Composite layup system with tailored modulus and method of use

Abstract

To provide an improved tailored modulus composite layup system and method.SOLUTION: The tailored modulus composite layup system 10 includes a composite layup assembly 16 having a structural ply layer 22 pre-impregnated with a structural resin 60. The system includes a face ply 12 applied to the composite laminate assembly, the face ply having a tailored modulus that is less than or equal to the ply modulus, a face ply coefficient of thermal expansion that is different than the ply coefficient of thermal expansion, and a face ply coefficient of hygroscopic expansion that is different than the ply coefficient of hygroscopic expansion. The system has at least one secondary coating 70 applied to the surface layer. The adjusted modulus of the surface layer is adjusted so as to be equal to or greater than the secondary film modulus. The surface layer provides stress relief and integrated damping between the secondary coating and the structural ply layer, thereby increasing the microcrack resistance of the tailored modulus composite laminate system and minimizing structural degradation.SELECTED DRAWING: Figure 2C

Description

[Technical Field]

[0001] FIELD OF THE DISCLOSURE This disclosure relates generally to composite structural systems and methods, and more particularly to co-cured composite laminate structural systems and methods for aircraft composite structures. [Background technology]

[0002] Composite structures, such as carbon fiber reinforced polymer (or plastic) (CFRP) structures and glass fiber reinforced polymer (or plastic) (GFRP) structures, have a high strength-to-weight ratio, corrosion resistance, and other excellent properties, making them suitable for a wide range of applications, including the manufacture of aircraft, spacecraft, rotorcraft, ships, automobiles, trucks, and other vehicles and structures. Composite structures, such as CFRP and GFRP structures, typically comprise a matrix material, such as a resin, reinforced with carbon fiber, glass fiber, aramid fiber, fiberglass, or other suitable fibrous materials. Examples of aircraft composite structures include aircraft skin panels, wings, vertical stabilizers, horizontal stabilizers, fuselages, and other aircraft composite structures.

[0003] Known composite structural systems and methods include a composite laminate made up of stacked structural ply layers, each of which is made of similar materials and has a similar coefficient of thermal expansion (CTE), coefficient of hygroscopic expansion (CME), and modulus. The composite laminate structure is cured at elevated temperatures during fabrication, thereby locking in thermally induced strains. After curing, a secondary coating, such as paint or primer, is applied to the surface of the composite laminate structure, typically at room or ambient temperature. The secondary coating has a different strain, coefficient of thermal expansion, coefficient of hygroscopic expansion, and modulus than the composite laminate, resulting in a strain mismatch between the secondary coating and the underlying composite laminate structure. When the composite laminate and secondary coating are repeatedly exposed to heat and moisture in an operational environment, because the composite laminate and secondary coating have different strains, coefficients of thermal expansion, coefficients of hygroscopic expansion, and moduli, the strains and stresses caused by the repeated heat and moisture exposure can cause cracks or microcracks in the secondary coating that can propagate into the composite laminate, resulting in increased repair and rework costs and potential degradation of the composite laminate.

[0004] Some known composite structural systems and methods involve modifying the surface of a composite laminate through mechanical, chemical, or energetic means to enhance adhesion of a secondary coating, such as a paint or primer, to the surface. However, these known systems and methods focus solely on the flexibility, e.g., extensibility, of the secondary coating. Furthermore, these known systems and methods do not address the modulus of the composite laminate in light of the secondary coating(s) applied to the composite laminate. As a result, excessive strain and stress may be generated in the secondary coating(s) applied to the composite laminate, potentially resulting in cracks or microcracks within the composite laminate.

[0005] Additionally, some known systems and methods use strain isolation pads between two materials to match the coefficients of thermal expansion between the two materials. However, strain isolation pads are used between ceramic and metallic materials, not between thermoset composites and secondary coatings such as paint.

[0006] Therefore, there is a need in the art for improved tailored modulus composite laminate systems and methods that tailor the modulus of the surface layers of a composite laminate structure to more closely match the modulus of the secondary coating, thereby avoiding strain mismatch, preventing or minimizing the occurrence of surface cracks and microcracking in the secondary coating to avoid repair and rework, minimizing structural degradation due to cracks and microcracking, improving lifetime thermal humidity cycling performance, and providing advantages over known composite structural systems and methods. Summary of the Invention

[0007] Exemplary embodiments of the present disclosure provide improved controlled modulus composite lamination systems and methods. As described below, various aspects of the improved controlled modulus composite lamination systems and methods provide advantages over known systems and methods.

[0008] In one aspect of the present disclosure, a tailored modulus composite laminate system is provided, comprising a cured composite laminate assembly including a plurality of structural ply layers pre-impregnated with a structural resin, each structural ply layer having the same ply modulus, ply thermal coefficient of expansion, and ply moisture coefficient of expansion.

