Elongated aircraft structural component for aircraft, aircraft structural assembly, aircraft and method for producing elongated aircraft structural component for aircraft

Abstract

The present invention relates to an elongate aircraft structural component for an aircraft, an aircraft structural assembly, an aircraft and a method of manufacturing an elongate aircraft structural component for an aircraft, the elongate aircraft structural component comprising a web and a flange wherein the flange extends from the web, the flange comprising a layer of composite material, the flange has a length in a length direction of the elongate aircraft structural component, a width perpendicular to the length of the flange, and a thickness less than and perpendicular to both the width and length of the flange, where the width of the flange varies along the length of the flange, whereby the flange comprises a wide flange region and a narrow flange region, where the width of the flange varies along the length of the flange. The width of the narrow flange region is smaller than the width of the wide flange region, and wherein the wide flange region comprises stitches interconnecting the layers of composite material, said stitches extending along a path in a direction having at least one component parallel to the width of the flange.

Description

Technical Field

[0001] The present invention relates to elongated aircraft structural components, such as stringers, and a method of manufacturing an elongated aircraft structural component. Background Art

[0002] Aircraft structural components, such as elongated aircraft structural components, are used in an aircraft to provide structural support and strength to an elongated portion of the aircraft. Examples of these elongated aircraft structural components include aircraft wing stringers used along the wingspan of an aircraft wing, or support elements in the fuselage of an aircraft. Summary of the Invention

[0003] A first aspect of the present invention provides an elongated aircraft structural component for an aircraft, the elongated aircraft structural component including a web and flanges, wherein the flanges extend from the web and include composite material layers, and have a length in the longitudinal direction of the elongated aircraft structural component, a width perpendicular to the length of the flange, and a thickness less than both the width and length of the flange and perpendicular to both the width and length of the flange, wherein the width of the flange varies along the length of the flange, such that the flange includes a wide flange region and a narrow flange region, the width of the narrow flange region being smaller than the width of the wide flange region, and wherein the wide flange region includes stitches connecting the composite material layers to each other, the stitches extending along a path in a direction having at least one component parallel to the width of the flange.

[0004] Stitches connecting the composite material layers to each other are provided to hold the composite material layers together, for example, for transportation, during manufacturing steps, during installation and / or in use. The stitches include one or more stitches in a straight line. The benefit provided by the stitches is to increase the strength of the flange along the stitch path. "Strength" may refer to, for example, the degree of resistance to yield, fracture, fatigue, delamination and / or deformation under tensile load, shear load, torsional load, compressive load and / or bending load. Thus, the stitches extending along a path in a direction having at least one component parallel to the width of the flange increase the strength of the flange in its width direction. The increased strength of the flange in its width direction can, for example, increase the resistance of the flange to loads in its width direction. The increased strength of the flange in its width direction can be customized by changing or appropriately selecting the magnitude of the component of the stitch path direction in the flange width direction.

[0005] Optionally, the composite material layer includes a fiber composite material layer. The fiber composite material may include pre-impregnated (also known as "prepreg") fibers or "dry" fibers. The dry fibers may be infused with a matrix material, such as a resin material, before or after the introduction of the stitches.

[0006] Optionally, the fiber composite layer comprises or consists of carbon fibers.

[0007] Optionally, the web comprises a composite layer.

[0008] Optionally, the fiber composite layer comprises a unidirectional fiber layer and a chopped strand mat layer, the unidirectional fiber layer having continuous fibers oriented along the length of the flange, and the chopped strand mat layer having discontinuous fibers oriented in multiple directions within the plane of each respective chopped strand mat layer.

[0009] The unidirectional fiber layer provides strength in the direction along the length of the flange and thus in the direction along the length of the elongated aircraft structural member. The chopped strand mat layer provides strength in multiple directions within the plane of the flange defined by the length and width of the flange. The chopped strand mat layer is used in place of a conventional secondary unidirectional fiber layer. Such a secondary unidirectional fiber layer can be oriented, for example, at 45° or 90° to the unidirectional fiber layer. The conventional secondary unidirectional fiber layer provides strength in the direction of the fibers it contains. The chopped strand mat layer can comprise recycled materials and can thus be more environmentally friendly and more economical than a conventional secondary unidirectional fiber layer that may require virgin fibers. The chopped strand mat layer can replace the secondary unidirectional fiber layer to provide strength in the width direction of the flange.

