Method for manufacturing a fuselage section, fuselage section and aircraft or spacecraft

Abstract

Method for manufacturing a fuselage section (1), in particular for an aircraft or spacecraft, comprising the following process steps: Welding a skin section (2) containing a thermoplastic material to a frame (3) containing a thermoplastic material in the area of ​​a predetermined welding zone (4); and Connecting a fastening element designed as a crack stopper (5) to the skin section (2) and the frame (3) in the area of ​​the welding zone (4); wherein the fastening element (5) is bonded to the skin section (2) and / or the frame (3).

Description

[0001] The present invention relates to a method for manufacturing a fuselage section, in particular for an aircraft or spacecraft, as well as a fuselage section and an aircraft or spacecraft with such a fuselage section.

[0002] Although applicable to any type of fuselage, the present invention and the underlying problem will be explained in more detail with regard to an aircraft fuselage.

[0003] Modern aircraft fuselages are often constructed using fiber-reinforced plastics, among other materials. An aircraft fuselage typically consists of skin sections and the stringers and frames that support these skin sections. The stringers run longitudinally along the fuselage, while the frames run circumferentially.

[0004] Riveting methods are typically used to fasten the frames, in particular, to the skin sections. For this purpose, connecting elements, so-called clips, and frame stabilizing elements, so-called cleats, are riveted to both the skin and the frame. Such a method and arrangement are described, for example, in DE 10 2014 103 438 A1.

[0005] Document US 2010 / 0 237 194 A1 describes a method in which a connecting plate is attached to an aircraft skin by means of a combination of welding and bolting.

[0006] Document US 2006 / 0 226 287 A1 describes a building plate with a skin having at least one offset surface.

[0007] Against this background, the object of the present invention is to provide an improved method for manufacturing a hull section and a corresponding improved hull section.

[0008] According to the invention, this problem is solved by a method for producing a hull section with the features of claim 1 and / or by a hull section with the features of claim 7.

[0009] Accordingly, the following is planned: - A method for manufacturing a fuselage section, in particular for an aircraft or spacecraft, comprising the following process steps: welding a skin section containing a thermoplastic material to a frame containing a thermoplastic material in the area of ​​a predetermined welding zone; and connecting a fastening element designed as a crack stop to the skin section and the frame in the area of ​​the welding zone; wherein the fastening element is materially bonded to the skin section and the frame. - A fuselage section, in particular for or in an aircraft or spacecraft, in particular manufactured according to a method according to the invention, comprising: a skin section; and a frame, wherein the skin section and the frame each contain a thermoplastic material and are welded together in a weld zone, wherein a fastening element connected to the skin section and the frame is provided in the area of ​​the weld zone, which is designed as a crack stop; and wherein the fastening element is formed to be materially bonded to the skin section and / or the frame. - An aircraft or spacecraft, with a fuselage comprising a fuselage section according to the invention or a fuselage section produced according to a method according to the invention.

[0010] The underlying idea of ​​the present invention is that a fuselage assembly is formed with a skin section and a frame, each containing a thermoplastic material and welded together in a weld zone for fastening. Additionally, according to the invention, fastening elements designed as crack stoppers are provided in the weld zone, which are connected to the frame and the skin section.

[0011] The skin section and the frame can be designed in a variety of ways and may contain a thermoplastic material. For example, it could be a fiber-reinforced composite material with a thermoplastic matrix or a multi-layered structure with at least one thermoplastic layer. In particular, the skin section and the frame do not need to be made entirely of thermoplastic material. Rather, only portions, only a single layer, or only a section of the weld zone can contain the thermoplastic material. It would also be conceivable to provide a thermoplastic top layer or an additional material of a different type for welding to the skin section and / or the frame.

[0012] A weld zone can have a variety of forms. For example, it can be a planar weld zone, a linear weld seam, an area surrounded by a weld seam, an area with a large number of weld spots or a weld pattern, or the like.

[0013] Thermoplastic welding joins the skin to the frame in the weld zone. This offers the advantage of joining the skin section and the frame in a comparatively simple and quick welding process, instead of the usual labor-intensive riveting. This minimizes manufacturing costs and lead times.

