Method for manufacturing a structural section of a vehicle

Abstract

Method for manufacturing a structural section (37, 55, 72) of a vehicle, comprising the steps: - Providing several separate skin panels (2) made of a fiber-reinforced plastic with an inner surface (10), an outer surface and a rim (8) surrounding the respective skin panel (2), - Arranging at least one stiffening component (12, 14) made of a fiber-reinforced plastic on each skin panel (2) on the respective inner side (10), - material-bonded connection of the respective at least one stiffening component (12, 14) with the relevant skin fields (2) to form a structural component (36, 38), - Arranging at least two structural components (36, 38) on a support (56) such that at least areas of the edges (8) of the relevant skin fields (2) of the structural components (36, 38) are in surface contact, and - material-bonding connection of the areas of the edges (8) that are in surface contact with each other.

Description

TECHNICAL AREA

[0001] The invention relates to a method for manufacturing a fuselage section of an aircraft and to a fuselage section or fuselage of an aircraft manufactured using such a method. BACKGROUND OF THE INVENTION

[0002] In the manufacture of commercial aircraft, large fuselage components are typically riveted together. Riveting results in a considerable number of riveted joints, thus enabling a force transmission between two interconnected skin panels over a large area and via many redundant connections. To reduce the number of rivets required, the size of the skin panels to be joined can be increased, or similar measures can be taken. It is known to use skin panels with lengths of 10 meters and more. While riveted joints originate from the manufacture of metallic components, they can also be used when working with fiber-reinforced plastics.

[0003] The production of riveted joints is relatively complex and expensive, as each rivet requires drilling, deburring and cleaning the hole, applying a sealant, and finally inserting and shaping the rivet. Joining large skin panels on commercial aircraft results in a very high number of riveted joints.

[0004] DE 10 2007 044 387 A1 discloses a method for manufacturing an aircraft fuselage, in which several large-format fuselage sections are joined together one after the other.

[0005] DE 10 2012 015 666 A1 discloses a method for mounting frames on a skin panel during the manufacture of a fuselage shell component for an aircraft.

[0006] DE 10 2011 085 937 B4 describes a lightweight structure which consists, at least in sections, of a large number of interconnected mosaic elements and / or skin mosaic elements.

[0007] DE 10 2009 002 697 B4 relates to a mold core and a method for manufacturing a fiber composite component for the aerospace industry.

[0008] DE 101 34 852 B4 describes a holding device for holding large-format components.

[0009] DE 36 05 256 C2 discloses a method for producing flat or shell-shaped curved components made of fiber-reinforced plastics and equipped with stringers, ribs or frames. SUMMARY OF THE INVENTION

[0010] It would be helpful to have an alternative method for manufacturing fuselage sections, fuselage assemblies, or structural components that could eliminate at least some rivet connections. This method should be particularly suitable for joining fuselage components or other structural sections made of fiber-reinforced plastics, and especially fiber-reinforced thermoplastics.

[0011] One object of the invention is therefore to propose an alternative method for manufacturing a structural section or structural assembly of a vehicle, which not only enables a reliable and firm connection between individual skin panels or the like, but also achieves a particularly homogeneous force flow and a further improved surface quality.

[0012] This problem is solved by a method with the features of independent claim 1. Advantageous further developments can be found in the dependent claims and the following description.

[0013] A method for manufacturing a structural section of a vehicle is proposed. The method comprises the steps of providing several separate skin panels made of a fiber-reinforced plastic, each with an inner surface, an outer surface, and a surrounding edge; arranging at least one stiffening component made of a fiber-reinforced plastic on the respective inner surface of each skin panel; bonding the respective stiffening component to the skin panels to form a structural component; arranging at least two structural components on a carrier such that at least areas of the edges of the skin panels are in surface contact; and bonding the areas of the edges in surface contact to each other.

[0014] A separate skin panel can be understood as a component, particularly a planar one, that can form a section of a larger structural element, such as a larger fuselage section. The size of a skin panel can be chosen according to several criteria. Firstly, selecting a smaller skin panel can improve its handling. However, this would increase the number of edges that need to be joined between multiple skin panels. Secondly, by integrating at least one stiffening component, individual skin panels can form a functional unit or module that can represent a specific area, such as a window with its associated stiffener or a door area with its associated stiffener. It can be advantageous to tailor the size of the skin panel to a specific function.

