Structural and / or acoustic panel with a U-shaped sealing flange directed towards the inside of the panel and method for producing such a panel

Description

Technical field

[0001] The invention relates to the field of manufacturing a structural and / or acoustic panel of the type comprising two skins and a honeycomb structure sandwiched between these skins. In particular, the invention relates to the aeronautical field, and more specifically concerns aircraft propulsion assembly components comprising at least one such panel. State of the prior art

[0002] An aircraft propulsion system comprises several structural panels commonly referred to as "sandwich panels", mainly formed by two skins and a honeycomb structure sandwiched between these skins. Such panels are, for example, used to form all or part of an ejection nozzle or a fixed internal structure of the propulsion system nacelle.

[0003] The honeycomb structure of the panel generally consists of transverse partitions connecting the skins and thus contributing to the structural strength of the panel. The transverse partitions delimit cells, i.e. spaces devoid of material, which in particular helps to reduce the mass of the panel.

[0004] Such a panel may also have an acoustic function to attenuate the noise generated by the propulsion unit. The skin intended to be oriented towards the source of the noise is for this purpose made permeable to air, typically using holes passing through this skin to guide the air into the cells and thus absorb acoustic energy.

[0005] Brazing is a conventional technique for attaching the constituent elements of such a structural and / or acoustic panel to each other during its manufacture. This technique consists of heating a filler metal to its melting temperature, interposed between the elements to be joined. The filler metal can take the form of a strip, a powder or a paste and is chosen so that its melting temperature is lower than the melting temperature of the elements to be joined.

[0006] To improve the structural strength of the panel, it generally includes closing flanges having a U-shaped section. The closing flanges are typically arranged on the lateral periphery of the honeycomb structure, according to the principle illustrated in the figure 2 of document WO 2018 / 087502 A1. Compared to a closing flange in the form of a solid bar, the U-shaped section allows the mass of the panel to be reduced.

[0007] Given the geometry of such a closing flange, the side perimeter of the panel includes a hollow, which can complicate its assembly with another structure.

[0008] Structural and / or acoustic panels according to the state of the art are presented in documents EP 2781728 A2, FR 3014011 A1, WO 2010 / 012900 A2, US 2009 / 019857 A1 and US 2016 / 312657 A1. Statement of the invention

[0009] The invention aims to provide a structural and / or acoustic panel making it possible to simplify the assembly of this panel with another structure.

[0010] For this purpose, the invention relates to a structural and / or acoustic panel for an aircraft propulsion system, this panel comprising an internal skin, an external skin, a honeycomb structure and at least one closing flange, this closing flange comprising a base and two lateral branches connected to the base, each lateral branch comprising a free end, the lateral branches and the base delimiting a cavity which opens at the free end of the lateral branches, the honeycomb structure and the closing flange being sandwiched between the internal skin and the external skin, one of the lateral branches of the closing flange being fixed to the internal skin, the other lateral branch of the closing flange being fixed to the external skin. According to the invention, the free end of each of the lateral branches is opposite the honeycomb structure.

[0011] In the following, reference is generally made to a single closing flange. The characteristics described for a closing flange can of course apply to each of the closing flanges of the panel when the panel comprises several.

[0012] The closing flange of the panel of the invention may have a conventional geometry, for example a section having a substantially U-shape. Thus, in one embodiment, the base may be substantially perpendicular to each of the lateral branches. In another embodiment, the base may have a non-planar geometry, for example a section having a V-shape pointing towards the outside of the panel. The invention is more specifically characterized by the arrangement of the closing flange relative to the other elements of the panel since, unlike conventional panels, the free ends of the lateral branches are not directed towards the outside of the panel but towards the inside thereof, i.e. towards the honeycomb structure.

[0013] In this way, the side perimeter of the panel is formed by the base of the closing flange.

[0014] This makes it easier to assemble the panel with other structures.

[0015] For example, where the closing flange has a flat base, the panel may be joined to another similar panel by welding the base of a respective closing flange of each of those panels to each other.

[0016] Furthermore, according to the invention, the panel comprises one or more reinforcements extending into said cavity so as to keep the lateral branches of the closing flange spaced apart from each other.