[0009] The controlled modulus composite laminate system further includes a surface layer applied directly to the composite laminate assembly either before or after curing, the surface layer applied either before or after curing having a controlled modulus less than or equal to the ply modulus, a surface layer coefficient of thermal expansion different from the ply coefficient of thermal expansion, and a surface layer coefficient of hygroscopic expansion different from the ply coefficient of hygroscopic expansion.

[0010] The controlled modulus composite laminate system further includes at least one secondary coating applied directly to the surface layer after curing of the composite laminate assembly, the at least one secondary coating having a secondary coating modulus different from the ply modulus, a secondary coating coefficient of thermal expansion different from the ply coefficient of thermal expansion, and a secondary coating coefficient of hygroscopic expansion different from the ply coefficient of hygroscopic expansion.

[0011] The tailored modulus of the surface layer is tailored to be equal to or greater than the modulus of the at least one secondary coating prior to application of the at least one secondary coating to the surface layer, and the tailored modulus surface layer provides stress relief and integrated damping between the at least one secondary coating and the plurality of structural ply layers of the composite laminate assembly, thereby enhancing the microcracking resistance and minimizing structural degradation of the tailored modulus composite laminate system when the tailored modulus composite laminate system is subjected to one or more of heat and moisture exposure in a heat and humidity cycle.

[0012] In yet another aspect of the present disclosure, there is provided an aircraft having one or more aircraft composite structures including a tailored modulus composite laminate system, the one or more aircraft composite structures including one or more of a fuselage, one or more wings, and a tail including one or more vertical and horizontal stabilizers.

[0013] The controlled modulus composite laminate system integrated into the one or more aircraft composite structures includes a cured composite laminate assembly including a plurality of structural ply layers pre-impregnated with a structural resin, each structural ply layer having the same ply modulus, ply coefficient of thermal expansion, and ply coefficient of moisture expansion.

[0014] The controlled modulus composite laminate system further includes a surface layer applied directly to the composite laminate assembly either before or after curing, the surface layer applied either before or after curing having a controlled modulus less than or equal to the ply modulus, a surface layer coefficient of thermal expansion different from the ply coefficient of thermal expansion, and a surface layer coefficient of hygroscopic expansion different from the ply coefficient of hygroscopic expansion.

[0015] The controlled modulus composite laminate system further includes at least one secondary coating applied directly to the surface layer after curing of the composite laminate assembly, the at least one secondary coating having a secondary coating modulus different from the ply modulus, a secondary coating coefficient of thermal expansion different from the ply coefficient of thermal expansion, and a secondary coating coefficient of hygroscopic expansion different from the ply coefficient of hygroscopic expansion.

[0016] The tuned modulus of the skin layer is tuned to be equal to or greater than the modulus of the at least one secondary coating prior to application of the at least one secondary coating to the skin layer, and the skin layer having the tuned modulus forms a stress relief and integrated damping between the at least one secondary coating and the plurality of structural ply layers of the composite laminate assembly, thereby enhancing microcracking resistance and minimizing structural degradation of the one or more aircraft composite structures having the tuned modulus composite laminate system when the one or more aircraft composite structures have the tuned modulus composite laminate system undergo one or more of heat and moisture exposure in a heat and humidity cycle.

[0017] In another aspect of the present disclosure, a method for enhancing microcracking resistance and minimizing structural degradation in a composite structure using a tailored modulus composite laminate system is provided, the method comprising providing the tailored modulus composite laminate system.

[0018] The tailored modulus composite laminate system includes a cured composite laminate assembly including a plurality of structural ply layers pre-impregnated with a structural resin, each structural ply layer having the same ply modulus, ply coefficient of thermal expansion, and ply coefficient of moisture expansion.

[0019] The controlled modulus composite laminate system further includes a surface layer applied directly to the composite laminate assembly either before or after curing, the surface layer applied either before or after curing having a controlled modulus less than or equal to the ply modulus, a surface layer coefficient of thermal expansion different from the ply coefficient of thermal expansion, and a surface layer coefficient of hygroscopic expansion different from the ply coefficient of hygroscopic expansion.

[0020] The controlled modulus composite laminate system further includes at least one secondary coating applied directly to the surface layer after curing of the composite laminate assembly, the at least one secondary coating having a secondary coating modulus different from the ply modulus, a secondary coating thermal coefficient of expansion different from the ply coefficient of thermal expansion, and a secondary coating moisture coefficient of expansion different from the ply moisture coefficient of expansion, the controlled modulus of the surface layer being adjusted to be equal to or greater than the at least one secondary coating modulus prior to application of the at least one secondary coating to the surface layer.

