Composite batten supports for beam panels

Abstract

A composite slat support for a stringer panel. The present invention provides a reinforced stringer panel with an integrated slat structure. An example composite panel includes a skin member having an inner surface, and a slat on the inner surface. The slat includes a set of layered laminate plies and extends from a first side to a second side. The size of each laminate ply in the set of layered laminate plies is set to form a desired geometric profile of each of the first side and the second side. The composite panel also includes a stringer, which includes a cap portion spanning from the first side of the slat to the second side of the slat to form a first flange portion and a second flange portion on the inner surface of the skin member, respectively. The base section of the cap portion conforms to the desired geometric profile of the respective first side and second side of the slat.

Description

Technical Field

[0001] The present disclosure relates generally to aircraft support structures and, more particularly, to stringer panels. Background Art

[0002] In the construction of aircraft and launch vehicles, various surfaces such as the skin of the aircraft may be attached to structural support members known as stringers or stiffeners. In a typical aircraft fuselage, stringers are attached to the fuselage skin and extend in the longitudinal direction of the aircraft. They are primarily responsible for transferring aerodynamic loads acting on the skin to internal structures including the frames. In an aircraft wing or horizontal stabilizer, stringers are attached to the wing / horizontal stabilizer skin and their primary function is to transfer bending loads acting on the wing to internal structures such as ribs and spars.

[0003] However, defects in the support structure can lead to existing methods of manufacturing reinforcement members on composite parts, such as deformation of the filler shape, resin bleed-out, cracking and wrinkling. Repairing or mitigating such defects can require additional labor and materials and is often tool-intensive. Given the scale of aircraft wings and fuselages, the forming and handling equipment used to form and handle these parts can also be expensive, cumbersome, and require a large amount of factory floor space.

[0004] Therefore, there is a need for improved systems and methods for manufacturing aircraft support structures that reduce manufacturing challenges and the need for labor, extensive tooling, factory space, and material handling equipment. Summary of the invention

[0005] A simplified summary of the present disclosure is provided below to provide a basic understanding of some examples of the present disclosure. This summary is not an extensive overview of the present disclosure, and it does not identify the key and important elements of the present disclosure, nor does it describe the scope of the present disclosure. Its sole purpose is to present some concepts disclosed herein in a simplified form as a prelude to a more detailed description presented later.

[0006] A reinforced composite panel for various aircraft structures and a method for manufacturing such a panel are provided. Specifically, a composite panel (100) includes a skin member (510) having an inner surface (510-A). The composite panel also includes a strip (400) on the inner surface. The strip includes a group of layered laminated plies (420, 422, 424) extending from a first side (402) to a second side (404). The size of each laminated ply in the group of layered laminated plies is set to form a desired geometric profile of each of the first side and the second side. The composite panel also includes a stringer (520), which includes a cap portion (521), wherein the cap portion spans from the first side of the strip to the second side of the strip to form a first flange portion (524-A) and a second flange portion (524-B) on the inner surface of the skin member, respectively, each flange portion extending from the cap portion.

[0007] The stringer may include a plurality of stringer plies. A first base section (522-A) of the cap portion contacts a first side of the slat and conforms to a desired geometric profile of the first side of the slat. A second base section (522-B) of the cap portion contacts a second side of the slat and conforms to a desired geometric profile of the second side of the slat.

[0008] The support tool (530) may be positioned on top of the uppermost laminate ply in the set of layered laminate plies such that the main section (526) of the cap portion joining the first base section and the second base section is supported by the support tool. The laths, stringers and skin members may be co-cured. The laths, stringers and skin members may be co-bonded. The support tool may be a silicon forming mandrel including 20% ​​silica microspheres uniformly mixed therein.

[0009] Each of the layered laminate plies may include composite fibers pre-impregnated with resin. The composite fibers of adjacent laminate plies may have different orientation angles.

[0010] Other embodiments of the disclosure include systems and methods corresponding to the described apparatus.For example, in another aspect that may include at least a portion of the subject matter of any of the foregoing and / or following examples and aspects, an aircraft is provided that includes one or more composite panels as described above.

[0011] A method for manufacturing the composite panel is also provided. The method includes providing (610) a skin member having an inner surface, and positioning (620) a strip on the inner surface of the skin member. The strip includes a set of layered laminate plies and extends from a first side to a second side. Each laminate ply in the set of layered laminate plies is sized to form a desired geometric profile for each of the first side and the second side. The method also includes placing (630) a stringer on a support tool (530), and positioning (640) the support tool and the stringer on the uppermost laminate ply in the set of layered laminate plies. The method also includes joining (650) the skin member, the strip, and the stringer.

[0012] These and other examples are further described below with reference to the figures. BRIEF DESCRIPTION OF THE DRAWINGS

[0013] Figure 1 Examples of reinforced stringer panels that may be implemented with various examples of the present disclosure are shown.

[0014] Figure 2 A cross-sectional view of an example of a prior art stringer panel assembly system is shown.

[0015] Figure 3 A cross-sectional perspective view of another prior art stringer panel assembly system having a manufacturing defect is shown.

[0016] Figure 4A , Figure 4B and Figure 4C Example slats are shown according to one or more examples.

[0017] Figure 5A and Figure 5B A cross-sectional view of a stiffened stringer panel incorporating battens is shown according to one or more examples.

[0018] Figure 6 An example process sequence for manufacturing a stiffened stringer panel is shown according to one or more examples.

[0019] Figure 7 An example process for manufacturing a reinforced stringer panel having battens is shown according to one or more examples.

[0020] Figure 8 is a schematic diagram of an aircraft that may include a stiffened stringer panel as described herein, according to one or more examples.

