METHOD FOR MANUFACTURING A HULL STRUCTURAL COMPONENT FOR A TRIANGLE AREA AND A SINGLE-PIECE HULL STRUCTURAL COMPONENT

Description

[0001] The invention relates to a method for manufacturing a fuselage structural component for a triangular section of an aircraft, as well as a one-piece fuselage structural component and an aircraft with a corresponding fuselage structural component. In particular, the invention relates to a method for manufacturing a one-piece fuselage structural component with integrally formed flanges, as well as a corresponding fuselage structural component and an aircraft.

[0002] In the lower deck of an aircraft, that is, in the area of ​​the fuselage below an intermediate floor, such as a cargo deck, there are typically two so-called triangular sections. In a cross-sectional view of the aircraft, these are defined by the outer structure of the fuselage (for example, the frames), the crossbeams of the intermediate floor (for example, the intermediate floor beams), and so-called Z-struts, which each support an intermediate floor beam on a frame and usually run vertically. See, for example, the disclosures in documents GB159351 A1 and US2013306795 A1. The Z-struts also serve to stiffen the aircraft structure below the intermediate floor.

[0003] Each of the Z-struts below the intermediate floor (usually two) is located on the outer edge of the cargo deck and is therefore difficult for personnel to access during installation. For example, the working height is already limited by the intermediate floor, while the rounded fuselage results in an increasingly steep floor slope. This creates ergonomically unfavorable conditions for the personnel.

[0004] The invention is therefore based on the objective of providing a manufacturing method, fuselage structure component and aircraft that enable simplified assembly of the triangular area.

[0005] This problem is solved by a method having the features of claims 1 and 7, a fuselage structure component having the features of claims 11 and 12, and an aircraft having the features of claims 13-15.

[0006] The dependent claims define exemplary embodiments.

[0007] According to a first aspect for a better understanding of the present disclosure, a method for manufacturing a fuselage structural component for a triangular section of an aircraft comprises providing a one-piece sheet of frame material, wherein the sheet has an opening. The one-piece sheet is made of a continuous material and therefore does not need to be assembled (from a multitude of components). Here, a sheet is understood to be a substantially flat component whose width and length are much greater than its thickness.

[0008] The opening in the panel is located in an arbitrary area of ​​the panel. The opening corresponds to a recess in the panel where no frame material is provided. Accordingly, the opening has an inner edge of the panel. Since it is a fuselage structural component forming a section of an aircraft's fuselage structure, the panel has a further inner edge that, when the fuselage structural component is installed, points into the interior of the aircraft. Accordingly, the panel has at least one outer edge that, when the fuselage structural component is installed, points away from a center point of the aircraft.

[0009] Furthermore, the method comprises deforming at least one inner edge of the sheet forming the opening into an inner flange, and deforming at least one outer edge of the sheet into an outer flange. A flange is thus understood to be a deformation of the sheet at the inner edge that protrudes from the plane of the sheet. The inner flange and the outer flange can point in the same direction or in different directions.

[0010] The single-piece, deformed panel forms a frame section, a floor beam, and a vertical support adjacent to the opening. In other words, the deformed panel forms a triangular area encompassing the triangular opening and, around it, a section of an aircraft frame, a section of a horizontal floor beam, and a vertical support. The vertical support functions as a Z-strut or vertical brace.

[0011] Since these components, which form the triangular section, are formed from a single piece of sheet metal, the entire assembly process of a conventional triangular section is eliminated. This process typically involves attaching crossbeams for the floor or entire floor elements to the frame. The Z-strut is then installed under the floor and mechanically fastened to the crossbeam and the frame to form the triangular section. The necessary and ergonomically challenging steps for this are completely eliminated. The number of components is also significantly reduced.

[0012] Furthermore, the deformation of the panel can take place outside the aircraft fuselage, so that the fuselage structural component can be manufactured independently in terms of time and in a truly independent manner.

[0013] Because the panel is a single piece, the resulting fuselage structure component is also lighter compared to previously used designs. Firstly, all fasteners (bolts, nuts, etc.) are eliminated. Secondly, overlapping areas are also unnecessary, such as the U-shaped or L-shaped ends at the top and bottom of the Z-strut that enclose and overlap the cross spar or frame.

[0014] The inner and outer flanges stiffen the corresponding section of the sheet because the flange extends in a different direction (or plane) than the sheet, both before and after deformation. This means that only the deformation steps are required to produce the entire triangular area, including any structurally necessary components.

[0015] The resulting hull structure component can be connected to further frame sections. Common joining methods can be used, such as welding, gluing, bolting, or riveting. A complete frame is usually made up of several frame sections, so the triangular hull structure component produced here can easily be integrated into the frame manufacturing process.