[0017] Such reinforcements make it possible to improve the manufacture of the panel, in particular the fixing of its constituent elements to each other, especially in the case where this fixing is carried out by brazing.

[0018] Indeed, the brazing of the panel can typically be carried out by pressurizing the elements to be assembled with gas in a vacuum furnace. This conventional technique allows one of the skins to be pressed against the other elements, under the action of a gas which expands when the temperature in the furnace increases.

[0019] One of the skins is thus pressed against the honeycomb structure and against one of the lateral branches of the closing flange, which causes the honeycomb structure and this closing flange to be pressed, via its base, against the other skin which itself rests on a counter-form.

[0020] The lateral branch located on the side of the skin resting on the counter-form is not pressed as such against this skin under the action of this gas pressure. Given the thermal expansion of the closing flange during brazing, this lateral branch is therefore likely to sag and not be correctly fixed to this skin, which can weaken the panel.

[0021] Since the cavity is oriented towards the inside of the panel, it is not possible to insert conventional tools into it to keep the side branches of the closing flange apart from each other in order to avoid such sagging.

[0022] The reinforcement(s) of the panel of the invention make it possible to achieve this support function.

[0023] What has just been described also applies when the plating of the elements to be assembled is obtained by applying mechanical pressure resulting for example from differential thermal expansion of tooling elements. In other words, the invention is compatible with different techniques for fixing the constituent elements of the panel to each other.

[0024] To some extent, reinforcements can also improve the structural strength of the panel during use.

[0025] Non-limiting examples of reinforcements are described below.

[0026] In one embodiment, the reinforcements may comprise transverse walls attached to the side legs of the closure flange and spaced apart from each other along a principal direction along which the closure flange extends.

[0027] According to a first variant, the transverse walls can extend between the free end of the lateral branches of the closing flange and the base of this closing flange, perpendicular to said main direction.

[0028] According to a second variant, the transverse walls may extend parallel to said main direction so that the transverse walls are fixed to the free ends of the lateral branches of the closing flange.

[0029] These variants can be combined with each other so that the panel includes both transverse walls perpendicular to the main direction and transverse walls parallel to the main direction.

[0030] The distance between two transverse walls can be between 10 mm and 30 mm, and preferably be approximately equal to 20 mm.

[0031] These spacing values are given as a non-limiting example and may in this case be suitable for a standard size exhaust nozzle panel.

[0032] In one embodiment, the reinforcements may comprise foils each comprising a base and two lateral branches connected to the base, each lateral branch of each foil comprising a free end facing the base of the closing flange, the base of each foil extending between the free ends of the closing flange.

[0033] Such foils can thus have a U-shaped section and be nested in the cavity so that the U of these foils is oriented in a direction opposite to the direction in which the U of the closing flange is oriented.

[0034] In another embodiment, the reinforcement(s) may comprise one or more portions of said honeycomb structure.

[0035] To do this, the honeycomb structure can be machined in stages.

[0036] In another embodiment, the reinforcement(s) may comprise one or more other honeycomb structures, i.e., one or more honeycomb structures different from the honeycomb structure attached to the two skins.

[0037] Of course, these different embodiments can be combined with each other so that one or more reinforcements are constituted by one or more of said side walls and / or one or more other reinforcements are constituted by one or more of said foils and / or one or more other reinforcements are constituted by one or more parts of the cellular structure and / or one or more other reinforcements are constituted by one or more of said other cellular structures.

[0038] In addition, such reinforcements may be fixed to the closing flange, for example by welding or brazing, before fixing this flange to the other elements of the panel.

[0039] Preferably, the inner and outer skins, the honeycomb structure and the closing flange may comprise metal.

[0040] The panel may have a generally flat or curved shape, or present a relatively complex geometry depending on its destination.

[0041] In one embodiment, the panel may have an axis of symmetry around which the inner and outer skins, the honeycomb structure and the closing flange extend.

[0042] In other words, the panel may have a substantially annular shape.

[0043] In a non-limiting manner, the invention also relates to a component for an aircraft propulsion unit, this component being able to be chosen from a list including an ejection cone, an ejection nozzle, a fixed internal structure and an air inlet lip, this component comprising one or more panels as described above.