[0021] Fig. 9 is a block diagram of an aircraft production and service method that may utilize the methods and assemblies described herein. DETAILED DESCRIPTION

[0022] In the following description, many specific details are set forth in order to provide a thorough understanding of the concepts presented. The concepts presented may be practiced without some or all of these specific details. In other cases, well-known process operations are not described in detail to avoid unnecessarily obscuring the concepts described. Although some concepts will be described in conjunction with specific examples, it should be understood that these examples are not intended to be limiting. On the contrary, it is intended to cover substitutions, modifications, and equivalents that may be included within the spirit and scope of the present disclosure defined by the appended claims.

[0023] For example, the technology of the present disclosure will be described in the context of a specific aircraft structure such as an aircraft panel. However, it should be noted that the technology and mechanism of the present disclosure can be applied to various other vehicles or various other panel assemblies of building structures. In the following description, many specific details are set forth in order to provide a thorough understanding of the present disclosure. Specific examples of the present disclosure can be implemented without some or all of these specific details. In other cases, well-known processing operations are not described in detail so as not to unnecessarily obscure the present disclosure. For clarity, various technologies and mechanisms of the present disclosure will sometimes be described in singular form. However, it should be noted that, unless otherwise stated, some examples include multiple iterations of the technology or multiple instantiations of the mechanism.

[0024] Overview

[0025] The present disclosure describes a novel reinforced stringer panel assembly for aircraft and other vehicles or industrial systems. As described herein, the terms "stringer," "stringer," "strengthening rib," "support structure," and "support member" may be used interchangeably. The reinforced stringer panel described herein includes a slat structure configured with sides that include a desired geometric profile for supporting and shaping the stringers during the manufacturing process. As used herein, the slat structure may be referred to simply as a slat. Support tools are used in the reinforced stringer panel during the build process to support the shape of various support structures, such as a reinforcing rib with an internal cavity, during the curing process.

[0026] The lath may include multiple layers of composite material stacked on top of each other. These plies may include composite fibers pre-impregnated with resin or prepreg. The lath may be multi-directional, where the composite fibers are arranged in different orientations in adjacent layers, which can increase the strength of the lath. The width of the laminated plies can be varied so that the lath includes a desired geometry corresponding to the parts of the stringers so as to support the stringer structure. The lath can be placed in the desired position on the skin member of the panel or other higher-level assembly components. The stringers (such as hat stringers) can then be placed on the lath together with the mandrel so that during the manufacturing process, the stringers are supported by the lath and the mandrel.

[0027] Existing manufacturing processes for reinforced stringer panels include the use of radius fillers, such as noodle-type fillers / facers, to support the space between assembly components and support tools. However, such facers are fragile and may crack or deform under the high temperatures and pressures of the curing process. This results in defects in the final panel assembly, which requires additional inspection, repair, or disposal / discarding, resulting in additional labor and materials.

[0028] Stronger, stiffer laths not only increase the structural integrity of the entire panel assembly, but also provide more stable support for smaller radius stringers under the extreme conditions of the curing process. Composite fibers in the filler are typically unidirectional, resulting in an undesirably different stiffness and strength of the filler structure relative to the surrounding structure compared to the lath. As a result, the filler often shifts or moves during the manufacturing process, forming voids between components. Such voids may be further exacerbated by restrictions on the filler geometry. Such voids can form low pressure areas, leading to resin seepage, layer wrinkling and deformation (such as bending), and radius thinning or thickening. Flat laths avoid the need for sharp vertical tips and allow for more precise assembly, eliminating voids and the problems caused by voids.

[0029] Thus, the described systems and assemblies are improvements over existing systems for making structural support members that rely on faceting structures as radius fillers. The described assemblies provide a more rigid support structure that improves component geometry and increases the overall strength of the panel assembly.

[0030] Example Device

[0031] refer to Figure 1 Examples of reinforced stringer panels 100 that may be implemented with various examples of the present disclosure are described. Figure 1 As shown, panel 100 is defined by a transverse axis (X axis), a longitudinal axis (Y axis), and a vertical axis (Z axis). In some examples, panel 100 may be a portion of a shell of an aircraft fuselage. However, in other examples, panel 100 may be a portion of various other structures. For example, various panels may include portions of structures in skin panels, wing structures, or horizontal and vertical stabilizers or control surfaces.

[0032] like Figure 1 As shown, panel 100 includes a shell member 110 having an inner surface 110-A. In various examples, the structure of panel 100 may include a frame 106. Panel 100 may be reinforced using one or more reinforcing support structures (such as stringers 120) coupled to panel 100 along inner surface 110-A. In some examples, stringers 120 may be configured such that the length of stringers 120 is oriented substantially parallel to flight direction 105, which is parallel to the Y axis, such as Figure 1 However, in various examples, the reinforcement support structure may be oriented at various other angles and orientations based on the load paths and deflection directions in order to reduce the loads and deflections of the primary structure. Additionally, the reinforcement support structure may include other orientations driven by other functions, such as required locations for routing of stringers or bracket attachment for exhaust gas.

[0033] The stringers 120 may be configured with various cross-sectional shapes for different structural properties. The various stringer types may include closed cross-sectional shapes such as hat stringers, or open cross-sectional shapes such as L-shaped stringers. Other formed stringer types may include blade stringers, Z-shaped stringers, C-shaped stringers, etc. Figure 2 A cross-sectional view of an example of an existing stringer panel assembly system 200 is shown. In some examples, the system 200 may be a panel 100, and Figure 2 Depicts the AA viewpoint corresponding to the longitudinal axis (at Figure 1 A cross section of a portion of the panel observed (marked in).

