Method for manufacturing aircraft panel and aircraft panel

Abstract

The invention relates to a method for manufacturing an aircraft panel and an aircraft panel. The invention relates to a method for manufacturing an aircraft panel comprising a composite skin (240) and a one-piece component (220) (stringer / frame) having an unsupported slot (220a), the method comprising: obtaining (110) a male tool having the one-piece component (220); laying (120) the composite skin directly over the unsupported slots of the one-piece component (220) using an automatic fiber placement, i.e., AFP machine, that provides tow deposition, where the one-piece component (220) is free of auxiliary support tools; and adjusting (130) a force applied to the tow with the AFP machine during deposition over the unsupported tank to compensate for the absence of the auxiliary support tool.

Description

Technical Field

[0001] The present invention relates to a method for the automated laying of composite materials to obtain aircraft panels, the aircraft panels comprising skin and integral components, such as stringers or frames.

[0002] This invention relates to the field of aerospace manufacturing, and more particularly to a method and apparatus for the automated laying of composite materials (carbon, glass, aramid, ceramic, etc.) used in the production of composite skins for aircraft panels. The invention particularly relates to the production of large, complex composite material structures in the aerospace industry, which includes both commercial and military aircraft. Background Technology

[0003] In the field of aircraft manufacturing, the use of composite materials to manufacture panels has become increasingly common due to their strength-to-weight ratio and other advantageous properties. Traditionally, the process of laying carbon fiber material over integral struts in fuselage / wing / VTP / HTP (vertical tail plane, horizontal tail plane) panels involves several auxiliary operations. These operations include using reaction mandrels, cladding, and tubular bags to support and resist the pressure applied by automated fiber placement (AFP) machines during tow deposition.

[0004] The standard process requires a reaction mandrel to be housed inside the hollow stringers to compensate for the pressure applied by the AFP machine. This mandrel must be removed before the vacuum bag is introduced and parts are prepared for the autoclave. Additionally, an auxiliary mandrel may be needed to ensure proper positioning of the vacuum bag, further increasing the complexity and time of the manufacturing process. These steps not only increase delivery time and recurring costs but also require precise coordination and handling, which can introduce variability and potential errors.

[0005] Given the competitive pressures in the aerospace industry to reduce production costs and increase efficiency, there is a significant need for improved manufacturing methods that minimize auxiliary operations while maintaining or improving the quality of the final product. Innovations that can simplify the laying-up process and reduce reliance on auxiliary tools and materials are highly anticipated.

[0006] This invention satisfies this need. Summary of the Invention

[0007] This invention relates to an advantageous method for automatically laying composite materials to obtain aircraft panels comprising composite skins and integral components (such as stringers or frames). This method eliminates the need for reaction mandrels and other auxiliary supports.

[0008] Therefore, a first aspect of the invention relates to a method for manufacturing an aircraft panel comprising a composite skin and at least one integral component (i.e., stringer or frame) having unsupported grooves. The method includes a first step of obtaining a convex tool having the integral component (e.g., an omega-shaped, triangular, or semi-circular stringer or frame) having unsupported grooves. The method includes a second step of directly laying the composite skin over the unsupported grooves of the integral component using an automated fiber placement (AFP) machine that provides filament deposition, wherein the integral component lacks auxiliary support tools. Furthermore, the method includes a third step of adjusting the force applied to the filaments using the AFP machine during deposition over the unsupported grooves to compensate for the absence of auxiliary support tools.

[0009] In a preferred example, the step of adjusting the force applied to the filament includes increasing the filament tension while reducing the compaction pressure applied to the filament using an AFP machine during deposition above an unsupported groove.

[0010] The key innovation lies in directly laying composite skin material over unsupported grooves in a monolithic component (hollow stringers or frame) without internal supports by optimizing (e.g., increasing) filament tension and controlling (e.g., reducing) the compaction pressure applied using AFP machine rollers. The proposed method ensures precise filament placement and bonding without additional support mechanisms.

