FIBER COMPOSITE COMPONENT AND PRODUCTION METHOD

Abstract

The invention relates to a fiber composite component (17) and to a method for producing a fiber composite component (17), for an aircraft, in particular for an aircraft cabin interior, a tabletop (21) or the like, the fiber composite component (17) being formed from a matrix composite material (19) and a support structure, wherein the matrix composite material (19) is formed from cut fibers, a curable resin, and a flame retardant, the support structure being formed from a dimensionally stable fiber composite (18) and/or a metal profile, the matrix composite material (19) together with the support structure being introduced into a component mold and cured to form the fiber composite component (17), the support structure being at least partially bonded with the matrix composite material (19).

Claims

1. A method for producing a fiber composite component (11, 17, 22, 27, 31, 37, 38), for an aircraft, in particular for an aircraft cabin interior, a tabletop (10, 21, 26, 30, 35, 39, 47) or the like, the fiber composite component being formed from a matrix composite material (14, 19, 24, 29, 36, 40, 49) and a support structure (12, 33, 48), characterized in that the matrix composite material is formed from cut fibers, a curable resin, and a flame retardant, the support structure being formed from a dimensionally stable fiber composite (13, 18, 23, 42) and/or from a metal profile (32, 43, 44, 45, 46), the matrix composite material together with the support structure being introduced into a component mold and cured to form the fiber composite component, the support structure being at least partially bonded with the matrix composite material.

2. The method according to claim 1, characterized in that the fiber composite (13, 18, 23, 42) is formed from textile fibers and/or unidirectional fibers.

3. The method according to claim 1, characterized in that the fiber composite (13, 18, 23, 42) is formed as a spatially oriented support structure (12, 33, 48) of the fiber composite component (11, 17, 22, 27, 31, 37, 38) which is adapted to a load condition of the fiber composite component.

4. The method according to claim 1, characterized in that the fiber composite (13, 18, 23, 42) is formed from carbon fibers, the carbon fibers being coated with pyrolytic carbon fibers so as to form the fiber composite.

5. The method according to claim 4, characterized in that the pyrolytic carbon is deposited onto the carbon fibers from the vapor phase.

6. The method according to claim 1, characterized in that the cut fibers are carbon fibers.

7. The method according to claim 4, characterized in that the fiber composite component (11, 17, 22, 27, 31, 37, 38) is formed in such a manner that it has a carbon fiber content of >35% by volume, preferably >50% by volume.

8. The method according to claim 4, characterized in that the fiber composite component (11, 17, 22, 27, 31, 37, 38) is formed in such a manner that the carbon fibers are distributed heterogeneously within the fiber composite component.

9. The method according to claim 1, characterized in that the matrix composite material (14, 19, 24, 29, 36, 40, 49) is a semi-finished fiber matrix product, in particular a sheet molding compound (SMC) or a bulk molding compound (BMC).

10. The method according to claim 1, characterized in that the matrix composite material (14, 19, 24, 29, 36, 40, 49) is compressed with the support structure (12, 33, 48) in the component mold at a pressure of 80 bar and 150 bar, in particular between 90 bar and 110 bar, and at a temperature between 125 C. and 150 C., in particular between 130 C. and 140 C.

11. The method according to claim 1, characterized in that the flame retardant is aluminum hydroxide.

12. The method according to claim 1, characterized in that the matrix composite material (14, 19, 24, 29, 36, 40, 49) contains at least 40 w %, in particular at least 50 w %, in particular at least 60 w %, in particular at least 70 w % of aluminum trihydroxide.

13. The method according to claim 1, characterized in that the fiber composite (13, 18, 23, 42) is disposed in a premold and is pre-stabilized, preferably pre-cured, by pressing.

14. The method according to claim 1, characterized in that the support structure (12, 48) is introduced into the component mold in such a manner that the matrix composite material (14, 19, 29, 49) completely surrounds the support structure.

15. The method according to claim 1, characterized in that the support structure (12, 33, 48) is formed in one piece or in multiple pieces, the support structure at least in sections forming a frame (50) which defines a frame inner surface, the frame inner surface being filled by the matrix composite material (14, 19, 29, 49).