[0010] Optionally, at least 50% of the composite layer in the flange is a unidirectional fiber layer. This can help provide sufficient strength in the direction along the length of the flange.

[0011] Optionally, the wide flange region includes holes for receiving fasteners for fastening the flange to a skin section of the aircraft.

[0012] The flange can comprise a plurality of wide flange regions spaced along the length of the flange, with narrow flange regions between each pair of adjacent wide flange regions. These wide flange regions can be spaced at regular or irregular intervals. The wide flange regions can correspond to, for example, such positions along the elongated aircraft structural member where the elongated aircraft structural member will pass through ribs of an aircraft structure such as a wing. The setting of holes can weaken the composite layer in the wide flange region. Embodiments of the present invention can help increase the strength of the wide flange region to compensate for and mitigate such weakening.

[0013] Optionally, the holes do not interrupt the stitches. Thus, the stitches can maintain their structure and integrity.

[0014] Optionally, the stitches surround the holes. This can help reinforce the holes and can compensate for the weakening of the composite layer caused by the setting of the holes.

[0015] Optionally, the stitches include a plurality of parallel stitches (i.e., multiple rows of stitches) that connect the composite material layers to each other. This arrangement can be relatively easy to manufacture compared to an arrangement that includes a plurality of non-parallel stitches. For example, to produce non-parallel stitches, it may be necessary to reset the stitching device with new instructions after each angle change, or more complex equipment may be required.

[0016] Optionally, the stitches include a plurality of intersecting stitches. Intersecting stitches can provide greater strength compared to non-intersecting stitches.

[0017] Optionally, the direction has equal components in the direction of the width of the flange and the direction of the thickness of the flange. This provides reinforcement along the width and thickness of the flange. In an example using a unidirectional fiber layer where the fibers are oriented in the direction of the length of the flange, the stitches provide reinforcement in the width direction and the thickness direction.

[0018] Optionally, the direction has a component parallel to the length of the flange. This can provide increased strength along the length of the flange.

[0019] Optionally, the elongated structural member is an aircraft spar. The aircraft spar can be used as an element to which skin sections of a wing or a fuselage will be attached for structural support. The aircraft spar can have different cross-sectional shapes, such as, for example, a T-shape, an L-shape, or an Ω-shape.

[0020] Optionally, the narrow flange region also includes stitches that connect the composite material layers to each other.

[0021] A second aspect of the present invention provides an aircraft structural assembly that includes an elongated aircraft structural member according to the first aspect and a skin section attached to a flange of the elongated aircraft structural member.

[0022] This can be achieved, for example, by attaching the flange of the aircraft structural member to the skin section at a face of the flange that is opposite to the face from which the flange extends from the web. Attaching the flange can be achieved by mechanical fasteners such as bolts or rivets, or by using a welding process (in the case where the flange is a weldable material), or an adhesive process, or any other suitable method. The attachment can be done at one or more wide flange portions. The aircraft structural assembly can be, for example, for an aircraft wing and / or fuselage.

[0023] A third aspect of the present invention provides an aircraft that includes an aircraft structural assembly according to the second aspect.

[0024] A fourth aspect of the present invention provides a method of manufacturing an elongate aircraft structural component for an aircraft, the method comprising: providing a web; and providing a flange extending from the web, the flange comprising a composite layer and having a length in the longitudinal direction of the elongate aircraft structural component, a width perpendicular to the length of the flange, and a thickness less than both the width and length of the flange and perpendicular to both the width and length of the flange, wherein the width of the flange varies along the length of the flange such that the flange comprises a wide flange region and a narrow flange region, the width of the narrow flange region being less than the width of the wide flange region; and providing stitches that interconnect the composite layers in the wide flange region, the stitches extending along a path in a direction having at least one component parallel to the width of the flange.

[0025] The method allows for the manufacture of an aircraft structural component according to the first aspect of the present invention.

