Aircraft structural component, aircraft structural assembly, aircraft, and method of manufacturing an aircraft structural component for an aircraft

Abstract

Aircraft structure components, aircraft structure assemblies, aircraft, and methods of manufacturing an aircraft structure component for an aircraft are disclosed. An aircraft structure component includes a web extending from a flange in a first direction to a distal end distal from the flange, wherein the web includes a plurality of layers, each layer having a plurality of fibers, wherein at least a first layer of the plurality of layers has a first portion of the fibers oriented at a first angle relative to the first direction and a second portion of the fibers oriented at a second angle relative to the first direction, wherein the first portion of the layer is closer to the flange than the second portion of the layer, and the second angle is greater than the first angle.

Description

Technical Field

[0001] The present invention relates to aircraft structural components such as stringers and methods for manufacturing aircraft structural components. Background Technology

[0002] Aircraft structural components, such as elongated components like stringers, are used in aircraft to provide structural support and strength to the elongated portions of the aircraft. For example, wing stringers are used in aircraft wings and extend along the wingspan. Stringers can also be present in the aircraft fuselage, where they extend longitudinally along the fuselage. Summary of the Invention

[0003] According to a first aspect of the invention, an aircraft structural component for an aircraft is provided, the aircraft structural component including a web and a flange, the web extending from the flange along a first direction to a distal end, the distal end being away from the flange, wherein the web includes a plurality of layers, each layer having a plurality of fibers, wherein at least a first layer of the plurality of layers has a first portion of fibers oriented at a first angle relative to the first direction, and a second portion of fibers oriented at a second angle relative to the first direction, wherein the first portion of the layer is closer to the flange than the second portion of the layer, and the second angle is greater than the first angle.

[0004] The orientation angle of fibers determines the stiffness of a structural component. "Stiffness" can refer to, for example, the degree of resistance to deformation under loading and / or bending. By providing a first portion of fibers with a first orientation and a second portion of fibers with a second orientation, an aircraft structural component can provide variable stiffness along its web. Changing the fiber orientation on the web, rather than using a uniform angle across the entire web, can help maintain the strength of the structural component in multiple directions, thereby meeting structural requirements under multiaxial loading conditions.

[0005] By positioning the second portion of the fiber, oriented at a large angle relative to the first direction, closer to the distal end of the web, the aircraft structural component can have lower stiffness in the second portion. This allows impact energy near the edge of the aircraft structural component to be more effectively distributed throughout the entire structural component, potentially improving performance compared to aircraft structural components with higher stiffness, such as those closer to the web. This can reduce the likelihood of the fiber buckling when subjected to impacts, for example, from a tool dropped from a height.

[0006] Composite materials may experience internal delamination due to impact damage during manufacturing, assembly, and / or transportation; this impact damage may be invisible externally. Post-impact edge compression (CAIE) testing can be used to determine the required dimensions of aircraft components. This process may include impacting a representative sample of the aircraft component and evaluating the remaining compressive strength of said sample to determine the required dimensions of the aircraft structural component. By reducing the stiffness of the fibers located in a second section adjacent to the distal end of the web, aircraft structural components can meet CAIE test requirements with reduced weight compared to aircraft structural components, for example, where the fibers are oriented in a single direction between the flange and the distal end of the web.

[0007] The second part may be adjacent to the distal end of the web. The distal end of the web may form the edge of an aircraft structural component. The fibers in the second part may be oriented at a large angle relative to the first direction.

[0008] The first part can be adjacent to the flange. The distance between the flange and the distal end of the web can be the height of the web.

[0009] The fibers can be continuous fibers oriented parallel to each other. By arranging the fibers parallel to each other, the fibers can provide strength in the direction along the length of the flange, and therefore in the direction along the length of the aircraft structural component.

[0010] Aircraft structural components may include a transition section located between a first part and a second part, the transition section being positioned at a distance from the distal end that is between five and ten times the thickness of the web, and includes both five and ten times the thickness of the web. This arrangement has been found to include areas of the web that are susceptible to impact, for example, from a tool falling from a height, or from collisions during manufacturing, assembly, and / or transportation. The transition section between the first and second parts may be positioned at a distance from the distal end that is between two and five times the thickness of the web, and includes both two and five times the thickness of the web.

[0011] The web can have a thickness of 5 mm to 25 mm.