[0016] In one implementation variant, providing the panel may involve removing the frame material from the panel in the area of ​​the opening.

[0017] According to the invention, the panel can be made of a metal or alloy. It can be produced by rolling, so that the panel resembles a sheet. By way of example only, the removal process can include milling or punching, whereby the opening is milled or punched into the metal or alloy. This manufacturing method is simple to carry out and therefore involves low costs and requires little labor time.

[0018] In another exemplary implementation, not related to the claimed subject matter, the panel can consist of a composite material. For example, the panel can consist of a fabric (glass fiber, carbon fiber, etc.) with a curable matrix (epoxy resin, thermoplastic, etc.). In this case, the opening can also be milled or punched. However, the composite material can already be manufactured with an opening.

[0019] In one implementation variant, the deformation (of both the inner and outer flanges) can be achieved by pressing. With a metal or thermoplastic material, the inner or outer edge can be easily produced by plastic deformation through pressing the sheet between a die and a negative mold. Depending on the material, the sheet can also be heated, at least partially, during this process.

[0020] In pressing, the negative mold can be made of any material, for example wood, metal, or plastic. The die can also be made of any material, for example wood, metal, plastic, rubber, or a combination thereof.

[0021] Pressing allows for the simple and particularly fast production of the fuselage structural component, as the statically necessary structures, such as the inner and outer flanges, can be manufactured in a single operation. Alternatively or additionally, the inner and / or outer flanges can be produced by folding. This may allow for a more precise bend in the flange. Folding also allows the flange to be cut, enabling the precise production of any desired flange length. However, folding is more time-consuming than pressing.

[0022] In a further implementation variant, the method can also include the insertion of a reinforcing element. The reinforcing element preferably rests against a planar section of the panel and against the inner flange. For example, the reinforcing element can form a doubling in a specific section of the fuselage structural component. The reinforcing element can have a cross-section that corresponds to that of the corresponding section of the fuselage structural component, so that in cross-sectional view, the reinforcing element largely follows the section of the fuselage structural component and rests against it over a planar area.

[0023] Another example, not related to the claimed subject matter, is the reinforcement element being a doubling on a flange. This statically strengthens the flange. In particular, one or more reinforcement elements can be attached to one or more flanges of the base beam and / or to the vertical support.

[0024] The reinforcing element can also be provided on a floor element or intermediate floor to be attached. The reinforcement of the flange (for example, by appropriate doubling) is then only achieved during the assembly of the floor element or intermediate floor and the hull structure component.

[0025] The reinforcement element can be attached to the fuselage structure component by welding, gluing, screwing or riveting.

[0026] In yet another implementation variant, the reinforcing element can have at least one extension that extends from one end of the inner flange and has an arc. In particular, a section of the reinforcing element that abuts the inner flange can extend beyond the end of the inner flange. This increases the stability of the entire fuselage structural component.

[0027] For example, the arc of the reinforcing element can run along a corner of the opening in the panel. This allows the at least one inner flange to be formed along straight inner edges of the opening, rather than in the area of ​​the corner, which is more difficult to achieve by deformation. The arc of the reinforcing element then takes on the function of reinforcing the corner.

[0028] Furthermore, the arc of the reinforcing element can overlap another internal flange and / or be attached to it. This allows two adjacent internal flanges to be mechanically connected via the arc, thus statically reinforcing and strengthening the corner.

[0029] The thickness of the reinforcing element material can be selected according to the component being reinforced. Naturally, multiple reinforcing elements can be attached to the fuselage structure component to structurally strengthen specific areas. By incorporating one or more reinforcing elements, the panel (and thus the rest of the fuselage structure component) can be manufactured with a uniform thickness. This simplifies the panel manufacturing process. Forming the edges to form a flange also requires the same amount of deformation energy. For example, all internal flanges can be formed simultaneously in a single operation (such as a pressing operation), which is facilitated by a uniform panel thickness.

[0030] In another implementation variant, not related to the claimed subject matter, providing the one-piece panel can include providing a panel with areas of varying thickness. For example, the panel (before the forming steps) can be manufactured and provided as a sheet or flat component with areas of different thicknesses (areas of varying strength). Thus, the panel can be provided already with an integrated reinforcing element or with reinforced areas.

[0031] These areas of varying thickness can be achieved, for example, by rolling or pressing with appropriate templates / molds / punches. Alternatively or additionally, sections of the sheet can also be sanded or milled to obtain a thinner layer. Also alternatively or additionally, during the manufacturing process, particularly with plastic materials, varying sheet thicknesses can be achieved using appropriate three-dimensional molds.

[0032] The resulting panel does not require additional reinforcing elements. However, the manufacturing effort for the panel is greater, so a careful assessment must be made regarding which type of reinforcement to use.