[0014] Depending on the type of welding process used, thermoplastic welding can be carried out with or without filler materials, thermoplastic adhesives or the like.

[0015] The internal pressure load in an aircraft fuselage can cause the frame to peel away from the skin section. Therefore, riveted connections are typically required in aviation for certification, as they prevent such peeling. According to the invention, the additional fastening elements, designed as crack stoppers, serve to increase the damage tolerance of the connection and prevent crack formation and / or crack propagation, thus preventing detachment. Advantageously, this allows the welded connection to be certified for aviation use.

[0016] According to the invention, the fastening elements can be selectively positioned locally at sections of the weld zone that are subject to higher stress and therefore more prone to cracking, particularly in the area of ​​a weld zone start or end. This advantageously prevents crack initiation from the outset.

[0017] Furthermore, unlike conventional riveted connections, the fasteners according to the invention can be placed individually and / or at comparatively large or generous intervals, which are sufficient to prevent crack propagation, since the actual fastening or connection is provided by the welded joint. Therefore, significantly fewer fasteners are required than with conventional riveted connections. Overall, producing the welded joint together with inserting the fasteners thus requires considerably less labor. The solution according to the invention therefore contributes to weight and cost optimization.

[0018] Advantageous embodiments and improvements of the invention are found in the dependent claims.

[0019] According to a preferred embodiment, the frame is formed with a through-hole at which the weld zone is interrupted. The fastening element is positioned in the area of ​​a weld zone extension or termination adjacent to the through-hole. In the area of ​​the weld zone extension or termination, a force flow often concentrates under load, leading to locally increased stresses. Thus, the fastening element positioned in the area of ​​the weld zone extension or termination reinforces that part of the weld zone subjected to the highest loads. In this way, even the initiation of a crack and therefore delamination is effectively prevented.

[0020] According to a non-inventive embodiment, the fastening element is connected to the skin section and the frame by a positive and / or non-positive connection. This can be achieved, for example, by rivets, screws, or other types of bolted connections. A positive connection on one side and a non-positive connection on the other side of the fastening element is also conceivable. Alternatively or additionally, combinations of a non-positive or positive connection with a material-bonded connection are also conceivable. Preferably, the connection of the fastening element to the skin section and the frame can be made after welding in the weld zone, so that the welding process is advantageously not affected.

[0021] According to the invention, the fastening element is bonded to the skin section and the frame in a material-bonded manner. In this way, more complex mechanical fastening steps can be advantageously avoided during manufacturing.

[0022] According to an advantageous embodiment, the fastener is melted into the weld zone during welding to create a metallurgical bond between the skin section and the frame. Advantageously, this eliminates the need for an additional step to connect the fastener to the skin section and the frame. For example, the fastener could be a pin. In particular, this pin could be inserted into a designated recess in the weld zone before welding. During the welding process, the thermoplastic material of the frame and skin section melts and bonds with the material of the fastener, creating a metallurgical bond.

[0023] It would also be conceivable to design the fastening element with an undercut, for example a head and / or a foot and / or a waist, which is filled with molten metal in the weld pool during welding, so that in addition to a material-bonded connection, a form-fit connection of the fastening element with the frame and / or the skin is formed during welding.

[0024] Furthermore, it is conceivable to countersink the fastener in the weld zone or in the thermoplastic material, so that it is encapsulated, particularly within the weld zone. The fastener therefore does not need to penetrate the material completely. In this way, for example, contact corrosion protection can be provided simultaneously with the connection for a carbon fiber reinforced plastic material used in the skin section and / or ribs. Advantageously, other contact corrosion protection measures can thus be eliminated or reduced.

[0025] According to one embodiment, the skin section and / or the frame is provided as a fiber composite component. In this case, the fastening element displaces fibers of the fiber composite component non-destructively during welding in the weld zone. For this purpose, the fastening element can be inserted into the weld zone in such a way that fibers are displaced but not cut. In this way, the fiber structure is advantageously not damaged.

