Aircraft structural component, method of providing aircraft structural component and aircraft fuselage

Abstract

The invention relates to an aircraft structural component, a method of providing an aircraft structural component and an aircraft fuselage. The invention relates to an aircraft structural component having a stacked arrangement of layers, where the layers are one of at least a first multiplier and a base layer of an outer skin of an aircraft fuselage, where the first multiplier and the base layer are bonded by means of a first intermediate bonding film layer to form a stepped configuration of the aircraft structural component, the first multiplied part comprises a first fibrous prepreg layer, in particular a glass fiber prepreg layer, which is arranged at the periphery of the first multiplied part, and wherein the first fibrous prepreg layer is bonded to the base layer by means of at least one first bonding film section, a first binding film section is disposed adjacent the first fibrous prepreg layer in a first wedge element at least partially bound to the first fibrous prepreg layer. The invention also provides a method for providing an aircraft structural component and an aircraft fuselage provided with an aircraft structural component.

Description

Technical Field

[0001] The present invention relates to an aircraft structural component and a method for providing such an aircraft structural component. Background Art

[0002] Aircraft, and in particular, aircraft fuselages, have various regions that are subject to specific load-bearing and load-transfer requirements. These regions or structural components are often located in inaccessible areas of the aircraft. Aircraft structural components must be designed to prevent cracks from occurring or to allow cracks to propagate in a controlled manner. Steps along longitudinal skin joints are generally avoided and rare because they are complex to design with regard to tolerances and require extreme care in fatigue calculations and certification testing. The components involved in these joints must often be stepped down with numerous small steps to meet the aerodynamic and optical requirements associated with the steps in the outer surface of the aircraft fuselage. Summary of the Invention

[0003] Although the present invention can be used in many applications, the present invention and the underlying problems will be explained in more detail in the context of aircraft. However, the structural components described can also be used in vehicles in all sectors of the transportation industry, for example in road vehicles, rail vehicles or ships.

[0004] Against this background, the object of the present invention is to provide an aircraft structural component which meets aerodynamic and optical requirements, provides load-bearing capacity, is not susceptible or is susceptible only to a small extent to crack formation and crack propagation, and can be manufactured in an efficient and cost-effective as well as material-saving manner.

[0005] This object is achieved by an aircraft structural component having the features of one main aspect, which is provided in a method having the features of another main aspect and which is integrated in an aircraft fuselage according to a further main aspect.

[0006] According to a first aspect of the present invention, an aircraft structural component is provided, comprising a stacked arrangement of layers, wherein the layers form an outer skin of an aircraft fuselage. The layers comprise at least a first doubler and a base layer, wherein the first doubler is bonded to the base layer via a first intermediate bonding film layer, thereby creating the stepped configuration of the aircraft structural component. The first doubler comprises a first fiber prepreg layer, particularly a glass fiber prepreg layer, positioned at the periphery of the first doubler and bonded to the base layer via at least one first bonding film segment, the first bonding film segment being positioned adjacent to the first fiber prepreg layer within a first wedge-shaped element. The wedge-shaped element is at least partially bonded to the first fiber prepreg layer. The wedge-shaped element is positioned at an extension of the first doubler, and the fiber prepreg layer is positioned below the periphery of the first doubler. Due to the fiber prepreg bonded within the wedge-shaped element, the aircraft structural component of the present invention prevents cracks that occur in the wedge-shaped element from propagating into the base skin. The aircraft structural component of the present invention has the advantage of providing a smooth transition between load-bearing metal sheet layers within a stepped longitudinal joint. The base layer and the first doubler are manufactured from a metallic material, preferably from aluminum.

[0007] Another aspect of the present invention is a method for providing an aircraft structural component, the method comprising: placing an intermediate bonding film layer on a surface facing an inner skin of an aircraft fuselage; placing a first fiber prepreg layer, particularly a glass fiber prepreg layer, adjacent to the first intermediate bonding film layer; placing a first doubler on the first intermediate bonding film layer and the fiber prepreg layer; placing at least a first bonding film segment adjacent to the first fiber prepreg layer; and autoclaving the aircraft structural component to form a first wedge-shaped element in an extended region of the first doubler, the first wedge-shaped element bonding to a region of the first fiber prepreg layer extending beyond a perimeter of the first doubler. This method avoids the time- and material-intensive milling process required to manufacture the aircraft structural component, improves load-bearing performance, and reduces or eliminates crack initiation and crack propagation into the inner skin or base layer.