[0044] The invention also relates to an aircraft comprising at least one such component.

[0045] More generally, the invention also relates to a nacelle for an aircraft propulsion unit and / or a propulsion unit for an aircraft and / or an aircraft comprising one or more panels as described above.

[0046] The invention also relates to a method of manufacturing a panel as described above, this method comprising a step of arranging the honeycomb structure and the at least one closing flange between the inner skin and the outer skin, and a step of brazing the honeycomb structure and the at least one closing flange with the inner skin and the outer skin.

[0047] In one embodiment, this method may comprise a step of arranging one or more of the reinforcements described above in the cavity of the at least one closing flange before carrying out said brazing step.

[0048] Of course, the reinforcement(s) can be placed in the cavity of a closing flange before positioning this closing flange between the internal and external skins of the panel.

[0049] Other advantages and characteristics of the invention will appear on reading the detailed, non-limiting description which follows. Brief description of the drawings

[0050] The following detailed description refers to the attached drawings in which: [ Fig. 1 ] is a schematic axial sectional view of an aircraft propulsion unit; [ Fig. 2 ] is a schematic perspective view of a nacelle for an aircraft propulsion system; [ Fig. 3 ] is a schematic perspective and axial sectional view of an ejection nozzle for an aircraft propulsion unit; [ Fig. 4 ] is a schematic perspective view of a fixed internal structure for an aircraft propulsion unit; [ Fig. 5 ] is a schematic perspective view of a central part of a panel according to the invention; [ Fig. 6 ] is a schematic perspective view in axial section of an end portion of a panel according to the invention; [ Fig. 7] is a schematic perspective and exploded view of a panel according to the invention, before fixing its constituent elements to each other, this figure showing sheets of solder interposed between said constituent elements; [ Fig. 8 ] is a schematic axial sectional view of a soldering tool receiving a panel according to the invention; [ Fig. 9 ] is a partial schematic perspective view of a closing flange and reinforcements for a panel according to a first embodiment of the invention; [ Fig. 10 ] is a partial schematic view in axial section of a panel according to the first embodiment of the invention; [ Fig. 11 ] is a partial schematic perspective view of a closing flange and reinforcements for a panel according to a second embodiment of the invention; [ Fig. 12] is a partial schematic view in axial section of a panel according to the second embodiment of the invention; [ Fig. 13 ] is a partial schematic perspective view of a closing flange and reinforcements for a panel according to a third embodiment of the invention; [ Fig. 14 ] is a partial schematic view in axial section of a panel according to the third embodiment of the invention; [ Fig. 15 ] is a partial schematic view of a reinforcement for a panel according to the second or third embodiment of the invention, this reinforcement comprising a first type of openings; [ Fig. 16 ] is a partial schematic view of a reinforcement for a panel according to the second or third embodiment of the invention, this reinforcement comprising a second type of openings; [ Fig. 17] is a partial schematic perspective view of a closing flange and reinforcements for a panel according to a fourth embodiment of the invention; [ Fig. 18 ] is a partial schematic view in axial section of a panel according to the fourth embodiment of the invention; [ Fig. 19 ] is a partial schematic view of a reinforcement for a panel according to the fourth embodiment of the invention; [ Fig. 20 ] is a partial schematic perspective view of a closing flange and reinforcements for a panel according to a fifth embodiment of the invention; [ Fig. 21 ] is a partial schematic view in axial section of a panel according to the fifth embodiment of the invention; [ Fig. 22 ] is a partial schematic perspective view of a closing flange and reinforcements for a panel according to a sixth embodiment of the invention; [ Fig. 23] is a partial schematic view in axial section of a panel according to the sixth embodiment of the invention; [ Fig. 24 ] is a partial schematic perspective view of a closing flange and reinforcements for a panel according to a seventh embodiment of the invention; [ Fig. 25 ] is a partial schematic view in axial section of a panel according to the seventh embodiment of the invention. Detailed description of embodiments

[0051] He is represented at the figure 1 an aircraft propulsion unit 1 comprising a turbomachine 2 shrouded by a nacelle 3. In this example, the turbomachine 2 is a twin-spool, twin-flow turbojet.