16. A fiber composite component (11, 17, 22, 27, 31, 37, 38) for an aircraft, in particular for an aircraft cabin interior, a tabletop (10, 21, 26, 30, 35, 39, 47) or the like, the fiber composite component being made of a matrix composite material (14, 19, 24, 29, 36, 40, 49) and a support structure (12, 33, 48), characterized in that the matrix composite material is made of cut fibers, a resin, and a flame retardant, the support structure being made of a dimensionally stable fiber composite (13, 18, 23, 42) and/or of a metal profile (32, 43, 44, 45, 46), the matrix composite material together with the support structure having been introduced into a component mold and cured to form the fiber composite component, the support structure being at least partially bonded with the matrix composite material.

17. The fiber composite component according to claim 16, characterized in that the fiber composite component (11, 17, 22, 27, 31, 37, 38) has a density of <2.7 g/cm.sup.3.

18. A use of a matrix composite material (14, 19, 24, 29, 36, 40, 49) having a support structure (12, 33, 48), for producing an aircraft cabin interior, in particular a tabletop (10, 21, 26, 30, 35, 39, 47), wherein the matrix composite material is made of cut fibers, a resin, and a flame retardant, the support structure being made of a dimensionally stable fiber composite (13, 18, 23, 42) and/or of a metal profile (32, 43, 44, 45, 46), the matrix composite material together with the support structure being introduced into a component mold and cured to form the fiber composite component, the support structure being at least partially bonded with the matrix composite material.

Description

[0036] Hereinafter, preferred embodiments of the invention will be explained in more detail with reference to the accompanying drawings.

[0037] FIG. 1 shows a side view of a first embodiment of a tabletop of an aircraft table;

[0038] FIG. 2 shows a top view of the tabletop;

[0039] FIG. 3 shows a partial section view of a second embodiment of a tabletop;

[0040] FIG. 4 shows a partial section view of a third embodiment of a tabletop;

[0041] FIG. 5 shows a partial section view of a fourth embodiment of a tabletop;

[0042] FIG. 6 shows a partial section view of a fifth embodiment of a tabletop;

[0043] FIG. 7 shows a partial section view of a sixth embodiment of a tabletop;

[0044] FIG. 8 shows a partial section view of a seventh embodiment of a tabletop;

[0045] FIG. 9 shows a perspective view of a first embodiment of a metal profile;

[0046] FIG. 10 shows a perspective view of a second embodiment of a metal profile;

[0047] FIG. 11 shows a perspective view of a third embodiment of a metal profile;

[0048] FIG. 12 shows a perspective view of a fourth embodiment of a metal profile;

[0049] FIG. 13 shows a perspective view of an eighth embodiment of a tabletop;

[0050] FIG. 14 shows a perspective view of the eighth embodiment of the tabletop.

[0051] A combined view of FIGS. 1 and 2 shows a tabletop 10 of an aircraft cabin interior, said tabletop 10 being made of a fiber composite component 11. The fiber composite component 11 has a support structure 12 formed by a dimensionally stable fiber composite 13 made of carbon fibers, which are coated with pyrolytic carbon and illustrated merely in outlines. A matrix composite material 14, which is made of cut fibers or carbon cut fibers (not illustrated), a cured resin, and a flame retardant, completely surrounds the fiber composite 13, the fiber composite 13 thus being located on the inside of the tabletop 10. In particular, the fiber composite component 11 has been formed by disposing the dimensionally stable fiber composite 13 in a component mold (not illustrated) together with the matrix composite material 14 and by subsequently curing the matrix composite material within the component mold. The support structure 12 is substantially configured in such a manner that a load condition of the fiber composite component 11 can be accommodated. For instance, the tabletop 10 has two fins 15 at which it can be inserted into a guide (not illustrated) of a galley and be slid into or pulled out of the guide along side edges 16.

[0052] FIGS. 3 to 9 show partial section views of different embodiments of tabletops in the area of a side edge. FIG. 3 shows a fiber composite component 17 having a fiber composite 18 within a cured matrix composite material 19. The matrix composite material 19 completely surrounds the fiber composite 18 and forms a side edge 20 and a tabletop 21.