[0008] Another aspect of the invention is an aircraft fuselage provided with an aircraft structural component, wherein the aircraft structural component is a load-carrying component of the aircraft fuselage which does not experience crack initiation or only experiences reduced crack initiation and limited crack propagation.

[0009] Advantageous embodiments and further developments are apparent from the other secondary aspects and from the description with reference to the accompanying drawings.

[0010] According to another aspect of the present invention, the aircraft structural component further includes at least one second doubler having a second fiber prepreg layer positioned at a periphery of the second doubler, wherein the second doubler is bonded to the first doubler via a second intermediate bonding film layer to form the stepped configuration of the aircraft structural component. The second fiber prepreg layer is bonded to the first doubler via at least one second bonding film segment positioned adjacent to the second fiber prepreg layer within a second wedge-shaped element, the second wedge-shaped element at least partially bonding to the second fiber prepreg layer. This provides the advantage of increasing the overall thickness of the aircraft structural component, thereby providing the same advantages with respect to crack formation and propagation as described above. The present invention is not limited to aircraft structural components having two doublers in the aforementioned arrangement, but also encompasses configurations including multiple doublers arranged in a stacked arrangement having a stepped design and disposed on a primary skin or primary layer of the aircraft fuselage. The second doubler and any subsequent doublers are manufactured from a metallic material, such as aluminum. The second doubler and any subsequent doublers may be provided with at least one step introduced by corresponding machining of the aluminium, in particular by grinding or milling the aluminium, in order to further reduce the thickness by stepping down the doubler.

[0011] According to another embodiment of the aircraft structural component, at least a first wedge-shaped element and a second wedge-shaped element are positioned in the extension region of the first doubler and the second doubler, respectively. This configuration also facilitates a smooth transition of the load-bearing sheet metal layers within the stepped longitudinal joint. Furthermore, due to the fiber prepreg layers positioned below the wedge-shaped elements, cracks forming in the wedge-shaped elements cannot propagate into the metal layers.

[0012] According to another embodiment of the aircraft structural component, the fibers of the fiber prepreg layer are oriented in the circumferential direction of the doubler. In addition to the advantages of the configuration of the invention outlined above, the invention has the following advantages: cracks that form, in particular in the region of one or more wedge-shaped elements, are deflected at an angle of essentially 90 degrees and, in particular, cannot propagate or spread into the metal layer, the base layer or base skin, the doubler or the fuselage of the aircraft structural component.

[0013] According to another embodiment of the aircraft structural component, each of the first and second fiber prepreg layers is arranged coplanar with a surface plane of the corresponding first and second doublers and projects beyond the surface plane of the corresponding doubler. This embodiment has the following advantages: cracks formed in the bonding membrane layer or the wedge-shaped element cannot propagate or diffuse into the metal layer, base layer, or skin of the aircraft structural component, or into the fuselage, and thus cannot affect the load-bearing performance of the structure.

[0014] According to another embodiment of the aircraft structural component, at least the wedge-shaped elements are formed during the autoclaving process. This embodiment has the advantage that, during the autoclaving process, the bonding film layer disposed adjacent to or on the fiber prepreg layer is bonded into the wedge-shaped elements, thereby achieving and / or improving the advantages described above with respect to the formation and propagation of cracks in the aircraft structural component or in the fuselage in which the aircraft structural component is disposed.

[0015] According to another embodiment of the method, a second bonding film section is placed to cover the area of the first fiber prepreg layer that extends beyond the periphery of the surface of the first doubler and the second bonding film section. This advantageously allows for improved bonding of the fiber prepreg layer into the wedge-shaped element during autoclaving.