[0033] As just one example, the plate can be thicker, particularly in the area where the frame section intersects the floor beam and / or the frame section intersects the vertical support. Similarly, the plate can be thicker in the area where the floor beam intersects the vertical support. In these areas, forces are transferred to the other section, and especially to the frame section. Therefore, a thicker plate in this area acts as a reinforcement, preventing deformation such as bulging and / or buckling.

[0034] As a further example, not related to the claimed subject matter, areas that are formed into a flange during deformation can be provided with a thicker sheet thickness. This makes it possible to already incorporate reinforced flanges on the hull structural component.

[0035] In another implementation variant, deforming an inner edge to form the inner flange can involve deforming several inner edges to create a single inner flange. A flange is thus formed at each deformed inner edge. The inner flange is interrupted at a corner between two inner edges. This simplifies the deformation process and allows for a sharp bend between the panel and the inner flange. It is also possible to deform the panel beyond 90°, which may be necessary, for example, for attaching additional components.

[0036] Alternatively, the deformation of an inner edge to form the inner flange can involve deforming several inner edges into a single inner flange that is formed continuously at at least one corner between two inner edges. In other words, the inner edge remains continuous even though, after deformation (viewed from above), it forms a continuous inner flange at two obliquely arranged inner edges. The continuously formed inner flange can be rounded at the corner. This increases the stiffness of the fuselage structural component in the area at the corner of the opening. Therefore, a separate reinforcing element can be omitted.

[0037] In yet another implementation variant, not related to the claimed subject matter, the continuously formed inner flange can have different heights. Here, the height of the inner flange is understood as the distance from the inner edge (now of the flange) to the plane of the (remaining) panel, or alternatively, the length of the flange from the flat panel to the deformed inner edge. A greater flange height signifies greater reinforcement in that area of ​​the panel by the resulting "L-beam." In areas where significant reinforcement is not statically necessary, the flange height can be reduced. By way of example, the inner edge along the frame section can be formed into a lower inner flange, while in the areas of the opening corners, as well as along the base beam and the vertical support, a higher inner flange is formed.

[0038] Alternatively, the inner flange can be of a lower height in the corner areas of the opening than in the (elongated) areas of the floor beam, the vertical support, and / or the frame section. This simplifies the deformation of the outer edge of the panel in the corner area.

[0039] An exemplary height of the inner flange is between 5 mm and 30 mm, preferably between 5 mm and 15 mm, and particularly preferably 10 mm or 10.4 mm. This applies especially to areas with the greater inner flange height. The areas with the lower height can depend on this and be chosen arbitrarily to comply with the static conditions and / or deformation possibilities.

[0040] In another implementation variant, the continuously formed inner flange can have an angle of 110° to 140°, preferably 120° to 130°, to the panel at a corner between two inner edges. In other words, the inner edge is bent by 70° to 40°, preferably 60° to 50°, at the corner of the opening to form the inner flange. Particularly in the area of ​​the corner of the opening, forming the inner edge into an inner flange is more difficult because a rounded flange must be formed. An angle between the panel and the flange that is too small (i.e., a large deformation) can lead to cracking of the frame material. An angle between the panel and the flange that is too large (i.e., a small deformation) can result in a flange that is too small and does not provide the desired reinforcement. By way of example only, the inner flange can also be further deformed in the area of ​​the bottom support, the vertical column, and / or the frame section.In these areas, the inner flange can have an angle of up to 90° to the panel.

[0041] In yet another implementation variant, not related to the claimed subject matter, the method can include deforming a section of the panel into a diagonal brace formed integrally with the panel. A conventional diagonal brace is usually positioned between and attached to two frames, with the diagonal brace arranged obliquely to each of the frames. This allows forces in the longitudinal direction of the aircraft (X-axis) to be transmitted between the frames. Torsional forces in the fuselage as well as vertical forces (Z-axis) can also be transmitted. Since the diagonal brace extends away from one of the frames, the material provided that forms the panel can already encompass the diagonal brace. This corresponding section of the panel can then be deformed, in particular bent out of the plane of the panel in the subsequent direction of extension.

[0042] In one implementation variant, the forming steps can be performed simultaneously. This allows all inner and outer flanges to be formed at once in a single process step, resulting in significant time savings and therefore reduced costs. Similarly, the step of forming a section of the sheet into a diagonal brace can be performed simultaneously with the forming of all inner and / or all outer flanges. According to a second aspect for a better understanding of the present disclosure, a fuselage structural component for a triangular area of ​​an aircraft comprises a one-piece sheet of frame material, wherein the sheet has an opening. This is a predominantly flat sheet made of the same material as the frame.