[0052] Subsequently, the terms “upstream”, “downstream”, “front” and “rear” are defined in relation to a direction D1 of flow of gases through the propulsion unit 1 when the latter is propelled.

[0053] The turbojet engine 2 has a longitudinal central axis A1 around which its various components extend, in this case, from upstream to downstream of the turbojet engine 2, a fan 4, a low-pressure compressor 5, a high-pressure compressor 6, a combustion chamber 7, a high-pressure turbine 8 and a low-pressure turbine 9. The compressors 5 and 6, the combustion chamber 7 and the turbines 8 and 9 form a gas generator.

[0054] Conventionally, during operation of such a turbojet 2, an air flow 10 enters the propulsion unit 1 through an air inlet 11 upstream of the nacelle 3, passes through the fan 4 and then divides into a central primary flow 10A and a secondary flow 10B. The primary flow 10A flows in a primary gas circulation vein 12A passing through the gas generator. The secondary flow 10B flows in a secondary vein 11B surrounding the gas generator and delimited radially outwards by the nacelle 3.

[0055] There figure 2 shows separately and in more detail the nacelle 3 of such a propulsion unit 1.

[0056] In particular, this figure shows an internal fixed structure 13 (“inner fixed structure” in English) surrounding the turbojet 2 and delimiting radially inwards a downstream part of the secondary vein 11B, as well as an external fixed structure 14 (“outer fixed structure” in English) delimiting this downstream part of the secondary vein 11B radially outwards.

[0057] There figure 2 also shows an ejection nozzle 15 comprising an ejection cone 16 (“plug” in English) and an ejection nozzle 17 (“nozzle” in English). The ejection nozzle 15 is configured to discharge downstream of the propulsion unit 1 the primary flow 10A exiting from the primary vein 12A of the turbojet 2 so as to generate thrust.

[0058] THE Figures 3 and 4 respectively show the ejection nozzle 15 and the fixed internal structure 13 of this nacelle 3.

[0059] In reference to the figure 3, the ejection cone 16 and the ejection nozzle 17 are both parts of revolution of axis A1, the ejection cone 16 being substantially conical, the ejection nozzle 17 being substantially annular.

[0060] In reference to the figure 4 , the fixed internal structure 13 is made of two parts 13A and 13B symmetrical with respect to each other with respect to a vertical median longitudinal plane passing through the longitudinal central axis A1. The part 13A comprises a semi-annular central part 18A forming a half-barrel as well as two end parts 19A and 20A forming islands respectively positioned vertically below and above the central part 18A and making it possible to connect the part 13A of the fixed internal structure 13 to the fixed external structure 14. The part 13B comprises a central part 18B and islands 19B and 20B similar to those of the part 13A.

[0061] The fixed internal structure 13 and the ejection nozzle 15, as well as other components of the nacelle 3 generally comprise structural and / or acoustic panels which may constitute all or part of such components.

[0062] The invention relates more specifically to the structure and manufacture of such a panel.

[0063] There Figure 5 shows a central part of a panel 30 according to the invention. This panel 30 comprises an inner skin 31, an outer skin 32 and a honeycomb structure 33 sandwiched between the inner 31 and outer 32 skins.

[0064] By way of non-limiting example, the inner 31 and outer 32 skins each have a thickness of between 0.2 mm and 2 mm, for example 0.6 mm, and the honeycomb structure 33 has a thickness of around ten mm.

[0065] The inner skin 31, the outer skin 32 and the honeycomb structure 33 comprise in this example metal such as titanium.

[0066] In this example, the honeycomb structure 33 comprises transverse partitions 34 delimiting between them hexagonal cells 35 forming a honeycomb structure. The cells 35 can have any other shape to avoid telegraphing.

[0067] In this example, the inner skin 31 is a solid, structuring skin, while the outer skin 32 comprises orifices 36 intended to guide air into the cells 35 in order to absorb acoustic energy.

[0068] Panel 30 of the Figure 5 is therefore an acoustic panel.

[0069] There figure 6 shows a peripheral part of a panel 30 according to the invention, the central part of which may correspond to that described above with reference to the Figure 5 . On this figure 6, no transverse partitions 34 of the alveolar structure 33 are shown, but the latter can of course be of the same type as the alveolar structure 33 of the Figure 5 .