[0015] Dividing a larger structural section into smaller parts can simplify the overall manufacturing process. These smaller parts, referred to here as structural components, can be manufactured in a highly automated process chain. Similar to automotive manufacturing, even complexly shaped structural components can be produced with high precision and speed, particularly using integral construction methods. The flexibility of the design and features of these structural components can be significantly increased by relatively simple adaptation of smaller production facilities. Due to the relatively small size of the structural components, improved warehousing and transportation, especially over long distances, can be achieved.This allows for the production of a large number of identical structural components and their intermediate storage before the process chain is switched to a different structural component or a different group of structural components.

[0016] The fiber-reinforced material of the skin panels can be a thermoplastic matrix material or a thermoset matrix material with embedded reinforcing fibers. The material composition of the skin panels is not limited to specific matrix materials. Rather, all conceivable matrix materials, especially those suitable for automated application, should be considered if they meet the strength requirements for the structural components. The reinforcing fibers could, in particular, include carbon fibers.

[0017] The layers of a skin panel arranged on top of each other do not necessarily have identical circumferential contours. Rather, differently designed layers can be used, resulting in a non-constant thickness of the skin panel. This allows, for example, the creation of thickened areas for the arrangement of circumferential door or window stiffeners.

[0018] The material-bonded joining of the at least one stiffening component and the skin panels, as well as the individual structural components, can be achieved by adhesive bonding when using a thermoset matrix material. However, welding should be considered for thermoplastic materials. Two components to be welded can be material-bonded by pressing them together and heating at least a localized area of ​​the joining surface. Tools can be used that can hold or press the components together and heat the joining surface to a welding temperature suitable for the thermoplastic material. These tools can also be designed to maintain the desired shape of the joining partners under the influence of heat.Alternatively, joining methods are also possible in which the skin panels and the relevant stiffening components or several structural components brought into surface contact are completely heated and welded together.

[0019] The material-bonded joining of the individual components of a structural part can also refer to an integral construction of the entire structural part. All components can be manufactured using a common layer structure and cured or co-consolidated together. The steps listed above can also be performed simultaneously or in a different sequence.

[0020] At least one stiffening component can be arranged on the inner surface of each skin panel. A stiffening component can be, for example, an elongated component that creates a pronounced cantilever on the inner surface of the respective skin panel, thereby increasing the area moment of inertia of the combination of skin panel and stiffening component. It is understandable that not only a single stiffening component, but also several stiffening components can be arranged on the respective skin panel. Depending on the function of the respective skin panel, the stiffening components can be spaced apart from each other and / or overlapping. Naturally, the method according to the invention is not limited to specific stiffening components, but can be implemented with any stiffening components with an open and / or closed profile cross-section.

[0021] By bonding the at least one stiffening component to the respective skin area, a stiffened skin area is created that is practically monolithic.

[0022] By arranging two structural components on a support, such as a storage frame, surface contact can be established between the skin surfaces of two adjacent structural components. The material-bonded joining of the contacting edges of the two structural components results in a group of structural components or a structural section of the vehicle.

[0023] In this way, a large number of such stiffened skin panels or structural components can be assembled and joined together in a highly automated manufacturing process. This eliminates the need for the one-piece production of individual large-format skin panels using complex manual or semi-automated methods. If the manufacturing method according to the invention is used to produce an aircraft fuselage, in particular a cylindrical fuselage, the individual structural components can be used to assemble barrel-shaped fuselage sections.

[0024] In an advantageous embodiment, providing a skin panel comprises placing at least two layers on top of each other in a mold and bonding the layers together, each layer having a circumferential contour and an offset between the circumferential contours of two overlapping layers. The edges of a skin panel, or of a fuselage component formed therewith, can therefore have a stepped shape. To join several structural components by surface contact of at least one area of ​​the edges, all structural components should therefore be designed to correspond with each other. The stepped design of the edges allows for particularly pronounced surface contact, which aligns the structural components with each other. Furthermore, improved local application of mechanical pressure can be achieved.By shaping the edges, overlapping fiber structures can be formed in the edge areas, leading to improved strength of the composite of several skin panels.

[0025] It should be noted here that a skin panel could also be used that is not composed of several offset layers. The edges of such a skin panel can be joined in a different way to allow alignment and large-area contact between multiple structural components.