[0034] As shown, the existing panel assembly system 200 includes a skin member 210 having a hat-shaped stringer 220 along an inner surface 210-A. In some examples, the stringer 220 may be positioned on the base charge 211 if desired for structural and manufacturing quality purposes. However, in other examples, the stringer 220 may be placed directly on the skin member 210.

[0035] In various examples, the assembly of the panels may require support tools to provide structural or positional support for the stringers while they are cured or otherwise set in place. For example, the support tool 230 may be a formed mandrel. Such mandrels may include any of various types of mandrels, such as solid rubber mandrels, expanded rubber mandrels, scour mandrels formed from clay or powder, and splash foam mandrels. In other examples, the support tool may be of the inflatable bladder type. In some examples, the support tool may include various radius fillers, such as caulking 240, which may function to fill gaps between the stringers 220, the support tool 230, and the base charge 211 or the skin member 210 in an attempt to prevent resin accumulation during curing.

[0036] In stringer panel assemblies, such support tools must be precisely aligned on the skin members or base charge in order to precisely control the position of the stringers and the positioning and shape of the stringers. In existing panel assemblies, the placement and / or alignment of the support tools and / or stringers is typically accomplished with large tool assist tools to lift, transport and place the support tools. Accurate placement may be additionally achieved by various methods, including the use of optical laser templates to guide manual placement within the boundaries defined by the laser. Various other alignment mechanisms may be implemented to guide the manual placement of the support tools. Efforts to ensure precise manual alignment of the support tools may further result in increased labor requirements and extended build times.

[0037] In addition, the support tool generally includes a long length, which can reach a length equal to the length of the stringer. As previously mentioned, the stringers used for the 777X aircraft can be up to 110 feet long. Thus, a 110-foot solid rubber mandrel can include considerable weight. Such large support tools may also require a large amount of handling tools and equipment to form, flip, position and transport the support tools. For example, handling and pick-and-place equipment can be used to transport and manipulate such tools. Due to their size, the size of the support tool placement will also change. This may further affect manufacturing efficiency and increase the cost of labor time and manufacturing productivity.

[0038] like Figure 2 As shown, the filler face is generally an elongated structure with a triangular or triangular cross-section. Radius fillers, such as fillers, are generally made of laminated tapes including pre-impregnated resin and carbon fiber or other substances with similar material properties in different proportions. However, such materials can be flexible and fragile. This characteristic, as well as the long length, can further exacerbate the difficulties in transporting, handling, and positioning such support tools. Such flexibility can also cause excessive movement between parts during the manufacturing process, which can damage the filler face or create gaps that lead to undesirable defects.

[0039] refer to Figure 3 , a cross-sectional perspective view of another prior art stringer panel assembly system 300 is shown illustrating a commonly occurring manufacturing defect. The prior art stringer panel assembly system 300 may include a stringer 320 positioned on a skin member 310 with a filler face 340. The prior art stringer panel assembly system 300 is shown with a forming mandrel, such as the support tool 230, removed. In various examples, the filler face 340 is typically used to fill in radii between the geometry of the stringer, the forming mandrel, and the skin member.

[0040] Visible Figure 3The filler 340 in the figure is deformed, which may be caused by the high pressure and high temperature conditions during the curing process. The stacked manufacturing of the filler also places restrictions on the geometry of the filler, especially at the apex angles of the triangular structure. The layers of the filler must be folded up to form the apex angles, which may make the vertical tips blunt. As a result, the geometry of the filler may not correspond accurately enough to the stringers, mandrels and / or skin components, leaving gaps. Such gaps form low-pressure areas, and resin from the stringer plies may flow into the low-pressure areas, resulting in resin seepage between the stringer plies and wrinkling of the layers of the stringer or skin component, resulting in surface unevenness. For example, resin seepage 356 between the stringer plies causes wrinkling of the stringer plies. In some examples, resin seepage between the plies of the stringer or skin component may cause wrinkling of such plies as shown in the figure. In addition, deformation of the filler during the curing process and differences in the coefficient of thermal expansion (CTE) between the materials of the support tool and the stringer may further exacerbate such irregularities during the manufacturing process.

[0041] Other irregularities that may occur in existing systems include cracking and wrinkling caused by the wrap made of fluorinated ethylene propylene (FEP) or other plastics around the support tooling (such as a mandrel). Wrap wrinkle 354 is an example of a wrinkle that may be caused by a portion of the FEP wrap moving into the inner surface of the stringer during the co-curing process under vacuum pressure. Deformation of the filler may also produce bow waves 352 in the layers of the skin member, which leads to deformation of the panel geometry and potential structural problems.

[0042] Such shim deformation may also result in thinning or thickening of the radius. As used herein, an upper radius refers to the curved portion at the top of a stringer, such as upper radius 321, while a lower radius refers to the curved portion at the bottom of a stringer, such as lower radius 323. The panel assembly may be vacuum bagged to provide uniform clamping pressure during the co-curing or co-bonding process. Due to the convex geometry of the upper radius 321, greater forces are applied from the vacuum bag to the upper radius against the mandrel, which typically results in thinning of the radius or reduced stringer thickness at these areas. In contrast, due to the concave geometry of the lower radius 323, less force is applied from the vacuum bag to the lower radius against the shim or other support structure, which typically results in thickening of the radius or thicker stringers at such areas.

[0043] The deformation of the filler can further reduce the pressure applied to the lower radius, thereby further thickening the radius. Figure 3The filler structure may be strong in the axial direction along the Y axis because the filler structure extends along the Y axis as shown. However, this causes the filler to have an undesirably different stiffness and strength relative to forces in other directions (along the X and Z axes) where the structure needs to be supported. This weakness, along with the CTE differences associated with the fiber directions, may cause further deformation and voids when forces are applied in other directions due to internal stresses during the curing process.