[0026] According to one embodiment of a hull assembly, a stringer is provided that runs transversely to the frame and is connected to the skin section. In a crossing area, the frame has a through-hole for the stringer, at which the weld zone is interrupted. The stringer also contains a thermoplastic material and is, for example, also welded to the skin section. In one embodiment, a fastening element connected to the skin and the stringer can also be provided in a weld zone of the stringer. Advantageously, this creates a hull structure stiffened by stringers and frames, which can be manufactured with comparatively low production costs and exhibits high damage tolerance.

[0027] According to a particularly advantageous embodiment, the fastening element is arranged in the area of ​​a run-out or start of the weld zone adjacent to the through-hole. Since the start or start and the run-out or end are often the most highly stressed sections of the weld zone, the highest stress often occurs there. The fastening element in the area of ​​the run-out or start effectively prevents crack initiation and thus delamination or separation.

[0028] According to a non-inventive embodiment of the hull section, the fastening element is connected to the skin section and / or the frame by a positive and / or non-positive connection. For example, it is a fastening element in the form of a rivet or a bolt. Advantageously, standard tools used for conventional rivet or bolt connections can be employed for this purpose. Alternatively or additionally, combinations of non-positive, positive, and / or material-locking connections are also conceivable. Accordingly, a positive connection can be provided only with the skin section or only with the frame, while a non-positive connection is provided on the other side of the fastening element.It would also be conceivable that only the skin section or the frame is positively and / or force-fit connected to the fastening element, with a different connection, for example a material-fit connection, being provided on the other side of the fastening element.

[0029] According to the invention, the fastening element is formed by a material bond with the skin section and / or the frame. In particular, it can be a pin fused into the weld zone. The fastening element can, in particular, be recessed within the weld zone. In this way, the fastening element is advantageously encapsulated, thus preventing contact corrosion problems.

[0030] In one embodiment, the pin may have an undercut and / or waist and / or the like to additionally form a positive fit.

[0031] According to one embodiment, the skin section and / or the frame is designed as a fiber composite component, wherein fibers of the fiber composite component are displaced non-destructively in the weld zone by the fastening element. In this way, the fiber structure is advantageously undamaged despite the fastening element.

[0032] According to an advantageous embodiment, a plurality of fastening elements are provided. The fastening elements are arranged at predetermined intervals within the weld zone. This predetermined interval is preferably greater than 10 cm. In particular, the interval is greater than 50 cm. A distance between 50 cm and 1 m is especially preferred. Advantageously, this method requires comparatively few fastening elements, which significantly reduces manufacturing costs. Nevertheless, with careful positioning, this interval is sufficient to effectively prevent cracking, delamination, or separation. For this purpose, the fastening elements are preferably positioned in the most highly stressed sections of the weld zone.

[0033] The above configurations and further developments can be combined with each other as appropriate. In particular, all features of the process for manufacturing the fuselage section can be transferred to the fuselage section, and vice versa.

[0034] Further possible embodiments, developments, and implementations of the invention also include combinations of features of the invention described previously or subsequently with regard to the exemplary embodiments, even if not explicitly mentioned. In particular, the person skilled in the art will also add individual aspects as improvements or additions to the respective basic form of the present invention.

[0035] The invention is explained in more detail below with reference to an exemplary embodiment and the accompanying figure. Unless otherwise stated, the same reference numerals denote identical or functionally equivalent components. Fig. Figure 1 shows a schematic cross-sectional view of a fuselage section. Fig. Figure 2 shows a perspective view of an aircraft or spacecraft.

[0036] The in Fig. The hull section 1 shown has a skin section 2 and a frame 3.

[0037] Skin section 2 corresponds to the round shape of a Fig. The hull 12 shown in section 2 is curved. Accordingly, frame 3, to which skin section 2 is attached, also has a corresponding curvature on its radial outer side.

[0038] The fuselage section 1 also has several stringers 8, which are essentially aligned in the longitudinal direction of the fuselage 12, which extends into the plane of the drawing in the cross-sectional view shown.