[0070] It is more specifically represented on the figure 6 a portion of a closing flange 40 (“close-out” in English) arranged in accordance with the invention.

[0071] This closing flange 40 comprises a base 41 and two lateral branches 42 and 43 giving the flange 40 a U-shaped section.

[0072] More precisely, the base 41 extends on the one hand in the direction of the thickness of the panel 30, in a transverse direction T, and in a so-called main direction P which can be a circumferential direction of the panel 30 when the latter is of the annular type.

[0073] The lateral branches 42 and 43 each extend in a so-called longitudinal direction L which can be parallel to the longitudinal central axis A1 when the panel 30 is mounted in the propulsion assembly 1. The lateral branches 42 and 43 are each connected to the base 41 so as to form a substantially right angle with the base 41.

[0074] The base 41, the side branch 42 and the side branch 43 have in this example a thickness of approximately 1.5 mm.

[0075] The side branches 42 and 43 each comprise a free end 44 and 45, respectively.

[0076] The lateral branches 42 and 43 and the base 41 delimit a cavity 46 which extends along the main direction P and which, in the longitudinal direction L, opens at the free ends 44 and 45. The opening of this cavity 46 is considered relative to the closing flange 40 independently of the other elements of the panel 30.

[0077] The closing flange 40 has a total thickness which corresponds to the thickness of the honeycomb structure 33 and is positioned against the latter so that the free end 44 / 45 of each of the lateral branches 42 / 43 faces the honeycomb structure 33.

[0078] Thus, the cavity 46 is in this example delimited transversely by the lateral branches 42 and 43, and longitudinally on the one hand by the base 41 and on the other hand by the alveolar structure 33.

[0079] The closing flange 40 is further sandwiched between the skins 31 and 32 and fixed to them in the same manner as the honeycomb structure 33 (see further below). The side branch 42 is fixed to the inner skin 31 and the side branch 43 is fixed to the outer skin 32.

[0080] In this way, the closing flange 40 is arranged at one end of the panel 30 so that the base 41 of this flange 40 defines a full lateral periphery of the panel 30, which facilitates the assembly of the panel 30 with another structure (not shown), for example by welding the base 41 to a part of this other structure.

[0081] Of course, a space can be left between the free ends 44 and 45 of the closing flange 40 and the cellular structure 33, in the longitudinal direction L. In addition, a part of the cellular structure 33 can penetrate into the cavity 46 so as to close all or part of this cavity 46 (see further below).

[0082] The panel 30 may have any shape, for example a substantially conical or annular shape, or a sector of substantially conical or annular shape, to form respectively all or part of the ejection cone 16 or the ejection nozzle 17 of the figure 3, or a half-barrel shape to form all or part of the central part 18A or 18B of the part 13A or 13B respectively of the fixed internal structure 13 of the figure 4 , or a substantially flat shape or having any other geometry making it possible to constitute a part of the propulsion unit 1.

[0083] In addition, the panel 30 may comprise several closing flanges 40. For example, when the panel 30 is annular, two annular closing flanges 40 may be used to close the two annular ends of the panel 30. When the panel 30 is not a piece of revolution, four closing flanges 40 may be used to close the four lateral ends of the panel 30.

[0084] According to the invention, the manufacture of the panel 30 comprises brazing of its constituent elements, preferably by interposition of brazing sheets between the respective elements to be assembled.

[0085] With reference to the illustration of the figure 7 in which the panel 30 comprises two closing flanges 40 to be assembled, a sheet of solder 50A is arranged between the internal skin 31 on the one hand, and on the other hand an internal face of the cellular structure 33 and one of the lateral branches of each of the closing flanges 40. Another sheet of solder 50B is arranged between the external skin 32 on the one hand, and on the other hand an external face of the cellular structure 33 and the other lateral branch of each of the closing flanges 40.

[0086] Conventionally, the elements of the panel 30 to be assembled are in this example placed in a brazing tool 60 allowing them to be placed under gas pressure within a vacuum furnace (not shown).

[0087] There figure 8 shows a conventional brazing tool 60 adapted for brazing a panel 30 intended to form a part of the ejector cone 16 of the figure 3 The implementation of such tooling is for example described in document WO 2014 / 020286 A1.