[0044] Such defects may interfere with inspection of panel assemblies during production and manufacturing. Variations in stringer thickness may interfere with ultrasonic pulse-echo techniques used to determine the porosity of the structure. Defects may make pulse-echo measurements more difficult, slowing the process. In cases where measurements are interfered with, safety requirements may dictate that porosity levels be assumed to exist, requiring additional labor and materials to reinforce the structure in specific areas.

[0045] Such manufacturing challenges can be reduced or eliminated by using solid plank laminates as support means and using radius fillers instead of long, fragile infills. Figure 4A , Figure 4B and Figure 4C , an example slat 400 is shown according to one or more examples. Figure 4A depicts a perspective view of a slat 400, and Figure 4B A front view of slat 400 is depicted.

[0046] The panel 400 may include a set of layered laminated plies of a composite material. Thus, a set of laminated plies forming a panel may be referred to herein as "panel plies" or "panel layers". The panel plies may include composite fibers pre-impregnated with resin, referred to as "prepregs". For example, a panel layer may include a composite fiber pre-impregnated with resin, referred to as a "prepreg". The composite fibers within a single ply may be configured to extend in a single orientation. However, in some examples, a single ply may be configured with composite fibers extending in multiple orientations. In some examples, the height of each ply may be as thin as approximately 0.007 inches.

[0047] like Figure 4B As shown, the slat is formed of a multilayer slat ply including a first layer 420, a second layer 422, and a third layer 424. However, in various examples, the composite slat 400 may include any number of layers to form a desired geometric profile. In some examples, the slat may include overlapping or discontinuous laminated plies. In some examples, different layers of the slat may include the same or different materials and configurations. In some examples, multiple layers may be combined to obtain various desired mechanical properties. For example, adjacent layers may include composite fibers that are configured perpendicular to the fibers in another layer to increase the strength of the overall composite material. However, the layers may include composite fibers configured in various different orientations relative to the composite fibers in adjacent layers.

[0048] In some examples, the width of each layer (in the direction of the X axis) can be varied to obtain the desired geometric profile of each side. Figure 4B As shown, the first layer 420 may include a greater width relative to the second layer 422, which in turn may include a greater width relative to the third layer 424. Thus, by gradually increasing the width of the layers from top to bottom, a curved geometric profile is formed.

[0049] In various examples, the shape of the slats is symmetrical with respect to the first side and the second side. In some examples, the shape of the slats is also symmetrical with respect to the front side and the rear side. However, depending on the geometry of the beams, the sides of the slats may not necessarily be symmetrical.

[0050] In some examples, the various sides of the planks may be further processed to obtain a desired geometric profile. For example, the edges of each layer may be smoothed such as by cutting, machining, grinding or other processes. Figure 4C An example slat 400-A is depicted having a first side 402-A and a second side 404-A that have been smoothed by eliminating ridges of layers on each side. In some examples, the smoothed sides can provide better support for the shape of the support structure, as well as reduce the spacing between portions of the slat and support tooling to further reduce resin bleed and other defects as described herein.

[0051] In some examples, the length of the layer (in the direction of the Y axis) may also vary to provide a geometric profile to the front side 410 and / or the back side 412 of the slat. Figure 4A As shown in the perspective view of FIG. 4 , each of the bottom layers can be slightly longer than the adjacent layers above, thereby forming a slightly curved geometric profile at the front side 410. Thus, in some examples, the front and / or rear slats of the bottom can also be configured to the desired shape of the support beams or other support structures. In some examples, these longer layers at the bottom of the slats can contribute additional strength and stability to the entire slat.

[0052] During the manufacture of the hardened stringer panels, the battens may be placed on the inner surface of the skin member along with other support means to support the shape of the stringers during curing or bonding of the components. As will be further described below, the geometric profiles of the various sides of the battens may replace existing filler faces to fill the space between the stringers and the skin member, thereby supporting the shape of the stringers.

[0053] refer to Figure 5A and Figure 5B, shows a cross-sectional view of a reinforced stringer panel 500 incorporating a lath 400 according to one or more examples. In various examples, the reinforced stringer panel 500 includes a skin member 510, whose lath 400 is located on an inner surface 510-A of the skin member 510. The lath 400 may be any of the various laths described herein, including a lath 400-A having a smooth side. In some examples, the skin member 510 may include multiple layers or multiple layers of composite material. As used herein, a composite material layer including a skin member may be referred to as a "skin member ply" or a "panel ply". In some examples, such composite materials of the skin member layer may be the same or different material as the ply of the lath. In some examples, in addition to or in place of the skin member 510, a base charge or other higher assembly structure may be included.

[0054] During manufacturing, the stringer 520 can be placed on the skin member 510. As shown, the stringer 520 is a hat-shaped stringer or omega stringer having a hat portion 521. However, it should be noted that the examples described herein can be implemented with various stringer shapes. In various examples, the stringer 520 can be placed on a support tool such as a forming mandrel 530. As used herein, the mandrel 530 can be referred to as a support tool 530. The mandrel 530 can support the stringer during transportation and positioning of the stringer on the skin member 510.

[0055] In various examples, mandrel 530 is made of silicon. In some examples, the mandrel may include a silicon formulation having 20% ​​silica microspheres uniformly mixed therein. Including silica microspheres may reduce the overall expansion of the mandrel during the curing process and may result in a more uniform expansion. In some examples, mandrel 530 may be wrapped with a FEP wrap to improve handling and removal of the mandrel during the manufacturing process. FEP wrap 532 is depicted as a thin shaded layer surrounding mandrel 530.