[0043] Furthermore, the fuselage structural component comprises at least one inner flange formed on an inner edge of the panel that forms the opening, and at least one outer flange formed on an outer edge of the panel. An inner flange is located closer to the center of a cross-section of the aircraft fuselage than an outer flange. For example, the at least one outer flange may be formed exclusively on an outer edge of the panel facing a subsequently attached outer skin of the aircraft. One or more inner flanges may be formed on an inner edge of the panel facing the center of the aircraft fuselage and / or on an inner edge of the panel's opening.

[0044] The one-piece panel forms a frame section, a base support, and a vertical support, all of which adjoin the opening. Thus, the frame section, the base support, and the vertical support are provided from a single, continuous piece of material, eliminating the need to assemble and connect / fasten different components.

[0045] The fuselage structure component can be manufactured using the method according to the first aspect or one or more of its implementation variants.

[0046] According to a third aspect for a better understanding of the present disclosure, an aircraft comprises a fuselage structural component for a triangular area according to the second aspect.

[0047] The aircraft can incorporate a large number of such fuselage structure components, which significantly reduces the overall construction time. This saves costs and simplifies the work for the relevant personnel, especially in the triangular area.

[0048] In one implementation variant, an intermediate floor can be attached to and secured on the floor support. This intermediate floor can be a standard aircraft cabin floor, separating a passenger / cabin area from a cargo area below. This creates the triangular area within the aircraft before the associated workspace is covered and restricted by the intermediate floor.

[0049] In another implementation variant, a stringer with an outer skin of the aircraft can be attached to the at least one outer flange. The outer skin of the aircraft can be attached to a multitude of stringers. Such a stringer can be attached to the at least one outer flange by welding, gluing, bolting, or riveting. A conventional method can be used, since standard frames have clips that are first attached to the frame, and the stringers are then attached to these clips. The outer flanges integrated into the frame according to the present disclosure also significantly reduce the aircraft's construction time, since conventional aircraft construction methods require attaching a multitude of clips to a multitude of frames.

[0050] The aspects, configurations, variants, and examples described above can, of course, be combined without this being explicitly stated. Each of the described configurations and each example is therefore optional to any of the aspects, configurations, variants, and examples, or even combinations thereof. This disclosure is thus not limited to the individual configurations and implementation variants in the described order or to any specific combination of aspects and implementation variants.

[0051] Preferred embodiments of the invention will now be explained in more detail with reference to the accompanying schematic drawings, wherein: Figure 1 schematically shows an aircraft with an intermediate floor and triangular section according to the prior art; Figure 2 schematically shows a fuselage structural component with reinforcing elements; Figure 3 schematically shows a fuselage structural component with reinforcing elements and a diagonal brace; Figure 4 schematically shows a detail view of a reinforcing element and the diagonal brace. Figure 3 Figure 5 shows a schematic representation of a hull structure component with reinforcement elements; Figure 6 shows a schematic representation of a hull structure component with a continuous inner flange; Figure 7 shows a schematic representation of a hull structure component with a continuous inner flange and diagonal brace; Figure 8 shows a schematic representation of a hull structure component with different material thicknesses and a diagonal brace; and Figure 9 shows a schematic flowchart of a process for manufacturing a hull structure component.

[0052] Figure 1Figure 1 schematically shows an aircraft 1 with an intermediate floor 38 and a triangular section 100 according to the prior art. The aircraft 1 is manufactured from a multitude of frames 10, all of which have the same or a very similar design. Such a frame element 10 comprises a circumferential bulkhead 30, an outer skin 20, and stringers 33 attached to an inner surface of the outer skin 20. Naturally, the outer skin 20 need not be provided in the length shown, but can span several bulkheads 30. The same applies to the stringers 33.

[0053] The intermediate floor 38 separates an overlying cabin area or passenger compartment from a baggage compartment or cargo deck below. The intermediate floor 38 is attached to a frame 30 on one side and supported on the frame 30 by a Z-strut 35 on the other. Thus, the frame 30, a section of the intermediate floor 38, and the Z-strut 35 form a triangular area 100. Cables, for example, can be routed along the longitudinal direction of the aircraft 1 (X-axis) within this triangular area 100.

[0054] Normally, the intermediate floor 38, or at least a crossbeam of the intermediate floor 38, is installed and connected to the frame 30 on both sides. Subsequently, the Z-struts 35 are installed and connected to the frame 30 at their lower end and to the intermediate floor 38 at their upper end. The necessary work for this takes place on the sloping (rounded) floor of the cargo deck. Furthermore, the working area in triangular section 100 is restricted by the intermediate floor 38 that has already been installed.

[0055] Figure 2 Figure 1 schematically shows a fuselage structure component 200 according to the present disclosure. Reinforcing elements 220 and 230 are also shown.

