Improved sound dampening device for an aircraft power unit

Description

Technical Field

[0001] The present invention relates to the general field of turbomachines, and applies more particularly to large-diameter aircraft turbojets and turboprops. The invention relates in particular to a propulsion unit for aircraft comprising such a turbojet and / or turboprop. Prior art

[0002] It is envisaged, for new generations of commercial aircraft, to use turbojets with a high bypass ratio (BPR), for example, bypass ratios greater than 10, and with a large diameter, i.e. with diameters greater than 1.5 m, in order to increase the propulsive efficiency of turbojets and reduce their fuel consumption, as well as the emissions of polluting gases. It should be noted that the term "turbomachine", encompassing "turbojets" and "turboprops", designates a gas turbine device providing, by reaction to the high-speed ejection of hot gas, a thrust which contributes to propulsion.

[0003] Different types of turbomachinery, including turbojets, both shrouded and unshrouded, can be used: Ultra High Bypass Ration (UHBR): A shrouded turbojet engine with a fixed-pitch fan wheel, with or without a fan speed reduction unit, as shown in the figure. figure 1A , “Variable Pitch Fan” (VPF): ducted turbojet with a variable-pitch fan, “Counter Rotating Open Rotor” (CROR): unducted turbojet with two (at least) counter-rotating variable-pitch propellers, as illustrated in the figure 1B , “Unducted Stator Fan” (USF): unducted turbojet with a variable-pitch propeller wheel and a fixed or variable-pitch rectifier (stator) wheel.

[0004] Unducted turboprop engines can also be used. Note that the term "turboprop" refers to a gas turbine engine whose thrust is primarily obtained by the traction of the variable-pitch propeller.

[0005] However, a major drawback of these integrated architectures on commercial aircraft is their acoustic impact, i.e., noise levels. The noise levels emitted by aircraft are subject to increasingly strict international regulations during takeoff and landing phases, in order to limit the acoustic footprint in the vicinity of airports.

[0006] One of the main sources of noise on large diameter turbojets and turboprops is linked to their integration or installation on aircraft, particularly on or under the wing. figures 1A et 1B illustrate two examples of turbojet engines (ducted and unducted) with a large diameter D>1.5 m, integrated on an aircraft wing 10. The turbojet engine 1 is fixed on the wing 10 of the aircraft, by means of a mast 40, the wing 10 itself being fixed to the fuselage 20 of the aircraft. The ducted turbojet engine 1 ( figure 1A ) comprises a fan 30 surrounded by a nacelle 32. The unducted turbojet 1 ( figure 1B ) comprises two counter-rotating propellers 31.

[0007] The integration of these large diameter turbojets and / or turboprops increases the interaction of the airflow accelerated by the rotating part, i.e. by the fan 30 or the propellers 31 of the turbojet (which produces a high-speed jet of air), with the wing 10 (in particular the leading edge 11, and in particular the leading edge line 11a), the engine pylon 40 (or pylon), high-lift devices 12, and other elements of the aircraft which generate additional sources of noise. Indeed, the dimensions of these turbojets are such that the rotating part is arranged upstream of the leading edge 11 of the wing 10, and opposite it, so that the turbulent flows generated in the wake of the fan 30 (or the propellers 31) directly impact the leading edge 11. The turbulent flows generated in the wake of the fan 30 or the propellers 31 are represented by the spirals S on the figures 1A et 1B , and the generated noise is represented by the dashed arcs. It should be noted that the jet-wing interaction produces broadband noise.

[0008] In addition, this turbojet and / or turboprop architecture has a short and thin nacelle 32, in the case of a ducted turbojet (UHBR or VPF type), and the nacelle may even be absent in the case of an unducted turbojet (for example, CROR type). This implies a significant reduction in the surfaces that can accommodate acoustic treatments, such as honeycomb-type resonators or absorbent materials, and therefore noise reduction means. WO 2020 / 079335 A1 discloses an example of a structure for the acoustic treatment of aircraft profiles.