[0011] The elimination of auxiliary operations not only simplifies workflows but also significantly reduces delivery times and recurring costs associated with the manufacturing process. This new strategy enhances flexibility, allowing vacuum bags to be positioned before or after skinning, thereby providing greater accessibility and adaptability for different manufacturing scenarios.

[0012] In some examples, the step of adjusting the force applied to the filament bundle is performed dynamically based on real-time feedback to maintain the desired filament bundle trajectory and position.

[0013] In the first alternative, the method further includes the step of positioning a vacuum bag (tubular bag) inside a cavity (wherein a cavity or hollow part means an unsupported groove) of an integral component (stringer or frame) without the use of an auxiliary mandrel after the composite skin material is laid.

[0014] In a second alternative, the method further includes the step of positioning a tubular vacuum bag inside the cavity of the integral component (stringer or frame) without the use of an auxiliary mandrel before laying the composite skin material.

[0015] In a preferred example, the skin is made of composite prepreg.

[0016] According to another preferred example, laying the composite skin material includes applying a bundle of filaments in multiple layer orientations, wherein one of the multiple layer orientations is equal to 0 degrees relative to a reference direction.

[0017] According to another preferred example, laying the composite skin involves making the layers continuous and cutting them in an unsupported groove (or cavity) of the monolithic component. Thus, once several layers have been laid, the method allows for detachment within the cavity of the monolithic component. This means that the wire bundle can be cut in the middle of the cavity of the monolithic component (stringer / frame) even though there is no reaction mandrel inside the cavity.

[0018] Furthermore, the applicability of this invention to both flat and curved parts, as well as its compatibility with various material specifications, fiber region weights, and layer orientations, underscores its versatility and potential industrial impact.

[0019] Some of the key features and advantages of this invention include:

[0020] Lay directly above the components (stringers or frame):

[0021] This process involves laying the skin material directly over the part groove, which is integrated into the convex tool, without requiring internal supports. This "flying" lay-up technique eliminates the need for any auxiliary tools to support machine pressure during filament deposition.

[0022] Reduced operations and costs:

[0023] By eliminating the use of reaction mandrels and other auxiliary materials, the number of operations is significantly reduced. This results in minimized delivery time, reduced process cycle time, and significant savings in recurring costs.

[0024] Enhanced flexibility and efficiency:

[0025] This process allows for the introduction and positioning of internal tubular vacuum bags before or after skin laying, providing complete accessibility and versatility. This change simplifies the workflow and reduces the need for multiple auxiliary mandrel placements.

[0026] Optimized machine parameters:

[0027] This method involves specifically adjusting the force applied to the filament by controlling the compaction pressure and filament tension of the AFP machine rollers above the unsupported stringers and frame grooves. This ensures correct filament trajectory and final position without additional supports.

[0028] Wide applicability:

[0029] This invention is applicable to both flat and curved parts, and can accommodate various materials and specifications, fiber area weights, and layer orientations. This invention provides a competitive alternative to traditional metal solutions in panel manufacturing.

[0030] Feasibility and quality assurance:

[0031] Extensive testing has been conducted to verify the feasibility of this method. The final product quality meets the same stringent requirements as the current manufacturing process.

[0032] In summary, this new laying strategy offers significant improvements in cost, time, and process efficiency for manufacturing composite panels, making it a valuable advancement in the field of aircraft production. Attached Figure Description

[0033] To better understand the above description and for illustrative purposes only, some non-limiting drawings are included, which schematically depict actual implementations.

[0034] Figure 1 The steps of a method for manufacturing an aircraft panel according to the present invention are shown.

[0035] Figure 2 The components involved in the method for manufacturing an aircraft panel according to the present invention are shown.

[0036] Figure 3 The image shows the use of an AFP machine with rollers to lay the composite layer onto a monolithic omega stringer.

[0037] Figure 4 Multiple first composite layers with a first layer orientation are shown.

[0038] Figure 5 A second set of composite layers with a second layer orientation is shown.

[0039] Figure 6 Several sets of composite layers with different layer orientations are shown.