[0056] The stringer 520 and the mandrel 530 may be placed on the inner surface 510-A such that the forming mandrel 530 is positioned on the uppermost laminate ply (upper surface 406) of the batten. The cap portion 521 of the stringer 520 may span from a first side to a second side of the batten and extend into flange portions 524-A and 524-B located on the inner surface 510-A of the skin member 510. In some examples, the stringer 520 includes a plurality of laminate plies, referred to herein as "stringer plies". In some examples, the stringer plies may include a composite material. In some examples, such composite materials of the stringer plies may be the same or different material as the plies of the skin member or batten. For purposes of illustration, as shown in FIG. Figure 5A and Figure 5B As shown, the height of the stringer plies and the panel plies may be exaggerated. In reality, the height of such laminate plies may be thinner than shown, and in some examples there may be more laminate plies.

[0057] The beam 520 of the portion may include a geometry corresponding to the geometric profile of the first and second sides of the slat. For example, the lower radius of the beam 520 may be defined by a base section of the cap portion that abuts the first and second sides of the slat. Figure 5A , the cap portion 521 includes a main section 526 that joins a first base section 522-A and a second base section 522-B of the cap portion, each defined by a dashed line. The first base section 522-A can abut against the first side 402, and the second base section 522-B can abut against the second side 404. Thus, the main section 526 of the cap portion can be flush with the support tool 530.

[0058] Thus, the lath can provide the same function as a radius filler to further support the shape and structure of the stringer 520 in conjunction with the mandrel 530. As previously described, the lath can be multi-directional relative to the composite fibers in adjacent layers that are configured in different orientations. This multi-directional configuration is much stronger in the uncured state than a filler that is unidirectional relative to the composite fiber configuration (if any). Moreover, the multi-directional shape of the lath allows the lath area to be designed with the desired strength and stiffness. Therefore, during the curing process, the lath provides a stiffer and more uniform support for smaller radii in all directions.

[0059] As shown, the space or gap between the base section and the various layers of the slat plies is shown. This may be due to the limitation of the thickness (height) of the composite material of the laminated plies used to form the slats. However, as shown, the height of the slat plies may be exaggerated, and in practice, such spaces or gaps are negligible and do not affect the ability of the board-bonded support tool 530 to support the stringer 520. In some examples, such spaces may only exist in a green / green uncured assembled state during the assembly of the panel. After the co-curing or co-bonding process, such spaces can be eliminated by the compaction of the parts and the flow of resin. As previously described, in some examples, the slats may be machined or processed to smooth the ridges caused by the thickness of the slat plies, such as in slat 400-A.

[0060] Figure 5B The reinforced stringer panel 500 is shown after the attachment and curing process, with the mandrel 530 removed leaving the cavity 570. The remaining structure is a reinforced panel assembly that exhibits higher quality than existing panel assembly systems (such as system 300). The strips do not require filler to fill the gaps between the stringers and support tools and the skin members. This reduces the time and labor costs and machinery costs required to transport long, heavy, but fragile filler to the panel assembly and place such filler.

[0061] The integration of the battens further increases the strength of the panel assembly. The multi-directional construction of the composite fibers of the battens is also much stronger than unidirectional fillers in the cured state. In addition, the battens add additional structure and rigidity to the skin members, especially in larger stringers where longer or taller battens are required. As a result, there is more material between the stringer fill cavities 570, and the entire box is stiffer and stronger than the existing stringer panel assembly system 300. The use of battens can also replace other reinforcing structures that are typically added to the panel assembly. Therefore, the integration of the battens does not significantly affect the weight of the panel assembly, and in some cases can reduce the overall weight.

[0062] The geometry of the lath is also more precise than that of the caulking, which can move after placement. Incorrect or imprecise caulking geometry can result in voids or gaps, creating low pressure areas during the curing process into which resin can migrate (resin bleed). The flat lay structure of the lath is stronger than caulking, which has layers that are folded to form the upper points or corners of a triangular structure. Because the upper surface of the lath is flat, sharp vertical tips are not required and the presence of voids between the interface of the mandrel and the lath is reduced.

[0063] Additionally, a single lath replaces two facer structures on either side of the mandrel, reducing the total number of components. Including fewer individual components further reduces the chance of movement during the manufacturing and curing process. As a result, overall resin bleed and ply wrinkling are reduced. Unlike facer sandwiched between three structures (stringers, skin members and mandrel), the single flat structure of the lath is subject to less strain, is less likely to crack or deform, and can withstand greater pressure.

[0064] Due to the improvement in the strength and geometry of the lath, the overall geometry of the stringer after curing is improved. The lath will deform less than a filler face, allowing a more constant pressure to be applied to the lower radius at the base section of the stringer, thereby reducing potential radius thickening. The described lath, combined with the use of an improved forming mandrel with silica microspheres, can provide more uniform internal pressure to the stringer and skin components, allowing for a cleaner cured stringer structure with more uniform thickness. In some examples, the straightness of the vertical and horizontal portions of the main section between radii is also improved, resulting in a stronger overall panel assembly.

[0065] Tests have shown that the geometry of the stringers is more uniform using such slats as described herein. Cross-sectional thickness and straightness were measured in various tests. In one example, the thickness T1 of the first base segment 522-A and the thickness T7 of the second base segment 522-B were 0.2120 inches and 0.2102 inches, respectively. In addition, the thicknesses T2 and T6 of the vertical sides of the main segment 526 were 0.2019 inches and 0.2013 inches, respectively. In addition, the thickness T3 and thickness T4 of the upper radius segment were 0.1864 and 0.1899 inches, respectively. Finally, the thickness T4 at the top of the main segment was 0.2117 inches. Therefore, the thickness uniformity of the entire stringer was maintained at an acceptable level.