[0056] The fuselage structure component 200 comprises a single-piece panel 211, 212, 213, which constitutes a continuous component. The panel 211, 212, 213 is, for example, made of a frame material, so that the fuselage structure component 200 can be easily integrated into the rest of the fuselage of aircraft 1. The panel 211, 212, 213 has an opening 250, which essentially corresponds to the opening of a typical triangular area 100.

[0057] The fuselage structural component 200 comprises at least one inner flange 215, 216, 217 formed on an inner edge of the panels 211, 212, 213 forming the opening 250. The at least one inner flange 215, 216, 217 is formed on inner edges of the panels 211, 212, 213 facing a center of the cross-section of the aircraft 1. Since they represent a reinforcement of the corresponding section of the fuselage structural component 200, the inner flanges 215, 216, 217 are formed particularly in those areas that require reinforcement.

[0058] Furthermore, the fuselage structural component 200 comprises at least one outer flange 214, which is formed on an outer edge of the panels 211, 212, 213. The outer edge of the panels 211, 212, 213 is an edge that faces away from a center of the cross-section of the aircraft 1. The at least one outer flange 214 is formed from the material of the panels 211, 212, 213, i.e., it is manufactured integrally with them. On the one hand, the outer flange 214 forms a reinforcement of the frame 210 (or the frame section 210).

[0059] On the other hand, the at least one outer flange 214 takes on the function of a conventional clip (not shown) that is applied to a conventional frame 30 in order to attach a stringer 33 to it. In the fuselage structural component 200 according to the present disclosure, a stringer 33 can be attached directly to the outer flange 214, which is formed integrally with the rest of the fuselage structural component 200. This significantly speeds up the manufacture of the aircraft, including the attachment of the outer skin 20, since no clip needs to be installed. The fuselage structural component 200 has recesses 252 on its outer edge through which the stringers 33 can pass.

[0060] The one-piece panel 211, 212, 213 forms a frame section 210, a bottom support 212, and a vertical support 213, all of which adjoin the opening 250. Since the hull structure component 200 is manufactured in one piece, this entire triangular area 100 is produced in a simple manner, and this can be done very quickly, for example, by a single pressing operation (deformation operation).

[0061] By way of example only, the one-piece plate 211, 212, 213 can be made of a material with a thickness of 1.0 mm to 4.0 mm, preferably 1.5 mm to 2.5 mm, and particularly preferably 1.8 mm. Thus, each flange 214, 215, 216, 217 also has the same material thickness.

[0062] In particular, the inner flanges 215, 216 reinforce the base support 212 and the vertical support 213. For example, the inner flanges 215, 216 can be formed on both sides of the base support 212 or on both sides of the vertical support 213, so that a C- or U-shaped one-piece support is formed in each case.

[0063] In the implementation variant according to Figure 2Several inner edges are deformed to form a respective inner flange 215, 216, 217, with the inner flange 215, 216, 217 being interrupted at a corner between two inner edges. This allows the individual inner flanges 215, 216, 217 to be produced easily and in a single deformation step. However, to reinforce the corners between the respective inner flanges 215, 216, 217 or the corners of the opening 250, reinforcing elements 220, 230 can be provided. These reinforcing elements 220, 230 can also reinforce the base support 212 (in the case of reinforcing element 220) and the vertical support 213 (in the case of reinforcing element 230).

[0064] The reinforcing element 220, 230 can thus abut a flat section of the panel 212, 213, for example with a section 221, 231 of the reinforcing element 220, 230. Furthermore, the reinforcing element 220, 230 can abut the inner flange 215, 216, 217, for example with a flange section 225, 236. In other words, the reinforcing element 220, 230 has a cross-section that corresponds to that of the base support 212 or the vertical column 213, but is slightly smaller. The reinforcing element 220, 230 can therefore be inserted or placed into the base support 212 or the vertical column 213 and fastened to it.

[0065] By way of example only, a reinforcing element 220, 230 can have a material thickness of 2.0 mm to 4 mm, preferably of 2.5 mm to 3.5 mm and particularly preferably of 2.8 mm.

[0066] As especially from Figure 5As can be seen, the reinforcing element 220, 230 can have at least one extension 226, 227, 237. Such an extension 226, 227, 237 can extend from a flange section 225, 236 of the reinforcing element and also from an end of the inner flange 215, 216, 217 (beyond this end) and have an arc. Furthermore, the extension 226, 227, 237 can also overlap and be attached to another inner flange 215, 216, 217 that is formed on an adjacent inner edge. Thus, the extension can connect and reinforce two adjacent inner flanges 215, 216, 217.

[0067] In Figure 5Sections 221 and 231 of the reinforcing element 220 and 230 are also shown in detail, resting on the planar section of plate 212 and 213. These sections 221 and 231 can also form an extension to create reinforcement in a transition area between frame section 210, floor beam 212, and vertical support 213.