[0026] Similarly, providing the at least one stiffening component can include placing at least two layers on top of each other in a mold, subsequently bonding the layers together to form the stiffening component, placing the stiffening component onto the skin panel, and then bonding it together with the skin panel. The production of a stiffening component can, in particular, be carried out separately from the step of producing a planar base for a skin panel. The profile of the stiffening component can have a significantly greater curvature than the layers of the skin panel, since the stiffening component is required to create a more pronounced overhang. The stiffening component to be placed on the skin panel can have at least one planar flange that is to be brought into surface contact with the inner surface of the respective skin panel.If the stiffening component in question is designed like a stringer, the stiffening element could be limited to the use of a single such flange. However, if a larger, frame-like stiffening component is considered, two parallel, flat flanges can also be provided, both of which must be in surface contact with the skin panel. These are, of course, merely examples; any number of stiffening component variants with one or more flanges are conceivable. Depending on the design, a cavity can be enclosed between the stiffening component and the inner surface of the skin panel. After the at least one stiffening component is placed on the skin panel, it is bonded to it, creating a strong, monolithic connection.

[0027] Analogous to the layers of the skin panel, the layers of the at least one stiffening component can be arranged such that an offset is formed between the circumferential contours of two superimposed layers. The stiffening components of two adjacent or to-be-joined structural components can therefore be brought into stepped surface contact, just like the skin panels themselves, in order to subsequently be welded together.

[0028] It should also be noted that the structural component can be manufactured in one piece, consisting of a skin panel and a stiffening component. This eliminates the need to separately join the skin panel to the at least one stiffening component.

[0029] In a particularly preferred embodiment, the at least one stiffening component is arranged offset from the skin panel such that one end of the stiffening component projects over an edge of the skin panel and an opposite end of the stiffening component is spaced from an opposite edge of the skin panel. Two structural components that are brought into surface contact at their edges thus allow for a partial overlap of the surface contact between the structural components and the stiffening component in question. This allows the stiffening component to be welded directly to another skin panel of the adjacent structural component, resulting in an even better overlap of the fiber orientations of all the components involved.

[0030] In a further advantageous embodiment, the structural components are manufactured directly in a handling tool. The handling tool can simplify handling for an automated or semi-automated joining process, since the handling tool can be configured to grip a suitable device and the structural component itself does not need to be gripped.

[0031] It goes without saying that, alternatively, the individual structural components can also be grasped without such a handling tool, for example by using suction cups or the like, which create a temporary force-fit connection with the relevant outer surface of the skin area.

[0032] In a particularly preferred embodiment, individual surface segments with a closed perimeter, especially as a fuselage section of a vehicle, are provided by arranging several structural components in series. For example, the formation of a closed chain of structural components can result in a closed circumferential contour. This would be particularly helpful for manufacturing an elongated fuselage of an aircraft. Depending on the width of the individual structural components, a closed circumferential contour can provide a type of ring that represents a defined axial fuselage section. On the other hand, arranging several structural components in series in the form of a matrix can lead to the formation of a half-shell or at least a larger skin segment of a vehicle body.

[0033] The method can further include arranging a support for the defined placement of individual structural components, as well as the successive placement of individual structural components onto the support and the subsequent bonding of adjacent structural components. The support can preferably be adjusted between a retracted and an extended position. It is preferred if, in an extended position, the support provides support points for the individual structural components in the desired final position and, in the retracted position, allows movement of the support within a finished structural section. For example, in the production of an aircraft fuselage section, the support could be compacted in a radial direction, thereby reducing the size of a circumferential contour formed by several support points.After the sequential placement and material bonding of the individual fuselage components, particularly of a continuous fuselage section, the support could be compacted and removed from the completed fuselage section, either to manufacture another fuselage section or to be completely removed from the fuselage area. This allows for fully automated production of at least one structural section. The support can be designed as a placement frame.

[0034] The individual structural components can be handled using a multi-axis robot, a suitably adapted multi-axis guide, or similar equipment. This allows an automated device to move each structural component to a designated location and place it onto corresponding support points on the carrier or support frame. An additional device for heating a joining zone can be provided, which can be integrated either into the carrier, the handling tool, or separately into the joining zone.

[0035] The fiber-reinforced plastic used to provide the skin panels and / or stiffening component has a matrix material consisting of a thermoset resin with embedded reinforcing fibers. All common resins, especially thermosetting ones, are suitable as matrix materials. The specific matrix material can depend on the intended use of the structural component. The reinforcing fibers can also be selected according to the intended use and may include, for example, carbon fibers, glass fibers, aramid fibers, Kevlar fibers, and the like.