[0026] Optionally, providing the flange comprises providing a unidirectional fiber layer and a chopped strand mat layer, the unidirectional fiber layer having continuous fibers oriented along the length of the flange and the chopped strand mat layer having discontinuous fibers oriented in multiple directions within the plane of each respective chopped strand mat layer.

[0027] Optionally, the method comprises providing holes in the wide flange region for receiving fasteners for fastening the flange to a skin section of the aircraft.

[0028] Optionally, providing holes in the wide flange region comprises providing the holes such that the stitches are not interrupted.

[0029] Optionally, providing holes in the wide flange region comprises providing the holes such that the stitches surround the holes.

[0030] Optionally, providing the stitches comprises providing a plurality of parallel stitches that interconnect the composite layers.

[0031] Optionally, providing the stitches comprises providing a plurality of interconnected stitches.

[0032] Optionally, providing the stitches comprises providing the stitches in a direction having equal components in the direction along the width of the flange and in the direction along the thickness of the flange.

[0033] Optionally, providing the stitches comprises providing the stitches in a direction having at least one component parallel to the length of the flange.

[0034] Where appropriate, the optional features of any one aspect of the present invention may be equivalently applied to any other aspect of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Embodiments of the present invention will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0036] Figure 1 A schematic view of an aircraft according to an exemplary embodiment of the present invention is shown.

[0037] Figure 2 An aircraft structural component according to an exemplary embodiment of the present invention is shown.

[0038] Figures 3A to 3C A side view, a cross-sectional view, and a top view of an aircraft structural member according to an exemplary embodiment of the present invention are respectively shown.

[0039] Figures 4A to 4C Examples of a unidirectional fiber composite layer, a chopped strand composite layer, and a composite laminate are respectively shown.

[0040] Figure 5 Shows a cross-sectional view of a laminate of a flange of an aircraft structural member passing through Figure 3C which includes an exemplary stitch according to an exemplary embodiment of the present invention.

[0041] Figure 6 Shows a cross-sectional view of a laminate of a flange of an aircraft structural member passing through Figure 3C which includes an alternative exemplary stitch according to an exemplary embodiment of the present invention.

[0042] Figure 7 Shows Figure 3C A top view of a section of an aircraft structural member, showing an example of a stitch according to an exemplary embodiment of the present invention.

[0043] Figure 8 Shows Figure 3C A top view of a section of an aircraft structural member, showing an example of a stitch according to an exemplary embodiment of the present invention.

[0044] Figure 9 An exemplary method for manufacturing an elongate aircraft structural member according to an exemplary embodiment of the present invention is shown. Detailed Description

[0045] Figure 1 An aircraft 100 according to an example of the present invention is shown. The aircraft has a plurality of structural components that benefit from improved elongate aircraft structural members. Such components are included in the fuselage 110, the wings 120, and the tail 130.

[0046] In Figure 2A cross-section through one of the wings 120 of the aircraft 100 is shown. The wing 120 is an aircraft structural component and includes a skin section 210 defining its outer surface, a spar 230 (shown in dashed lines in Figure 2 ), and ribs 240 ( Figure 2 only one of the ribs 240 is visible in ), and a series of stringers 300. The rib 240 has an opening 242 through which the stringer 300 passes. The stringer 300 is attached to the inner surface 212 of the skin section 210 at its respective flange 320 (as described below). This can be achieved by mechanical connection, for example, by mechanical fasteners such as rivets or bolts (not shown). Alternatively, other connection methods such as adhesion or welding can be envisaged. The stringer is an example of an elongated aircraft structural component.

[0047] In Figure 3A , an elongated aircraft structural component according to an embodiment of the present invention is shown. In this exemplary embodiment, the component refers to Figure 2 one of the stringers 300 shown. In other embodiments, the aircraft structural component can be another component, such as a spar. The stringer 300 (and the remaining stringers in each stringer 300) is disposed in the wing 120 to provide structural support and must therefore be adapted to withstand the loads to which the wing 120 is subjected on the ground and in flight.

