reducing stress concentrations by its mechanical flexibility. This local optimization of the rigidity of the material limits the appearance of cracks under cyclic loading, thus increasing the life of the metal protection of the leading edge, at 61. The porous material 85 will define a part of the corrugated shape 63.Thus, together, the porous material 85 and the surrounding body of the blade (metal part 73 or composite body 71, marked 71 / 73 figures 27,28) will then define the undulations 63 at the BA of the blade. The porous material 85 is integrated into the blade, as a replacement. of a part of the metal reinforcement 73 or of the composite body 71. The porous material 85 will usefully occupy bottoms or hollows 634 of undulations.

[0188] Twenty-first consideration: We could also seek to reduce the interaction of a vortex (or a separation) at the free end 51 of the blade.

[0189] It is also proposed that the blades 90 or 60 of the upstream and / or downstream propellers have a radius Re greater at the leading edge (BA) than at the trailing edge (BF), that is to say: Re,BA > Re,BF, when the lines of BA and BF are not merged, as shown diagrammatically in the example of [Fig.29].

[0190] The three following configurations are shown diagrammatically in figures 36 and 37:

[0191] Twenty-second consideration: We could also seek to reduce the interaction of a vortex (or a separation) at the free end 51 of the blade by reducing the load at the free end 51 of the blade 90 of the first series of blades 9.

[0192] It is also proposed (as for example in [Fig.36]) that the blades 90 of the upstream propeller 9 have the greatest deflection at 0.4x(Re-Ri) from its free end 51, where the incident relative speeds are higher. Indeed, increasing the deflection makes it possible to reduce the incident speed which is perceived by the blade, which makes it possible to reduce its load and therefore the noise. In a particular embodiment, the increase in the deflection near the free end 51 can be achieved by defining a lower radius Re at the leading edge (BA) than at the trailing edge (BF), that is to say: Re,BA < Re,BF, when the lines of BA and BF are not merged. This can be combined with blades of the downstream propeller 6 having a Re,BA>Re,BF for the reasons indicated in the twenty-second consideration.

[0193] In [Fig.36], the sweep angle (marked by double points) of the multi-radius blade. The sweep of the blade increases near its free end 51, on the blade 9 taken as an example. ReBA and ReBF mark the aforementioned radius Re, therefore between the central longitudinal axis (X) and a location on the free end 51 at the leading edge and the trailing edge of the blade, respectively.

[0194] Twenty-third consideration: We may also seek to decorrelate the noise sources along the leading edge 61 and trailing edge 91, as well as their interactions between the propellers 9,6.

[0195] It is also proposed that the trailing edge 93 and leading edge 63 undulations have an amplitude (h(r)) and / or a spacing (X(r)) which vary inversely in the radial direction towards the free end 51 of the blade, as in [Fig.37].

[0196] In other words, it is proposed that

[0197] - the trailing edge undulations 93 on the blades 90 of the first series of blades 9 decrease (or increase) in amplitude, h(r), and / or spacing, X(r), towards the free end 51,

[0198] - the leading edge undulations 63 on the blades 60 of the second series of blades 9 increase (or decrease) in amplitude, h(r), and / or spacing, X(r), towards the free end 51.

[0199] Twenty-fourth consideration: To further decorrelate noise sources and reduce interactions between propellers, it is also proposed: - that, towards the free end 51, the trailing edge undulations 93 on the blades 90 of the first series of blades 9 decrease in amplitude h(r), and / or in spacing, X(r), and - that the leading edge undulations 63 of the blades 60 of the second series of blades 6 increase towards the free end 51 of these blades 60.

[0200] Twenty-fifth consideration: We may also seek to reduce the noise sources at the location where the acoustic radiation is at its maximum, that is to say, near the free end 51.

[0201] It is also proposed that near the free ends of the blades 90 of the first series of blades 9 and the blades 60 of the second series of blades 6, there are trailing edge 93 and leading edge 63 undulations, and that these undulations begin with a tooth or a peak 630,635, not with a trough.

[0202] Concerning each blade, and at least in the intermediate part of its span L, it will be preferable: - that its leading edge is convex, or curved upstream, and - that its leading edge is concave, or hollowed out towards the downstream. And we will also note that, among the possible configurations of the 90 and / or 60 blades, we can find: - certain blades with a smooth leading edge, over the entire span L, and a trailing edge at least partly corrugated, and - other blades with a smooth trailing edge, over the entire span L, and with a leading edge that is at least partly corrugated.