[0039] The stringers 8 cross the frame 3 at intersection points 9, for which passage openings 6 are provided in the frame 3. The passage openings 6 are shown here in an arched shape for illustrative purposes only.

[0040] Both skin section 2 and frame 3 are exemplary examples made of a fiber composite material with a thermoplastic matrix. In particular, skin section 2 can be manufactured using a tape laying process.

[0041] The attachment of skin section 2 to frame 3 is realized by welding zones 4, in which skin section 2 is thermoplastically welded to frame 3.

[0042] The welding zones are interrupted in the area of ​​the passage openings 6.

[0043] Within each of the welding zones 4, additional fastening elements 5 are provided, which are designed as crack stoppers. These are each arranged in the area of ​​a start or end 7 of a welding zone 4.

[0044] A start or end point refers to the end of a continuous weld zone. For example, it can be the beginning or end of a weld seam.

[0045] The welding zone 4 preferably has a width that exceeds the width or diameter of the fastening element 5.

[0046] To manufacture fuselage section 1, skin section 2 and frame 3 are thermoplastically welded together in the predetermined welding zones 4. This can be done, for example, by conduction welding, ultrasonic welding, or hot gas welding. Other welding processes, such as laser beam welding, are also conceivable.

[0047] For example, in hot gas welding, the surfaces to be joined are heated to welding temperature with hot air and welded under pressure. Specifically, the welding process includes the following steps: preparing the welding surfaces in the area of ​​the predetermined welding zone; heating the welding zone; joining the skin section to the stringer while simultaneously applying welding pressure; cooling the weld while maintaining welding pressure; and relieving the welded material of the welding pressure.

[0048] Furthermore, in the welding zones 4, the fastening elements 5 are inserted in the area of ​​a start or end 7 of each welding zone 4.

[0049] For example, the fasteners 5 can be inserted into weld zone 4 after welding during manufacturing. In this case, the fasteners 5 could be rivets inserted into the solidified weld zone 4. Alternatively, they could also be various conventional types of bolts.

[0050] As a further alternative according to the invention, the fastening elements 5 can be melted into the weld zone 4 during welding. For example, in this case, the fastening elements 5 can be pins that are melted into the weld zone during welding. The pins can be positioned within the predetermined weld zone 4 before welding. In particular, corresponding positioning recesses, for example bores, can be provided in the predetermined weld zone before welding. When the thermoplastic material melts, the fastening elements are melted into or embedded in it and thus metallurgically bonded to the thermoplastic material.

[0051] The fibers of the fiber composite material of the skin section and the frame are not damaged in the process. Rather, fiber displacement by the fastening elements is possible in the melt bath, so that advantageously no fiber damage occurs.

[0052] The fastening elements themselves preferably remain fixed, so that after the thermoplastic melt has solidified, they are embedded in their original shape in the intended position and are bonded to the skin section 2 and the frame 3.

[0053] The welded connection between the frame and the skin section is realized in a damage-tolerant manner by means of the fastening elements 5. In the event of detachment, any crack or crack propagation is always stopped at the nearest fastening element 5. In this way, the fuselage section 1 according to the invention can be certified for aviation despite a welded connection between the skin section and the frame.

[0054] The stringers 8, for example, are designed as omega stringers and are also made of a fiber composite material with a thermoplastic matrix. Accordingly, the stringers are preferably also welded to the skin section.

[0055] Fig. Figure 2 shows an aircraft or spacecraft 11 in the form of a passenger aircraft.

[0056] The aircraft or spacecraft 11 has a fuselage 12 which has at least one according to Fig. 1 has a formed hull section 1.

[0057] Preferably the entire hull 12 is in the according to Fig. 1. It is manufactured in the construction method described above and therefore contains a large number of such hull sections.

[0058] Although the present invention has been described here with reference to preferred embodiments, it is not limited to these, but can be modified in many ways.

[0059] For example, additional fastening elements 5 can also be arranged within a welding zone 4 between the inlet and outlet 7. This is particularly useful if the stringers 8 or the through-hole 6 are spaced more than 1 m apart in order to stop any cracks at an early stage. The additional fastening elements 5 are then preferably also arranged at a distance of between 50 cm and 1 m from the fastening elements provided at the inlet or outlet 7.