[0053] FIG. 4 shows a fiber composite component 38 in which a plate 39 made of a matrix composite material 40 is formed, the plate 39 being inserted into an edge profile 41, the edge profile 41 being made of a fiber composite 42 which is bonded with the plate 39.

[0054] FIG. 5 shows a fiber composite component 22 having a fiber composite 23 and a matrix composite material 24, the fiber composite 23 being connected to the matrix composite material 24 in sections only and disposed on an underside 25 of a tabletop 26.

[0055] FIG. 6 shows a fiber composite component 27 which differs from the fiber composite component of FIG. 3 in that a dimensionally stable fiber composite 28 is plate-shaped and extends substantially within a matrix composite material 29 across an entire surface of a tabletop 30.

[0056] FIG. 7 shows a fiber composite component 31 having a U-shaped metal profile 32 which forms a support structure 33 and which is bonded with a matrix composite material 36 along side edges 34 of a tabletop 35.

[0057] FIG. 8 shows a fiber composite component 37 which combines the support structure illustrated in FIG. 3 with the support structure in FIG. 7.

[0058] FIGS. 9 to 12 show metal profiles 43 to 46, respectively, which can form a support structure. In principle, profiles made of a fiber composite can be used instead of the metal profiles 43 to 46.

[0059] In FIGS. 13 and 14, a part of an insertable or folding aircraft table, in particular a tabletop 47, is shown. The tabletop 47 has a table surface 52 which is limited by a frame 50. The frame 50 further comprises two frame protrusions 51 which serve to connect the tabletop 47 to an inserting or folding mechanism of the insertable or folding aircraft table.

[0060] The tabletop 52 is preferably rectangular and has a flat or plane surface. It is also possible that the tabletop 47 has recesses, i.e. areas of reduced wall thickness of the table surface, or through-holes in the area of the tabletop 52. Recesses or through-holes of this kind may serve as cup holders, for example.

[0061] As is clearly visible in FIG. 14, the table surface 52 is sunk or lowered in relation to the frame 50. Thus, the table surface 52 has a reduced wall thickness compared to the frame 50. It is preferably envisaged that an underside of the table surface 52 is flush with an underside of the frame 50. Overall, the entire tabletop 47 thus has a plane underside. A lowered portion 53 merely shows on the upper side of the tabletop 47, the lowered portion 53 being formed by the lowered table surface 52.

[0062] A cutout 54 is provided between the frame protrusions 51. Said cutout 54 serves in particular to offer freedom of movement for an inserting or folding mechanism. At the same time, the cutout 54 allows sufficient free space to remain below the insertable or folding table or tabletop 47 in the inserted or folded state, such as for a newspaper holder on a backrest of an aircraft seat.

[0063] For the sake of stability, it is envisaged that the frame protrusions 51 taper toward their free ends 55. The outer edge of the frame protrusions 51 aligns with the outer edge of the frame 50, resulting in a straight or plane side surface of the tabletop 47. The cutout 54 has a substantially trapezoidal contour, the cutout 54 being broader between the free ends 55 of the frame protrusions 51 than along the frame 50.

[0064] The frame 50 comprises four frame sections 50a, 50b, the frame section 50b connecting the frame protrusions 51 having a larger web width than the free other frame sections 50. This increases the stability of the tabletop 47.

[0065] The tabletop 47 comprises an inner support structure 48, which is indicated by dashed lines in FIG. 14. The inner support structure 48 is provided with a matrix composite material 49, which forms the complex outer shape of the tabletop 47. In particular, the support structure 48 comprises rods, tubes or profiles which are made of a fiber composite material reinforced by endless fibers, in particular a carbon fiber composite material reinforced by endless fibers. As is clearly visible in FIG. 14, the rods of the support structure 48 extend along the free frame sections 50a as far as into the frame protrusions 51. The rods, tubes or profiles of the support structure 48 preferably have a rectangular crosssectional contour. The endless fibers preferably extend in the longitudinal direction of the rods, tubes or profiles of the support structure 48.

