• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to a method for manufacturing a composite sandwich panel, the panel comprising a core formed of a plurality of cells that extend vertically between a first skin and a second skin, characterized in that it comprises at least one at least one step of producing at least a first band (20a) and a second band (20b) in a fugitive material, each band (20a, 20b) having at least one impression (30a, 30b, 31b) which comprises a succession of aligned half-cells, a step of upholstering which consists in lining at least one fibrous fold (34) on the impression (30a) of the first band (20a), a step of assembling the first band (20a) and the second band (20b) which consists in interleaving the imprint (30a) of the first band (20a) with the imprint (31b) of the second band (20b), by trapping said fibrous fold (34), and a trimming step of removing excess fungible material from all the strips (2 0a, 20b) formed during the preceding assembly step, so as to form a new cavity (38) which forms a succession of aligned half-cells (40).
    公开号:FR3044963A1
    申请号:FR1562300
    申请日:2015-12-14
    公开日:2017-06-16
    发明作者:Pierre Caruel；Herve Hurlin
    申请人:Aircelle SA；
    IPC主号:
    专利说明:

    The present invention relates to a method for manufacturing a composite matrix sandwich panel, in particular for an aircraft propulsion assembly.
    As is known per se, an aircraft propulsion unit conventionally comprises a turbojet engine housed inside a nacelle.
    The nacelle generally has a tubular structure comprising an air inlet upstream of the turbojet, a median section intended to surround a fan of the turbojet engine and its casing, and a downstream section intended to surround the combustion chamber of the turbojet engine and sheltering where appropriate. thrust reverser means. It can be terminated by an ejection nozzle whose output is located downstream of the turbojet engine.
    In order to reduce mass and increase performance, especially mechanical performance, the aeronautics industry is increasingly using composite materials to produce structural parts and / or panels.
    Such elements are generally made from sets of fibrous folds embedded in a matrix of organic or ceramic materials. These elements can be monolithic or contain a soul made of another material.
    This is for example the acoustic panels used to reduce noise emissions generated by the turbojet and which is common to equip several parts of the nacelle and / or turbojet.
    These parts of the nacelle and / or the turbojet engine may be, for example, an internal surface of the air circulation duct for a double-flow turbojet engine, a nozzle or ejection cone surface, an internal air intake panel air, etc.
    Such acoustic panels are in the form of a sandwich-type panel comprising at least one honeycomb core, for example a honeycomb-type structure, covered with a perforated or acoustically porous outer skin (called acoustic skin) and a full inner skin (called skin support or structuring). This set forms a resonator.
    The skins of these panels may be made of composite material comprising a set of fibers (superposition of fibrous folds), embedded in a matrix of organic or ceramic origin.
    More particularly, ceramic matrix composites, also known by the acronym "CMC", are composite materials that are formed by a set of often ceramic fibers associated with a also ceramic matrix.
    Ceramic composites have interesting properties, including tensile strength, bending and breaking properties, heat resistance properties, and anti-corrosion properties.
    Also, ceramic composites are lightweight materials whose density is close to that of aluminum.
    Among the ceramic composites, ceramic composites based on oxides (aluminas, aluminosilicates in particular) are known which form good electrical insulators and which also have an interesting heat insulating capacity, and good resistance to heat and humidity. mechanical fatigue.
    Although ceramic composites have many advantages, their manufacture is delicate.
    The present invention aims in particular to provide a manufacturing method that allows for a sandwich composite composite honeycomb core type oxide.
    For this purpose, the invention proposes a method of manufacturing a composite sandwich panel, the panel comprising a core formed of a plurality of cells that extend vertically between a first skin and a second skin, characterized in that it comprises at least: - a step of producing at least a first band and a second band in a fugitive material, each band having at least one footprint which comprises a succession of aligned half-cells, - a step of upholstering which consists in lining at least one fibrous fold on the impression of the first band, - a step of assembling the first band and the second band which consists in interleaving the impression of the first band with the imprint of the first band. second band, by trapping said fibrous fold, - a trimming step which consists in removing the surplus of fungible material from the set of strips formed during the assembly step p recede, so as to form a new imprint which forms a succession of aligned half-cells, said steps of realization, upholstering, assembly and trimming being repeated to form the cellular core of the panel.