[0009] There is therefore a need for a device to at least partially overcome the aforementioned drawbacks. Statement of the invention

[0010] The present disclosure relates to a propulsion assembly for an aircraft comprising a turbomachine having at least one rotating part configured to rotate about an axis of rotation of the turbomachine, a mounting pylon, and an aircraft wing supporting the turbomachine via the mounting pylon, the at least one rotating part being arranged upstream of the wing and the mounting pylon such that a jet of air exiting the rotating part, in the wake thereof, impacts said wing and said mounting pylon, a leading edge of said wing locally comprising at least one acoustic attenuation device arranged at least partly in the wake of the rotating part, the acoustic attenuation device being a local modification of the structure and / or the profile of the leading edge, and extending over a distance L along the leading edge of the wing, such that L<2.5D, where D is the diameter of the rotating part of the turbomachine.

[0011] In some embodiments, the turbomachine is a turbojet or a turboprop. In the present disclosure, the terms “upstream”, “downstream” and their derivatives are defined according to a normal flow direction of the air through the turbomachine, along the axial direction of the turbomachine, i.e. along the axis of rotation thereof. For example, when the rotating part is a fan of the turbomachine, and the structural element is an aircraft wing, the fan is arranged upstream of the wing, in particular the leading edge of the wing, such that the air is first sucked in by the fan, then expelled downstream of the fan, towards the wing.

[0012] Furthermore, "arranged in the wake" means that a portion of the leading edge of the wing is arranged in the path of the air expelled by the fan and flowing downstream of the latter, and consequently impacting said portion of the leading edge.

[0013] Furthermore, by "locally" is meant that the acoustic attenuation device does not extend over the entire length of the leading edge of the structural element, but only over a portion of this leading edge. In other words, the structure and / or the profile of the leading edge is only modified locally, in particular in the wake of the rotating part, downstream of it. Preferably, at least 30% of the length of the leading edge of the structural element does not include the acoustic attenuation device.

[0014] Preferably, a large portion of the acoustic attenuation device is disposed in the wake of the rotating portion, for example at least 70% of the length of said device along the leading edge.

[0015] Arranging this acoustic attenuation device, at least in part, in the wake of the rotating part makes it possible to attenuate the noise generated by the air impacting the leading edge of the structural element and / or the suspension pylon. In particular, the propulsion unit according to the present disclosure makes it possible to reduce the acoustic impact of a ducted or unducted turbomachine, when the latter is installed under or on the wing of an aircraft. In addition, modifying the leading edge of the structural element, for example an aircraft wing, only locally, makes it possible to limit the influence of this device on the aerodynamics of the wing and on the performance of the aircraft, in particular when these modifications are on a moving part of the wing. In addition, this local modification makes it possible to adapt the noise reduction devices to the local properties of the flow (or jet) which impacts the leading edge of the wing.Furthermore, modifying the structure and / or profile of the leading edge makes this device a passive noise reduction device, not requiring dedicated air samples or active sensors with an energy cost.

[0016] In some embodiments, the local modification of the leading edge structure and / or profile is not uniform along a leading edge line.

[0017] It is thus possible to adapt the shape and / or structure of the leading edge locally, depending on the local properties of the flow, or air jet which impacts the leading edge.

[0018] In some embodiments, the acoustic attenuation device includes undulations along the leading edge, the undulations having a succession of troughs and peaks.

[0019] Preferably, a corrugation comprises at least two troughs and at least two peaks. It is understood that the “troughs” and “peaks” are considered according to the direction of the chord of the structural element, the latter being substantially parallel to the axis of rotation of the turbomachine. Consequently, the troughs are regions where the chord of the structural element is weaker than at the peaks. These corrugations, or serrations, constitute a local modification of the profile of the leading edge, and therefore a passive noise attenuation solution.

[0020] In some embodiments, an amplitude h(z) of the undulations and / or a spacing λ(z) between two successive peaks of the undulations varies along the leading edge as a function of a distance E, the distance E being a distance between a position z along the leading edge and the rotating portion, in a direction parallel to the axis of rotation.

[0021] It is understood that an amplitude h(z) of the undulations corresponds, at a given position z along the leading edge, to the distance between the bottom of the trough of the undulation and its peak closest to this position z, this distance being considered in a direction parallel to the axis of rotation. Similarly, the spacing λ(z) corresponds to the distance, at a given position z along the leading edge, between two peaks (or two troughs) of a undulation closest to this position z.