[0048] The stringer 300 includes a web 310. Flanges 320 extend perpendicularly from a first end 314 of the web 310, and the web 310 also has a second end 312 remote from the flange 320. The flange 320 has a length L and a thickness T perpendicular to the length L. In Figure 3B , a cross-sectional view of the stringer 300 is depicted, in which it is shown that the web 320 also has a width W. The width W is perpendicular to both the thickness T and the length L and is also perpendicular to the web 310. The length L is parallel to the length of the entire stringer 300.

[0049] In Figure 3C , another view of the stringer 300 looking down from above in the direction of the thickness T of the flange 320 is shown, in which the web 310 is shown in dashed lines. As best understood from Figure 3B , in this embodiment, the web 310 and the flange 320 provide an inverted "T" - shaped cross-section for the stringer 300. Other arrangements are possible and other arrangements may be desirable for different structural applications. In other examples, the cross-section can be, for example, "L" - shaped or "Ω" - shaped.

[0050] The flange 320 includes a plurality of wide flange regions 400 - also referred to as "grow - outs" ( Figure 3COnly one outgrowth in the outgrowths is shown (——), and the narrow flange regions 402 on both sides of the outgrowth 400. At each outgrowth in the outgrowth 400, the flange 320 has a width W2, which is greater than the width W1 of the flange 320 at the narrow flange region 402. The outgrowths 400 are spaced apart such that the flange 320 alternates along the length L between sections having a width W1 and sections having a width W2. In this example, W2 is about 1.7 times W1. In other embodiments of the flange, the width W2 can be, for example, 1.2 times the width W1, 1.3 times the width W1, 1.5 times the width W1, or 2 times the width W1.

[0051] The outgrowths 400 are spaced at regular intervals, reflecting the spacing of the ribs 240 in the aircraft wing 120. The outgrowths 400 are provided to facilitate the attachment of the flange 320 to the inner surface 212 of the skin section 210 by increasing the available surface area. If, for example, bolt holes are provided in the flange of the stringer without the outgrowths, the flange may become too weak to function properly. In other examples, other configurations of one or more outgrowths are possible.

[0052] In this exemplary embodiment, the stringer 300 is made of a carbon fiber composite material. Other materials that can be used to manufacture the stringer 300 in other embodiments include, for example, other fiber composite materials such as glass fiber, or other composite materials. Composite materials are used for aircraft components because they are lightweight and have favorable mechanical properties, such as strength and stiffness.

[0053] Composite materials such as carbon fiber composite materials or glass fiber composite materials can be formed by layers stacked to form a laminated structure. In Figure 4A An example of a composite material layer is shown, and the example of the composite material layer is a unidirectional fiber layer 430. The unidirectional fiber layer 430 includes carbon fibers 435 arranged side by side and aligned parallel to each other. The unidirectional fiber layer 430 is stronger in the direction along the fibers 435 than in other directions. The composite laminated structure or laminated structure is generally composed only of such unidirectional fiber layers in different relative orientations to provide sufficient strength in all directions.

[0054] In Figure 4BA second example of a composite layer is shown, and the second example of the composite layer is the chopped strand layer 440. The chopped strand layer 440 includes carbon fibers 445 that are relatively short compared to the fibers 435 in the unidirectional fiber layer 430. In this example, the length of each fiber in the fibers 445 is from 10 mm to 100 mm. The fibers 445 are arranged in the plane of the layer 440 but are aligned in random directions or at least in a plurality of directions relative to each other. Thus, a single chopped strand layer 440 has substantially equal strength in all directions in the plane of the chopped strand layer 440.

[0055] As Figure 4C shown, a hybrid laminate 420 can be provided, and the hybrid laminate 420 includes each of a plurality of unidirectional fiber layers 430 and chopped strand layers 440. Entering Figure 4C the page, the unidirectional fibers 435 in all the unidirectional fiber layers 430 are aligned in one direction, and the chopped strand fibers 445 within the chopped strand layer 440 provide strength in other directions. The hybrid laminate 420 includes one unidirectional fiber layer 430 or two unidirectional fiber layers 430 located between pairs of chopped strand layers 440 and terminates at corresponding chopped strand layers in the chopped strand layer 440 at each of the two opposite faces of the hybrid laminate 420. The total number of layers in the hybrid laminate 420 is 13.