    The method according to the invention makes it possible to produce a ceramic composite panel comprising a lightweight cellular core.
    In addition, the use of strips of fungible material allows to "mold" the cells of the core of the panel with great freedom of shape, the cells may be in particular of hexagonal or circular section.
    According to another characteristic, the method comprises a closing step which consists in draping the first skin and the second skin on a first face and a second face respectively of said cellular core formed during the previous steps of production, upholstering, assembly and trimming. At the end of the closing step, the first skin and the second skin are sintered on the alveolar core to assemble the cellular core and said skins to form the composite panel.
    This characteristic makes it possible to mechanically connect the alveolar core to the skins, which allows a passage of shearing forces between the skins and the core, in the plane of the skins of the panel.
    In addition, each fibrous fold is made of ceramic fibers and is impregnated with a slip containing metal or mineral oxide powder for producing a ceramic matrix during the sintering operation.
    Also, the method comprises a drying step which is carried out following each lining step and which consists in drying the resin impregnated fibrous fold and upholstered on the associated band during the upholstering step.
    According to another characteristic, during the upholstering step, the fibrous fold is lined on the imprint of the associated band by means of a toothed wheel which has a shape complementary to the shape of said imprint.
    The toothed wheel, meshing with the impression, makes it possible to line the impression with the fibrous fold so that the fibrous fold takes the form of the impression.
    The shape of the wheel is complementary to that of the impression, which allows a wide variety of cell shapes.
    According to an exemplary embodiment, the toothed wheel is a heating wheel which is designed to heat and dry the fibrous fold simultaneously.
    According to an exemplary embodiment, each cell of the core of the panel has a hexagonal shape.
    According to another exemplary embodiment, each cell of the core of the panel has a generally circular shape.
    The circular cells make it possible to reduce the stresses in the fibrous folds during draping and make it possible to improve the flexibility to accommodate deformations during the placement of the cellular core on the skins.
    According to an exemplary embodiment, each strip of fugitive material has a curved shape, so that the panel produced has a curved shape.
    Also, each fibrous fold which forms the cells of the panel has at least a first protrusion and a second protrusion which are folded on each side of the associated band and which are adapted to cooperate with the first skin and the second skin respectively.
    These protuberances allow a mechanical connection with the skins to promote the passage of efforts between the core and the skins of the panel.
    Finally, the panel obtained by the method is adapted to equip a nacelle with a propulsion unit of aircraft. Other features and advantages of the invention will appear on reading the detailed description which follows for the understanding of which reference will be made to the appended drawings in which: FIG. 1 is a perspective view which illustrates a first strip of material fungible having a first imprint; FIG. 2 is a perspective view which illustrates the first band of FIG. 1 during a step of lining a fibrous fold on the imprint; FIG. 3 is a perspective view illustrating the step of drying the fibrous fold; FIG. 4 is a perspective view illustrating a second strip of fungible material having two indentations; FIG. 5 is a perspective view which illustrates the assembly step of the first band and the second band; Fig. 6 is a perspective view illustrating the step of trimming excess fungible material to form a new imprint; Figure 7 is a perspective view illustrating the cellular core of the composite panel; FIG. 8 is a perspective view which illustrates the composite panel manufactured according to the manufacturing method according to the invention; FIG. 9 is a perspective view which illustrates the second strip following a lining step according to the second embodiment of the invention; FIG. 10 is a perspective view which illustrates a step of assembling the second band and the third band according to the second embodiment of the invention; FIG. 11 is a perspective view which illustrates a trimming step according to the second embodiment of the invention; - Figures 12 and 13 are perspective views which illustrate a step of assembling the assembly formed by the second band and the third band on the first band, according to the second embodiment of the invention; FIG. 14 is a perspective view which illustrates a trimming step according to the second embodiment of the invention; FIG. 15 is a perspective view which illustrates the protuberances formed by the fibrous fold to form a fastening flap on the skins of the panel; FIG. 16 is a perspective view illustrating cutouts made on the protuberances of FIG. 15 formed by the fibrous fold; - Figure 17 is a perspective view which illustrates an alternative embodiment of the cellular core having substantially cylindrical cells.