[0022] Since the distance between the rotating part and the leading edge is not constant along the leading edge line, the impact of the air on the leading edge and therefore the local properties of the flow therefore also vary as a function of this distance. Modifying the amplitude of the undulations and / or the spacing between two successive peaks of the undulations thus makes it possible to adapt the device as a function of this distance, and therefore as a function of the properties of the flow.

[0023] In some embodiments, the amplitude h(z) of the undulations is such that 0.005c(z) < h(z) < 0.5c(z), c(z) being the value of a chord of the structural element or said suspension mast as a function of a position z along the leading edge. This relationship makes it possible to limit the mechanical impact of the undulations, or serrations.

[0024] In some embodiments, the acoustic attenuation device is such that the leading edge locally comprises a porous material.

[0025] The porous material may be a metal foam. Porous materials such as metal foams have the advantage of having good acoustic properties, and of being effective in reducing broadband noise, particularly at low frequencies. This configuration corresponds to a local modification of the structure of the leading edge, also making this device a passive acoustic attenuation device. Preferably, the dimensions and positions of the region(s) of the leading edge having this material are similar to those defined in the case of corrugations, in particular the dimensions L, corresponding to the length, along the leading edge line, over which the porous material extends, and h, corresponding to the length of this porous material, in the direction of the chord of the support element.

[0026] In some embodiments, the porous material has a variable porosity ratio t(z) along the leading edge.

[0027] The porosity rate is the ratio of the volume occupied by air to the total volume of the porous material. In the same way as for corrugations, these variations in the porosity rate make it possible to adapt the absorbent properties of the material according to the properties of the flow of the jet / wake generated by the blades / vanes of the rotating part arranged upstream of the structural element. The porosity rate is preferably lower for the porous materials closest to an aircraft fuselage, than for the porous materials furthest from the fuselage, when the acoustic attenuation device is applied to an aircraft wing.

[0028] In some embodiments, the spaces between two peaks of the undulations are at least partially filled with a porous material.

[0029] This configuration allows for the combination of a modification of the profile, and a modification of the structure of the leading edge, thus further improving the effectiveness of the device's acoustic attenuation device.

[0030] The acoustic attenuation device extending over a distance L along the leading edge of the wing, such that L<2.5D, where D is a diameter of the rotating part of the turbomachine.

[0031] In certain embodiments, the spacings λ(z) between two successive peaks and / or the amplitudes h(z) of the undulations are smaller for the undulations closest to a fuselage of the aircraft, than for the undulations furthest from the fuselage.

[0032] When the structural element is a wing attached to a fuselage, the distances between the leading edge of the wing and the rotating part are generally smaller near the fuselage, i.e. near the root, or embedding of the wing. Consequently, reducing the spacing between two successive peaks, in other words the wavelength of the undulations, and / or the amplitudes of the undulations, allows the acoustic attenuation device to adapt to this configuration.

[0033] In some embodiments, the at least one acoustic attenuation device comprises a single hollow, the hollow extending a distance L along the leading edge of the airfoil, such that L ≤ 2.5D.

[0034] According to this configuration, the modification of the leading edge is not in the form of a corrugation (comprising at least one trough and one peak), but comprises a single trough, extending over a greater distance, along the leading edge line, than a trough of a corrugation. This configuration makes it possible to locally modify the aerodynamic profile of the wing. It will be noted that the leading edge may comprise two acoustic attenuation devices, therefore two of these troughs, arranged in different regions along the line of the leading edge, preferably at the location where the tip vortices of the propeller blade impact the wing. Alternatively, it is also possible to locally modify the sweep of the wing, thus locally increasing the distance between the leading edge and the rotating part.

[0035] Increasing this distance helps reduce the flow turbulence before interacting with the leading edge of the airfoil, effectively reducing noise. In addition, increasing the sweep decreases the correlation of noise sources along the leading edge. Similarly to undulations, the maximum amplitude of the trough h(z), i.e. the maximum depth of the trough, can verify the relationship 0.005c(z) < h(z) < 0.5c(z).

