Method For Forming A Structural Component For An Airframe Of An Aircraft Or Spacecraft And Structural Component For An Airframe Of An Aircraft Or Spacecraft

Abstract

A method for forming a structural component for an airframe of an aircraft or spacecraft, includes: providing a prefabricated shell element comprising a thermoplastic substrate; and applying a stiffening structure to the shell element by additive manufacturing, wherein a plurality of continuous thermoplastic filaments filled with reinforcing fibers are continuously heated and three dimensionally formed such that the filaments are crossing and bonded to each other at a plurality of crossing points to form a three dimensional grid truss integrally formed on the shell element. A corresponding structural component and an aircraft or spacecraft including such a structural component are also described.

Claims

1. A method for forming a structural component for an airframe of an aircraft or spacecraft, the method comprising: providing a prefabricated shell element comprising a thermoplastic substrate; and applying a stiffening structure to the shell element by additive manufacturing, wherein a plurality of continuous thermoplastic filaments filled with reinforcing fibers are continuously heated and three dimensionally formed such that the filaments cross and bond to each other at a plurality of crossing points to form a three dimensional grid truss integrally formed on the shell element.

2. The method according to claim 1, wherein the filaments comprise a thermoplastic matrix material filled with continuous fibers.

3. The method according to claim 1, wherein prior to the application the filaments are at least locally heated to a melting temperature by a first laser beam for bonding with each other and/or with the shell element.

4. The method according to claim 1, wherein the thermoplastic substrate of the shell element prior to the application of the filaments is locally heated by a second laser beam for bonding with the filaments.

5. The method according to claim 1, wherein subsequent to the heating the filaments are consolidated by at least one ironing pad.

6. The method according to claim 1, wherein the filaments are arranged to together with the shell element form a truss section with a hat section, a hexagonal section or a T-section.

7. The method according to claim 1, wherein between two crossing points each filament forms an independent truss strut.

8. The method according to claim, wherein for applying the stiffening structure: a first filament is bonded to the shell element at a first end, then three dimensionally formed free standing along a predetermined trajectory of the grid truss and bonded to the shell element at a second end; and a second filament is bonded to the shell element at a first end, then three dimensionally formed along a predetermined trajectory of the grid truss crossing the first filament, bonded to the first filament, at a crossing point and bonded to the shell element at a second end.

9. The method according to claim 8, wherein a third filament and/or further filaments are three dimensionally formed along a predetermined trajectory of the grid truss crossing the first filament and the second filament and bonded to the first filament and to the second filament at respective crossing points until the grid truss is complete.

10. A structural component for an airframe structure comprising: a shell element comprising a thermoplastic substrate; and a stiffening structure comprising a three dimensional grid truss integrally formed on the shell element, wherein the grid truss comprises a plurality of continuous thermoplastic filaments filled with reinforcing fibers, which filaments are arranged crossing each other and bonded to each other at a plurality of crossing points and bonded to the shell element at contacting points or lines between the filaments and the shell element.

11. The structural component according to claim 10, wherein the filaments comprise a thermoplastic matrix material filled with continuous fibers.

12. The structural component according to claim 10, wherein the filaments are cranked at some of the plurality of crossing points and straight between the crossing points to form a three dimensional contour of the truss.

13. The structural component according to claim 10, wherein the filaments are arranged in a triangular grid configuration.

14. The structural component according to claim 10, wherein the filaments are arranged to together with the shell element form a truss section with a hat section, hexagonal section or T-section.

15. An aircraft or spacecraft comprising an airframe, wherein the airframe comprises a structural component according to claim 10.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0036] FIG. 1 schematically illustrates an exemplary conventional grid truss.

[0037] FIG. 2A shows a cross section of a conventional full shell stringer.

[0038] FIG. 2B shows a cut-out stringer internally known by the applicant.

[0039] FIG. 3 schematically illustrates a cross section of a structural component comprising a hexagonal grid truss according to an embodiment of the present invention.