[0060] Instead of Omega stringers, other types or geometries of stringers 8 can also be used. Accordingly, the through-holes 6 of the rib 3 do not necessarily have to be arc-shaped, but can also have a different shape adapted to the stringer geometry.

[0061] Furthermore, to realize the concept according to the invention, the stringers do not necessarily also need to contain a thermoplastic material or be welded to the skin.

[0062] Only frame 3 and skin section 2 require thermoplastic material to form the weld zone. However, it would also be possible to use a thermoplastic material only locally, as a layer, outer coating, or possibly as an additive material to join the frame to the skin section, so that the frame and / or skin section might only contain the thermoplastic material in the weld zone area once welded. Reference symbol list 1 hull section 2 skin section 3 frame 4 welding zones 5 Fastening element 6 Passage opening 7 Outlet or approach 8 Stringer 9 Intersection area 10 distance 11 aircraft 12 Hull

Claims

[1] Method for manufacturing a fuselage section (1), in particular for an aircraft or spacecraft, comprising the following process steps: Welding a skin section (2) containing a thermoplastic material to a frame (3) containing a thermoplastic material in the area of ​​a predetermined welding zone (4); and Connecting a fastening element designed as a crack stopper (5) to the skin section (2) and the frame (3) in the area of ​​the welding zone (4); wherein the fastening element (5) is bonded to the skin section (2) and / or the frame (3). [2] Method according to claim 1, characterized by , that the frame (3) is formed with a through-hole (6) at which the weld zone (4) is interrupted, wherein the fastening element (5) is arranged in the area of ​​an outlet or projection (7) of the weld zone (4) adjacent to the through-hole (6). [3] Method according to one of claims 1 or 2, characterized by , that the fastening element (5) is melted into the welding zone (4) during welding to create a material-bonded connection with the skin section (2) and the frame (3). [4] Method according to claim 3, characterized by , that the skin section (2) and / or the frame (3) are provided as a fiber composite component, wherein the fastening element (5) displaces fibers of the fiber composite component non-destructively during welding in the welding zone (4). [5] Fuselage section (1), in particular for or in an aircraft or spacecraft, comprising: a section of skin (2); and a frame (3), wherein the skin section (2) and the frame (3) each contain a thermoplastic material and are welded together in a welding zone (4), wherein in the area of ​​the sweat zone (4) a connection with the skin section (2) and A fastening element (5) connected to the frame (3) is provided, which is designed as a crack stop; and wherein the fastening element (5) is formed to be materially bonded to the skin section (2) and / or the frame (3). [6] Hull section according to claim 5, characterized by , that a stringer (8) running transversely to the frame (3) and connected to the skin section (2) is provided, wherein the frame (3) has a passage recess (6) for the stringer in a crossing area (9) at which the weld zone (4) is interrupted. [7] Hull section according to claim 6, characterized by , that the fastening element (5) is arranged in the area of ​​an outlet or projection (7) of the welding zone (4) adjacent to the through-hole (6). [8] Hull section according to at least one of claims 5 to 7, characterized by, that the fastening element (5) is formed in the form of a pin fused into the weld zone (4). [9] Hull section according to at least one of claims 5 to 8, characterized by , that the skin section (2) and / or the frame (3) is designed as a fiber composite component, wherein fibers of the fiber composite component are displaced non-destructively in the weld zone (4) by the fastening element (5). [10] Hull section according to any one of claims 5 to 9, characterized by that a plurality of fastening elements (5) is provided, wherein the fastening elements (5) are each arranged at a predetermined distance (10) from each other, in particular at a distance greater than 10 cm, preferably greater than 50 cm, particularly preferably between 50 cm and 1 m, in the welding zone (4). [11] Aircraft or spacecraft (11) comprising a fuselage (12) which has a fuselage section (1) according to any one of claims 5 to 10 and / or manufactured by a method according to any one of claims 1 to 4.

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