[0066] Furthermore, significant portions of the vertical and horizontal sides are geometrically maintained to the desired straightness. In the example above, the first vertical portion is measured to have a continuous length S1 of 1.415 inches of straightness within a deviation of 0.002 inches. The second vertical portion is measured to have a continuous length S3 of 1.417 inches of straightness within a deviation of 0.002 inches. Finally, the horizontal portion of the main segment is measured to have a continuous length S2 of 0.6227 inches of straightness within a deviation of 0.002 inches. Furthermore, as Figure 5B As shown, cross-sectional imaging also shows that the skin member has fewer bow waves and there is almost no deformation of the cured strips.

[0067] How to operate

[0068] A method of making a reinforced composite panel is also provided. Figure 6 , shows an example method 600 for manufacturing a reinforced stringer panel according to one or more examples. The operations of the method 600 may be combined with Figure 7 To describe, Figure 7 An example process sequence 700 is shown for manufacturing a stiffened stringer panel having battens according to one or more examples. Figure 7 A simplified cross-section of the assembled components is depicted.

[0069] At operation 610, a skin component is provided. Figure 7 As shown in step (A) of , an assembly component 710 may be prepared for stringers and other support structures at a first position (position 0). Such an assembly component 710 may be a skin panel of an aircraft. In some examples, the assembly component 710 may be other structures, such as base loading, or structures in horizontal and vertical stabilizers or control surfaces. For example, component 710 may be a skin member 510 having an inner surface 510-A. In some examples, multiple panel plies may be laminated to form a skin member.

[0070] At operation 620, such as Figure 7In step (B) shown, the skin component is transported to a second position (position 1) in the assembly line, where the lath is positioned on the inner surface of the skin component. In some examples, the lath may be lath 400, or any of the various laths described herein. In some examples, various guide systems and / or pick-and-place systems may be used to place the lath at the desired location on the skin component. The prepreg composition of the lath plies may exhibit stickiness, thereby allowing the component to adhere and remain in place during the manufacturing process. However, in the case where the lath is co-bonded to the skin component, an adhesive film may be used to bond the lath to the skin component. In some examples, a robotic end effector may be used to automatically position the lath.

[0071] At operation 630, such as Figure 7 In step (C), the stringer is placed on a support tool. The stringer may be a hat-shaped stringer 520, and the support tool may be a forming mandrel, such as mandrel 530. As previously described, the stringer may include a curved hat portion including a main section joining a first base section and a second base section. The stringer may include one or more plies of stringers and may be preformed before being placed on the forming mandrel. Each base section of the hat portion may further extend into a flange portion including a first flange portion from the first base section and a second flange portion from the second base section. One or more portions of the hat portion, particularly the main section, may be flush against the forming mandrel having a geometry corresponding to the stringer. The support tool 530 supports the stringer 520 to prevent damage to the stringer during transportation, handling, and positioning. Step (C) may occur at a separate location (location 2).

[0072] At operation 640, such as Figure 7 In step (D), the support tool and the stringer are positioned on the uppermost laminate ply of the batten. Then, in step (D), the stringer and the mandrel are positioned on the batten. This can occur in the position of assembling the components (position 1). As described herein, the mandrel can be positioned on the upper surface of the batten. In addition, the base section of the cap portion of the stringer can rest on the corresponding side of the batten, and the flange portion of the stringer can be located on the inner surface of the skin member.

[0073] At operation 650, the laths, stringers, and skin members are bonded. In order to cure the prepreg laminate, a combination of pressure and heat may be used. As described, the skin members, laths, and stringers may be co-cured or co-bonded. The co-curing or co-bonding process may also bond the laths to the skin members and / or stringer plies. In some examples, the panel including the assembly may be placed in a vacuum bag to apply additional pressure to hold the composite layers in place for curing or bonding. Figure 7In step (E), the assembly is placed in a vacuum bag 760 to apply additional pressure to hold the composite layers in place for curing in step (F). In some examples, the assembly can be moved to a separate location (position 3) on the assembly line for steps (E) and (F). For example, the panel assembly can be cured in an oven or autoclave at 350 degrees Fahrenheit (90 PSI) for several hours.

[0074] In various examples, one or more structures described may be automatically formed and deposited by a robotic end effector for automatically placing composite materials with minimal tools during manufacturing. In some examples, the slats are positioned on the skin member by a robotic end effector. In some examples, one or more laminated plies of the slats may be deposited on the skin member via an automated end effector to form the slats. For example, a series of laminated plies may be laminated to build the slats. In some examples, the slats may be formed by depositing a single laminated ply at a time. However, in some examples, multiple laminated plies may be deposited at a time. For example, such pre-laminated plies may be manufactured using laminating equipment such as Fives Forest-Liné ATLAS or ACCESS machines. The laminated plies may then be co-cured or co-bonded with the laminated plies of the skin member. In examples of slats and skin members or co-bonding, an adhesive film may be placed between the laminate stack and the skin member to help attach the laminate stack to the skin member.

[0075] In various examples, such composite strips may also be used to manufacture one or more layers of skin members 510, as described above. In some examples, composite strips may also be used to manufacture one or more layers of stringers 520. For example, once the support tool 530 is properly positioned on the slat 400, the robotic arm assembly may form a composite strip into one or more stringer plies for deposition onto the mandrel, slats, and skin members. Such automated placement and formation of components may reduce the floor space and manpower required to store, transport, and place such components, which may be very heavy and large in size.