[0036] In some embodiments, the leading edge locally comprises at least one first and at least one second acoustic attenuation device extending respectively over a distance L1 and L2 along the leading edge, such that L1 + L2 <2.5D.

[0037] Each of the first and second sound attenuating devices may comprise undulations or a single trough. Furthermore, the number of sound attenuating devices along the leading edge is not limited to two, but may be greater. For example, a number n of sound attenuating devices may be arranged locally along the leading edge, such that L1 + L2 + ... + Ln < 2.5D.

[0038] In some embodiments, a cover is configured to transition from a closed position in which the acoustic attenuation device is covered by the cover, to an open position in which the acoustic attenuation device is uncovered.

[0039] The cover can be a movable cowl, a leading edge morphing system, or a leading edge slat. This cover allows the undulations to be visible or active only during certain phases of flight, such as takeoff and landing. This allows the undulations to be hidden, and therefore a smooth leading edge to be found in cruise, i.e., outside of the takeoff and landing phases. The benefit of hiding the undulations in cruise is to eliminate their impact on overall aerodynamic performance when noise reduction is not required.

[0040] In some embodiments, a diameter D of the rotating part of the turbomachine is such that D>1.5 m.

[0041] In some embodiments, the turbomachine comprises a fixed portion disposed downstream of the rotating portion, the suspension mast being attached to the fixed portion. The fixed portion may be a straightener for the flow exiting the rotating portion.

[0042] The present disclosure also relates to an aircraft comprising the propulsion assembly according to any of the preceding embodiments. Brief description of the drawings

[0043] The invention and its advantages will be better understood upon reading the detailed description given below of different embodiments of the invention given as non-limiting examples. This description refers to the appended pages of figures, in which: [ Fig. 1A-1B ] THE figures 1A et 1B represent front and top views of propulsion units according to the prior art, [ Fig. 2 ] There figure 2 represents a top view of a propulsion unit according to a first embodiment of the invention, [ Fig. 3 ] There figure 3 represents a top view of a propulsion assembly according to a modified example of the first embodiment of the invention, [ Fig. 4 ] There figure 4 represents a top view of a propulsion assembly according to another modified example of the first embodiment of the invention, [ Fig. 5 ] There figure 5 represents a top view of a propulsion unit according to a second embodiment of the invention, [ Fig. 6 ] There figure 6 represents a top view of a propulsion assembly according to a modified example combining the first and second embodiments of the invention, [ Fig. 7 ] There figure 7 represents a top view of a propulsion unit according to a third embodiment of the invention, [ Fig. 8 ] There figure 8 represents a top view of a propulsion assembly according to a modified example of the third embodiment of the invention, [ Fig. 9 ] There figure 9 represents a top view of a propulsion assembly according to an embodiment not forming part of the claimed invention, [ Fig. 10A-10C ] THE figures 10A à 10C represent different examples of aircraft, seen from the front, comprising a propulsion unit according to any one of the embodiments of the invention. Description of the embodiments

[0044] A first embodiment of the present disclosure will be described with reference to the figures 2 à 4 . It will be noted that the remainder of the description refers to a propulsion unit comprising a turbojet. This example is however not limiting, the invention can also be applied to a propulsion unit comprising a turboprop.

[0045] The propulsion system illustrated on the figure 2 comprises an unducted turbojet 1 having a rotating part, and being fixed to a structural element by means of a pylon or attachment mast 40 (not visible on the figure 2 ). In this example, the structural element is an aircraft wing 10, and the turbojet engine 1 is fixed below the wing 10. The rotating part of the turbojet engine 1 comprises the counter-rotating propellers 31, movable around an axis of rotation A. A spatial reference frame comprises an axis X corresponding to the main axis of the fuselage 20, substantially parallel to the axis of rotation A, and an axis Z perpendicular to the axis X, corresponding to the longitudinal direction in which the wing 10 extends, from its root 10a fixed to the fuselage 20, to its opposite end 10b. The plane XZ, corresponding to a horizontal plane, which comprises the leading edge line 11a of the wing 10, in the absence of a dihedral angle on the wing. In the presence of a dihedral angle, the XZ plane would not be horizontal, but would include the leading edge line 11a. A front view is a view parallel to the X axis, and a top view is a view perpendicular to the XZ plane.Furthermore, an upstream-downstream direction corresponds to a direction of flow of the air entering at a speed U into the rotating part, then leaving the rotating part, towards the leading edge 11 of the wing 10, along the axis of rotation A, in other words along the axis X. The propellers 31 are arranged upstream of the leading edge 11, and opposite it along the axis X.