[0021] The method further includes incorporating the tailored modulus composite laminate system into the composite structure, and using the composite structure with the tailored modulus surface layer to form stress relief and integrated damping between the at least one secondary coating and the plurality of structural ply layers of the composite laminate assembly, thereby increasing the microcracking resistance and minimizing structural degradation of the composite structure incorporating the tailored modulus composite laminate system when the composite structure incorporating the tailored modulus composite laminate system is subjected to one or more of heat and moisture exposure in a heat and humidity cycle.

[0022] The above-described features, functions, and advantages may be achieved individually in various aspects of the present disclosure or may be combined in yet other aspects, details of which will become apparent with reference to the following description and drawings. [Brief explanation of the drawings]

[0023] The present disclosure will be better understood from the following detailed description in conjunction with the accompanying drawings, which, although not necessarily to scale, illustrate preferred and exemplary embodiments and are for illustrative purposes only and are not intended to limit the scope of the embodiments or claims.

[0024] [Figure 1A] FIG. 1 is a block diagram illustrating an exemplary tailored modulus composite laminate system of the present disclosure. [Figure 1B] FIG. 1B is a block diagram illustrating an exemplary surface layer used in the controlled modulus composite laminate system of the present disclosure. [Figure 2A] 1 is a cross-sectional front view of an exemplary controlled modulus composite laminate assembly of the present disclosure, with a surface layer in the form of a surface film applied to the uncured composite laminate assembly. [Figure 2B]FIG. 2B is a cross-sectional front view of the tailored modulus composite laminate assembly of FIG. 2A, showing the assembly being co-cured in an autoclave. [Figure 2C] FIG. 2C is a cross-sectional front view of an exemplary controlled modulus composite laminate system of the present disclosure showing the controlled modulus composite laminate assembly of FIG. 2B after a secondary coating has been applied and co-cured. [Figure 2D] FIG. 2D is a cross-sectional front view of the controlled modulus composite laminate system of FIG. 2C, showing the system exposed to heat and moisture. [Figure 2E] FIG. 2E is a cross-sectional front view of the tailored modulus composite laminate system of FIG. 2D, showing the microcracks propagating through the secondary coating layer and the surface layer. [Figure 3A] 1 is a cross-sectional front view of an exemplary tailored modulus composite laminate assembly of the present disclosure, with a surface layer in the form of a lightning strike protection assembly applied to the uncured composite laminate assembly. [Figure 3B] FIG. 3B is a cross-sectional front view of the tailored modulus composite laminate assembly of FIG. 3A, showing the assembly being co-cured in an autoclave. [Figure 3C] FIG. 3C is a cross-sectional front view of an exemplary controlled modulus composite laminate system of the present disclosure showing the controlled modulus composite laminate assembly of FIG. 3B after a secondary coating has been applied and co-cured. [Figure 3D] FIG. 3D is a cross-sectional front view of the controlled modulus composite laminate system of FIG. 3C, showing the system exposed to heat and moisture. [Figure 3E] FIG. 3E is a cross-sectional front view of the tailored modulus composite laminate system of FIG. 3D showing microcracks propagating through the secondary coating layer and the surface layer. [Figure 4A] FIG. 2 is a cross-sectional front view of an exemplary uncured composite laminate assembly of the present disclosure; [Figure 4B] FIG. 4B is a cross-sectional front view of the composite laminate assembly of FIG. 4A, showing the assembly being cured in an autoclave. [Figure 4C]FIG. 4C is a cross-sectional front view of an exemplary controlled modulus composite laminate system of the present disclosure showing the composite laminate assembly of FIG. 4B after a surface layer in the form of a resin repair layer has been applied and cured, and a secondary coating has been applied to the surface layer. [Figure 4D] FIG. 4D is a cross-sectional front view of the controlled modulus composite laminate system of FIG. 4C, showing the system exposed to heat and moisture. [Figure 4E] FIG. 4E is a cross-sectional front view of the tailored modulus composite laminate system of FIG. 4D, showing the microcracks propagating through the secondary coating layer and the surface layer. [Figure 5] 1 is a flowchart illustrating an exemplary method of the present disclosure. [Figure 6] 1 is a perspective view of an aircraft having one or more composite structures incorporating an exemplary controlled modulus composite laminate system of the present disclosure; FIG. [Figure 7] 1 is a flowchart of an exemplary aircraft manufacturing and service method. [Figure 8] FIG. 1 is an exemplary block diagram of an aircraft. DETAILED DESCRIPTION OF THE INVENTION

[0025] Hereinafter, embodiments of the present disclosure will be described in more detail with reference to the accompanying drawings. Note that the drawings illustrate some, but not all, of the disclosed embodiments. In fact, several different embodiments are possible, and the present disclosure should not be construed as being limited to the embodiments described herein. Rather, these embodiments are presented to provide a sufficient disclosure to enable those skilled in the art to fully understand the scope of the present disclosure.