[0076] Aircraft Example

[0077] Examples of the present disclosure can be found at Figure 8 The aircraft 800 shown and Fig. 9 The illustration is described in the context of aircraft manufacturing and service method 800 . Figure 8 FIG. 8 is a schematic diagram of an aircraft 800 that may include the reinforced stringer panels described herein, according to one or more examples. Figure 7As shown, aircraft 800 includes a fuselage 850 having an interior 880. Aircraft 800 includes wings 820 coupled to fuselage 850. Aircraft 800 may also include engines 830 supported by wings 820. According to the illustrative example, aircraft 800 is an example of a vehicle in which the described systems and methods, such as reinforced stringer panel 100, may be implemented and operated. Although an aviation example is shown, the principles disclosed herein may be applied to other industries, such as the automotive industry. Therefore, in addition to aircraft 800, the principles disclosed herein may be applicable to other vehicles, such as land vehicles, marine vehicles, aerospace vehicles, etc.

[0078] Fig. 9 800 . 900 is a block diagram of an aircraft production and service method that may utilize the methods and assemblies described herein. During pre-production, the illustrative method 900 may include specification and design of the aircraft 800 (block 904) and material procurement (block 906). During production, component and subassembly manufacturing (block 908) and inspection system integration (block 910) of the aircraft 800 may be performed. The described apparatus and corresponding operating methods may be implemented in any of the specification and design of the aircraft 800 (block 904), material procurement (block 906), component and subassembly manufacturing (block 908), and / or inspection system integration (block 910) of the aircraft 800.

[0079] Thereafter, the aircraft 800 may be certified and delivered (block 912) to be placed in service (block 914). While in service, the aircraft 800 may be scheduled for routine maintenance and service (block 916). Routine maintenance and service may include modification, reconfiguration, refurbishment, etc. of one or more inspection systems of the aircraft 800. The described apparatus and corresponding operating methods may be implemented in any of certification and delivery (block 912), placement in service (block 914), and / or routine maintenance and service (block 916).

[0080] Furthermore, the present disclosure includes embodiments according to the following clauses:

[0081] Clause 1. A composite panel (100) comprising:

[0082] A skin member (510) having an inner surface (510-A);

[0083] a panel (400) on the inner surface, the panel comprising a set of layered laminate plies (420, 422, 424), the panel extending from a first side (402) to a second side (404), wherein each laminate ply in the set of layered laminate plies is sized to form a desired geometric profile of each of the first side and the second side; and

[0084] A beam (520) includes a cap portion (521), wherein the cap portion spans from a first side of the strip to a second side of the strip to form a first flange portion (524-A) and a second flange portion (524-B) on an inner surface of the skin member, respectively, each flange portion extending from the cap portion.

[0085] Clause 2. The composite panel of Clause 1, wherein the stringer comprises a plurality of stringer plies.

[0086] Clause 3. A composite panel according to clause 1 or clause 2, wherein a first base segment (522-A) of the cap portion contacts a first side of the slat and conforms to a desired geometric profile of the first side of the slat, and wherein a second base segment (522-B) of the cap portion contacts a second side of the slat and conforms to a desired geometric profile of the second side of the slat.

[0087] Clause 4. A composite panel according to clause 3, wherein a support tool (530) is positioned on top of an uppermost laminate ply of the group of layered laminate plies such that a main section (526) of a cap portion joining the first base section and the second base section is supported by the support tool.

[0088] Clause 5. The composite panel of Clause 4, wherein the support tool is a silicon forming mandrel comprising 20% ​​silica microspheres uniformly mixed therein.

[0089] Clause 6. A composite panel according to any one of Clauses 1 to 5, wherein the laths, stringers and skin members are co-cured.

[0090] Clause 7. A composite panel according to any one of Clauses 1 to 5, wherein the laths, stringers and skin members are co-bonded.

[0091] Clause 8. A composite panel according to any one of Clauses 1 to 7, wherein each laminate ply in the set of layered laminate plies comprises composite fibres pre-impregnated with resin.

[0092] Clause 9. The composite panel of Clause 8, wherein the composite fibers of adjacent laminate plies have different orientation angles.

[0093] Clause 10. A method (700) of manufacturing a reinforced composite panel (100), the method comprising:

[0094] Providing (710) a skin member (510) having an inner surface (510-A);

[0095] Positioning (720) a panel (400) on an inner surface of the skin member, the panel comprising a set of layered laminate plies (420, 422, 424), the panel extending from a first side (402) to a second side (404), wherein each laminate ply in the set of layered laminate plies is sized to form a desired geometric profile of each of the first side and the second side;

[0096] placing (730) the stringer (520) onto the support tool (530);

[0097] positioning (740) the support tool and the stringers thereon on an uppermost laminate ply in the set of layered laminate plies; and

[0098] The skin members, laths and stringers are joined (750).

[0099] Clause 11. A method according to Clause 10, wherein the beam includes a cap portion (521) spanning from a first side of the strip to a second side of the strip to form a first flange portion (524-A) and a second flange portion (524-B) on the inner surface of the skin member, respectively, each flange portion extending from the cap portion.

[0100] Clause 12. A method according to clause 11,

[0101] wherein the first base section (522-A) of the cap portion contacts the first side of the slat and conforms to a desired geometric profile of the first side of the slat,

[0102] wherein the second base section (522-B) of the cap portion contacts the second side of the slat and conforms to the desired geometric profile of the second side of the slat, and

[0103] A main section (526) of the cap portion joining the first base section and the second base section is supported by and conforms to the shape of the support tool.

[0104] Clause 13. The method of any one of Clauses 10 to 12, wherein the laths, stringers and skin members are joined by a co-curing process or a co-bonding process.