[0046] The propulsion assembly according to this embodiment differs from the assemblies illustrated in the figures 1A et 1B , in that it comprises an acoustic attenuation device, the acoustic attenuation device comprising undulations 50, or serrations, locally along the leading edge line 11a. These undulations extend over a length L along the leading edge line 11a, so that L<2.5D, D being the diameter of the turbojet 1, i.e. the diameter of the propellers 31 in the case of an unducted turbojet, or the diameter of the air inlet of the nacelle 32 in the case of a ducted turbojet, equivalent to the diameter of the rotating part. It should be noted that the figures do not illustrate machines at real scales and proportions, the dimensions of the turbojet, in particular its diameter D, being deliberately exaggerated, in order to facilitate visibility and description. For example, the wingspan of a 200-passenger commercial aircraft is around 15 to 20 m, with a 2 m diameter engine.

[0047] Apart from the undulations 50, the leading edge line 11a has an unmodified profile, corresponding to a usual profile, or neutral profile, not including an acoustic attenuation device. The neutral profile is represented by the broken line on the figure 2 The length L is such that a large part of the undulations 50, preferably at least 70% of the length L, are located in the wake of the rotating part of the turbojet, that is to say in this example the propellers 31. In other words, a fictitious cylinder of diameter D and axis A, extending downstream of the propellers 31, comprises a part of the leading edge 11, and in particular comprises at least 70% of the length L of the acoustic attenuation device.

[0048] The undulations 50 are characterized by a succession of troughs 51 and peaks 52. A trough 51 corresponds to a local depression of the leading edge 11 towards the downstream, compared to an unmodified profile of the leading edge (see broken lines on the figure 2 ) in the direction of the chord c of the wing, corresponding to the direction of the X axis. A vertex 52 corresponds to a relief, or advance of the leading edge upstream, compared to an unmodified profile of the leading edge (see broken lines on the figure 2 ) in the direction of the chord c of the sail, corresponding to the direction of the X axis.

[0049] A height h is a distance, in a direction perpendicular to the leading edge neutral line 11a, or in a direction parallel to the engine axis A, between a successive trough 51 and a peak 52, more precisely between the bottom of the trough 51 and the end of the peak 52. In other words, the height h corresponds to the amplitude of the undulations 50. A length λ is a distance, in a direction parallel to the leading edge neutral line 11a, between two successive troughs 51 (or between two successive peaks 52). In other words, the length λ corresponds to the wavelength of the undulations 50.

[0050] In the example illustrated on the figure 2 , the propulsion assembly comprises a single acoustic attenuation device, extending over a length L along the leading edge line 11a. The figure 3 illustrates an alternative example of the first embodiment, in which the propulsion unit comprises two acoustic attenuation devices, extending respectively over a distance L1 and L2 along the leading edge line 11a, such that L1 + L2 < 2.5D. According to this alternative, the turbojet 1 is a ducted turbojet comprising a rotating part, in this case, a fan 30, and a nacelle 32. The turbojet 1 is fixed above the wing 10 by means of a mounting pylon 40. The respective undulations 50a, 50b of the two acoustic attenuation devices are arranged on either side of the pylon 40, in the wake of the fan 30, such that the air jet (not shown again on the figure 3 and the following figures, for the sake of clarity) emerging from the blower 30, impacts the corrugations 50a, 50b. In this example, the corrugations 50a, 50b each comprise two troughs 51 and two peaks 52. However, more troughs 51 and peaks 52 are also possible for each acoustic attenuation device 50a, 50b.

[0051] In the examples of the first embodiment, illustrated in the figures 2 And 3 , the undulations are uniform along the entire length L, and are also identical along the lengths L1 and L2. By "uniform" or "identical" we understand that the amplitudes h and the wavelengths λ are constant along the leading edge line 11a.