    In the description and the claims, the terms "upper" and "lower" will be used in a nonlimiting manner with reference to the upper part and the lower part respectively of the panel illustrated in FIG. 8.
    Moreover, to clarify the description and the claims, the longitudinal, vertical and transverse terminology will be adopted in a nonlimiting manner with reference to the L, V, T trihedron indicated in the figures.
    FIGS. 1 to 6 show the steps of a method of manufacturing a ceramic composite panel 10, illustrated in FIG. 8, according to a first embodiment of the method.
    As can be seen in FIG. 1, the panel 10 comprises a core 12 formed of a plurality of cells 14 which are arranged in a honeycomb and which extend vertically between a first skin 16 and a second skin 18 below.
    The method according to the invention comprises a step of producing a plurality of bands, including at least a first band 20a shown in FIG. 1, and a second band 20b represented in FIG.
    To facilitate the understanding of the description, the elements relating to the first band 20a are indicated by a numerical reference followed by the letter "a" and the elements relating to the second band 20b are indicated by a numerical reference followed by the letter "b".
    Each strip 20a, 20b extends in length in a longitudinal direction, each strip 20a, 20b being delimited transversely between a front face 22a, 22b respectively and a rear face 24a, 24b respectively, and each strip 20a, 20b being delimited vertically between an upper face 26a, 26b respectively and a lower face 28a, 28b respectively.
    The front face 22a of the first band 20a forms a first cavity 30a before.
    Also, the front face 22b of the second band 20b forms a first front footprint 30b and the rear face 24b of the second band 20b forms a second rear footprint 31b which is offset transversely with respect to the first footprint 30b by a distance equivalent to the width of a cell 14.
    It will be noted that the first band 20a has only a footprint 30a because the first band 20a is intended to form a peripheral edge of the cellular core 12.
    Each imprint 30a, 30b, 31b has a crenellated shape which comprises a succession of half-cells 32 aligned longitudinally.
    Each strip 20a, 20b is made of a fugitive material, that is to say a material which is adapted to disappear during the high temperature sintering phase. According to a preferred embodiment, the fugitive material is a fungible material that is designed to liquefy from a certain temperature.
    The strips 20a, 20b are for example shaped by mechanical cutting, or by means of a laser or by means of a hot wire. Following the step of producing the strips 20a, 20b, the method comprises a lining step, illustrated in Figure 2, which consists in lining a fibrous fold 34, or several fibrous folds 34, on the first footprint 30a of the first band 20a.
    The fibrous ply 34 is composed of a fabric of ceramic fibers, such as alumina or alumino-silicate fibers, for example.
    As can be seen in Figure 2, the fibrous ply 34 is lined on the first cavity 30a of the first band 20a by means of a toothed wheel 36 which has a shape complementary to the shape of the first cavity 30a. For this purpose, the toothed wheel 36 is rotated about a vertical axis A, so that the toothed wheel 36 rolls and meshes the first cavity 30a to press the fibrous ply 34.
    According to a preferred embodiment, the fibrous ply 34 is prepreg of slip. The slip may consist of a liquid and alumina powder or alumino-silicate suspension, slip required for the realization of the matrix.
    However, in a nonlimiting manner, the fibrous ply 34 may be lined dry, in which case the fibrous ply 34 will be impregnated with slip following the upholstering step, for example by spraying or dipping. The upholstering step is followed by a drying step, illustrated in Figure 3, which consists of drying the fibrous fold 34 upholstered during the upholstering step. Following the drying step, the method comprises a step of assembling the first band 20a and the second band 20b, which consists in interleaving the first imprint 30a of the first band 20a with the second imprint 31b of the second band 20b, trapping the fibrous fold 34 between the first band 20a and the second band 20b, as can be seen in Figure 5. The assembly step is followed by a trimming step of removing the surplus fungible material of the assembly formed by the strips 20a, 20b during the preceding assembly step, so as to form a third cavity 38 which forms a succession of half-cells 40.