[0068] Figure 3 Figure 1 schematically shows a fuselage structural component 200 with reinforcing elements 220, 230 and also with a diagonal brace 240. The diagonal brace 240 primarily serves to connect two frames 30 in the longitudinal direction of the aircraft (X-axis) and preferably runs at an angle to the X-axis.

[0069] To avoid having to connect and fasten the diagonal brace 240 to frame 30 (at least on one side of the diagonal brace 240), the panels 211, 212, 213 can also have a section designated for the diagonal brace 240. Thus, the diagonal brace 240 can also be manufactured in one piece with the rest of the panels 211, 212, 213. The diagonal brace 240 can be formed in a corresponding deformation step during the manufacture of the hull structural component 200.

[0070] In the detailed views of the Figures 3 and 4The diagonal brace 240 is integrally connected in a first section 243 to an inner flange 215 of the base support 212. This first section 243 is deformed (bent) relative to the inner flange 215. To form the entire diagonal brace 240, further sections 242, 241 can be produced by deformation. Finally, flanges 244, 245 can also be produced on the diagonal brace 240 by deformation.

[0071] Although such a diagonal brace 240 can be manufactured through several forming steps, it is integrated into the fuselage structure component 200 (formed in one piece). This eliminates all connection elements and associated fastening work on at least one side of the diagonal brace 240.

[0072] Figure 6 Figure 1 schematically shows a fuselage structural component 200 with a continuous inner flange 215, 216, 217. Further views and details of this fuselage structural component 200 are in Figure 7As illustrated, several inner edges can be deformed into a single inner flange 215, 216, 217, which is formed continuously in at least one corner between two inner edges. This eliminates the need for a separate reinforcement element, since the fuselage structural component 100 is already reinforced in the respective corner.

[0073] If the continuously formed inner flange 215, 216, 217 is formed by pressing, it is advantageous to round the corner. The rounding here is to be understood as meaning that, in the plane of the sheet, the corner has a radius R that rotates about an axis perpendicular to the plane of the sheet. Otherwise, i.e., if a sharp edge were formed (as is the case in the Figures 2 to 5(in the case of) deformation of the inner edge of an inner flange 215, 216, 217 could lead to cracking. By way of example only, the inner flange 215, 216, 217 can have a minimum radius R that depends on the thickness t of the material of the sheet 211, 212, 213. Thus, with a material thickness t of 1.0 mm to 2.8 mm, a minimum radius R of between 8 cm and 12 cm, preferably 10 cm, can be provided at a corner, and with a material thickness t of >2.8 mm to 3.0 mm, a minimum radius R of approximately 12 cm can be provided.

[0074] Furthermore, the continuously formed inner flange 215, 216, 217 can have an angle α to the plate 211, 212, 213 of 110° to 140°, preferably 120° to 130°, at a corner between two inner edges. To also prevent cracking, the inner flange 215, 216, 217 is not further deformed, for example not up to 90° as in the variants of Figures 2 to 5. However, the inner flange outside the corners (i.e. in the straight areas 215, 216) can be further deformed up to and including 90°.

[0075] In Figure 6 Furthermore, a reinforcing element 260 is shown, which can be attached to the base support 212. This allows, in particular, the load-bearing area on the upper side of the base support 212 to be reinforced. The reinforcing element 260 can be attached to the inner flange 215 (for example, by welding, bonding, bolting, or riveting). By way of example only, the reinforcing element 260 can have a material thickness of 1.0 mm to 4.0 mm, preferably 1.5 mm to 3.0 mm, and particularly preferably 2.0 mm.

[0076] On the other hand, it is also possible to manufacture the reinforcing element 260 integrally with the inner flange 215 and thus integrally with the flat section 212 of the bottom support 212 and the remaining hull structural component 200. In this case, a section of the sheet can be deformed to form the reinforcing element 260 after the inner edge has been deformed to the inner flange 215, as shown. Alternatively, the reinforcing element 260 can also be molded simultaneously with the inner flange 215. In this case, the reinforcing element 260 has the same material thickness as the inner flange 215 and the flat section 212 of the bottom support.

[0077] Alternatively, it is also possible to provide reinforcement element 260 on the floor element or intermediate floor (not shown) to be attached to the floor support 212. During the assembly of the floor element / intermediate floor and the hull structure component 200, the reinforcement of the floor support 212 is achieved.