[0052] There figure 4 illustrates a modified example of the first embodiment, in which the propulsion unit also comprises two acoustic attenuation devices. The turbojet 1 is furthermore an unducted turbojet, fixed above the wing 10. Unlike the example illustrated in the figure 3 , the undulations 50a and 50b are not identical, but have different amplitudes h and / or wavelengths λ between these two devices. More precisely, the amplitude h(z) and the wavelength λ(z) vary depending on the position z along the leading edge line 11a, in the Z direction.

[0053] It will be noted that the distances E1 and E2 each correspond to a distance between a point along the leading edge line 11a, for example a peak 52 of a corrugation, and the rotating part, in the X direction. More precisely, the distance E2 corresponds to the distance between the rotating part and the leading edge 11, at the trailing edge of a first radial end of the rotating part, corresponding to the end closest to the fuselage 20, and the distance E1 corresponds to the distance between the rotating part and the leading edge 11, at a second radial end of the rotating part, diametrically opposite the first radial end, and corresponding to the end furthest from the fuselage 20.

[0054] Taking into account the profile of the wing 10, the section of which decreases from the root 10a to the end 10b, and the sweep angle of the wing 10, the distance E2 is smaller near the fuselage 20, than the distance E1, further away from the fuselage 20. Consequently, the size of the vortices, characterized by an integral scale Λ (represented on the figure 2 ), generated by the turbulent flow downstream of the rotating part, and impacting the leading edge 11, differs depending on the position z along the leading edge line 11a. Indeed, near the trailing edge of the rotating part, the width of the wake is small, and characterized by small vortices. The turbulent flow then develops downstream, gradually increasing the size of the vortices, i.e. the integral scale Λ. When the distance between the rotating part 30, 31 and the leading edge 11 is smaller, the size of the vortices impacting the leading edge 11 is also smaller, since these vortices were able to develop to a lesser extent than when this distance is greater.

[0055] In this context, it is therefore preferable that the wavelengths λ(z) and / or the amplitudes h(z) are lower for the undulations closer to the fuselage 20 of the aircraft. In the example illustrated in the figure 4 , the wavelengths λ 1 (z) of the undulations 50a and the wavelengths λ 2 (z) of the undulations 50b are such that λ 1 (z) > λ 2 (z). Indeed, the smaller the distance between the rotating part 30, 31 of the turbojet and the leading edge 11 of the wing 10, the smaller the integral scale Λ (represented on the figure 2 ) of the turbulent flow downstream of the rotating part, decreases. However, the optimal amplitudes and wavelengths of the undulations depend on this integral scale Λ of the turbulence. Preferably, the wavelength values λ, as a function of the position z along the leading edge line 11a, approximately verify λ(z) = Λ(z) / 2, in order to maximize the acoustic gains.

[0056] The example of the figure 4 illustrates a case in which the propulsion assembly comprises two acoustic attenuation devices 50a, 50b, the undulations of each of them being different. This example is however not limiting, a configuration according to which the propulsion assembly comprises a single device, as in the example of the figure 2 , but with non-uniform undulations over the length L, also being possible. The wavelengths λ can be for example increasing, from the end of the undulations 50 closest to the fuselage 20, to the end of the undulations 50 furthest from the fuselage 20. Furthermore, the amplitude h of the undulations 50 preferably verifies 0.005c(z) < h(z) < 0.5c(z), where c(z) corresponds to the value of the chord c of the wing 10, at a position z along the leading edge line 11a.

[0057] A second embodiment of the present disclosure will be described with reference to the figure 5 .

[0058] The propulsion system illustrated on the figure 5 comprises a ducted turbojet 1 comprising a nacelle 32 surrounding a fan 30 and being fixed above a wing 10 by means of a pylon or attachment mast 40. The propulsion unit according to the second embodiment differs from the first embodiment in that the acoustic attenuation device is such that the leading edge 11 of the wing 10 locally comprises a porous material 60, preferably a metal foam. Unlike the presence of undulations, the profile of the leading edge line 11a is not modified, but the structure of the leading edge 11, in particular the material, is modified locally.