[0088] In particular, to prevent the sagging of one of the lateral branches 42 and 43 of the closing flange 40 during brazing, reinforcements are placed in the cavity 46 of this closing flange 40 so as to keep these lateral branches 42 and 43 spaced apart from each other.

[0089] Different types of reinforcements 70 are described below with reference to the figures 9 to 23 .

[0090] THE figures 9 and 10 show a first embodiment, in which reinforcements 70 are produced in the form of transverse walls each extending perpendicularly to the main direction P.

[0091] In this example, each transverse wall 70 has a thickness X1 of approximately 1 mm and two consecutive transverse walls 70 are spaced from each other, in the main direction P, by a distance P1 of approximately 20 mm.

[0092] Each transverse wall 70 also has a dimension X2 along the transverse direction T so as to extend in the cavity 46 substantially from the lateral branch 42 to the lateral branch 43 of the closing flange 40.

[0093] Each transverse wall 70 also has a dimension X3 in the longitudinal direction L so as to extend in the cavity 46 substantially from the base 41 to the free ends 44 and 45 of the lateral branches 42 and 43 of the closing flange 40.

[0094] In this example, each transverse wall 70 is flush with the surface of the lateral branches 42 and 43 opposite the alveolar structure 33.

[0095] THE figures 11 and 12show a second embodiment, in which reinforcements 70 are produced in the form of transverse walls each extending parallel to the main direction P.

[0096] In this second embodiment, each of the transverse walls 70 has a thickness X1 of approximately 1 mm, and a dimension X2 in the transverse direction T so as to extend in the cavity 46 substantially from the lateral branch 42 to the lateral branch 43 of the closing flange 40.

[0097] More precisely, each of the transverse walls 70 is in this example positioned at the level of the free ends 44 and 45 of the lateral branches 42 and 43 so as to be flush with the surface of these lateral branches 42 and 43 opposite the alveolar structure 33.

[0098] Each of the transverse walls 70 has in this example a dimension X3 along the main direction P of approximately 60 mm, and two consecutive transverse walls 70 along this direction are spaced from each other by a distance P1 of approximately 20 mm.

[0099] It is represented on the figures 24 and 25 a variant of this second embodiment making it easier to put the transverse walls 70 in place.

[0100] This variant differs from the second embodiment in that each of the lateral branches 42 and 43 comprises, at its free end 44 or 45, a groove extending in the main direction P. It also differs in that the dimension X2 of the transverse walls 70 is greater than the distance between the lateral branches 42 and 43 in the transverse direction T, so that the ends of the transverse walls 70, in the transverse direction T, are housed in the grooves of the lateral branches 42 and 43. Thus, each of the transverse ends of the transverse walls 70 is located longitudinally between the lateral branch 42 or 43 and the cellular structure 33 while extending in the transverse direction T between shoulders of the lateral branches 42 and 43, these shoulders being formed by the grooves described above.

[0101] THE figures 13 and 14show a third embodiment which differs from the second embodiment in that the free ends 44 and 45 of the lateral branches 42 and 43 comprise a chamfer, and the ends of the transverse walls 70, in the transverse direction T, each comprise a corresponding chamfer making it possible to wedge these walls 70 onto the lateral branches 42 and 43.

[0102] In each of the embodiments described above, it is preferable to fix the transverse walls 70 to the closing flange 40, for example by welding or brazing, taking into account in particular the thickness X1 of these walls 70, in order to keep them in position at least during the brazing of the panel 30.

[0103] Many variations can be made to these embodiments without departing from the scope of the invention. For example, the transverse walls 70 of the second and third embodiments, and of the embodiment of the figures 24 and 25, may include openings 71 as illustrated in the figures 15 and 16 , in order to reduce the mass of the panel 30.

[0104] THE figures 17 and 18 show a fourth embodiment, in which reinforcements 70 are foils having a U-shaped section.

[0105] Such a 70 foil is represented in isolation at the figure 19 This foil 70 comprises a base 80 and two lateral branches 81 and 82 connected to the base 80. Each of the lateral branches 81 and 82 comprises a free end 83 and 84 respectively.