[0016] According to another embodiment of the method, a second intermediate bonding film layer and a second fiber prepreg layer positioned adjacent to the second intermediate bonding film layer are placed on a first doubler, and at least one second doubler is placed on the second intermediate bonding film layer and the second fiber prepreg layer, wherein the second fiber prepreg layer extends beyond the periphery of the surface of the second doubler, and a third bonding film segment is positioned adjacent to the second fiber prepreg layer. After placement, the aircraft structural component is autoclaved to form second wedge-shaped elements in the extended region of the second doubler. The second wedge-shaped elements bond to the region of the second fiber prepreg layer that extends beyond the periphery of the surface of the second doubler. This has the advantage that the total number of doublers arranged in a stacked configuration can be increased, and the placement area and load-bearing requirements of the aircraft structural component can be adapted. The method of the present invention is not limited to a method in which two doublers are arranged in the configuration described above; configurations comprising multiple doublers in a stacked configuration and manufactured using the method of the present invention are also contemplated.

[0017] According to another embodiment of the method, a fourth bonding film segment is positioned to cover the area of the second fiber prepreg layer that extends beyond the periphery of the second doubler, as well as the third bonding film segment. This embodiment has the advantage that, during high pressure, the bonding film layer positioned adjacent to or on the fiber prepreg layer is bonded into the wedge-shaped element at the extension area of the second doubler, thereby achieving and / or ameliorating the previously described advantages with respect to the formation and propagation of cracks in the aircraft structural component or in the fuselage in which the aircraft structural component is provided. The method of the present invention is not limited to methods in which two doublers are positioned in the arrangement described above, but also encompasses configurations comprising a plurality of doublers in a stacked arrangement and manufactured using the method of the present invention.

[0018] According to another embodiment of the method, at least one separator foil and a forming sheet are placed on the extension area of at least one of the first doubler and the second doubler before the autoclaving, in order to form at least one of the first wedge-shaped element and the second wedge-shaped element during the autoclaving. This has the advantage that the final shape of the wedge-shaped element can be determined during the autoclaving and can be adapted to the requirements imposed by the envisaged configuration of the aircraft structural component and the area in which it is to be placed.

[0019] According to another embodiment of the method, an aircraft structural component is positioned on an inner skin of an aircraft fuselage and fastened to the inner skin by means of a plurality of fastener elements, wherein the fastener elements extend through the base skin and the inner skin and clamp the aircraft structural component and the inner skin together, thereby causing the base skin and at least one doubler to bend at least partially toward the inner skin. This has the advantage that, due to the configuration of the aircraft structural component, in which the added fiber prepreg layer is positioned below the end region of the doubler, crack initiation in the bonding film layer and crack propagation into the metal layer are prevented. The fiber prepreg layer also prevents cracks that have occurred in the wedge-shaped elements or bonding film layers from growing and / or propagating toward the base layer of the aircraft structural component or the base skin of the fuselage.

[0020] According to a further embodiment of the method, the gap formed between the aircraft structural component and the inner skin is filled with a sealant, in particular an interference sealant. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] The invention is explained in more detail below with reference to the embodiments shown in the schematic diagram:

[0022] Figure 1 schematically depicts a view of a prior art aircraft structural component;

[0023] Figure 2 schematically depicts a side cross-sectional view of an aircraft structure component in a pre-pressurized state according to an embodiment of the present invention;

[0024] Figure 3 schematically depicts a side cross-sectional view of an aircraft structure component in a post-pressurized state according to an embodiment of the present invention; and

[0025] Figure 4 A side cross-sectional view of an aircraft structural component according to an embodiment of the invention clamped to an inner skin of an aircraft fuselage is schematically depicted. DETAILED DESCRIPTION

[0026] The accompanying drawings are included to provide a further understanding of the present invention and are incorporated into and constitute a part of this specification. The accompanying drawings illustrate embodiments of the present invention and, together with the description, serve to explain the principles of the present invention. Other embodiments of the present invention and many of the expected advantages of the present invention will be readily understood as they become better understood by reference to the detailed description. The elements of the accompanying drawings are not necessarily drawn to scale relative to each other. In the accompanying drawings, unless otherwise indicated, similar reference numerals represent similar or functionally similar components.

[0027] Although specific embodiments are illustrated and described herein, it will be understood by those skilled in the art that various alternative implementations and / or equivalent implementations may be substituted for the specific embodiments shown and described without departing from the scope of the invention. Generally, this application is intended to cover any modifications or variations of the specific embodiments discussed herein.