[0078] Again with reference to Figure 7The inner flanges 215, 216, and 217 can have different heights. In the illustrated implementation variant, the inner flanges 215 and 216 of the floor beam 212 and the vertical support 213 are designed with a first height H1. In the area of ​​the frame section 210, the inner flange 217 is designed with a second height H2, which can be less than the first height H1. This saves material and weight. For example, the reinforcing effect of the inner flange 217 can be less along the frame section 210 than in the area of ​​the floor beam 212 and the vertical support 213, which have to bear larger loads. These components (floor beam 212, vertical support 213) have a greater risk of buckling, which is why the inner flanges 215 and 216 should be stronger in these areas. In the area of ​​frame section 210, the inner flange 217 can also be largely omitted. Thus, the inner flange 215, 216 can be separated from the bottom support 212 or 216.Vertical support 213 tapers to a height of zero after the corner. As a further example, the inner flange at the corners can be formed with a lower height than the inner flange in the straight sections 215, 216.

[0079] In the Figure 7 and 8 Details of a one-piece manufactured diagonal brace 240 are also shown. This diagonal brace 240 can be manufactured in the same way as in the implementation variant according to Figures 3 and 4 , which is why the description of the diagonal brace 240 is not repeated here.

[0080] Figure 8Figure 1 schematically shows a hull structural component 200 with varying material thicknesses. Areas 211b of the panel, into which loads are introduced, can be made of a thicker material than the other areas 211a of the panel. These reinforced areas 211b are, in particular, the transition areas between frame section 210 and floor beam 212, as well as between frame section 210 and vertical support 213. The thicker material 211b can, for example, have a thickness of 1.5 mm to 4.0 mm, preferably 1.5 mm to 3.0 mm, and most preferably 2.8 mm. The thinner areas 211a, on the other hand, can consist of a material with a thickness of 1.0 mm to 3.0 mm, preferably 1.2 mm to 2.0 mm, and most preferably 1.4 mm or 1.8 mm.

[0081] Optionally, a third material thickness can also be provided in the single-piece panel. For example, a section of the panel forming the base support 212 and / or the vertical support 213 can be made of an even thicker material. By way of example only, the material thickness here can be between 2.0 mm and 5.0 mm, preferably between 2.5 mm and 4.5 mm, and particularly preferably 4.0 mm.

[0082] Figure 9 schematically shows a flowchart of a process for manufacturing a fuselage structure component 200.

[0083] The process begins in step 300 with the provision of a one-piece sheet 211, 212, 213 of frame material, the sheet having an opening 250.

[0084] Depending on the material, the opening 250 can be produced in one step 302 by removing the frame material from the panel in the area of ​​the opening 250. For example, milling or punching can be used for this purpose.

[0085] In step 311, this sheet 211, 212, 213 is processed such that at least one inner edge of the sheet forming the opening 250 is deformed into an inner flange 215, 216, 217. In step 312, at least one outer edge of the sheet 211, 212, 213 is deformed into an outer flange 214.

[0086] The one-piece and deformed plate 211, 212, 213 thus forms a frame section 210, a floor support 212 and a vertical support 213 adjacent to the opening 250.

[0087] In an optional step 313, a section of the panel can be deformed to form a diagonal brace 240, so that the diagonal brace 240 is formed integrally with the panel 211, 212, 213.

[0088] The deformation steps 311, 312, 313 can be carried out in a single process step. For example, all flanges 214, 215, 216, 217 and at least part of the diagonal brace 240 can be formed in a single pressing operation.

[0089] Optionally, in step 320 a reinforcing element 220, 230 can be inserted, which rests against a planar section of the panel 212, 213 and against an inner flange 215, 216, 217.

Claims

1. Method for producing a fuselage structural component (200) including a triangular region (100) for an aircraft (1), wherein the method comprises: providing (300) an integral board (211, 212, 213) of a former material, wherein the board is made of a metal or an alloy and has an opening (250), deforming (311) a plurality of internal edge of the board (211, 212, 213) that form the opening (250) to a single internal flange (215, 216, 217) formed continuously in at least one corner between two internal edges; and deforming (312) at least one external edge of the board (211, 212, 213) to an external flange (214), wherein the integral and deformed board (211, 212, 213) forms, in each case adjacent to the opening (250), a former portion (210), a floor beam (212) and a vertical support (213) in such a way that the opening (250) is triangular and the integral and deformed board (211, 212, 213) forms the triangular region about the opening (250), and wherein the continuously formed internal flange (215, 216, 217) in a corner between two internal edges has an angle of 110° to 140°, preferably of 120° to 130°, in relation to the board (211, 212, 213).

2. Method according to Claim 1, wherein providing (300) the board (211, 212, 213) comprises removing (302) the former material from the board (211, 212, 213) in the region of the opening (250), wherein removing (302) preferably comprises milling or punching.

3. Method according to Claim 1 or 2, furthermore comprising: incorporating (320) a reinforcement element (220, 230), wherein the reinforcement element (220, 230) bears on a planar portion of the board (212, 213) and on at least one portion of the internal flange (215, 216, 217).