[0066] The matrix composite material 49 is formed by an SMC material. Said SMC material preferably comprises a carbon fiber composite material, the carbon fibers being embedded non-directionally in a polymer matrix as long fibers. The SMC material does not only completely envelop the support structure 48, but also forms the table surface 52 and the connecting frame section 50b. Moreover, the outer contour of the frame protrusions 51 is defined by the SMC material.

[0067] For example, the tabletop 47 of an aircraft table illustrated here by way of example is produced by a method that involves the following steps:

[0068] The support structure 48 is produced using a pultruding method, in particular by pultrusion, for example, or by wet winding or pre-preg lamination or vacuum infusion or another RTM method, from fiber-reinforced plastic or fiber composite materials. The fiber-reinforced plastic preferably comprises carbon fibers which are embedded in a matrix made of an epoxy resin, a vinyl ester resin, or a fire-resistant phenolic resin. The carbon fibers are endless fibers and can be oriented in a common main orientation direction.

[0069] The support structure 48 can be produced in one piece or comprise multiple pieces which are at least temporarily connected to one another by corresponding joining techniques. In particular, the support structure 48 can be formed by multiple rods, tubes or profiles which are glued together. The support structure 48 is pre-cured and, in a next step, is either embedded into the SMC material or placed into a component mold, which is preferably already filled or lined with an SMC material.

[0070] The support structure 48 can be embedded into the SMC material by placing the support structure 48 on a layer of the SMC materials, the support structure 48 covering only part of the layer of the SMC material. An overlapping part of the SMC materials can be folded over and laid on the support structure 48. Preferably, the support structure 48 is thus sandwiched between two portions of the layer of the SMC material and, together with the SMC material, forms a preform.

[0071] The preform is subsequently placed in a pressing tool. Alternatively, it may also be envisaged for the preform to be formed in the pressing tool itself. For this purpose, a layer of the SMC material can be placed into a tool half of the pressing tool, a portion of the layer extending beyond the tool half. The support structure 48 is placed into the pressing tool on top of the layer of the SMC material. The portion of the layer sticking out of the tool half is then folded over and laid on the support structure 48 so that the preform forms directly in the component mold.

[0072] In general, the support structure 48 can be embedded into multiple layers of the SMC material. In particular, the support structure 48 can be placed on a first layer of the SMC material, and an independent second layer of the SMC material can be placed on the first layer and on the support structure 48 so that the support structure 48 is covered by a separate layer of the SMC material on either side.

[0073] It may be advantageous if the component mold has holding devices for positioning the support structure 48. The SMC material can have glass fibers, carbon fibers, and/or aramid fibers which are embedded in a polymer matrix. The polymer matrix can comprise epoxy resin and/or vinyl ester resin and/or phenolic resin.

[0074] Preferably, the SMC material is fire-resistant pursuant to the aviation regulations. For example, the SMC material can have a polymer matrix which is filled with flame-retardant aluminum trihydroxide. In its raw state, the aluminum trihydroxide is preferably a powder and is admixed to the polymer matrix. The polymer matrix of the SMC material can also contain epoxy resin and/or vinyl ester resin and/or phenolic resin.

[0075] By embedding the support structure into the SMC material, a preform is formed, which is compressed in a pressing tool. The pressing tool preferably has a tool shape that corresponds to a negative shape of the component to be produced.

[0076] For example, compression takes place at a pressure between 80 bar and 150 bar and at a temperature between 125 C. and 150 C. in the pressing tool. The SMC material fills the mold geometry of the pressing tool and structurally bonds with the pre-cured support structure 48. Thus, a bonded connection is formed between the support structure 48 and the matrix composite material 49. Hence, the finished aircraft component has a monolithic sandwich structure which has high mechanical stability owing to the embedded support structure 48. The support structure 48 mainly serves to transmit loads and to absorb mechanical forces, whereas the matrix composite material 49, which is formed by an SMC material, presents the complex outer component contour.

[0077] The compressed component preferably cures after few minutes, in particular within a period of 1 minute to 10 minutes, in the hot pressing tool.

[0078] Once the curing time has elapsed, the finished aircraft component, in particular the tabletop 47 described above, is removed from the hot pressing tool.