[0012] The transition section may include a portion of the web in which the fibers are bent. By providing bent fibers, it is not necessary to overlap fibers with different orientations, which can provide a lighter aircraft structural component compared to an arrangement where, for example, the first part is stacked on top of the second part.

[0013] The curved portion of the transition section can have a radius of curvature between 2 mm and 100 mm between the first and second portions of the fiber. This arrangement has been found to be suitable for manufacturing methods and provides a desired variation in the orientation of the fibers extending from the flange to the distal end of the web.

[0014] The first part may include a first group of fibers oriented parallel to each other, and the second part may include a second group of fibers oriented parallel to each other. The first part may include multiple groups, each group having fibers oriented parallel to each other, and the second part may include multiple groups, each group having fibers oriented parallel to each other.

[0015] The first fiber group can terminate within the transition section, and the second fiber group can begin within the transition section. This provides a relatively simple arrangement for changing fiber orientation without requiring a complex manufacturing process.

[0016] The transition may include an incremental step change between the first angle and the second angle. This can provide a gradual change in fiber orientation, and thus a gradual change in stiffness. Fibers may be arranged in groups, wherein each group has an incremental change in orientation between preceding and subsequent groups.

[0017] The incremental step size of the angle between the first and second sections of the fiber varies between 5 and 45 degrees. Each group can have an incremental change of 5 degrees in orientation between the preceding and subsequent groups. This arrangement reduces the risk of buckling compared to arrangements with an incremental step size greater than 45 degrees, and is easier to manufacture compared to arrangements with an incremental step size less than 5 degrees.

[0018] Arrangements of multiple fiber groups with different orientations allow designers to customize the mechanical properties of the stringers to address varying load conditions that may be encountered in different regions of the aircraft structure. For example, the stiffness at the edges of the web can be reduced compared to the stiffness at the portion of the web adjacent to the flange.

[0019] The first portion of the fiber can be oriented at an angle of approximately 0 degrees relative to a first direction, and the second portion of the fiber can be oriented at an angle of approximately 45 degrees. The fiber at an angle of approximately 0 degrees relative to the first direction can act as a reinforcing column in the web. The first direction can be equal to the height of the web.

[0020] The first portion of the fiber can be oriented at an angle of approximately 45 degrees relative to a first direction, and the second portion of the fiber can be oriented at an angle of approximately 75 degrees. This arrangement has been found to provide the desired strength for aircraft structural components.

[0021] Fiber composites can include preimpregnated (also known as "prepreg") fibers or "dry" fibers. Dry fibers can be infused with a matrix material, such as a resin material.

[0022] The fiber can include or be composed of carbon fibers. Carbon fibers can have superior properties compared to other types of fibers that can be used in composite layers, such as glass fibers. These superior properties can include, for example, higher strength and a higher strength-to-weight ratio. This can be particularly advantageous in aerospace applications, where high strength and low weight of structural components contribute to performance and efficiency.

[0023] The web may include a total number of fibers, wherein at least eight percent of the total number of fibers are oriented in each of four directions: a first and a second direction that are mutually perpendicular in a plane, and a third and a fourth direction that are mutually perpendicular in a plane and oriented at 45 degrees to the first and second directions.

[0024] Aircraft structural components can be stringers. Aircraft stringers can serve as elements to which wing or fuselage skin sections are attached for structural support. Aircraft stringers can have different cross-sectional shapes, such as T-shaped, L-shaped, or Ω-shaped.

[0025] A second aspect of the invention provides an aircraft structural assembly comprising an aircraft structural component according to the first aspect and a skin segment attached to a flange of the aircraft structural component.

[0026] This can be achieved, for example, by attaching the flange of an aircraft structural component to the skin section on the side of the flange opposite to the side from which the flange extends from the web. Attaching the flange can be achieved using mechanical fasteners such as bolts or rivets, or by welding (if the flange is made of a weldable material), or by adhesive bonding, or any other suitable method. The aircraft structural component can be, for example, the wing and / or fuselage of an aircraft.

[0027] A third aspect of the present invention provides an aircraft comprising aircraft structural components according to the second aspect.