[0040] Figure 7 Several sets of composite layers with different layer orientations are shown, wherein one of the orientations is 0°. Detailed Implementation

[0041] Figure 1 The steps of a method 100 for manufacturing an aircraft panel according to the present invention are shown.

[0042] Method 100 for manufacturing includes the step of obtaining a convex tool 110, the convex tool comprising at least one integral component (i.e., stringer or frame) having an unsupported groove. The convex tool can be designed to incorporate the stringer / frame prior to the laying process, thereby eliminating the need for additional positioning operations for the reaction core.

[0043] Method 100 for manufacturing includes the step of directly laying 120 of the composite skin over an unsupported groove of a monolithic component using an AFP machine that provides fiber layers or fiber strips as filament deposits. During the laying 120 step, no auxiliary support tools are used on the monolithic component.

[0044] Conventional manufacturing processes use reaction mandrels during the laying of the skin over the stringers or frame already integrated into the tool. The proposed method avoids the use of any type of auxiliary tool, such as internal cladding, reaction mandrels, internal mandrels, or any tool that supports or resists the pressure of the AFP machine when placing the wire bundle during the process of laying the skin over the stringers / frame already integrated into the tool via the AFP machine.

[0045] The manufacturing method 100 includes a step of adjusting the force applied to the filament bundle using an AFP machine 130 during deposition above an unsupported tank to compensate for the absence of auxiliary support tools.

[0046] The step of adjusting the 130 force involves reducing the compaction pressure applied over the unsupported slot using the AFP machine (e.g., reducing the pressure to 0 Pa) while increasing the tension applied to the filament bundle already positioned in the tool. Therefore, this step is a combination of locking and unlocking the compaction of the AFP machine rollers along the Z-axis along the filament bundle deposition and / or filament bundle trajectory, primarily focused on the unsupported slot or hollow section of the stringer / frame where there is no auxiliary support from the tool.

[0047] Figure 2 The components involved in the method for manufacturing an aircraft panel according to the present invention are shown. Figure 2 The diagram shows a convex tool 210, an integral component 220 with an unsupported groove 220a (for this example, an omega-shaped stringer), a tubular vacuum bag 230, and a final composite skin 240.

[0048] First, the convex tool 210 is designed to engage component 220 prior to the laying process, thereby eliminating the need for additional positioning operations on the reaction core during step 110.

[0049] Next, during step 110, the integral component 220 having the unsupported groove 220a is attached to the convex tool 210.

[0050] Optionally, a tubular vacuum bag 230 is included in the unsupported groove 220a. The purpose of the tubular vacuum bag 230 is only to ensure the form (configuration) of the integral part during the curing cycle, rather than to act as a reaction mandrel.

[0051] Fourth, the composite skin is directly laid on the unsupported groove 220a of the integral component 220 using an AFP process (AFP machine) that provides filament deposition, wherein the integral component 220 is not supported by any auxiliary support tool during step 120.

[0052] Fifth, during the laying process, the force applied to the filament using an AFP machine during deposition above the unsupported trench 220a is adjusted to compensate for the absence of auxiliary support tools. Specifically, the force applied to the filament includes increasing the filament tension while reducing the compaction pressure applied to the filament using the AFP machine during deposition above the unsupported trench 220a. In a preferred example, the AFP machine does not apply any pressure to the unsupported trench 220a at all.

[0053] This adjustment can be performed dynamically based on real-time feedback to maintain the desired filament trajectory and filament position.

[0054] Additional steps may include positioning a vacuum bag over the skin material after laying the 120 composite skin material, without using an auxiliary mandrel. The introduction and positioning of the vacuum bag can be completed before skin laying and is fully accessible without using any auxiliary mandrels or additional components. During the laying process, reduced compaction pressure can be applied to the filament bundle on the unsupported groove 220a using an AFP machine, or no compaction pressure can be applied at all.

[0055] Alternatively, the step of positioning the tubular vacuum bag inside the cavity above the integral component can be performed before laying the composite skin material, without using an auxiliary mandrel. The vacuum bag can be placed after the composite skin is laid, if needed, since the laying can be done while the slots in the component are empty.