[0040] FIG. 4 schematically illustrates a cross section of a structural component comprising a T-section grid truss according to an embodiment of the present invention.

[0041] FIG. 5 schematically illustrates a cross section of a structural component comprising a hat section grid truss according to an embodiment of the present invention.

[0042] FIG. 6 shows application of a filament to a shell element.

[0043] FIG. 7 shows free forming of a filament extending from a shell element.

[0044] FIGS. 8A-8D schematically illustrate steps of applying a stiffening structure according to an embodiment of the present invention.

[0045] FIG. 9 schematically illustrates a structural component according to an embodiment of the present invention.

[0046] FIG. 10 schematically illustrates an aircraft or spacecraft.

[0047] Although specific embodiments are illustrated and described herein, it will be appreciated by those of ordinary skill in the art that a variety of alternate and/or equivalent implementations may be substituted for the specific embodiments shown and described without departing from the scope of the present invention. Generally, this application is intended to cover any adaptations or variations of the specific embodiments discussed herein.

DETAILED DESCRIPTION

[0048] FIG. 3 schematically illustrates a cross section of a structural component 1 comprising a hexagonal grid truss 6 according to an embodiment of the present invention.

[0049] The structural component 1 is configured for an airframe structure, in for particular for a fuselage structure, of an aircraft or spacecraft and comprises a shell element 2 comprising a thermoplastic substrate 8 and a stiffening structure 3 comprising a three dimensional grid truss 6 integrally formed on the shell element 2. The grid truss 6 is formed by a plurality of continuous thermoplastic filaments 4, which are filled with reinforcing fibers, preferably continuous carbon fibers, and bonded to the shell element 2.

[0050] The filaments are arranged along the trajectory of the grid truss, which in the present example has a hexagonal cross-section. In other embodiments, other cross sectional shapes are possible, such as hat section, T-section or L-section.

[0051] The filaments 4 have predefined different orientations such that they cross each other and meet the shell element 2. Furthermore, at least some filaments 4 can be arranged offset to the others. In the present example, at least some of the crossing points 5 form nodes of the hexagonal cross-section and the filaments 4 are cranked at the crossing points 5 to form the three dimensional hexagonal contour of the grid truss 6.

[0052] At the crossing points 5, the filaments are bonded to each other, preferably by thermoplastic welding. Furthermore, the filaments 4 are bonded to the shell element 2 at contacting points 15 between the filaments 4 and the shell element 2. Preferably, the filaments 4 are also welded to the thermoplastic substrate 8 of the shell element 2.

[0053] The contact points 15 are arranged along two parallel lines of nodes of the hexagonal structure along the truss. The contacting points 15 can form an end of a respective filament 4, which is bonded to the shell element. Alternatively, a filament may be bonded to the shell element 2 along a line in a middle part and run meander-like on the trajectory of the grid truss 6, in particular along the complete length of the grid truss 6. With such an arrangement of the filaments, the shell element 2 forms one side of the hexagonal cross-section of the grid truss 6, which accordingly is formed integrally on the shell element 2.

[0054] FIG. 4 schematically illustrates a cross section of a structural component 1 comprising a T-section grid truss 6 according to an embodiment of the present invention.

[0055] According to this embodiment, the filaments 4 are arranged in such a manner that they follow the trajectory of a T grid truss. Accordingly, there is only one line of nodes of the T section forming the contact points 15 between the filaments 4 and the shell element 2.

[0056] FIG. 5 schematically illustrates a cross section of a structural component comprising a hat section grid truss 6 according to an embodiment of the present invention.

[0057] In contrast to the embodiment of FIG. 3, the grid truss 6 according to this embodiment has a trapezoid cross section, which is also called hat section. Similar to FIG. 3, two parallel lines of nodes form the contacting points 15 between filament 4 and shell element 2. However, there are only two additional nodes forming crossing points 5 of the filaments between the contacting points 15.

[0058] FIG. 6 shows an application of a filament 4 to a shell element 2.