[0028] A fourth aspect of the invention provides a method of manufacturing an aircraft structural component for an aircraft, the method comprising: providing a flange; and providing a web extending from the flange along a first direction to a distal end portion remote from the flange, wherein the web comprises a plurality of layers, each layer having a plurality of fibers, wherein at least a first layer of the plurality of layers has a first portion of fibers oriented at a first angle relative to the first direction and a second portion of fibers oriented at a second angle relative to the first direction, wherein the first portion of the layer is closer to the flange than the second portion of the layer, and the second angle is greater than the first angle.

[0029] This method allows for the manufacture of aircraft structural components according to the first aspect of the invention.

[0030] This method may include using rapid tow shearing technology to guide fiber orientation. Rapid tow shearing can provide a cost-effective method for changing the orientation of fibers in each layer of the web. Rapid tow shearing technology enables the fiber orientation to be designed into a specific configuration according to the desired structural characteristics of the aircraft component.

[0031] This method may include using rapid tow redirection technology to guide the orientation of fibers.

[0032] The method may include infusing the layer with resin.

[0033] Other features and advantages of the invention will become apparent from the following description of preferred embodiments of the invention, given by way of example only, with reference to the accompanying drawings. Attached Figure Description

[0034] Embodiments of the invention will now be described by way of example only with reference to the accompanying drawings, in which:

[0035] Figure 1 A schematic diagram of an aircraft according to an exemplary embodiment of the present invention is shown.

[0036] Figure 2 An aircraft structural component according to an exemplary embodiment of the present invention is shown.

[0037] Figures 3A to 3C A side view, a cross-sectional view, and a top view of an aircraft structural component according to an exemplary embodiment of the present invention are shown respectively.

[0038] Figure 4 A side view of an aircraft structural component according to an exemplary embodiment of the present invention is shown.

[0039] Figure 5 A side view of an aircraft structural component according to an exemplary embodiment of the present invention is shown.

[0040] Figure 6 A side view of an aircraft structural component according to an exemplary embodiment of the present invention is shown.

[0041] Figure 7 A side view of an aircraft structural component according to an exemplary embodiment of the present invention is shown.

[0042] Figure 8 A side view of an aircraft structural component according to an exemplary embodiment of the present invention is shown.

[0043] Figure 9 A side view of an aircraft structural component according to an exemplary embodiment of the present invention is shown.

[0044] Figures 10A to 10C Examples of continuous fiber composite layers, chopped filament composite layers, and composite stacks are shown respectively.

[0045] Figure 11 An example method for manufacturing an aircraft structural component according to an exemplary embodiment of the present invention is shown. Detailed Implementation

[0046] Figure 1 An example of an aircraft 100 according to the present invention is shown. The aircraft 100 includes one or more structural components, which include aircraft structural parts as described below. Figure 1 (Not shown in the image). Such components are included in the fuselage 110, wings 120, and tail 130.

[0047] exist Figure 2 The image shows a cross-section of one of the wings 120 passing through the aircraft 100. The wing 120 is an aircraft structural component and includes a skin section 210 defining its outer surface and spars 230 supporting the skin section 210. Figure 2 (shown in dashed lines) and rib 240 (in Figure 2 Only one rib of the rib section 240 is visible, along with a series of stringers 300. The rib section 240 has an opening 242 through which the stringers 300 pass. The stringers 300 are attached at their respective flanges 320 to the inner surface 212 (described below) of the skin section 210. 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 adhesive or welding can be used. The stringers are examples of aircraft structural components according to the invention.

[0048] exist Figure 3A The image shows an aircraft structural component according to an embodiment of the present invention. In this example embodiment, the component is... Figure 2 One of the stringers 300 shown. In other embodiments, the aircraft structural component may be another component, such as an elongated component, like a wing spars. The stringers 300 are disposed in the wing 120 to provide structural support to the wing 120 against loads borne by the wing 120 on the ground and in flight.

[0049] The stringer 300 includes a web 310 and a flange 320 located at a first end of the web 310. The web 310 extends perpendicularly from the flange 320 from the first end 314 to a second end 312 along a first direction (which may be a direction corresponding to the height H of the web 310), the second end 312 being located away from the flange 320. The web has a height H extending from the first end 314 to the second end 312. The flange 320 has a length L and a thickness T perpendicular to the length L. The stringer 300 in… Figure 3B The figure is depicted in cross-sectional view, where it is shown that the flange 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.