[0028] In the figures of the accompanying drawings, identical elements, features and components having the same function and the same effect are each given the same reference numerals unless otherwise specified.

[0029] Figure 1 A view of a prior art aircraft structural component 100 is depicted, schematically showing a single milled component. In such a milled component design, a milling ramp or small step 112 is used to produce a reduction in thickness, wherein the small step 112 reflects current standards in the manufacture of the aircraft structural component 100, which has disadvantages related to the formation and propagation of cracks.

[0030] Figure 2A schematic side cross-sectional view of an aircraft structural component 100 in a pre-autoclave state according to an embodiment of the present invention is depicted. Aircraft structural component 100 is constructed from a total of three metal substrates, particularly aluminum substrates, arranged in a stacked configuration: a base skin layer 101, a first doubler 103, and a second doubler 105. Base layer 101 may be the outer skin of an aircraft fuselage. To bond first doubler 103 to base layer 101, a first intermediate bonding film layer 102 is used, arranged between the two elements. First doubler 103 includes a first fiber prepreg layer 106a, particularly a glass fiber prepreg layer, arranged at a peripheral edge 111a of first doubler 103 and below. First fiber prepreg layer 106a is bonded to base layer 101 by means of a first bonding film segment 107 positioned adjacent to first fiber prepreg layer 106a and covered by a second bonding film segment 108. A second doubler 105 is positioned above the first doubler 103 and bonded to the first doubler 103 via a second intermediate bonding film layer 104. Second doubler 105 is shorter than first doubler 103, thereby forming a stepped configuration of aircraft structural component 100. A second fiber prepreg layer 106b is arranged below the periphery 111b of second doubler 105 and bonded to the first doubler 103 via a third bonding film segment 109 positioned adjacent to the second fiber prepreg layer 106b and covered by a fourth bonding film segment 110. During the manufacture of second doubler 105, a small step 112 is introduced by milling the second doubler 105 to further reduce its thickness. Second doubler 105 may also be provided with more than one step 112 to further step down the second doubler 105.

[0031] Prior to autoclaving, a separator foil 401a and a shaped sheet 400a are placed on an arrangement positioned in the extension region 200a of the first doubler 103, and another separator foil 401b and a shaped sheet 400b are placed on an arrangement positioned in the extension region 200b of the second doubler 105. During autoclaving, the shaped sheet 400a forms a first wedge-shaped element 500a in the extension region 200a of the first doubler 103, and the shaped sheet 400b forms a second wedge-shaped element 500b in the extension region 200b of the second doubler 105. The wedge-shaped elements 500a, 500b at least partially bond with the fiber prepreg layers 106a, 106b and extend each extension region 200a, 200b of the first doubler 103 and the second doubler 105 toward the respective preceding layer.

[0032] Figure 3A schematic side cross-sectional view of an aircraft structural component 100 in a post-pressurization state according to an embodiment of the present invention is depicted. In the extension regions 200a, 200b, wedge-shaped elements 500a, 500b are formed, and the metal layers comprising the base layer 101, the first doubler 103, and the second doubler 105 are bonded to one another. The stepped configuration of the aircraft structural component 100 is achieved without requiring extensive milling of the component. The thickness of the second doubler 105 is further reduced by providing at least one step 112 in the surface of the element. The fiber material layer, created by fiber prepreg layers 106a, 106b positioned beneath the first doubler 103 and the second doubler 105, respectively, prevents cracks formed in the wedge-shaped element 500a, 500b formed adjacent to the first doubler 103 and the second doubler 105 from propagating in the metal layers and growing toward the base layer 101 and, therefore, the metal outer skin.