[0036] It is understood that the individual layers of a skin panel and / or a stiffening component can be joined by curing together.

[0037] When using a thermoset with reinforcing fibers, the material-bonded joining of two structural components can include gluing.

[0038] In a further advantageous embodiment, the fiber-reinforced plastic for providing the skin panels and / or the stiffening component can comprise a matrix material made of a thermoplastic polymer with embedded reinforcing fibers. Suitable thermoplastic materials for the matrix can include PPS (polyphenylene sulfide), PEEK (polyetheretherketone), PEKK (polyetherketoneketone), or others. The use of a thermoplastic material further increases the flexibility in the design and outfitting of the structural components, as the addition and integration of outfitting elements by local fusion is significantly easier compared to the production of a resin-infiltrated preform from dry or pre-impregnated fibers.

[0039] Similarly, in the production of structural components from a thermoplastic material, the material-bonded joining of structural components can include welding by at least locally heating a joining zone of the structural components.

[0040] The invention further relates to a vehicle with a structural section consisting of several structural components and manufactured according to a method as described above.

[0041] The vehicle can be an aircraft. Furthermore, the structural section can be a closed fuselage segment of an aircraft fuselage. BRIEF DESCRIPTION OF THE FIGURES

[0042] Further features, advantages, and applications of the present invention will become apparent from the following description of the exemplary embodiments and the figures. All features described and / or illustrated, individually and in any combination, constitute the subject matter of the invention, irrespective of their composition in the individual claims or their cross-references. In the figures, the same reference numerals denote identical or similar objects. Fig. Figures 1a to 1c show the construction of a fuselage component with several details in different representations. Fig. Figure 2 shows a fuselage component as an exemplary functional unit. Fig. Figures 3a to 3b show the automated production of a fuselage section. Fig. Figure 4 shows the possibility of modifying a fuselage by inserting a different number of additional fuselage sections. Fig. Figure 5 shows different functional units as fuselage components. DETAILED PRESENTATION OF EXEMPLARY EXECUTION FORMS

[0043] Fig. Figure 1 shows an exemplary skin panel 2, which in this example is made of several layers 4a, 4b, and 4c of a thermoplastic, fiber-reinforced polymer. For the sake of simplicity, the three layers 4a, 4b, and 4c, with identical dimensions, are shown stacked on top of each other and welded together. Each layer 4a, 4b, and 4c has a circumferential contour 6a, 6b, and 6c, which, due to the exemplary rectangular shape, is composed of four edges. The layers 4a, 4b, and 4c are slightly offset from each other, forming a stepped structure at one edge 8 of the skin panel 2. The edge 8 is therefore scarfed.

[0044] In Fig. 1b, skin field 2 is removed Fig. Figure 1a shows an example of a skin panel 2, on the inner surface 10 of which two stiffening components 12 and 14 are arranged. These are also made of a thermoplastic, fiber-reinforced material. In this illustration, both stiffening components 12 and 14 have two flanges 16 and 18, and 20 and 22, respectively, with which the stiffening components 12 and 14 are in direct, flush contact with the inner surface 10 of the skin panel 2. The stiffening components 12 and 14 are welded to the skin panel 2 at these points.

[0045] The stiffening component 14 has, by way of example, a recess 24 through which the stiffening component 12 extends. As can be seen from the curvature, the stiffening component 14 can be part of a rib, while the stiffening component 12 can be part of a longitudinal stiffening element (stringer). However, the recess 24 is not strictly necessary. The stiffening components 12 and 14 can also be manufactured together in a coordinated manner. For example, both stiffening components 12 and 14 could be integrally manufactured using injection molding and have an integral intersection point.

[0046] The individual stiffening components 12 and 14 can also be made from several layers 26a, 26b and 26c, which are positioned offset from each other. This can be implemented in an analogous manner to the layers 4a, 4b and 4c of the skin field 2.

[0047] Furthermore, the stiffening components 12 and 14 are positioned with a slight offset on the inner side 10, such that a portion of each stiffening component 12 and 14 projects beyond the edge 8 of the skin field 2 or is spaced apart from it on the inner side 10. Thus, stiffening component 12 forms an overhang 28 and a defect 30 on the opposite side of stiffening component 12. Similarly, stiffening component 14 forms an overhang 32 and a defect 34 on the opposite side. Overhangs 28 and 32 should be dimensioned correspondingly to the respective defects 30 and 34.