    As can be seen in Figure 6, at the end of the trimming step, the second band 20a is reduced to a series of hexagonal columns 41, each column 41 forming a mold adapted to be lined with a fibrous fold to to form a cell 14.
    Note also that the columns 41 are spaced longitudinally in a regular manner to allow entanglement of the cells 14 staggered, in the manner of a honeycomb. The trimming step is for example carried out by means of a laser or by means of a hot wire.
    Also, the trimming step is followed by a new lining step (not shown) which consists in lining a fibrous fold on the third imprint 38 formed during the preceding trimming step to form a series of cells 14, then a third band is assembled on the third footprint 38, this third band being in turn cropped during a new trimming step.
    The steps of upholstering, assembly and trimming are repeated to form the cells 14 of the cellular core 12 of the panel 10, as many times as necessary to obtain the desired panel size. The alveolar core 12 thus obtained is illustrated in FIG. 7.
    Also, the method comprises a closing step which consists in draping the first skin 16 on a tool, positioning a first face 42 of the alveolar core 12 on the first skin 16, and draping the second skin 18 on a second face 44 of the alveolar core 12. The two skins 16, 18 are both made from the same ceramic fibers impregnated with the same slip as for the realization of the honeycomb structure. At the end of the closing step, the first skin 16 and the second skin 18 are sintered on the cellular core 12 to assemble the cellular core 12 and the two skins 16, 18, to form the composite panel 10 .
    The previously described manufacturing method can be performed substantially differently according to a second embodiment described below.
    According to the second embodiment, the method comprises the same steps of production, upholstering and drying as in the first embodiment, these steps leading to the first strip 20a lined by the fibrous ply 34, as illustrated in FIG. FIG. 3. Unlike the first embodiment, the method according to the second embodiment comprises a new upholstering step which consists in lining a fibrous fold 34, or several fibrous folds, on the first impression 30b before the second strip 20b, as can be seen in Figure 9. The upholstering step is followed by a drying step fibrous fold 34 carpeted previously.
    With reference to FIG. 10, following the drying step, the method according to the second embodiment comprises a first step of assembling the second band 20b and a third band 20c, which consists of interleaving the first footprint 30b before the second band 20b with a second footprint 31c back of the third band 20c, trapping the fibrous fold 34 between the second band 20b and the third band 20c.
    Still according to the second embodiment, the first assembly step is followed by a trimming step illustrated in FIG. 11, which consists in removing the surplus of fungible material from the assembly formed by the strips 20b, 20c at during the first preceding assembly step, so as to form a new cavity 46 which forms a succession of half-cells 48.
    According to FIG. 12, the second assembly step is followed by a second assembly step which consists in rotating the assembly formed by the second strip 20a and the third strip 20c, and assembling this assembly on the first strip 20a. , as can be seen in Figure 13.
    The third assembly step is followed by a new trimming step illustrated in FIG. 14, which consists in removing the excess fungible material from the second strip 20b so as to form a new imprint 50.
    Thus, the fibrous folds 34 covered on the first band 20a and on the second band 20b are arranged facing each other and staggered to form a row of cells 14.
    According to another aspect of the invention, each cell 14 described later has a hexagonal shape.
    However, the manufacturing method according to the invention makes it possible to produce other cell shapes, for example cavities 52 of substantially circular cross-section, as can be seen in FIG. 17. For this purpose, the shape of the indentations formed the strips of fungible material are adapted as well as the complementary shape of the toothed wheel 36.
    The cavities 52 of substantially circular shape make it possible to reduce the stresses in the fibrous folds during draping and make it possible to improve the flexibility to accommodate deformations during the placement of the cellular core 12 on the first skin 16.
    In addition, the method according to the invention makes it possible to produce a shaped panel 10, that is to say a panel 10 which has a concave or curved shape for example.
    For this purpose, the upper face and the lower face of each strip 20a, 20b, 20c of fugitive material has a curved shape, so that the panel 10 also has a curved shape.
    According to an embodiment not shown, the toothed wheel 36 is a heating wheel which is designed to heat the fibrous plies 34 prepreg slip and simultaneously dry the fibrous folds 34, during the draping step.