[0033] Figure 4 A side cross-sectional view of an aircraft structural component 100 according to an embodiment of the present invention is schematically depicted, clamped to an inner skin 700 of an aircraft fuselage. Figure 3 The depicted aircraft structural component 100 is placed on an inner skin 700 of an aircraft fuselage and is fastened to the inner skin 700 by means of a plurality of fastener elements 600 , wherein the fastener elements 600 extend through the base layer 101 forming the outer skin of the aircraft fuselage and the inner skin 700 . Figure 4 A configuration is depicted in which a base skin 101 and an inner skin 700 are connected by means of rivets serving as fastener elements 600, wherein an upper first rivet head 602 is flushly sunken into the base skin 101 and a lower second rivet head 603 rests on the inner skin 700, with a rivet shaft 601 connecting the first rivet head 602 and the second rivet head 603. The fastener elements 600 clamp the aircraft structural component 100 and the inner skin 700 together, thereby allowing the layers of the aircraft structural component 100 of the present invention to be at least partially bent toward the inner skin 700. The stacked configuration prevents cracks from forming in the metal layers during bending, and the fiber prepreg layers 106a, 106b prevent cracks formed in the wedge-shaped elements 500a, 500b formed adjacent to the first doubler 103 and the wedge-shaped elements 500b formed adjacent to the second doubler 105 from propagating into the metal layers. Thus, a smooth transition of the load-bearing metal layer along the longitudinal direction of the stepped joint is achieved, and crack propagation beyond the wedge-shaped elements 500a, 500b is prevented. The thickness of the second doubler 105 is further reduced by providing at least one step 112 in the surface of this element. After the connection, the gap 800 remaining between the aircraft structural component 100 and the inner skin 700 is filled with a sealant (not shown), preferably an interference-type sealant.

[0034] Reference Signs List

[0035] 100 Aircraft structural components

[0036] 101 Basic Layer

[0037] 102 first intermediate bonding film layer

[0038] 103 First Multiplier

[0039] 104 second intermediate bonding film layer

[0040] 105 Second multiplier

[0041] 106a, 106b fiber prepreg layers

[0042] 107 First conjunctival segment

[0043] 108 Second conjunctival segment

[0044] 109 Third conjunctival segment

[0045] 110 Fourth conjunctival segment

[0046] 111a, 111b periphery

[0047] 112 steps

[0048] 200a, 200b extension area

[0049] 400a, 400b formed sheets

[0050] 401a, 401b separator foil

[0051] 500a, 500b wedge-shaped elements

[0052] 600 Fastener Elements

[0053] 601 rivet shaft

[0054] 602 first rivet head

[0055] 603 Second rivet head

[0056] 700 inner skin

[0057] 800 gap

Claims

1. An aircraft structural component (100), said aircraft structural component (100) having a stacked arrangement of layers, wherein: The layer is one of at least a first doubler (103) and a basic layer (101) of an outer skin of an aircraft fuselage, wherein the first doubler (103) and the basic layer (101) are bonded by means of a first intermediate bonding film layer (102) to form the stepped configuration of the aircraft structural component (100), wherein the first doubler (103) comprises a first fiber prepreg layer (106a), in particular a glass fiber prepreg layer, which is arranged at a peripheral edge (111a) of the first doubler (103), and wherein the first fiber prepreg layer (106a) is bonded to the basic layer (101) by means of at least one first bonding film segment (107), which is placed adjacent to the first fiber prepreg layer (106a) in a first wedge-shaped element (500a) at least partially bonded to the first fiber prepreg layer (106a).

2. The aircraft structural component (100) according to claim 1, comprising at least one second doubler (105), said second doubler (105) having a second fiber prepreg layer (106b) arranged at a peripheral edge (111b) of said second doubler (105), wherein: The second doubler (105) is bonded to the first doubler (103) by means of a second intermediate bonding film layer (104) to form a stepped configuration of the aircraft structural component (100), and wherein the second fiber prepreg layer (106b) is bonded to the first doubler (103) by means of at least one second bonding film segment (108), the second bonding film segment (108) being placed adjacent to the second fiber prepreg layer (106b) in a second wedge-shaped element (500b) at least partially bonded to the second fiber prepreg layer (106b).

3. The aircraft structural component (100) according to any one of claims 1 or 2, wherein: At least the first wedge-shaped element (500a) and the second wedge-shaped element (500b) are positioned in the extension region (200a) of the first doubler (103) and the extension region (200b) of the second doubler (105), respectively.

4. The aircraft structural component (100) according to any one of claims 1 to 3, wherein: The fibers of the fiber prepreg layers (106a, 106b) are oriented along the circumferential direction of the first doubler (103) and the second doubler (105).