[0105] Clause 14. The method according to any one of clauses 10 to 13, wherein the support tool is a forming mandrel.

[0106] Clause 15. The method according to any one of Clauses 10 to 14, wherein each laminate ply in the set of layered laminate plies comprises composite fibers pre-impregnated with resin.

[0107] Clause 16. The method of Clause 15, wherein the composite fibers of adjacent laminate plies have different orientation angles.

[0108] Clause 17. An aircraft (800) comprising one or more composite panels (100), the one or more composite panels comprising:

[0109] A skin member (510) having an inner surface (510-A);

[0110] a panel (400) on the inner surface, the panel comprising a set of layered laminate plies (420, 422, 424), the panel extending from a first side (402) to a second side (404), wherein each laminate ply in the set of layered laminate plies is sized to form a desired geometric profile of each of the first side and the second side; and

[0111] A beam (520) includes a cap portion (521), wherein the cap portion spans from a first side of the strip to a second side of the strip to form a first flange portion (524-A) and a second flange portion (524-B) on an inner surface of the skin member, respectively, each flange portion extending from the cap portion.

[0112] Clause 18. The aircraft of Clause 17, wherein the stringer comprises a plurality of stringer plies.

[0113] Clause 19. An aircraft according to clause 17 or clause 18, wherein a first base segment (522-A) of the cap portion contacts a first side of the slat and conforms to a desired geometric profile of the first side of the slat, and wherein a second base segment (522-B) of the cap portion contacts a second side of the slat and conforms to a desired geometric profile of the second side of the slat.

[0114] Clause 20. The aircraft of any of Clauses 17 to 19, wherein each of the set of layered laminate plies comprises composite fibers pre-impregnated with resin.

[0115] Each process of illustrative method 900 may be performed or carried out by an inspection system integrator, a third party, and / or an operator (e.g., a customer). For purposes of this description, an inspection system integrator may include, but is not limited to, any number of aircraft manufacturers and major inspection system subcontractors; a third party may include, but is not limited to, any number of suppliers, subcontractors, and vendors; an operator may be an airline, a leasing company, a military entity, a service organization, etc.

[0116] in conclusion

[0117] In the above description, numerous specific details are set forth to provide a thorough understanding of the disclosed concepts, which concepts may be practiced without some or all of these details. In other instances, details of well-known devices and / or processes are omitted to avoid unnecessarily obscuring the present disclosure.

[0118] Although the present disclosure has been specifically shown and described with reference to specific examples of the present disclosure, it will be understood by those skilled in the art that the form and details of the disclosed examples may be changed without departing from the spirit or scope of the present disclosure. Therefore, it is intended that the present disclosure be interpreted as including all variations and equivalents that fall within the true spirit and scope of the present disclosure. Therefore, the present examples are considered to be illustrative rather than restrictive.

[0119] Although many components and processes are described above in singular form for convenience, those skilled in the art will appreciate that multiple components and repeated processes may also be used to practice the techniques of this disclosure.

Claims

1. A composite panel (100), comprising: A skin member (510) having an inner surface (510-A); a single strip (400) on the inner surface, the strip being comprised of a set of layered laminate plies (420, 422, 424), the single strip extending from a first side (402) to a second side (404), wherein each laminate ply in the set of layered laminate plies is sized to form a desired geometric profile of each of the first side and the second side; and A beam (520) comprising a cap portion (521), wherein the cap portion spans from the first side of the strip to the second side of the strip to form a first flange portion (524-A) and a second flange portion (524-B) on the inner surface of the skin member, respectively, each flange portion extending from the cap portion.

2. The composite panel of claim 1, wherein the stringer comprises a plurality of stringer plies.

3. A composite panel according to claim 1, wherein the first base segment (522-A) of the cap portion contacts the first side of the strip and conforms to the desired geometric profile of the first side of the strip, and wherein the second base segment (522-B) of the cap portion contacts the second side of the strip and conforms to the desired geometric profile of the second side of the strip.

4. A composite panel according to claim 3, wherein a supporting tool (530) is positioned on top of an uppermost laminate ply of the group of layered laminate plies so that a main section (526) of the cap portion joining the first base section and the second base section is supported by the supporting tool.

5. The composite panel according to claim 4, wherein the support tool is a silicon forming mandrel comprising 20% ​​silica microspheres uniformly mixed therein.

6. A composite panel according to any one of claims 1 to 5, wherein each laminate ply in the set of layered laminate plies comprises composite fibres pre-impregnated with resin.

7. A composite panel according to claim 6, wherein the composite fibers of adjacent laminate plies have different orientation angles.

8. A method (700) of manufacturing a reinforced composite panel (100), the method comprising: Providing (710) a skin member (510) having an inner surface (510-A); positioning (720) a single strip (400) on the inner surface of the skin member, the single strip being comprised of a set of layered laminate plies (420, 422, 424), the single strip extending from a first side (402) to a second side (404), wherein each laminate ply in the set of layered laminate plies is sized to form a desired geometric profile of each of the first side and the second side; placing (730) the stringer (520) onto the support tool (530); positioning (740) the support tool and the stringer thereon on an uppermost laminate ply of the set of layered laminate plies of the single plank; as well as The skin members, the laths, and the stringers are joined (750).

9. The method according to claim 8, wherein the beam includes a cap portion (521) spanning from the first side of the strip to the second side of the strip to form a first flange portion (524-A) and a second flange portion (524-B) on the inner surface of the skin member, respectively, each flange portion extending from the cap portion.

10. The method according to claim 8 or 9, wherein the laths, the stringers and the skin members are joined by a co-curing process or a co-bonding process.

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