[0106] In reference to the figures 17 and 18 , for each foil 70, the free end 83 and 84 of each of the lateral branches 81 and 82 is opposite the base 41 of the closing flange 40, and the base 80 of the foil 70 extends between the free ends 44 and 45 of the closing flange 40 so as to be flush with the surface of these free ends 44 and 45 opposite the cellular structure 33.

[0107] In this fourth embodiment, each part of the foils 70, that is to say the base 80 and the lateral branches 81 and 82, has a thickness X1 of approximately 1 mm.

[0108] The base 80 has a dimension X2 along the transverse direction T so as to extend in the cavity 46 substantially from the lateral branch 42 to the lateral branch 43 of the closing flange 40.

[0109] The lateral branches 81 and 82 each have a dimension X3 in the longitudinal direction L so as to each extend in the cavity 46 substantially from the free ends 44 and 45 of the closing flange 40 to the base 41 of this flange 40.

[0110] Each of the foils 70 has in this example a dimension X4 along the main direction P of approximately 8 mm, and two consecutive foils 70 along this direction are spaced from each other by a distance P1 of approximately 20 mm.

[0111] The shims 70 may be sized to be held in position by spring action due to their U-shape. Alternatively, the shims 70 may be secured to the closing flange 40, for example by welding or brazing.

[0112] THE figures 20 and 21 show a fifth embodiment, in which the reinforcements 70 are formed by blocks extending in the main direction and having a section substantially similar to that of the cavity 46, that is to say a substantially square section.

[0113] In this example, the blocks 70 are made of a honeycomb structure, for example of the same type as the honeycomb structure 33. The cells (not shown) of the blocks 70 reduce their mass.

[0114] In this fifth embodiment, each of the blocks 70 has a dimension X1 in the longitudinal direction L so as to extend in the cavity 46 substantially from the base 41 to the free ends 44 and 45 of the closing flange 40, that is to say in this case to the cellular structure 33 so as to be flush with the surface of the lateral branches 42 and 43 opposite the cellular structure 33.

[0115] Each of these blocks 70 has a dimension X2 in the transverse direction T so as to extend in the cavity 46 substantially from the lateral branch 42 to the lateral branch 43 of the closing flange 40.

[0116] Finally, each of these blocks 70 has in this example a dimension X3 along the main direction P of approximately 60 mm, two consecutive blocks 70 along this direction being spaced from each other by a distance P1 of approximately 20 mm.

[0117] For each of the embodiments described above, each reinforcement 70 may be permanently attached to the closure flange 40, for example by welding or brazing. In this case, the reinforcements 70 contribute to some extent to the structural strength of the panel 30 during its use.

[0118] Alternatively, when the closing flange 40 is not annular and the cavity 46 remains, after manufacture of the panel 30, open at least at one end in the main direction P, the reinforcements 70 can be removed in order to reduce the mass of the panel 30.

[0119] The technique for removing the reinforcements 70 after manufacturing the panel 30 depends on the type of connection between the reinforcements 70 and the closing flange 40.

[0120] When the reinforcements 70 have been welded or brazed, their removal requires breaking the corresponding weld points. Of course, their attachment to the closing flange 40 must in this case be sufficiently weak to allow such a break while being strong enough to remain in position during the brazing of the constituent elements of the panel 30.

[0121] Maintaining the reinforcements 70 in the cavity 46 during the manufacture of the panel 30 can also be achieved by thermal expansion of the reinforcements 70, which are then sized and adjusted with a tightening allowing, at room temperature, to insert them into the cavity 46 and to remove them after manufacture of the panel 30.

[0122] Whatever the type of connection between reinforcements 70 and closing flange 40, it is preferable to use an anti-diffuser to prevent the reinforcements 70 from welding to the closing flange 40 and to the honeycomb structure 33, by diffusion, during the brazing of the panel 30.

[0123] Generally, the reinforcements 70 preferably comprise a material having a coefficient of expansion identical to or greater than that of the closure flange 40, so that the reinforcements 70 keep the lateral branches 42 and 43 of the closure flange 40 spaced apart from each other during brazing of the panel 30.

[0124] THE figures 22 and 23 show a sixth embodiment, in which a reinforcement 70 is constituted by a respective part of the honeycomb structure 33 of the panel 30.