[0056] In other embodiments, fewer or more unidirectional fiber layers 430 and / or chopped strand layers 440 can be provided in the hybrid laminate, and / or there can be different proportions of unidirectional fiber layers 430 and chopped strand layers 440. Preferably, at least half of the total number of layers is the unidirectional fiber layer 440. In some examples, different stacking orders of the unidirectional fiber layer 430 and the chopped strand layer 440 are possible. Preferably, pairs of adjacent chopped strand layers in the chopped strand layer 440 are separated by at least one unidirectional fiber layer in the unidirectional fiber layer 430. In other examples, not all of the unidirectional fibers 435 in all the unidirectional fiber layers 430 may be aligned in a single direction; they can be oriented in two or more directions. The layers in the hybrid laminate can have different thicknesses. For example, some or each chopped strand layer in the chopped strand layer 440 can have a smaller thickness than each unidirectional fiber layer in the unidirectional fiber layer 430, or some or each chopped strand layer in the chopped strand layer 440 can have a greater thickness than each unidirectional fiber layer in the unidirectional fiber layer 430.

[0057] The stringer 300 includes a hybrid laminate 420 such that the fibers 435 of the unidirectional fiber layer 430 are aligned with the length L of the flange 320 of the stringer 300, and the layers 430 and 440 of the hybrid laminate 420 are stacked in the thickness direction of the flange 320. In other examples of elongated aircraft structural components, other hybrid laminates are possible, as are conventional unidirectional fiber layer laminates or other laminated composites or structures.

[0058] Figure 5 A cross-sectional view of the hybrid laminate 420 through the Figure 3C flange 320 is shown. For clarity, only half of the flange 320 and the edge of the first end 314 of the web 310 are shown. The flange 320 includes Figure 4C the hybrid laminate 420 shown in Figure 5 . Entering the

[0059] page, in this example, the fibers of the unidirectional fiber layer of the hybrid laminate 420 are placed perpendicular to both the thickness T and the width W of the flange. It is not shown for clarity. Figure 5 The flange 320 includes a stitch 500 (i.e., one stitch) passing through the hybrid laminate 420. The stitch 500 connects the layers of the hybrid laminate 420. The stitch 500 is arranged at an angle θ with respect to an axis A parallel to the thickness T of the flange 320. Thus, the stitch 500 is arranged or extends along a path in a direction that has a component in the direction of the width W and a component in the direction of the thickness T. By changing the angle θ, the component in the direction of the width W can be changed. A larger component in the direction of the width W results in greater strength for the flange 320 in that direction. In

[0060] the example, the angle θ is approximately 30°. In other examples, the angle θ can be any other angle, such as greater than 0° and less than 90°, for example 45°. Figure 6 An example is shown in Figure 4C . This example is a cross-sectional view through the flange. In this example, the flange 320 also includes Figure 6As shown by the arrow in , the angle θ is approximately 45° clockwise, and the angle θ' is approximately 45° counterclockwise. Thus, the first type of suture 510 is arranged perpendicular to the second type of suture 512.

[0061] In alternative examples, the angles θ and θ' can have unequal magnitudes. For example, each of the angles θ and θ' can have any magnitude greater than 0° and less than 90°, and can be clockwise or counterclockwise. In other examples, there can be more than two types of sutures. In other examples, there can be only one series of parallel sutures. In other examples, the sutures do not intersect each other. In other examples, the sutures are not parallel to each other.

[0062] Figure 7 A top view of the outgrowth 400 of the flange 320 is shown. This top view is Figure 3C a close-up view of a part, where only one side of the outgrowth 400 is visible. The outgrowth 400 includes a hole 410 for receiving a fastener, such as a bolt for example, when the stringer 300 is attached to the skin section 210, as described with reference to Figure 2 above. When the stringer 300 is in place in the wing 120, the outgrowth 400 can be positioned, for example, in the rib opening 242 or along the stringer 300 between the ribs 240.