[0059] The filament 4 is formed and applied to the shell element 2 by means of an additive layer manufacturing (ALM) head 16 mounted on a robotic arm 17. The shell element 2 is a prefabricated thermoplastic fiber reinforced panel, for example a skin panel, prefabricated by fast large scale technologies such as press forming or automated tape laying (ATL). The shell element comprises a thermoplastic substrate 8, e. g. a thermoplastic matrix.

[0060] According to the present invention, a stiffening structure 3 is applied to the prefabricated shell element 2 by means of additive manufacturing, wherein a plurality of filaments 4 are formed as an integral grid truss 6 on the shell element 2. The ALM head 16 therefore heats and melts the thermoplastic fiber reinforced filaments 4, which are preferably filled with a tow of continuous carbon fibers.

[0061] According to the present example, such filaments 4 are provided as a prefabricated filament coil 23 and heated by means of a first laser beam 7. However, filaments could also be conductively heated or even directly printed from raw fibers and a thermoplastic material, which includes heating.

[0062] The welding head 16 comprises a laser source 18 and a corresponding laser optic for generating and guiding the first laser beam 7 onto the filament 4. In addition, the laser source 18 is also used to provide a second laser beam 9 with another laser optic, which second laser 9 beam is used for heating and locally melting the substrate 8 for bonding with the filaments 4. In this way, the thermoplastic materials of the filament 4 and the substrate 8 can be easily bonded.

[0063] Furthermore, the ALM head 16 comprises a plurality of ironing pads 10 provided for consolidation of the filament 4. FIG. 6 shows a first state of the ALM head 16, in which one ironing pad 10 is pressing the molten filament 4 onto the molten substrate 8 in order to create a strong connection.

[0064] The ALM head 16 for example comprises two such ironing pads 10 which can be used together or independently from each other for consolidation. In the first state shown in FIG. 6, only one of the ironing pads 10 is in use for pressing the filament 4 against the substrate 8 and the other pad 10 is pivoted to a side.

[0065] Furthermore, the ALM head 16 comprises guide rollers 19 to guide the filament 4 in a molten state in the desired direction and to the desired location.

[0066] FIG. 7 shows free forming of a filament 4 extending from a shell element 2.

[0067] In order to realize a free forming of the filament, both ironing pads 10 are used for consolidation in a second state to steer the filament in a desired direction. The consolidated material thus will be standing free in its desired position as the ALM head 16 moves along.

[0068] The second laser beam 9 can be switched of at this stage. The filament 4, which is molten by the first laser beam 7, can be formed starting from the substrate 8 along a three dimensional working path of the ALM head 16 moved by the robotic arm 17. According to the present invention, the ALM head is moved along a predetermined trajectory of a grid truss 6. The three dimensional form of the filament is maintained by consolidation of the filament with both ironing pads 10 contacting the filament from two opposing sides. In this way, the filament 4 can be built free standing. A three dimensional grid truss 6 can be formed out of a plurality of such filaments 4 crossing and bonded to each other at a plurality of crossing points.

[0069] FIGS. 8A-8D schematically illustrate steps of applying a stiffening structure 3 according to an embodiment of the present invention.

[0070] In particular, FIGS. 8A to 8D show a sequence of displacement of a plurality of filaments 4a, 4b, 4c, . . . , 4n on a shell element 2 for forming an integral grid truss 6.

[0071] The ALM head 16 is configured to displace thermoplastic filaments 4 filled with reinforcing fibers in such a manner to manufacture a strut of the grid truss 6 by one filament 4 containing all the needed fibers for reinforcement of the strut.

[0072] In this way, a structural component 1 for an airframe of an aircraft or spacecraft can be formed according to a method of the present invention. Therefore, in a first step, a prefabricated shell element 2 comprising a thermoplastic substrate 8 is provided. In the present example, the shell element 2 is configured as a fuselage skin panel. In a second step, a stiffening structure 3 is applied to the shell element 3 by additive manufacturing, wherein a plurality of continuous thermoplastic filaments 4a, 4b, 4c, . . . , 4n filled with reinforcing fibers are continuously heated and three dimensionally formed by an ALM head, as explained with regard to FIGS. 6 and 7. The filaments 4a, 4b, 4c, . . . , 4n are formed crossing and bonded to each other at a plurality of crossing points 5 to form a three dimensional grid truss 6 integrally formed on the shell element 2.