[0050] exist Figure 3C Another view of the stringer 300, viewed from top to bottom along the thickness T of the flange 320, is shown, in which the web 310 is shown in dashed lines. (See image below.) Figure 3B As is best understood in this embodiment, the web 310 and flange 320 provide an inverted "T" shaped cross-section for the stringer 300. Other arrangements are possible and may be desirable for different structural applications. In other examples, the cross-section may, for example, be "L" or "Ω" shaped.

[0051] The web 310 of the stringer 300 may include multiple layers. Each layer has multiple fibers. Figure 4 and Figure 5 A side view of an example first layer 400 of a plurality of layers is shown. Layer 400 has a first portion 430 of fibers oriented at a first angle θA relative to a first direction (shown as dashed lines) and a second portion 440 of fibers oriented at a second angle θB relative to the first direction. The first portion 430 of the layer is closer to the flange 320 than the second portion 440 of the layer, and the second angle θB is greater than the first angle θA. The first portion 430 of the layer may be adjacent to the flange. The second portion 440 of the layer may be adjacent to the second end 312. A transition region 450 connects the first and second portions, in which, in this example, the fibers 410 gradually bend to transition between the two different orientation angles.

[0052] The first direction is along the height H of the web. The fiber orientation is configured such that the second angle θB is greater than the first angle θA, thereby providing varying stiffness along the web in the height direction. Figure 4 In the example shown, θA is approximately 0 degrees relative to height H, and θB is approximately 45 degrees relative to height H. Figure 5 An alternative configuration is shown, where θA is approximately 45 degrees and θB is approximately 75 degrees.

[0053] The transition region 450 is characterized by a curved fiber path with a radius of curvature of approximately 25 mm. Other radius values ​​between 2 mm and 100 mm are also contemplated in various embodiments.

[0054] In alternative examples, angles θA and θB can have different sizes. For example, each of angles θA and θB can have any size between 0° and 90° relative to the stringer height H. In all embodiments, θB is greater than θA, and the first portion 430 is positioned closer to the flange than the second portion 440.

[0055] Figures 6 to 9 An alternative embodiment of the stringer 300 is shown, wherein the fibers are not curved, but arranged in groups. Each group consists of unidirectional fibers comprising fibers arranged side-by-side and aligned parallel to each other, the fibers extending in the same direction relative to the length of the stringer 300.

[0056] Figure 6 and Figure 7 Layer 700 is shown, comprising: a first fiber group 710 oriented at a first angle θC relative to the height H of the stringer 300; and a second fiber group 610 oriented at a second angle θD relative to the height H of the stringer 300. The first fiber group 710 and the second fiber group 610 may overlap each other at their adjacent edges to create an overlapping region that provides continuity between the groups while maintaining different orientation angles.

[0057] Figure 8 and Figure 9 An example of a layer 800 comprising multiple fiber groups arranged in sequence is shown. Layer 800 includes a first fiber group 705, a second fiber group 715, and a third fiber group 725. Layer 800 includes a transition region 850 that includes the end portion of the first group 705, the entire second fiber group 715, and the beginning portion of the third group 725.

[0058] In transition region 850, there are incremental step changes in the orientation angles of the fibers relative to the length of the stringer 300. The orientation angle for the second group is θF, and the orientation angle for the third group is θG. The orientation angle can decrease by approximately 15 degrees with each successive group, thus creating a gradual transition from one orientation to another along the length of the stringer.

[0059] This stepped arrangement provides an alternative to the aforementioned curved fiber implementation. The incremental variation in fiber orientation allows for controllable variations in mechanical properties along the length of the stringers. Any number of groups can be envisioned, with a range of incremental variations between adjacent groups.

[0060] The web 310 may include multiple layers 700 and / or 800, which may be arranged such that overlapping regions between fiber groups are staggered between successive layers 700 and 800. Alternatively, the first fiber groups 710, 705 and the second fiber groups 610, 715 may abut against each other without any overlapping regions. The positions of transition regions may be staggered between successive layers 700, 800 of the web 310 to avoid potential weakening lines aligning across the web 310.

[0061] In this embodiment, the stringer 300 is made of carbon fiber composite material. In other embodiments, the stringer 300 may be made of other materials, including, for example, other fiber composite materials, such as glass fiber or other composite materials. Composite materials are used in aircraft components because of their light weight and advantageous mechanical properties, such as strength and stiffness.