[0059] In the same way as for the first embodiment, a single acoustic attenuation device can be arranged along the leading edge over a distance L, in which the porous material may not be uniform, for example having a variable porosity rate along the leading edge line. The porosity rate is defined as the ratio of volume occupied by air to the total volume of the porous material. Alternatively, and according to the example illustrated in the figure 5 , two acoustic attenuation devices are arranged on either side of the mast 40, in the wake of the fan 30. These devices are such that the respective porous materials 60a, 60b of the acoustic attenuation devices each have a porosity rate t 1 (z) and t 2 (z), respectively, different from each other, and depending on the position z along the leading edge line 11a. In particular, the porosity rate t 2 (z) of the porous material 60b closer to the fuselage 20, is lower than the porosity rate t 1 (z) of the porous material 60a, further from the fuselage 20, for the same reasons as in the case of undulations. In particular, the variation of these porosity rates makes it possible to adapt the noise attenuation as a function of the distance between the leading edge 11 and the rotating part.Furthermore, the thickness h(z) of the porous material, i.e. the distance over which the porous material extends along the X axis, preferably verifies the same relationship as the amplitude of the undulations, as a function of the chord, i.e. 0.005c(z) < h(z) < 0.5c(z).

[0060] There figure 6 illustrates a configuration combining the first and second embodiments of the present disclosure. According to this configuration, the propulsion assembly comprises undulations 50a, 50b similar to the undulations shown in the figure 4 , in which the spaces between the teeth formed by the undulations are filled, partially or completely, with porous material, preferably the porous materials 60a, 60b as described in the embodiment illustrated in the figure 5 and having the same characteristics. This configuration makes it possible not to modify, or to minimize the modifications of the profile of the leading edge 11 of the wing 10.

[0061] A third embodiment of the present disclosure will be described with reference to figures 7 et 8 .

[0062] According to this embodiment, the profile of the leading edge 11 of the wing 10 is modified locally, in order to locally increase the distances E1, E2 between the rotors (or stators) of the rotating part of the turbojet 1 and the leading edge 11 of the wing 10. Unlike the undulations according to the first embodiment, comprising at least one trough and at least one peak, the attenuation device according to this third embodiment locally comprises a single trough 70. In the example illustrated on the figure 7 , the propulsion unit comprises two acoustic attenuation devices each comprising a single hollow 70a, 70b respectively, on either side of the attachment mast 40. The hollows 70a, 70b can extend respectively over a distance L1 and / or L2 of up to 2 times the diameter D of the turbojet. Furthermore, as for the undulations, the maximum amplitude h(z) of the hollow can verify the relationship 0.005c(z) < h(z) < 0.5c(z).

[0063] The local increase in this distance makes it possible to reduce the turbulence of the flow before it interacts with the leading edge 11 of the wing 10, and thus to improve the effectiveness of the noise reduction.

[0064] Alternatively, the local increase in the distance between the leading edge 11 and the rotating part can be induced by a local increase 80 of the arrow α of the wing 10, as illustrated in the figure 8 . According to this example, two local increases 80a, 80b of the sweep of the wing 10 are made on either side of the mast 40, in the wake of the fan 30. These increases 80a, 80b of the sweep are of a value α" and α' respectively, such that the increased sweep is equal to α" + α and α' + α respectively. The broken line represents the leading edge line in the absence of an acoustic attenuation device, with an unmodified sweep. The values α' and α" are preferably different, in order to adapt the modification of the sweep to the distance between the fan 30 and the leading edge 11. The local increase in the sweep makes it possible to reduce the correlation of the noise sources along the leading edge.

[0065] The sound attenuation devices according to the various embodiments of the invention can be used with defrosting systems, which can be reinforced at the hollows and porous materials.

[0066] There figure 9 illustrates a modified example of the present disclosure not forming part of the claimed invention, in which the acoustic attenuation device is arranged on a mounting pylon 40, making it possible to fix a turbojet engine 1 to the fuselage 20, and not on a wing 10. According to this configuration, the structural element is the portion of the fuselage 20 on which the mounting pylon 40 is fixed. This configuration is suitable for turbojet engines installed at the rear tip of the fuselage 20, in a so-called “puller” configuration, in which the propellers 31 (or the fan 30 in the case of a ducted turbojet engine) are arranged upstream of the mounting pylon 40, in the direction of flow of the air flowing along the axis of rotation A at a speed U. In the example illustrated on the figure 9 , the acoustic attenuation device comprises corrugations 50, and a porous material 60.