Claims

Claims

1. An aeronautical thruster along which a gas flow can circulate from upstream to downstream, the thruster having a central longitudinal axis (X), and comprising: - a first series of blades (9), - a second series of blades (6) positioned downstream of the first series of blades (9), - drive means (3, 18, 21, 23, 40, 94, 104, 920) for driving in rotation about the central longitudinal axis the blades of at least one of the first series of blades (9) and the second series of blades (6), - a nacelle (5) which has an aerodynamic external surface (50) relative to which the first series of blades (9) and the second series of blades (6) project radially from the central longitudinal axis (X), each blade of the first series of blades and of the second series of blades having: — a free end (51) opposite a connection end (53) forming a blade root close to the nacelle (5), — an intrados face (55) and an extrados face (57), — at a defined radius centered on the central longitudinal axis (X), a chord C, — a radius (Ri) between the central longitudinal axis (X) and a location, on the blade or a wedging arm (39,36) of the blade, which is level with the external surface (50) of the nacelle (5), — a radius (Re) between the central longitudinal axis (X) and a location on the free end (51) of the blade furthest from the central longitudinal axis (X), along a direction transverse to the central longitudinal axis (X), and — a span (L) defined, radially to the central longitudinal axis (X), between the free end (51) and the connecting end (53), in said transverse direction, at least some of the blades of the first series of blades (9) being variable pitch, so that each of them can pivot around a said pitch arm (39) to which said blade is fixed, around a pitch axis (390) which passes through the blade (90), and / or at least some of the blades of the second series of blades (6) being variable pitch, so that each of them can pivot around a said pitch arm (36) to which said blade is fixed, around an axis of pitch (360) which passes through the blade (60), in which propeller at least one of the blades of the first series of blades (9) has a trailing edge (91) along which undulations (93) extend, and / or at least one of the blades of the second series of blades (6) has a leading edge (61) along which undulations (63) extend, and in which propeller, a variation in pitch angle (Ay) on the blade, between a profile containing a tooth tip (635) and a profile containing a tooth trough (634), adjacent to each other, - along the span (L) of the blade (90,60), or - radially to the central longitudinal axis (X), is less than 45°.

2. An aeronautical propeller according to claim 1, wherein the drive means comprise a gas turbine (21,23,24,96,102) for said rotational drive about the central longitudinal axis (X) of the blades of at least one of the first series of blades (9) and second series of blades (6).

3. An aeronautical propeller according to claim 1 or 2, wherein the drive means comprise a speed reducer (104) engaged with the blades of at least one of the first series of blades (9) and second series of blades (6), to adapt the speed of rotation of said blades around the central longitudinal axis (X).

4. 4. An aeronautical propeller according to any one of the preceding claims, wherein the undulations (93,63) on one of said blades have a maximum amplitude h(r) between an adjacent apex and trough, h(r), such that: 0.0005xCmax < h(r) maximum < 0.5xCmax, where Cmax is the maximum chord of the blade and h(r) corresponds to the difference in chord between a profile at an adjacent apex and a profile at a trough in the direction of the span (L) of the blade, or radially to the central longitudinal axis (X).

5. An aeronautical propeller according to any one of the preceding claims, wherein at least some of the blades (90, 60) comprise a composite material and, on the intrados face (55) and / or the extrados face (57), a metal reinforcing cover (73) fixed with the composite material and extending along at least part of the leading edge and / or the trailing edge.

6. An aeronautical propeller according to claim 5 wherein the corrugations (93, 63) are formed only on the reinforcing cover (73). metallic, not on the composite material.

7. An aeronautical propeller according to any one of the preceding claims, in which, on at least some of the blades, the undulations have an amplitude between a peak and an adjacent trough, h(r), and a spacing between two successive peaks of undulations, X(r), which vary radially.

8. Aeronautical thruster according to claim 7, in which the laws of evolution of the amplitude, h(r), and of the spacing, X(r), of undulations are functions defined piecewise along the span (L) of the blade (90,60).

9. An aeronautical propeller according to claim 7 or 8, wherein, on at least some of the blades: - a first part of the leading edge (61) and / or the trailing edge (91) is smooth, without undulations, - a second part of the leading edge (61) and / or the trailing edge (91) has said undulations (63, 93), and - the amplitude of the undulations decreases where the second part joins the first part, so that the connection is progressive.

10. An aeronautical propeller according to any one of the preceding claims, wherein the undulations (63,93) are located only towards the free end (51) of several said blades (90,60) and / or at a radial position where the chord is greatest, on the blade (90,60).

11. An aeronautical propeller according to any one of the preceding claims, wherein, along the leading edge and / or the trailing edge, the undulations extend along a cumulative length H which is limited to: H / (Re-Ri) <0.

8.

12. An aeronautical propeller according to any one of the preceding claims, wherein the leading edge (63) and / or trailing edge (93) undulations are located: - beyond 0.4x(Re-Ri), starting from the side of the connection end (53), and / or - where the chord (C) is greatest.

13. An aeronautical propeller according to claim 2, or any one of claims 3 to 12 attached thereto, wherein the gas turbine (21,23,24,96,102) belongs to an engine (3) for driving the blades in rotation around the central longitudinal axis (X) and the first series of blades (9) and the second series of blades (6) are located towards an upstream end of the motor (3).

14. Aeronautical propeller according to claim 13, in which: - the engine (3) has an air inlet (35), for bringing an air flow (Fp) into the engine, and - the air inlet (35) of the flow is located axially between the setting axes (390, 360) of the blades of the first series of blades (9) and of the second series of blades (6).