[0063] In other examples, there can be no holes in the outgrowth 400, or there can be multiple holes for receiving corresponding fasteners. The hole 410 in the present embodiment is threaded, but in other embodiments it can be unthreaded, depending on what type of fastener it is configured to receive.

[0064] The outgrowth 400 includes a plurality of parallel sutures 520, and each suture of the plurality of parallel sutures 520 extends or elongates along a path in the following direction: the direction having only components along the width W and thickness T of the flange 320.

[0065] In other examples, such as Figure 8 the example shown in , each suture of the plurality of sutures is arranged along a path extending in the following direction: the direction also having a component in the direction of the length L of the flange 320. In the Figure 8 example of , when viewed in a top-down perspective, the suture 530 is arranged along two orientations, and the suture 530 intersects itself. In this example, the suture 530 surrounds the hole 410 for receiving the fastener, such that the rows of the suture 530 are not interrupted by the hole 410. Thus, the edge of the hole 410 is strengthened to prevent mechanical failure.

[0066] In other embodiments of the present invention, the flange 320 can include one or more sutures outside the area of the outgrowth 400.

[0067] In other embodiments of the present invention, the web 310 may further include a hybrid laminate and / or stitches as described with respect to the above examples.

[0068] The stitches described with respect to the above examples include polyester material threads with a thickness of 0.05 mm. In other embodiments of the present invention, other flexible polymer materials such as polyamide, carbon fiber, or glass fiber, and other thicknesses, for example, from 0.05 mm to 0.2 mm, are possible.

[0069] Figure 9 An example method according to the present invention is shown. Method 1000 is a method of manufacturing an elongated aircraft structural component. In this example, method 1000 is a method of manufacturing the spar 300 described above.

[0070] Method 1000 includes disposing 1010 the web 310.

[0071] Method 1000 further includes disposing 1020 the flange 320.

[0072] Disposing 1010 the web and disposing 1020 the flange include forming the respective web and flange from a hybrid laminate 420. In this embodiment, disposing 1010 the web and disposing 1020 the flange occur simultaneously. In other embodiments, disposing 1010 the web may occur before or after disposing 1020 the flange.

[0073] Method 1000 includes disposing 1030 stitches in the extension 400 of the flange 320. In this example, the stitches are arranged as Figure 8 shown. Disposing 1030 the stitches is achieved by automated sewing, where the tufting needles are positioned at the angle of the stitches. In other embodiments, the sewing may be performed manually or in any other manner.

[0074] Method 1000 includes disposing 1040 threaded holes 410 in the extension 400 by machining, the holes 410 being for receiving fasteners. In this embodiment, disposing 1040 the holes occurs after disposing 1030 the stitches, but in other embodiments, disposing 1040 the holes may occur after disposing 1030 the stitches or simultaneously with disposing 1030 the stitches. In other examples, disposing 1040 the holes may include disposing a plurality of holes in the extension 400, where the plurality of holes may be the same or different. In some examples, such holes 410 may not be disposed.

[0075] In other exemplary methods, other methods can be contemplated. For example, a resin infusion step can be performed, in which the dry fibers of the composite material layer are infused with a resin material. The resin infusion can occur before or after setting the 1030 sutures and / or before or after setting the 1040 holes. In some such embodiments, the sutures can additionally assist in stabilizing the layers during the resin infusion step. When performing such a resin infusion process, it can be followed by a resin curing process. The resin curing process can also be used in embodiments where prepreg-type composite materials are used instead of dry fibers and where there is no resin infusion process.

[0076] It should be noted that unless otherwise explicitly stated, the term "or" as used herein shall be interpreted to mean "and / or".

Claims

1. An elongated aircraft structural component for an aircraft, the elongated aircraft structural component comprising a web and a flange, in, the flange extending from the web and comprising a composite material layer and having a length in the length direction of the elongated aircraft structural component, a width perpendicular to the length of the flange, and a thickness less than both the width and the length of the flange and perpendicular to both the width and the length of the flange, wherein the width of the flange varies along the length of the flange, whereby the flange comprises a wide flange region and a narrow flange region, the narrow flange region having a smaller width than the wide flange region, and Wherein the wide flange region comprises stitches connecting the composite material layers to each other, the stitches extending along a path in a direction having at least one component parallel to the width of the flange.