[0026] References herein may be made to "one embodiment" or "an embodiment." The phrases "one embodiment" and "an embodiment" do not necessarily refer to the same embodiment. Particular features, structures, and characteristics may be combined in any manner consistent with the description herein. All features disclosed in this specification, including the claims, abstract, and drawings, and all steps in the disclosed methods and processes, may be combined in any manner, unless at least some of such features and / or steps are mutually exclusive. Each feature disclosed in this specification, including the claims, abstract, and drawings, may be replaced by alternative features serving the same, equivalent, or similar purpose, unless expressly stated otherwise.

[0027] As used herein, "comprising" is open-ended and does not exclude the presence of additional features or steps when used in a claim.

[0028] As used herein, "configured to" describes or defines that various parts or components are "configured to perform a given task or tasks." In this context, "configured to" is used to define structure by indicating that the part or component includes a structure for performing the given task or tasks during operation. Thus, such a part or component may be configured to perform a given task even if it is not currently operational (e.g., not turned on).

[0029] As used herein, the terms "first," "second," and the like are used merely as labels, unless expressly stated otherwise, and do not imply any ordering (e.g., spatial, temporal, or logical) of the items referred to by these terms.

[0030] As used herein, the singular reference to an element or step does not necessarily exclude a plurality of elements or steps. As used herein, the term "and / or" includes any and all combinations with one or more of the listed items. Also, as used herein, the term "combinations thereof" includes combinations with at least one of the listed items, and such combinations may include additional items, such as unlisted items.

[0031] As used herein, the phrase "at least one," when used in conjunction with a list of items, means that one or more of the listed items can be used in different combinations, and that only one of each listed item may be required. In other words, "at least one" means that any number of the listed items can be used in any combination, and does not require the use of all the listed items. An item in this context may be, for example, a particular object, thing, or category.

[0032] Reference is now made to FIGS. 1A-1B. FIG. 1A is a block diagram illustrating an exemplary controlled modulus composite laminate system 10 of the present disclosure, and FIG. 1B is a block diagram illustrating an exemplary surface layer 12 used in the controlled modulus composite laminate system 10 of the present disclosure. The blocks in FIGS. 1A-1B represent various elements, and the lines connecting these blocks do not imply any particular subordinate relationships between the elements. Furthermore, the connecting lines shown in the figures herein indicate exemplary functional relationships and / or physical connections between the various elements, although alternative or additional functional relationships and physical connections may be included in embodiments of the present disclosure. It should be noted that the illustrated blocks may be combined, divided, or combined and then divided into multiple blocks when implemented in an exemplary embodiment. Furthermore, the illustration of the controlled modulus composite laminate system 10 in FIG. 1A and the surface layer 12 in FIG. 1B does not imply any physical or structural limitations on the manner in which the exemplary embodiment may be implemented. Other components may be used in addition to or in place of the components described. Also, some components may not be necessary.

[0033] Referring to FIG. 1A, as shown, a tailored modulus composite layup system 10 includes a structural assembly 14, which may take the form of, for example, a cured or configured to be cured composite layup assembly (CLA) 16. As shown in FIG. 1A, the composite layup assembly 16 may be in the form of an uncured composite layup assembly (CLA) 16a, a partially cured composite layup assembly (CLA) 16b, and / or a cured composite layup assembly (CLA) 16c. As shown in FIG. 1A, in one example, the structural assembly 14, such as the composite layup assembly 16, includes a panel 18 (see also FIG. 6). Examples of panels include a wing panel 18a (see also FIG. 6) on a wing 204 (see FIG. 6) of an aircraft 200a (see FIG. 6), a fuselage panel 18b on a fuselage 202 (see FIG. 6) of an aircraft 200a, a horizontal stabilizer (HORIZ.STAB.) panel 18c on a horizontal stabilizer 212 (see FIG. 6) of an aircraft 200a, a vertical stabilizer (VERT.STAB.) panel 18d on a vertical stabilizer 210 (see FIG. 6) of an aircraft 200a, or any other suitable panel 18 or structural assembly 14.