15. Aeronautical propeller according to claim 14, in which: - the air inlet (35) of the flow has a nozzle located at a radius Rb from the central longitudinal axis (X), and - the trailing edge (91) of the first series of blades (9) has undulations located at the radius Rb.

16. An aeronautical propeller according to claim 2, or any one of claims 3 to 15 attached thereto, wherein the turbine (21,23,96,102) is connected to the first series of blades (9) to drive in rotation about the central longitudinal axis (X) only the blades (90) of the first series of blades (9), the blades of the second series of blades (6) defining air rectifier blades (60).

17. An aeronautical propeller according to claim 16 wherein, circumferentially around the central longitudinal axis (X), only the air straightener blades (60) located in a first angular range of + / -60° relative to 3 o'clock, and in a second angular range of + / -60° relative to 9 o'clock, have undulations (63).

18. An aeronautical propeller according to any one of the preceding claims, wherein the undulations (93,63) are located at least on the trailing edge (91) of the first series of blades (9) and on the leading edge (61) of the second series of blades (6).

19. Aeronautical propeller according to any one of the preceding claims, in which: - the undulations (93,63) at the leading edge (61) and at the trailing edge (91) have different geometries or patterns from each other, and / or - according to at least one of the pattern, amplitude, spacing, radial positioning of the undulations along the span, the undulations (93,63) of some of the blades of the first series of blades (9) are different from the undulations (93,63) of some of the blades of the second series of blades (6).

20. An aeronautical propeller according to any preceding claim, wherein at least some of the undulations (93,63) have an amplitude between a peak and an adjacent trough (h(r)) and / or a spacing between two successive peaks of undulations (X(r)) which varies monotonically or strictly monotonically in the radial direction towards the free end (51) of the blade.

21. An aeronautical propeller according to claim 20, wherein the trailing edge (93) and leading edge (63) undulations have the amplitude (h(r)) and / or the spacing (X(r)) vary(s) respectively inversely in the radial direction towards the free end (51) of the blade.

22. An aeronautical propeller according to any preceding claim, wherein the trailing edge undulations (93) on the blades (90) of the first series of blades (9) decrease in amplitude (h(r)) between a peak and an adjacent trough, and / or in spacing (X(r)) between two successive peaks of undulations, towards the free end (51), and wherein the leading edge undulations (63) of the blades (60) of the second series of blades (6) increase towards the free end (51).

23. Aeronautical propeller according to any one of claims 5, 6 in which the cover (73) of at least one of the blades connects the leading edge to the trailing edge, by the free end (51) of the blade.

24. Aeronautical propeller according to any one of the preceding claims, in which the free end (51) of at least one of the blades (90, 60): - is elliptical in shape (79) or rounded in shape (80), and / or - has an abrupt change in the bank angle (77).

25. 25. Aeronautical propeller according to any one of the preceding claims, in which the leading and / or trailing edge of at least some of the blades (90, 60) is formed locally by a porous material (85), at the location of at least some of the undulations.

26. Aeronautical propeller according to any one of the preceding claims, in which, on a leading edge (61) or trailing edge (91) zone of a blade having undulations (63, 93), a variation in skeleton angle (A[31) at the leading edge or (A[32) at the trailing edge, between a tooth tip (635) and a tooth trough (634), adjacent to each other, - along the span (L) of the blade (90, 60), or - radially to the central longitudinal axis (X), is less than 45°. O

27. ​​An aeronautical propeller according to any one of the claims preceding, in which, on a leading edge (61) or trailing edge (91) zone of a blade having undulations (63,93), a variation in skeleton angle (A[31) at the leading edge or (A[32) at the trailing edge, between two tooth peaks (630,631) and / or two tooth troughs (632,634), adjacent to each other, - along the span (L) of the blade (90,60), or - radially to the central longitudinal axis (X), is less than 45°.

28. Aeronautical propeller according to any one of the preceding claims, in which a variation in pitch angle (Ay) on the blade, between two adjacent profiles containing tooth peaks (630,631) and / or between two profiles containing tooth troughs (632,634), adjacent to each other, - along the span (L) of the blade (90,60), or - radially to the central longitudinal axis (X), is less than 45°.

29. Aeronautical propeller according to any one of the preceding claims, in which a variation in pitch angle (Ay), between two adjacent profiles containing tooth peaks (630,631) and / or between two profiles containing tooth troughs (632,634), adjacent to each other, - along the span (L) of the blade (90,60), or - radially to the central longitudinal axis (X), is less than 45°.

30. An aeronautical propeller according to any one of the preceding claims, wherein at least one of the blades (90,60) of one of the propellers (9,6) has a greatest sweep at a radial position located over a radial length of 0.4x(Re-Ri) of the free end (51).

31. Aeronautical propeller according to any one of the preceding claims, in which at least one of the blades (60) of the downstream series of blades (6) has a radius (Re) greater at the leading edge (BA) than at the trailing edge (BF), when the leading edge and trailing edge lines are not merged, and in which at least one of the blades (90) of the first series of blades has a radius Re smaller (preferably) or greater at the leading edge (BA) than at the trailing edge (BF), when the leading edge and trailing edge lines are not merged.

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