[0048] The advantages gained through this are reflected in the Fig. Figure 1c illustrates this. The stepped or scarfed designs of the edges 8 of several skin panels 2 can be executed in a corresponding manner, so that the edges 8 of two skin panels 2 lie flush against each other, utilizing the stepped structure. Since fibers extend in each of the individual layers 4a, 4b, and 4c, these can be brought into overlap with adjacent layers 4a, 4b, and 4c by means of the stepped design.

[0049] Furthermore, the overhangs 28 can be used to create an overlap with an adjacent skin field 2. This can positively influence the arrangement of the fiber structure and the size of the connection area. Consequently, a structural section 37 can be produced from several practically monolithic structural components 36 that are joined together.

[0050] Fig. Figure 2 shows a larger structural component 38, which can be used as a finished, functional module for manufacturing a structural section. This structural component 38 can extend over a larger area than, for example, the structural component in Fig. 1a extend. However, several stiffening components 12 and 14 are also provided here, which stiffen the underlying skin field 2.

[0051] Structural component 38 features, for example, cutouts 40 for emergency exits and window cutouts 42. Analogous to the one in Fig. In the example shown in Figure 1b, the edges 8 can have a step-like structure. Furthermore, a projection 44 and a recess 46 are provided in the plane of the skin field 2, which are largely corresponding to each other. When connected to a structural component that follows in the axial direction (not shown), an overlap can be created, as shown in Figure 1b. Fig. 1c already shown. A particular advantage lies in the subdivision of a larger structural section into several structural components 38, which are easier to manufacture and handle. Furthermore, the use of a fiber-reinforced thermoplastic material allows for simplified one-piece construction of stiffened areas of the structural component. It is also conceivable that the structural component 38 could already consist of several structural components 36 of the Fig. 1a to 1c is composed of.

[0052] Fig. 3a and Fig. Figure 3b schematically shows necessary or advantageous equipment for the fully automatic production of a fuselage section. Fig. Figure 3a initially shows a transport carrier 48, which could be used to transport and temporarily store several structural components 36. This is merely an example; vertically oriented carriers or holding frames are also conceivable, enabling transport between different production departments. The structural components 36 stored here can, for example, be gripped by a multi-axis robot 50 using a gripping tool 52. As in Fig. As shown in Figure 3a, several structural components 36 can be joined together to produce circumferential structural sections 55 of a hull 54. These could also be referred to as ring-shaped shell sections, circumferential segments, or hull segments.

[0053] The individual structural components 36 are heated at least locally in the joining zones, so that the thermoplastic material containing reinforcing fibers melts and two contacting structural components 36 are bonded together. The joining area between the structural components 36 can be heated by inductive methods, the use of heating resistors, by transmission, or other methods.

[0054] To place the individual structural components 36 in a predefined position, a carrier 56 designed as a placement frame can be used, which provides several holding points 58 on which the individual structural components 36 can be placed. In particular, the joining zones should be arranged on adjacent holding points 58 so that at least the joining zones are located in a spatially very precisely defined position in order to be connected there to the adjacent parts. Of course, the use of several multi-axis robots 50, which simultaneously position several structural components 36 accordingly, would also be conceivable.

[0055] It may still be advantageous not only to select the smaller structural components 36 from Fig. not only to use 1a to 1c, but also larger structural components 38, which could be composed of several smaller structural components 36.

[0056] In addition to simplified, automatable manufacturing and the production of an essentially monolithic hull 54 by using a multitude of individual, easily handled and automatically manufactured structural components 38, the method according to the invention also offers further advantages. Fig. Figure 4a shows three aircraft fuselages 60, 62 and 64 as examples, which have identical core dimensions, but differ in length and equipment of a middle fuselage section.

[0057] The aircraft fuselage 60, for example, is the shortest of the three fuselages and has only three fuselage sections 66, 68, and 70 – at least in this illustration. These can be manufactured using the method according to the invention or by other means. The aircraft fuselage 62, shown vertically in the middle, differs from the shorter aircraft fuselage 60 by an extension of the central fuselage section 68, exemplified by a total of six circumferential structural sections 55 made of adjoining structural components 36. These can be manufactured using the method according to the invention and serve to modify the fuselage 60 to achieve a greater length and a greater number of passenger seats. These additionally inserted shell sections 55 can be designed to withstand the load and, where feasible, allow for the individualization of an aircraft.Another example is shown with the underlying aircraft fuselage 64, which has only two additional circumferential shell sections 55 and thus a shorter extension. Overall, the method according to the invention can therefore also be used in addition to other, established methods.