[0034] The structural assembly 14 is a composite layup assembly 16, e.g., in the form of an uncured composite layup assembly 16a, a partially cured composite layup assembly 16b, and a cured composite layup assembly 16c, and includes a plurality of structural (STRUCT.) layers 20 (see FIGS. 1A, 2A, 3A, and 4A), e.g., a plurality of structural (STRUCT.) ply layers 22 (see FIGS. 1A, 2A, 3A, and 4A). Each structural ply layer 22 includes a ply 24 (see FIGS. 1A, 2A, 3A, and 4A), and the plurality of structural ply layers 22 includes a plurality of plies 24. As shown in FIGS. 2A, 3A, and 4A, the plurality of plies 24 includes a top ply 24a, a bottom ply 24b, and two middle plies 24c. However, the plurality of plies 24 may include more than two middle plies 24c. As shown in Figures 2A, 3A, and 4A, the plurality of plies 24 are in the form of structural plies 24d.

[0035] 1A, each structural ply layer 22 has a modulus 25, such as ply modulus 25a, a coefficient of thermal expansion (CTE) 26, such as ply coefficient of thermal expansion (CTE) 26a, and a coefficient of moisture expansion (CME) 28, such as ply coefficient of moisture expansion (CME) 28a. Ply modulus 25a, ply coefficient of thermal expansion 26a, and ply coefficient of moisture expansion 28a are the same for all ply layers 22.

[0036] 1A, each structural ply layer 22 has stresses 30, such as ply stress 30a, and strains 32, such as ply strain 32a, that are generated by the heat 36 and elevated temperature 38 applied during a heating period 40 in a cure 34, such as a snap cure process (PROC.) 35. Because the composite material (COMP. MAT.) 42 comprising the plies 24 of each structural ply layer 22 is anisotropic, each structural ply layer 22 will tend to expand to different amounts during cure 34. However, because the structural ply layers 22 are bonded together, the expansion results in stresses 30, such as ply stress 30a.

[0037] 1A, each of the plurality of structural layers 20, such as the plurality of structural ply layers 22, and plies 24 includes a composite material 42 having structural fibers 44, such as composite (COMP.) fibers 46. As shown in FIG. 1A, the composite material 42 includes one or more of a carbon fiber reinforced polymer (CFRP) 48 or carbon fiber reinforced plastic having carbon fibers 50, a glass fiber reinforced polymer (GFRP) 52 or glass fiber reinforced plastic having glass fibers 54 or fiberglass fibers, an aramid polymer 56 or aramid plastic having aramid fibers 58, or other suitable composite material 42 having structural fibers 44, such as composite fibers 46. Note that "aramid" refers to an aromatic polyamide.

[0038] The composite laminate assembly 16 includes multiple structural ply layers 22 (see FIG. 1A) in an uncured, partially cured, or cured state that are pre-impregnated with a structural resin (SR) 60, such as a structural prepreg resin. As shown in FIG. 1A, the structural resin 60 can take the form of an uncured structural resin (SR) 60a, a partially cured structural resin (SR) 60b, and a cured structural resin (SR) 60c.

[0039] The composite material 42 comprises a matrix structural resin 60 (see FIG. 1A) reinforced with composite fibers 46. Each structural layer 20, such as the structural ply layer 22, includes a plurality of structural fibers 44, such as composite fibers 46, that are at least partially or completely encapsulated in the structural resin 60, e.g., uncured structural resin 60a (see FIG. 1A), partially cured structural resin 60b (see FIG. 1A), and / or cured structural resin 60c (see FIG. 1A). Examples of the plurality of structural fibers 44, such as composite fibers 46, include a plurality of carbon fibers 50, a plurality of glass fibers 54 or fiberglass fibers, and / or a plurality of aramid fibers 58, or other suitable structural fibers 44. In a preferred embodiment, the structural layer 20 and / or structural fibers 44, such as composite fibers 46, comprise or are arranged in a stack of multiple structural ply layers 22, including structural plies 24d comprised of composite material 42, which includes structural resin 60, for example, in the form of uncured structural resin 60a, partially cured structural resin 60b, and / or cured structural resin 60c.

[0040] The plurality of structural layers 20 may include a plurality of structural ply layers 22 comprised of a composite material 42 including a structural resin 60, such as an uncured structural resin 60a. In each of the structural layers 20, such as the plurality of structural ply layers 22, a plurality of structural fibers 44, such as composite fibers 46, are at least partially encapsulated in the structural resin 60, e.g., the uncured structural resin 60a, the partially cured structural resin 60b, and / or the cured structural resin 60c. In each of the structural layers 20, such as the plurality of structural ply layers 22, a plurality of carbon fibers 50 are at least partially encapsulated in the structural resin 60, e.g., the uncured structural resin 60a, the partially cured structural resin 60b, and / or the cured structural resin 60c.