[0062] Composite materials such as carbon fiber composites or glass fiber composites can be formed by stacking layers to form a laminated structure. Web 310 may include such a laminated structure, for example comprising a continuous fiber layer 400 as described above and a chopped filament layer 500 as described in detail below.

[0063] Figure 10A The illustration shows an example of a continuous fiber layer 400 as described above. Figure 10B A second example of a composite layer is shown, and this second example of a composite layer is a chopped fiber layer 500. The chopped fiber layer 500 comprises carbon fibers 510 that are relatively short compared to the fibers 410 in the continuous fiber layer 400. In this example, the length of each fiber 510 is between 10 mm and 100 mm. The fibers 510 are arranged in the plane of the layer 500, but aligned relative to each other in random or at least multiple directions. Therefore, a single chopped fiber layer 500 has substantially equal strength in all directions within the plane of the chopped fiber layer 500.

[0064] In this embodiment, the web 310 is composed of a hybrid laminate 420, which includes a plurality of continuous fiber layers 400 and a plurality of chopped fiber layers 500, such as... Figure 10C As shown, the hybrid stack 420 includes one or two continuous fiber layers 400 located between pairs of chopped strand layers 500, and terminates at each of the corresponding continuous fiber layers 400 on each of the two opposite sides of the hybrid stack 420. The total number of layers in the hybrid stack 420 is 13.

[0065] In other embodiments, fewer or more continuous fiber layers 400 and / or chopped filament layers 500 may be provided in the hybrid laminate, and / or different proportions of continuous fiber layers 400 and chopped filament layers 500 may be present. Additionally or alternatively, the hybrid laminate 420 may include as described above regarding... Figures 6 to 9 The one or more layers mentioned above.

[0066] Preferably, at least half of the total number of layers are continuous fiber layers 400. In some examples, different stacking orders of the continuous fiber layers 400 and the chopped strand layers 500 are possible. Preferably, pairs of adjacent chopped strand layers in the chopped strand layers 500 are separated by at least one continuous fiber layer in the continuous fiber layers 400.

[0067] In this embodiment, the stringer 300 includes a hybrid laminate 420 such that the fibers 410 of the continuous fiber layer 400 are aligned with the height H of the web 310 of the stringer 300, and the layers 400 and 500 of the hybrid laminate 420 are stacked in the thickness direction of the web 310. In other embodiments of the aircraft structural component according to the invention, other hybrid laminates 420, as well as conventional continuous fiber layer laminates or other layered composite materials or structures, are possible.

[0068] In some embodiments, at least eight percent of the total number of fibers in the web 310 are oriented in each of four directions: a first and a second direction that are mutually perpendicular in a plane, and a third and a fourth direction that are mutually perpendicular in a plane and oriented at 45 degrees to the first and second directions.

[0069] Figure 11 An example method according to the present invention is shown. Method 1000 is a method for manufacturing an aircraft structural component. In this example, method 1000 is a method for manufacturing the aforementioned stringer 300.

[0070] Method 1000 includes providing a 1010 flange 320.

[0071] Method 1000 also includes providing a web 310 1020.

[0072] Providing the 1010 flange and the 1020 web includes forming the respective flange and web by a hybrid laminate 420. In this embodiment, providing the 1020 web and providing the 1010 flange occur simultaneously. In other embodiments, providing the 1010 flange may be performed before or after providing the 1020 web.

[0073] Providing a web 310 includes extending the web 310 from a flange 320 along a first direction to a distal end, the distal end being remote from the flange. The web 320 includes at least a first layer of a plurality of layers, the first layer having a first portion of fibers oriented at a first angle relative to the first direction and a second portion of fibers oriented at a second angle relative to the first direction. The first portion of the layer is closer to the flange than the second portion of the layer, and the second angle is greater than the first angle.

[0074] Method 1000 can use rapid fiber shearing technology to guide the orientation of the fibers, such that a first portion of the fiber adjacent to the flange 320 is oriented at a first angle relative to a first direction, and a second portion of the fiber adjacent to the distal end of the web 310 is oriented at a second angle relative to the first direction, such that the second angle is greater than the first angle. Rapid fiber shearing enables the manufacture of the aforementioned curved fiber 410.

[0075] Alternatively, method 1000 can use rapid tow redirection technology to guide the orientation of the fibers.