[0073] According to FIG. 8A, a first filament 4a is bonded to the shell element 2 at a first end 11 in the manner as explained with regard to FIG. 6. Then the first filament 4a is three dimensionally formed free standing along a predetermined trajectory of the grid truss 6. This trajectory is predetermined by the desired path and the cross sectional form of the grid truss 6, which is depicted here with dashed lines.

[0074] The forming of the first filament 4a is realized in the way as explained with regard to FIG. 7 and the form of the filament 4a is defined by the movement path of the ALM head 16.

[0075] At a second end 12, the filament 4 is bonded to the shell element 2 again in the manner according to FIG. 6. In the present example, the first end 11 of the first filament 4a is positioned on a first side 25 of a hat section and a first longitudinal end 24 of the grid truss 6 and a second end 12 is positioned on a second side 27 of the hat section and a second longitudinal end 26 of the grid truss 6.

[0076] According to FIG. 8B, a second filament 4b is bonded to the shell element 2 at a first end 13. The first end 13 is positioned opposing to the first end 11 of the first filament 4a and in parallel thereto on the second side 27 of the hat section at the first longitudinal end 24.

[0077] The second filament 4b is then three dimensionally formed along the predetermined trajectory of the grid truss 6 and thereby crosses the first filament 4a. At a crossing point 5, the second filament 4b is bonded to the first filament 4a. Furthermore, a second end 14 of the second filament 4b is bonded to the shell element 2 at a position opposing to the second end 12 of the first filament 4a and in parallel thereto on the first side 25 of the hat section at the second longitudinal end 26.

[0078] FIG. 8C shows a third filament 4c applied three dimensionally formed along the predetermined trajectory of the grid truss 6. The third filament 4c starts on the first longitudinal end 24 in a central part of the hat section. In contrast to the first and second filaments 4a, 4b, the third filament 4c is bonded to the shell element 2 at a contacting point 15 which is not at a longitudinal end but in a middle part of the third filament 4c. Furthermore, the third filament 4c crosses the first filament 4a and the second filament 4b at new crossing points 5 and is bonded to the first and second filament 4a, 4b. Accordingly, the third filament 4c is formed on the trajectory of the truss 6 offset with respect to the first and second filaments 4a, 4b. In the same way, further filaments 4n are applied offset thereto until the grid truss 6 is complete.

[0079] The number of filaments 4n necessary to complete the truss 6 depends on the actual geometry and size of the grid truss 6 and density and pattern of the desired grid. At all contact points 15 with the shell element 2 and crossing points 5, preferably both the newly applied and previously applied material is heated so as to obtain a strong welded connection.

[0080] FIG. 8D shows the complete grid truss 6 with the filaments 4a, 4b, 4c, . . . , 4n. In addition, longitudinal 0 filaments 24 and transverse 90 filaments 23 have been added crossing and bonded to the filaments 4a, 4b, 4c, . . . , 4n at crossing points. All filaments 4a, 4b, 4c, . . . , 4n, 23 and 24 together with the shell element 2 form a grid truss 6. In the present example, the grid truss is formed with a hollow cross section. However, a grid truss is not necessarily formed hollow. According to other embodiments, a grid truss 6 may also comprise diagonal struts crossing a hollow cross sectional truss profile, connecting for example opposing nodes.

[0081] In the present example the cross section forms a hat section as explained with respect to FIG. 5. However, other truss sections may be formed in a similar way, for example a hexagonal section or a T-section as shown in FIGS. 3 and 4.