    According to an exemplary embodiment of the invention shown in FIG. 15, the fibrous fold 34 which is lined on a strip 20a of fungible material has a first upper protrusion 54 and a second lower protrusion 56 which are folded on the upper face and on the lower face respectively of the associated band.
    The protrusions 54, 56 folded can significantly increase the excess length of the fibrous ply 34 relative to the fungible material forming the strips, to allow an increase in the contact area between the fibers of the fibrous ply 34 and the fibers of the skins 16, 18 of the panel 10.
    Preferably, the protuberances 54, 56 are not dried during the drying phase so that they retain the ability to deform.
    Also, according to a complementary exemplary embodiment shown in FIG. 16, the protuberances 54, 56 are cut along a cutting line at each nodal joint in order to allow deformation of the fibrous fold 34 in order to fold the protuberances 54, 56 onto the fungible material forming the associated band.
    The method according to the invention is particularly suitable for manufacturing an oxide-based ceramic matrix composite sandwich panel, the panel obtained by the method being adapted to equip a nacelle with an aircraft propulsion unit, in particular to form a panel. acoustically resistant to heat.
    The present description of the invention is given by way of non-limiting example.
    It will be understood that the steps of the method can be performed in a substantially different order than in the order described in the present description.
    权利要求:
    Claims (12)
    [1" id="c-fr-0001]
    A method of manufacturing a composite sandwich panel (10), the panel (10) having a core (12) formed of a plurality of cells (14) which extend vertically between a first skin (16) and a second skin (18), characterized in that it comprises at least: - a step of producing at least a first strip (20a) and a second strip (20b) in a fugitive material, each strip (20a, 20b ) having at least one impression (30a, 30b, 31b) which comprises a succession of aligned half-cells (32), - a lining step which consists in lining at least one fibrous fold (34) on the impression (30a) of the first band (20a), - a step of assembling the first band (20a) and the second band (20b) which consists in interleaving the impression (30a) of the first band (20a) with the impression (31b) of the second web (20b), trapping said fibrous ply (34), and - a trimming step of removing the excess web fongible web of the set of strips (20a, 20b) formed during the preceding assembly step, so as to form a new footprint (38) which forms a succession of aligned half-cells (40), said steps embodiment, upholstering, assembly and trimming being repeated to form the core (12) of the panel (10) alveolar.
    [2" id="c-fr-0002]
    2. A method of manufacturing a panel (10) composite sandwich according to claim 1, characterized in that it comprises a closing step which consists of draping the first skin (16) and the second skin (18) on a first face (42) and a second face (44) respectively of said core (12) formed alveolar during the previous steps of production, upholstering, assembly and trimming.
    [3" id="c-fr-0003]
    3. A method of manufacturing a composite sandwich panel (10) according to claim 2, characterized in that at the end of the closing step, the first skin (16) and the second skin (18) are sintered. on the cellular core (12) for assembling the cellular core (12) and said skins (16, 18) to form the composite panel (10).
    [4" id="c-fr-0004]
    4. A method of manufacturing a composite sandwich panel (10) according to any one of the preceding claims, characterized in that each fibrous ply (34) is made of ceramic fibers and is impregnated with a slip containing powder of metal or mineral oxide for producing a ceramic matrix during the sintering operation.
    [5" id="c-fr-0005]
    5. A method of manufacturing a composite sandwich panel (10) according to claim 4, characterized in that it comprises a drying step which is carried out following each lining step and which consists in drying the fibrous fold ( 34) impregnated with slip and upholstered on the strip (20a, 20b) associated during the upholstering step.
    [6" id="c-fr-0006]
    6. A method of manufacturing a composite sandwich panel (10) according to any one of the preceding claims, characterized in that during the upholstering step, the fibrous fold (34) is upholstered on the impression (30a). , 30b, 31b) of the associated band by means of a toothed wheel (36) which has a shape complementary to the shape of said cavity.
    [7" id="c-fr-0007]
    A method of manufacturing a composite sandwich panel (10) according to claim 6, characterized in that the gear wheel (36) is a heating wheel which is adapted to heat and dry the fibrous ply (34).