[0058] As in Fig. As shown in Figure 5, the inventive method also allows for simple individualization of a structural section 72 by adapting emergency exit hatches 74, arranging windows 76 and the like.

[0059] It should be further noted that "having" does not exclude any other elements or steps, and "a" or "an" does not exclude a plurality. It should also be noted that features described with reference to one of the above embodiments may also be used in combination with other features of other embodiments described above. Reference numerals in the claims are not to be considered as a limitation.

Claims

[1] Method for manufacturing a structural section (37, 55, 72) of a vehicle, comprising the steps: - Providing several separate skin panels (2) made of a fiber-reinforced plastic with an inner surface (10), an outer surface and a rim (8) surrounding the respective skin panel (2), - Arranging at least one stiffening component (12, 14) made of a fiber-reinforced plastic on each skin panel (2) on the respective inner side (10), - material-bonded connection of the respective at least one stiffening component (12, 14) with the relevant skin fields (2) to form a structural component (36, 38), - Arranging at least two structural components (36, 38) on a support (56) such that at least areas of the edges (8) of the relevant skin fields (2) of the structural components (36, 38) are in surface contact, and - material-bonding connection of the areas of the edges (8) that are in surface contact with each other. [2] Method according to claim 1, where providing a skin field (2) comprises placing at least two layers (4a, 4b, 4c) on top of each other in a forming tool and joining the layers (4a, 4b, 4c) together in a material-bonded manner, wherein each layer (4a, 4b, 4c) has a circumferential contour (6a, 6b, 6c) and an offset is formed between the circumferential contours (6a, 6b, 6c) of two superimposed layers (4a, 4b, 4c). [3] Method according to claim 1 or 2, wherein the provision of the at least one stiffening component (12, 14) comprises placing at least two layers (26a, 26b, 26c) on top of each other in a mold, the subsequent joining of the layers (26a, 26b, 26c) to form the stiffening component (12, 14), and placing the stiffening component (12, 14) onto the skin field (2) and the subsequent joining of the layers. [4] Method according to claim 3, wherein each layer (26a, 26b, 26c) of the stiffening component has a circumferential contour and an offset is formed between the circumferential contours of two superimposed layers (26a, 26b, 26c). [5] Method according to one of the preceding claims, wherein the at least one stiffening component (12, 14) is arranged in an offset to the skin field (2) such that one end of the stiffening component (12, 14) projects over an edge (8) of the skin field (2) and an opposite end of the stiffening component (12, 14) is spaced apart from an opposite edge of the skin field (2). [6] Method according to one of the preceding claims, further comprising the joining of several structural components (36, 38) to form individual surface segments (55) with a closed perimeter. [7] Method according to any one of the preceding claims, furthermore, the arrangement of a support (56) for the defined placement of individual structural components (36, 38), and successive placement of individual structural components (36, 38) onto the support (56) and the subsequent material-bonding connection of adjacent structural components (36, 38). [8] Method according to claim 7, wherein at least the placing of the structural components (36, 38) onto the carrier (56) is carried out by a multi-axis robot (50). [9] Method according to one of the preceding claims, wherein the fiber-reinforced plastic for providing the skin panels (2) and / or the stiffening component (12, 14) comprises a matrix material of a thermoset with reinforcing fibers embedded therein. [10] Method according to claim 9, wherein the material-bonding joining of two structural components (36, 38) comprises gluing. [11] Method according to one of the preceding claims, wherein the fiber-reinforced plastic for providing the skin panels (2) and / or the stiffening component (12, 14) comprises a matrix material made of a thermoplastic polymer with reinforcing fibers embedded therein. [12] Method according to claim 11, wherein the material-bonding joining of structural components (36, 38) comprises welding by at least local heating of a joining zone of the structural components (36, 38). [13] Vehicle with at least one structural section (37, 55, 72) consisting of several structural components (36, 38) and manufactured according to a method according to claims 1 to 12. [14] Vehicle according to claim 13, wherein the vehicle is an aircraft. [15] Vehicle according to claim 14, wherein the structural section (37, 55, 72) is a closed fuselage segment (55) of an aircraft fuselage (54).

Images