[0041] 1A , in one embodiment, structural resin (SR) 60 preferably comprises a thermosetting structural resin (SR) 62, such as epoxy structural resin (SR) 62a, phenolic structural resin (SR) 62b, polyimide structural resin (SR) 62c, bismaleimide structural resin (BMI SR) 62d, such as bismaleimide structural resin, polyurethane structural resin (SR) 62e, fluoropolymer structural resin (SR) 62f, cyanate ester structural resin (SR) 62g, or other suitable thermosetting structural resin 62. In other embodiments, structural resin 60 comprises any suitable resin having, defining, or exhibiting, for example, a given structural resin viscosity, structural resin cure profile, structural resin rheology, structural resin gel point temperature, and / or structural resin gel time.

[0042] 1A, the controlled modulus composite laminate system 10 further includes one or more surface layers 12 applied directly to a structural assembly 14, such as a composite laminate assembly 16. The surface layer 12 or surface layers 12 may be formed prior to curing 34 or co-curing 34a (see FIG. 1A) of the structural assembly 14, such as the composite laminate assembly 16, or may be formed after curing 34 or co-curing 34a of the structural assembly 14, such as the composite laminate assembly 16.

[0043] 1A, in one embodiment, the surfacing layer 12 or layers 12 include a surfacing film 64. The surfacing film 64 is preferably applied directly to a structural assembly 14, such as a composite laminate assembly 16, prior to curing 34 or co-curing 34a of the structural assembly 14, such as a composite laminate assembly 16.

[0044] 1A, in another embodiment, one or more of the surface layers 12 includes a lightning strike protection (LSP) assembly 66. The lightning strike protection assembly 66 is preferably applied directly to a structural assembly 14, such as a composite layup assembly 16, prior to curing 34 or co-curing 34a of the structural assembly 14, such as a composite layup assembly 16.

[0045] 1A, in yet another embodiment, one or more surface layers 12 include one or more resin repair layers 68. The one or more resin repair layers 68 are preferably applied directly to a structural assembly 14, such as a composite layup assembly 16, after curing 34 of the structural assembly 14, such as a composite layup assembly 16.

[0046] In other embodiments, the surface layer 12 or layers 12 include other suitable types of surface layers 12, such as a surface film 64, a lightning strike protection assembly 66, and one or more resin repair layers 68, as described in more detail below with reference to FIG. 1B.

[0047] As shown in FIG. 1A, the controlled modulus composite laminate system 10 further includes at least one or more secondary coatings (SC) 70. The secondary coating, for example, at least one or more secondary coating (SC) layers 70a, is applied directly to the surface layer 12 on a first surface 150 (see FIG. 2C), such as the upper surface 150a (see FIG. 2C), of the surface layer 12 after curing 34 of the structural assembly 14, such as the composite laminate assembly 16. The at least one secondary coating 70 has a modulus 25 (see FIG. 1A), such as secondary coating (SC) modulus 25d, that is different from the ply modulus 25a. The controlled modulus 25c of the surface layer 12 is adjusted to be equal to or greater than the secondary coating modulus 25d of the at least one secondary coating 70 prior to application of the at least one secondary coating 70 to the surface layer 12. Therefore, the surface layer modulus 25b is already adjusted to the adjusted modulus 25c before the surface layer 12 is covered with one or more secondary coatings 70.

[0048] The at least one secondary coating 70 further has a coefficient of thermal expansion (CTE) 26, such as secondary coating coefficient of thermal expansion (SC CTE) 26c, that is different from ply coefficient of thermal expansion (CTE) 26a. The at least one secondary coating 70 further has a coefficient of moisture expansion (CME) 28, such as secondary coating coefficient of moisture expansion (SC CME) 28c, that is different from ply coefficient of moisture expansion 28a. The at least one secondary coating 70 has a stress 30 (see FIG. 1A), such as secondary coating (SC) stress 30c (see FIG. 1A), and a strain 32 (see FIG. 1A), such as secondary coating (SC) strain 32c (see FIG. 1A), when the secondary coating 70 is exposed to thermal moisture cycling (TM) 100 (see FIG. 1A) or a thermal moisture cycling event.

[0049] As used herein, "thermo-hygroscopic cycling" or "thermo-hygroscopic cycling event" refers to a process in which a material during its life is subjected to changes in heat and / or moisture, such as endothermic or thermal exposure, or hygroscopic or moisture exposure, such as when an aircraft is repeatedly moved between parked positions on the ground at ambient or high temperatures and in-flight stages in the air at low or freezing temperatures.

[0050] As shown in FIG. 1A, the at least one secondary coating 70 may include one or more of a paint coating 72, a basecoat paint coating 72a, a topcoat paint coating 72b, a primer coating 74, an appliqué 76, a decal 78, or a repair past coating 80, or other suitable secondary coating 70.