[0056] Figure 3 Parts of a system 1000 for manufacturing an aircraft panel are shown, the aircraft panel including a composite skin 240 and an integral component 220.

[0057] System 1000 includes a convex tool 210 configured to engage component 220 with an unsupported groove 220a. Figure 3 An AFP machine 250 is shown, which has a roller 260 that provides a filament bundle 270 and is configured to lay a composite skin 240.

[0058] Figure 3The image shows an AFP machine 250 depositing a filament bundle 270 directly over an unsupported groove 220a. As shown, the omega-shaped stringers are free of auxiliary support tools.

[0059] Figure 4 , Figure 5 and Figure 6 Different layer orientations for different groups of composite layers are shown during the laying process.

[0060] In particular, Figure 4 A first set of composite layers or filament bundles 270 with a first layer orientation is shown. Figure 5 A second set of composite layers or filament bundles 270 with a second layer orientation is shown. Figure 6 Three sets of composite layers or filament bundles 270, each with different layer orientations, are shown.

[0061] These previous figures illustrate components of a system 1000 for manufacturing an aircraft fuselage panel, which includes a composite skin 240 and a monolithic component 220. These figures also show a convex tool 210 configured to engage the monolithic component 220 with an unsupported groove 220a.

[0062] Figure 7 The diagram illustrates several sets of composite layers with different layer orientations applied during installation, as well as another set of layers (A) with a layer orientation of 0°, wherein the composite fibers are aligned with a reference direction to provide maximum strength and stiffness in that direction. After applying composite layers with, for example, three or four different layers (depending on the material), the applied laminate behaves like a conventional laminate obtained using a reaction mandrel. Thereafter, it is feasible to apply another set of layers (A) with a layer orientation of 0° to obtain a laminate capable of withstanding the ordinary (conventional) compaction pressure applied by an AFP machine.

Claims

1. A method (100) for manufacturing an aircraft panel, the aircraft panel comprising a composite skin (240) and at least one integral component (220) having an unsupported groove (220a), the integral component (220) being a stringer or frame, the method comprising: Obtain (110) a convex tool comprising the integral component (220) having the unsupported groove (220a); The composite skin (240) is directly laid (120) over the unsupported groove (220a) of the integral component (220) using an automated fiber placement machine (250) that provides filament deposition, wherein the integral component (220) has no auxiliary support tools. as well as During deposition above the unsupported trough (220a), the adjustment (130) utilizes the force applied to the filament by the automatic fiber placement machine (250) to compensate for the absence of the auxiliary support tool.

2. The method according to claim 1, wherein, The force applied to the filament by adjustment (130) includes reducing the compaction pressure applied to the filament by the automatic fiber placement machine (250) and increasing the filament tension during deposition above the unsupported groove.

3. The method according to the preceding claim, wherein, Reducing the compaction pressure applied to the filament bundle includes providing a compaction pressure of 0 Pa.

4. The method according to claims 1 to 3, wherein, The force applied to the filament by adjustment (130) is dynamically performed based on real-time feedback to maintain the desired filament trajectory and filament position.

5. The method according to claims 1 to 4, further comprising the step of positioning a vacuum bag above the composite skin (240) without using a mandrel after laying (120) the composite skin (240).

6. The method according to claims 1 to 5, further comprising the step of positioning a tubular vacuum bag (230) above the integral component (220) inside the unsupported groove (220a) without using a mandrel before laying (120) the composite skin (240).

7. The method according to any one of the preceding claims, wherein, Laying (120) the composite skin (240) includes applying the filament bundle in multiple layer orientations, wherein one of the multiple layer orientations is equal to 0 degrees relative to a reference direction.

8. The method according to any one of the preceding claims, wherein, Laying (120) the composite skin (240) includes making the layers continuous and cutting the layers in the unsupported groove (220a) of the integral component (220).

9. An aircraft panel obtained by the method according to any one of claims 1 to 8, the aircraft panel comprising a composite skin (240) and an integral component (220), the integral component (220) being a stringer or frame.

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