[0076] The method 1000 of this embodiment includes resin infusing layers in the web 310 and flange 320 of the stringer 300. In some embodiments, the resin infusion component 1030 may be omitted, for example, when using prepreg fibers instead of dry fibers. When performing this resin infusion process, a resin curing process may follow.

[0077] It should be noted that, unless otherwise expressly stated, the term “or” as used herein shall be interpreted as “and / or”.

Claims

1. An aircraft structural component, used in an aircraft, the aircraft structural component comprising a web and a flange, the web extending from the flange along a first direction to a distal end, the distal end being located away from the flange. in, The web comprises multiple layers, each layer having multiple fibers. Wherein, at least a first layer of the plurality of layers has a first portion of fibers oriented at a first angle relative to the first direction, and a second portion of fibers oriented at a second angle relative to the first direction. Wherein, the first portion of the layer is closer to the flange than the second portion of the layer, and the second angle is greater than the first angle.

2. The aircraft structural component according to claim 1, wherein, The second portion is adjacent to the distal end of the web.

3. The aircraft structural component according to claim 1 or claim 2, wherein, The first portion is adjacent to the flange.

4. The aircraft structural component according to any one of the preceding claims, wherein, The fibers are continuous fibers oriented parallel to each other.

5. The aircraft structural component according to any one of the preceding claims, comprising a transition portion located between the first portion and the second portion, the transition portion being positioned at a distance from the distal end portion, the distance being between five and ten times the thickness of the web, and including five and ten times the thickness of the web.

6. The aircraft structural component according to claim 5, wherein, The thickness of the web is 5 mm to 25 mm.

7. The aircraft structural component according to claim 5 or claim 6, wherein, The transition section includes the portion of the web where the fibers are bent.

8. The aircraft structural component according to claim 7, wherein, The radius of curvature between the first portion and the second portion of the fiber is between 2 mm and 100 mm.

9. The aircraft structural component according to any one of claims 1 to 6, wherein, The first portion includes a first fiber group oriented parallel to each other, and the second portion includes a second fiber group oriented parallel to each other.

10. The aircraft structural component according to claim 9 when dependent on claim 5 or claim 6, wherein, The transition section includes an incremental step change between the first angle and the second angle.

11. The aircraft structural component according to claim 10, wherein, The incremental step size of the angle between the first and second portions of the fiber varies between 5 degrees and 45 degrees.

12. The aircraft structural component according to any one of the preceding claims, wherein, The first angle is approximately 0 degrees.

13. The aircraft structural component according to claim 12, wherein, The second angle is approximately 45 degrees.

14. The aircraft structural component according to any one of claims 1 to 11, wherein, The first angle is approximately 45 degrees.

15. The aircraft structural component according to claim 14, wherein, The second angle is approximately 75 degrees.

16. The aircraft structural component according to any one of the preceding claims, wherein, The fiber includes carbon fiber.

17. The aircraft structural component according to any one of the preceding claims, wherein, The web comprises a total number of fibers, wherein at least eight percent of the total number of fibers are oriented in each of four directions: a first and a second direction that are mutually perpendicular in a plane, and a third and a fourth direction that are mutually perpendicular in the plane and oriented at 45 degrees to the first and second directions.

18. The aircraft structural component according to any one of the preceding claims, wherein, The structural component of the aircraft is a stringer.

19. An aircraft structural component, comprising: Aircraft structural components according to any one of claims 1 to 18; as well as A skin section, the skin section being attached to the flange of the aircraft structural component.

20. An aircraft comprising the aircraft structural components according to claim 19.

21. A method for manufacturing an aircraft structural component for an aircraft, the method comprising: Provide flanges; as well as A web is provided, the web extending from the flange in a first direction to a distal end, the distal end being located away from the flange. The web comprises multiple layers, each layer having multiple fibers. Wherein, at least a first layer of the plurality of layers has a first portion of fibers oriented at a first angle relative to the first direction, and a second portion of fibers oriented at a second angle relative to the first direction. Wherein, the first portion of the layer is closer to the flange than the second portion of the layer, and the second angle is greater than the first angle.

22. The method according to claim 21, wherein, The method includes using rapid tow shearing technology to guide the orientation of the fibers.

23. The method according to claim 21, wherein, The method includes using rapid tow redirection technology to guide the orientation of the fibers.

24. The method according to any one of claims 21 to 23, wherein, The method includes infusing the layer with resin.

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