[0051] 1A , one or more facing layers 12, for example in the form of a facing film 64 or lightning strike protection assembly 66 in an uncured state, are laid up, such as by layup 82 in a layup process 84, to a composite layup assembly 16, such as an uncured composite layup assembly 16a, to form a layup assembly (LA) 86. Alternatively, a structural assembly 14, such as a composite layup assembly 16, for example an uncured composite layup assembly 16a, may be laid up to a facing layer 12, such as a facing film 64 or lightning strike protection assembly 66.

[0052] As shown in FIG. 1A , the layup assembly (LA) 86 can take the form of an uncured layup assembly (LA) 86a, a partially cured layup assembly (LA) 86b, and / or a cured layup assembly (LA) 86c, depending on the stage of curing 34 or co-curing 34a, specifically, whether before, during, or after curing 34 or co-curing 34a. Co-curing refers to simultaneous curing of multiple layers. The uncured layup assembly 86a includes an uncured facing film 64 or uncured lightning protection assembly 66 as the facing layer 12, which is then laminated and bonded to an uncured composite layup assembly 16a to form a tailored modulus composite layup assembly (ASSY.) 11 (see FIG. 1A ).

[0053] In one embodiment, when the surface layer 12 includes one or more resin repair layers 68, curing the uncured composite layup assembly 16a results in a cured composite layup assembly 16c, with the surface layer 12 in the form of the resin repair layer 68 applied to the cured composite layup assembly 16c after curing 34.

[0054] With regard to curing 34 or co-curing 34a, the controlled modulus composite layup assembly 11, in its uncured state, including a surface layer 12 in the form of a facing film 64 or lightning strike protection assembly 66, bonded or attached to the uncured composite layup assembly 16a, is placed in a heating device 88 (see FIG. 1A), such as an autoclave 90 (see FIG. 1A) or other suitable heating device 88, to effect curing 34 (see FIG. 1A) or co-curing 34a (see FIG. 1A). Preferably, a snap-cure process (PROC.) 35 (see FIG. 1A) is performed on the uncured layup assembly 86a, including the surface layer 12, such as a facing film 64 or lightning strike protection assembly 66, laminated and bonded to the uncured composite layup assembly 16a. The snap-cure process is disclosed, for example, in U.S. Pat. No. 11,752,708, which is incorporated herein by reference in its entirety.

[0055] In the curing 34 or co-curing 34a in a heating device 88, such as an autoclave 90, a layup assembly 86, which is an uncured layup assembly 86a, is heated. The layup assembly includes an uncured surface layer 12a laminated and bonded to an uncured composite laminate assembly 16a, where the surface layer 12 is a surface film 64, such as a resin-based composite surface film 64a (see FIG. 1B) in the form of an uncured prepreg surface film 64b (see FIG. 1B), or a lightning protection assembly 66, such as an uncured lightning protection assembly 66a (see FIG. 1B). The heating is, for example, an initial heating, in which the layup assembly 86 is heated to a temperature (TEMP.) 38a (see FIG. 1A) by heat 36 (see FIG. 1A), which is an initial heat 36a (see FIG. 1A). This results in a partially cured layup assembly 86b (see FIG. 1A) in which a partially cured surface layer 12b (see FIG. 1B), such as a partially cured prepreg surface film 64c (see FIG. 1B) or a partially cured lightning strike protection assembly 66b (see FIG. 1B), is laminated and bonded to a partially cured composite laminate assembly 16b (see FIG. 1A). The initial heat 36a and the initial temperature 38a resulting from the initial heating are preferably equal to or greater than 88°C (88 degrees Celsius, 190 degrees Fahrenheit (°F)) and equal to or less than 140°C (140 degrees Celsius, 284 degrees Fahrenheit (°F)). The initial heating includes heating for a heating time 40a (see FIG. 1A), for example, between 2 minutes and 30 minutes. The initial heating by the initial heat 36a is sufficient to gel the uncured injection resin 136, such as the prepreg thermosetting resin 118 (see FIG. 1B), to a state of gel 140 (see FIG. 1B), but the initial heating by the initial heat 36a is not sufficient to gel the uncured structural resin 60.