[0082] In the complete grid truss 6, each filament 4a, 4b, 4c, . . . , 4n between two crossing points 5 forms an independent truss strut. In this way, the filaments 4a, 4b, 4c, . . . , 4n are arranged in a regular triangular grid configuration, for example. According to further embodiments, the grid structure may have any other configuration, such as a square grid or a honey come grid. Furthermore, such a grid is not necessarily regular, but may be locally optimized to actual strength requirements, for example by locally providing a higher or lower density of struts by adding or removing filaments and/or by locally spreading or concentrating the filaments.

[0083] FIG. 9 schematically illustrates a structural component 1 according to an embodiment of the present invention.

[0084] The structural component 1 comprises a shell element 2 formed as a fuselage skin shell and comprising a stiffening structure 3 with a plurality of axial integral grid trusses 6 with a hat section as of FIG. 8D.

[0085] Furthermore, a radial integral grid truss 6 with an L-section is provided and connected to the axial grid trusses 6, in particular by thermoplastic welding.

[0086] FIG. 10 schematically illustrates an aircraft or spacecraft 20.

[0087] The aircraft 20 comprises an airframe 21, wherein the airframe 21 comprises a structural component 1 according to one of the preceding FIGS. 3 to 9. For example, the fuselage 22 of the airframe 21 is formed with a fuselage skin shell according to FIG. 9.

[0088] Although specific embodiments of the invention are illustrated and described herein, it will be appreciated by those of ordinary skill in the art that a variety of alternate and/or equivalent implementations exist. It should be appreciated that the exemplary embodiment or exemplary embodiments are examples only and are not intended to limit the scope, applicability, or configuration in any way. Rather, the foregoing summary and detailed description will provide those skilled in the art with a convenient road map for implementing at least one exemplary embodiment, it being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope as set forth in the appended claims and their legal equivalents. Generally, this application is intended to cover any adaptations or variations of the specific embodiments discussed herein.

[0089] It will also be appreciated that in this document the terms comprise, comprising, include, including, contain, containing, have, having, and any variations thereof, are intended to be understood in an inclusive (i.e. non-exclusive) sense, such that the process, method, device, apparatus or system described herein is not limited to those features or parts or elements or steps recited but may include other elements, features, parts or steps not expressly listed or inherent to such process, method, article, or apparatus. Furthermore, the terms a and an used herein are intended to be understood as meaning one or more unless explicitly stated otherwise. Moreover, the terms first, second, third, etc. are used merely as labels, and are not intended to impose numerical requirements on or to establish a certain ranking of importance of their objects.

LIST OF REFERENCE SIGNS

[0090] 1 structural component [0091] 2 airframe [0092] 3 stiffening structure [0093] 4 filament [0094] 4a first filament [0095] 4b second filament [0096] 4c third filament [0097] 4n further filaments [0098] 5 crossing point [0099] 6; 6; 6 grid truss [0100] 7 first laser beam [0101] 8 shell element [0102] 9 second laser beam [0103] 10 ironing pad [0104] 11 first end [0105] 12 second end [0106] 13 first and [0107] 14 second end [0108] 15 contacting point [0109] 16 ALM head [0110] 17 robotic arm [0111] 18 laser source [0112] 19 guide roller [0113] 20 aircraft [0114] 21 airframe [0115] 22 fuselage [0116] 23 filament coil [0117] 24 first longitudinal end [0118] 25 first side [0119] 26 second longitudinal end [0120] 27 second side

[0121] While at least one exemplary embodiment of the present invention(s) is disclosed herein, it should be understood that modifications, substitutions and alternatives may be apparent to one of ordinary skill in the art and can be made without departing from the scope of this disclosure. This disclosure is intended to cover any adaptations or variations of the exemplary embodiment(s). In addition, in this disclosure, the terms comprise or comprising do not exclude other elements or steps, the terms a or one do not exclude a plural number, and the term or means either or both. Furthermore, characteristics or steps which have been described may also be used in combination with other characteristics or steps and in any order unless the disclosure or context suggests otherwise. This disclosure hereby incorporates by reference the complete disclosure of any patent or application from which it claims benefit or priority.