4. Method according to Claim 3, wherein the reinforcement element (220, 230) has at least one appendage (226, 227, 237) which extends from one end of the internal flange (215, 216, 217) and has an arc.

5. Method according to one of Claims 1 to 4, wherein providing (300) the integral board (211, 212, 213) comprises providing a board (211, 212, 213) with regions of different thicknesses, wherein the board (211b) is configured to be thicker in particular in a region in which the former portion (210) intersects the floor beam (212) and / or the vertical support (213).

6. Method according to Claim 1, wherein the continuously formed internal flange (215, 216, 217) has different heights.

7. Method for producing a fuselage structural component (200) including a triangular region (100) for an aircraft (1), wherein the method comprises: providing (300) an integral board (211, 212, 213) of a former material, wherein the board is made of a metal or an alloy and has an opening (250), deforming (311) a plurality of internal edges to a respective internal flange (215, 216, 217), wherein the respective internal flanges (215, 216, 217) are interrupted in a corner between two internal edges; or deforming (312) at least one external edge of the board (211, 212, 213) to an external flange (214); incorporating (320) a reinforcement element (220, 230), wherein the reinforcement element (220, 230) bears on a planar portion of the board (212, 213) and on one of the internal flanges (215, 216, 217), and wherein the reinforcement element (220, 230) has at least one appendage (226, 227, 237) which extends from one end of the internal flange (215, 216, 217), has an arc, and overlaps a further one of the internal flanges and is fastened to the latter, wherein the integral and deformed board (211, 212, 213) forms, in each case adjacent to the opening (250), a former portion (210), a floor beam (212) and a vertical support (213) in such a way that the opening (250) is triangular and the integral and deformed board (211, 212, 213) forms the triangular region about the opening (250).

8. Method according to one of Claims 1 to 7, furthermore comprising: deforming (313) a portion of the board (211, 212, 213) to a diagonal brace (240) which is formed to be integral to the board (211, 212, 213).

9. Method according to one of Claims 1 to 8, wherein the steps of deforming (311, 312, 313) are carried out simultaneously.

10. Method according to one of Claims 1 to 9, wherein the floor beam (212) is specified for attaching thereto a cabin floor (38) of the aircraft (1).

11. Fuselage structural component (200) including a triangular region (100) for an aircraft (1), wherein the fuselage structural component (200) comprises: an integral board (211, 212, 213) of a former material, wherein the board is made of a metal or an alloy and has an opening (250), an internal flange (215, 216, 217) which is formed on a plurality of internal edges of the board (211, 212, 213) that form the opening (250) and in at least one corner is formed continuously between two internal edges; and at least one external flange (214) which is formed on an external edge of the board (211, 212, 213), wherein the integral board (211, 212, 213) forms a former portion (210), a floor beam (212) and a vertical support (213), all being in each case adjacent to the opening (250), so that the opening (250) is triangular and the integral board (211, 212, 213) forms the triangular region about the opening (250), and wherein the continuously formed internal flange (215, 216, 217) in a corner between two internal edges has an angle of 110° to 140°, preferably of 120° to 130°, in relation to the board (211, 212, 213).

12. Fuselage structural component (200) including a triangular region (100) for an aircraft (1), wherein the fuselage structural component (200) comprises: an integral board (211, 212, 213) of a former material, wherein the board is made of a metal or an alloy and has an opening (250), wherein the integral board (211, 212, 213) forms a former portion (210), a floor beam (212) and a vertical support (213), all being in each case adjacent to the opening (250), so that the opening (250) is triangular and the integral board (211, 212, 213) forms the triangular region about the opening (250), a plurality of internal flanges (215, 216, 217) which are in each case formed on an internal edge of the board (211, 212, 213) that forms the opening (250), wherein in a corner between two internal edges the respective internal flanges (215, 216, 217) are interrupted; at least one external flange (214) formed on an external edge of the board (211, 212, 213); and a reinforcement element (220, 230) which bears on a planar portion of the board (212, 213) and on one of the internal flanges (215, 216, 217), and wherein the reinforcement element (220, 230) has at least one appendage (226, 227, 237) which extends from one end of the internal flange (215, 216, 217), has an arc, and overlaps a further one of the internal flanges and is fastened to the latter.

13. Aircraft (1), comprising: a fuselage structural component (200) including a triangular region (100) according to Claim 11 or 12.

14. Aircraft (1) according to Claim 13, wherein an intermediate floor (38) is attached and fastened to the floor beam (212).

15. Aircraft (1) according to Claim 13 or 14, wherein fastened to the at least one external flange (214) is a stringer (33) having an external skin (20) of the aircraft (1).