5. The aircraft structural component (100) according to any one of claims 1 to 4, wherein: Each of the first fiber prepreg layer (106a) and the second fiber prepreg layer (106b) is arranged coplanar with a surface plane of the respective first doubler (103) and the second doubler (105) and protrudes beyond the surface plane of the respective doubler (103, 105).

6. The aircraft structural component (100) according to any one of claims 1 to 5, wherein: At least the wedge-shaped elements (500a, 500b) are formed in a high pressure process.

7. A method for providing an aircraft structure component (100) according to any one of claims 1 to 6, wherein: The method comprises: placing a first intermediate bonding film layer (102) on a surface of the base layer (101) of the aircraft structural component (100) facing the inner skin (700) of the aircraft fuselage; placing a first fiber prepreg layer (106a), in particular a glass fiber prepreg layer, adjacent to the first intermediate bonding film layer (102); placing a first doubler (103) on the first intermediate bonding film layer (102) in such a way that a region of the first fiber prepreg layer (106a) extends beyond a peripheral edge (111a) of the first doubler (103); a bonding film layer (102) and the fiber prepreg layer (106a); placing at least a first bonding film segment (107) adjacent to the first fiber prepreg layer (106a); and autoclaving the aircraft structural component (100) to form a first wedge-shaped element (500a) in the extended region (200a) of the first doubler (103), the first wedge-shaped element (500a) bonding to a region of the first fiber prepreg layer (106a) extending beyond a peripheral edge (111a) of the first doubler (103).

8. The method according to claim 7, wherein: A second bonding film section (108) is positioned to cover the area of the first fiber prepreg layer (106a) extending beyond the periphery (111a) of the first doubler (103) and the first bonding film section (107).

9. The method according to claim 7 or 8, wherein A second intermediate bonding film layer (104) and a second fiber prepreg layer (106b) arranged adjacent to the second intermediate bonding film layer (104) are placed on the first doubler (103), and at least one second doubler (105) is placed on the second intermediate bonding film layer (104) and the second fiber prepreg layer (106b), wherein the second fiber prepreg layer (106b) extends beyond the periphery (111b) of the second doubler (105), a third bonding film segment (109) is placed adjacent to the second fiber prepreg layer (106b), and the aircraft structural component (100) is autoclaved to form a second wedge-shaped element (500b) in the extended region (200b) of the second doubler (105), the second wedge-shaped element (500b) being bonded to the region of the second fiber prepreg layer (106b) extending beyond the periphery (111b) of the second doubler (105).

10. The method according to any one of claims 7 to 9, wherein: A fourth bonding film section (110) is positioned to cover the area of the second fiber prepreg layer (106b) extending beyond the periphery (111b) of the second doubler (105) and the third bonding film section (109).

11. The method according to any one of claims 7 to 10, wherein: In the method according to any one of claims 9 or 10, at least one further doubler is placed on the second doubler (105).

12. The method according to any one of claims 7 to 11, wherein: Prior to autoclaving, at least one separator foil (401a, 401b) and a forming sheet (400a, 400b) are placed on the extension area (200a, 200b) of at least one of the first doubler (103) and the second doubler (105) to form at least one of the first wedge-shaped element (500a) and the second wedge-shaped element (500b) during autoclaving.

13. The method according to any one of claims 7 to 12, wherein: The aircraft structural component (100) is placed on the inner skin (700) of the aircraft fuselage and is fastened to the inner skin (700) by means of a plurality of fastener elements (600), wherein the fastener elements (600) extend through the basic layer (101) and the inner skin (700) and clamp the aircraft structural component (700) and the inner skin (700) together, thereby causing the basic layer (101) and at least one of the doublers (103) to bend at least partially towards the inner skin (700).

14. The method according to any one of claims 7 to 13, wherein: A gap (800) formed between the aircraft structural component (100) and the inner skin (700) is filled with a sealant, in particular an interference sealant.

15. An aircraft fuselage provided with an aircraft structural component (100) according to any one of claims 1 to 6, wherein: The aircraft structural component (100) is a load-bearing component of the aircraft fuselage.

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