[0067] THE figures 10A à 10C illustrate different types of aircraft 100, on which the propulsion unit according to the present description is applicable. The invention can in particular be used for wings 10 having non-conventional profiles and shapes, such as low wings whose dihedral Δ is scalable in span, in particular stronger (dihedral Δ') on an internal side of the wing 10 ( figure 10A ). This type of wing provides ground clearance, i.e. the distance between the engine and the ground, defined in the standards for large diameter engines. figure 10B shows an aircraft having a high wing 10, and the figure 10C shows an aircraft whose wing 10 has an inverted arrow α.

[0068] Although the present invention has been described with reference to specific exemplary embodiments, it is evident that modifications and changes may be made to these examples without departing from the general scope of the invention as defined by the claims. In particular, individual features of the various illustrated / mentioned embodiments may be combined in additional embodiments. In particular, features related to the parameters described in one embodiment, for example the heights h and the lengths L, L1 and L2, may also be applicable to other embodiments described herein. Therefore, the description and the drawings should be considered in an illustrative rather than restrictive sense.

Claims

1. A propulsion assembly for an aircraft comprising a turbomachine (1) having at least one rotating part (30, 31) configured to rotate about an axis of rotation (A) of the turbomachine (1), an attachment strut (40), and an aircraft airfoil (10) carrying the turbomachine (1) via the attachment strut (40), the at least one rotating part (30, 31) being disposed upstream of the airfoil (10) and of the attachment strut (40) such that an air jet emerging from the rotating part (30, 31), in the wake thereof, impacts said airfoil (10) and said attachment strut (40), a leading edge (11) of said airfoil (10) locally comprising at least one acoustic attenuation device (50, 60, 70, 80) disposed at least partly in the wake of the rotating part (30, 31), the acoustic attenuation device being a local modification of the structure and / or of the profile of the leading edge (11), and extending over a distance L along the leading edge of the airfoil (10), such that L<2.5D, where D is the diameter of the rotating part of the turbomachine (1).

2. The assembly according to claim 1, wherein the acoustic attenuation device comprises corrugations (50) along the leading edge (11), the corrugations (50) having a succession of valleys (51) and peaks (52).

3. The assembly according to claim 2, wherein an amplitude h(z) of the corrugations (50) and / or a spacing λ(z) between two successive peaks (52) of the corrugations (50) varies along the leading edge (11) as a function of a distance E, the distance E being a distance between a position z along the leading edge (11) and the rotating part (30, 31), along a direction parallel to the axis of rotation (A).

4. The assembly according to claim 3, wherein the amplitude h(z) of the corrugations (50) is such that 0.005c(z) < h(z) < 0.5c(z), c(z) being the value of a chord of the airfoil (10) as a function of a position z along the leading edge (11).

5. The assembly according to any one of claims 1 to 4, wherein the acoustic attenuation device is such that the leading edge (11) locally comprises a porous material (60).

6. The assembly according to claim 5, wherein the porous material (60) has a porosity rate t(z) variable along the leading edge (11).

7. The assembly according to claim 2 or 3, wherein the spaces between two peaks (52) of the corrugations (50) are at least partially filled with a porous material (60).

8. The assembly according to any one of claims 1 to 7, wherein the at least one acoustic attenuation device comprises a single valley (70), the valley (70) extending over a distance L along the leading edge (11) of the airfoil (10), such that L<2.5D.

9. The assembly according to any one of claims 1 to 8, wherein the leading edge (11) locally comprises at least a first and at least a second acoustic attenuation device extending respectively over a distance L1 and L2 along the leading edge (11), such that L1 + L2 < 2.5D.

10. The assembly according to any one of claims 1 to 9, wherein the turbomachine comprises a fixed part disposed downstream of the rotating part (30, 31), the attachment strut (40) being fixed to the fixed part.

11. An aircraft comprising the propulsion assembly according to any one of the preceding claims.