[0125] To do this, the cellular structure 33 is previously machined so that the end E1 of this structure 33 housed in the cavity 46 of the closing flange 40 to form the reinforcement 70 has a dimension X1 in the transverse direction T less than the thickness X2 of the cellular structure 33 on the central part E2 of this structure 33.

[0126] The embodiments described above are given as non-limiting examples. In particular, to the extent that the reinforcements 70 perform their function of maintaining the spacing of the lateral branches 42 and 43 during the brazing of the panel 30, the shape, dimensions or even the number of these reinforcements 70 and their positioning relative to each other may be different from what has just been described, without departing from the scope of the invention.

Claims

1. Structural and / or acoustic panel (30) for an aircraft propulsion assembly (1), this panel (30) comprising an inner skin (31), an outer skin (32), an alveolar structure (33) and at least one sealing flange (40), this sealing flange (40) comprising a base (41) and two lateral arms (42, 43) connected to the base (41), each lateral arm (42, 43) comprising a free end (44, 45), the lateral arms (42, 43) and the base (41) delimiting a cavity (46) which opens at the free end (44, 45) of the lateral arms (42, 43), the alveolar structure (33) and the sealing flange (40) being enclosed between the inner skin (31) and the outer skin (32), one of the lateral arms (42) of the sealing flange (40) being fastened to the inner skin (31), the other lateral arm (43) of the sealing flange (40) being fastened to the outer skin (32), the free end (44, 45) of each of the lateral arms (42, 43) facing the alveolar structure (33) such that the alveolar structure (33) delimits said cavity (46), said panel (30) being characterised in that it comprises one or more reinforcements (70) extending into said cavity (46) so as to keep the lateral arms (42, 43) of the sealing flange (40) separated from one another, the reinforcements (70) comprise transverse walls fastened to the lateral arms (42, 43) of the sealing flange (40) and spaced apart along a principal direction (P) along which the sealing flange (40) extends.

2. Panel (30) according to claim 1, wherein said transverse walls (70) extending preferably: - between the free end (44, 45) of the lateral arms (42, 43) of the sealing flange (40) and the base (41) of this sealing flange (40), perpendicularly to said principal direction (P), or - parallel with said principal direction (P) such that the transverse walls (70) are fastened to the free ends (44, 45) of the lateral arms (42, 43) of the sealing flange (40).

3. Panel (30) according to claim 1 or 2, wherein the reinforcements (70) comprise foils each comprising a base (80) and two lateral arms (81, 82) connected to the base (80), each lateral arm (81, 82) of each foil (70) comprising a free end (83, 84) facing the base (41) of the sealing flange (40), the base (80) of each foil (70) extending between the free ends (44, 45) of the sealing flange (40).

4. Panel (30) according to any one of claims 1 to 3, wherein the reinforcement(s) (70) comprise one or more portions (E1) of said alveolar structure (33) and / or one or more other alveolar structures.

5. Panel (30) according to any one of claims 1 to 4, wherein the inner (31) and outer (32) skins, the alveolar structure (33) and the sealing flange (40) comprise metal.

6. Panel (30) according to any one of claims 1 to 5, this panel (30) having an axis of symmetry (A1) about which the inner (31) and outer (32) skins, the alveolar structure (33) and the sealing flange (40) extend.

7. Aircraft propulsion assembly (1) component, this component being selected from a list including an ejection plug (16), an ejection nozzle (17), an inner fixed structure (13) and an air inlet lip (11), this component comprising one or more panels (30) according to any one of claims 1 to 6.

8. Aircraft, comprising at least one component according to claim 7.

9. Method for manufacturing a panel (30) according to any one of claims 1 to 6, this method comprising a step of disposing the alveolar structure (33) and the at least one sealing flange (40) between the inner skin (31) and the outer skin (32), and a step of brazing the alveolar structure (33) and the at least one sealing flange (40) with the inner skin (31) and the outer skin (32).

10. Method according to claim 9, for manufacturing a panel (30) according to any one of claims 1 to 4, this method comprising a step of disposing one or more of said reinforcements (70) in the cavity (46) of the at least one sealing flange (40) before implementing said brazing step.