    [8" id="c-fr-0008]
    8. A method of manufacturing a composite sandwich panel (10) according to any one of the preceding claims, characterized in that each cell (14) of the core (12) of the panel (10) has a hexagonal shape.
    [9" id="c-fr-0009]
    9. A method of manufacturing a panel (10) composite sandwich according to any one of claims 1 to 7, characterized in that each cell (14) of the core (12) of the panel (10) has a generally circular.
    [10" id="c-fr-0010]
    10. A method of manufacturing a composite sandwich panel (10) according to any one of the preceding claims, characterized in that each strip (20a, 20b) of fugitive material has a curved shape, so that the panel (10) realized has a curved shape.
    [11" id="c-fr-0011]
    11. A method of manufacturing a composite sandwich panel (10) according to any one of the preceding claims, characterized in that each fibrous fold (34) which forms the cells (14) of the panel (10) has at least a first protrusion (54) and a second protrusion (56) which are folded on each side of the associated band and which are adapted to cooperate with the first skin (16) and the second skin (18) respectively.
    [12" id="c-fr-0012]
    12. A method of manufacturing a composite sandwich panel (10) according to any one of the preceding claims, characterized in that the panel (10) obtained by the method is adapted to equip a nacelle with an aircraft propulsion system. .
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    同族专利:
    公开号 | 公开日
    US20180290330A1|2018-10-11|
    US10239227B2|2019-03-26|
    WO2017103390A1|2017-06-22|
    FR3044963B1|2018-01-19|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
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    WO2014118216A1|2013-01-29|2014-08-07|Herakles|Method for the production of a curved ceramic sound attenuation panel|WO2020016520A1|2018-07-17|2020-01-23|Safran Ceramics|Method for manufacturing an acoustic panel|US5549773A|1990-02-05|1996-08-27|Northrop Grumman Corporation|Woven preform/magnetic press process for thermoplastic honeycomb cores|
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    US5139596A|1990-05-31|1992-08-18|Basf Structural Materials, Inc.|Continuous process for the preparation of thermoplastic honeycomb|FR3070308B1|2017-08-25|2021-04-09|Safran Nacelles|STRUCTURAL AND / OR ACOUSTIC PANEL OF AN AIRCRAFT PROPULSIVE NACELLE ASSEMBLY AND RELATED MANUFACTURING PROCESS|
    FR3071245B1|2017-09-21|2019-09-20|Safran Ceramics|METHOD FOR INJECTING A BARBOTIN CHARGED WITH FIBROUS TEXTURE|
    EP3674081B1|2018-12-31|2022-02-23|Ansaldo Energia Switzerland AG|High-temperature resistant tiles and manufacturing method thereof|
    法律状态:
    2016-10-20| PLFP| Fee payment|Year of fee payment: 2 |
    2017-06-16| PLSC| Publication of the preliminary search report|Effective date: 20170616 |
    2017-11-23| PLFP| Fee payment|Year of fee payment: 3 |
    2018-03-02| CD| Change of name or company name|Owner name: SAFRAN NACELLES, FR Effective date: 20180125 |
    2019-11-20| PLFP| Fee payment|Year of fee payment: 5 |
    2020-11-20| PLFP| Fee payment|Year of fee payment: 6 |
    2021-11-17| PLFP| Fee payment|Year of fee payment: 7 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1562300|2015-12-14|
    FR1562300A|FR3044963B1|2015-12-14|2015-12-14|METHOD FOR MANUFACTURING A COMPOSITE SANDWICH PANEL FOR A PROPELLANT AIRCRAFT ASSEMBLY|FR1562300A| FR3044963B1|2015-12-14|2015-12-14|METHOD FOR MANUFACTURING A COMPOSITE SANDWICH PANEL FOR A PROPELLANT AIRCRAFT ASSEMBLY|
    PCT/FR2016/053303| WO2017103390A1|2015-12-14|2016-12-09|Method of manufacturing a composite sandwich panel for an aircraft propulsion unit|
    US16/008,208| US10239227B2|2015-12-14|2018-06-14|Method of manufacturing a composite sandwich panel for an aircraft propulsion unit|
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