2. The elongated aircraft structural component according to claim 1, wherein: The composite material layer comprises a fiber composite material layer.

3. The elongated aircraft structural component according to claim 2, wherein: The fiber composite material layer includes carbon fibers.

4. An elongated aircraft structural component according to claim 2 or claim 3, wherein: The fibrous composite layer includes a unidirectional fiber layer having continuous fibers oriented along the length of the flange and a chopped strand layer having discontinuous fibers oriented in multiple directions within the plane of each respective chopped strand layer.

5. The elongated aircraft structural component according to claim 4, wherein: At least 50% of the composite material layers in the flange are unidirectional fiber layers.

6. An elongated aircraft structural component according to any one of the preceding claims, wherein: The wide flange region includes holes for receiving fasteners for fastening the flange to a skin section of the aircraft.

7. The elongated aircraft structural component according to claim 6, wherein: The holes do not interrupt the suture.

8. An elongated aircraft structural component according to claim 6 or 7, wherein: The suture surrounds the hole.

9. An elongated aircraft structural component according to any one of the preceding claims, wherein: The stitching includes a plurality of parallel stitching lines connecting the composite material layers to each other.

10. An elongated aircraft structural component according to any one of claims 1 to 9, wherein: The suture includes a plurality of intersecting sutures.

11. An elongate aircraft structural component according to any one of the preceding claims, wherein: The direction of the path has equal components in the direction of the width of the flange and in the direction of the thickness of the flange.

12. An elongate aircraft structural component according to any one of the preceding claims, wherein: The direction of the path has a component parallel to the length of the flange.

13. An elongate aircraft structural component according to any one of the preceding claims, wherein: The elongated structural component is an aircraft stringer.

14. An elongate aircraft structural component according to any one of the preceding claims, wherein: The narrow flange area also includes stitching that interconnects the composite material layers.

15. An aircraft structural assembly comprising: An elongated aircraft structural component according to any one of claims 1 to 14; as well as A skin section is attached to the flange of the elongated aircraft structure component.

16. An aircraft comprising an aircraft structure assembly according to claim 15.

17. A method of manufacturing an elongated aircraft structural component for an aircraft, the method comprising: Set up the belly plate; as well as providing a flange extending from the web, the flange comprising a composite material layer and having a length in the length direction of the elongated aircraft structural component, a width perpendicular to the length of the flange, and a thickness less than both the width and the length of the flange and perpendicular to both the width and the length of the flange, wherein the width of the flange varies along the length of the flange, whereby the flange comprises a wide flange region and a narrow flange region, the narrow flange region having a width less than the width of the wide flange region; as well as A seam is provided connecting the composite material layers to each other in the region of the wide flange, the seam extending along a path in a direction having at least one component parallel to the width of the flange.

18. The method according to claim 17, wherein: The composite material layer includes a unidirectional fiber layer having continuous fibers oriented along the length of the flange and a chopped strand layer having discontinuous fibers oriented in multiple directions within the plane of each of the respective chopped strand layers.

19. The method according to claim 17 or claim 18, wherein: The method includes providing the wide flange region with holes for receiving fasteners for fastening the flange to a skin section of the aircraft.

20. The method according to claim 19, wherein: Providing the wide flange area with holes includes providing the holes so that the holes do not interrupt the suture.

21. The method of claim 19 or claim 20, wherein: Providing the wide flange area with the hole includes arranging the hole such that the stitching surrounds the hole.

22. The method according to any one of claims 17 to 21, wherein: Providing the stitches includes providing a plurality of parallel stitches connecting the composite material layers to each other.

23. The method according to any one of claims 17 to 22, wherein: Providing the sutures includes providing a plurality of intersecting sutures.

24. The method according to any one of claims 17 to 23, wherein: Providing the stitches includes providing the stitches in a direction having equal components in a direction along the width of the flange and in a direction along the thickness of the flange.

25. The method according to any one of claims 17 to 24, wherein: Positioning the stitches includes positioning the stitches in a direction having at least one component parallel to the length of the flange.

Images