[0056] The partially cured layup assembly 86b continues to be heated for curing 34 or co-curing 34a in a heating device 88, such as an autoclave 90. The partially cured layup assembly, which comprises a partially cured surface layer 12b (see FIG. 1B), such as a partially cured prepreg surface film 64c (see FIG. 1B) or a partially cured lightning strike protection assembly 66b (see FIG. 1B), laminated and bonded to a partially cured composite laminate assembly 16b (see FIG. 1A), is subsequently heated to a desired subsequent temperature (TEMP.) 38b (see FIG. 1A) using a subsequent heat 36b (see FIG. 1A). The subsequent heating is continued until the heat reaches a final heat 36c (see FIG. 1A), which is a final temperature (TEMP.) 38c (see FIG. 1A) that is higher than the initial temperature 38a. This produces a cured layup assembly 86c (see FIG. 1A) in which a cured surface layer 12c (see FIG. 1B), such as a cured prepreg surface film 64d (see FIG. 1B) or a cured lightning strike protection assembly 66c (see FIG. 1B), is laminated and bonded to the cured composite layup assembly 16c. The combination of the initial, subsequent, and final heating is sufficient to fully cure both the prepreg thermoset resin 118, such as infusion resin 136, and the structural resin 60.

[0057] Subsequent heating by subsequent heat 36b includes heating for a subsequent heating time 40b (see FIG. 1A), for example, 15 minutes or more and 120 minutes or less. Final heat 36c (see FIG. 1A) and final temperature 114c by final heating are, for example, final curing temperatures, preferably 120°C or more (120 degrees Celsius, 248 degrees Fahrenheit (°F)) and 260°C or less (260 degrees Celsius, 500 degrees Fahrenheit (°F)). The difference between final temperature 38c, such as the final curing temperature, and initial temperature 38a (see FIG. 1A) is at least 20°C (20 degrees Celsius, 68 degrees Fahrenheit (°F)). Initial heating involves initial heating for initial heating time 40a, and subsequent heating involves heating for subsequent heating time 40b, with the difference between subsequent heating time 40b and initial heating time 40a being at least 10 minutes.

[0058] Curing 34 or co-curing 34a of layup assembly 86 includes, in addition to heating uncured facing layer 12a (see FIG. 1B), such as uncured prepreg facing film 64b (see FIG. 1B) or uncured lightning protection assembly 66a (see FIG. 1B) laminated and bonded to uncured composite layup assembly 16a (see FIG. 1A) with heat 36 in heating device 88, such as autoclave 90, at least one of applying pressure 92 (see FIG. 1A) to uncured composite layup assembly 16a during initial heating with initial heat 36a and applying pressure to partially cured composite layup assembly 16b during subsequent heating with subsequent heat 36b. In some embodiments, applying pressure 92 includes applying high pressure (ATM.) 92a between 0 and 150 psi (0 and 150 pounds per square inch) (0 and 10.2 atm) to the uncured composite layup assembly 16a and / or the partially cured composite layup assembly 16b. In some embodiments, applying pressure 92 includes using a heating device 88, such as an autoclave 90, autoclaving the uncured composite layup assembly 16a in an initial heating and / or autoclaving the partially cured composite layup assembly 16b in a subsequent heating.

[0059] As shown in FIG. 1A, once the final cure 34 or co-cure 34a is complete, e.g., after the snap cure process 35 is complete, the adjusted modulus composite layup assembly 11 is in a cured state, resulting in a cured layup assembly 86c. In this embodiment, one or more secondary coatings 70 are disposed on a cured facing layer 12c (see FIG. 1B), such as a cured prepreg facing film 64d or a cured lightning strike protection assembly 66c, thereby resulting in the adjusted modulus composite layup system 10 (see FIG. 1A). In another embodiment, one or more resin repair layers 68 are applied to the cured composite layup assembly 16c as the facing layer 12, and then one or more secondary coatings 70 are applied to the one or more resin repair layers 68.

[0060] As shown in FIG. 1A, the tailored modulus composite laminate system 10 is incorporated into or constitutes a composite structure (STRUC.) (CS) 94, such as a wing 204 (see FIG. 6) of an aircraft 200a (see FIG. 6), a fuselage 202 (see FIG. 6) of the aircraft 200a, a horizontal stabilizer 212 (see FIG. 6) of the aircraft 200a, or any other suitable composite structure 94 or other suitable aircraft composite structure 94a.

[0061] When the adjusted modulus composite laminate system 10 or a composite structure 94, such as an aircraft composite structure 94a, including the adjusted modulus composite laminate system 10, is subjected to one or both of heat exposure 96 (see FIG. 1A) and moisture exposure 98 (see FIG. 1A) in a heat and humidity cycle 100 (see FIG. 1A), the surface layer 12 having the adjusted modulus 25c, i.e., the adjusted modulus surface layer 12d (see FIG. 1B), forms a stress relief 31 (see FIG. 1B), such as a layered stress relief 31a (see FIG. 1B), across the thickness of the adjusted modulus composite laminate system 10. The surface layer 12 having the tailored modulus 25c provides integrated dampening 102 (see FIG. 1B) and integrated cushioning 104 (see FIG. 1B) between the secondary coating 70 and the structural ply layer 22 of the composite laminate assembly 16, thereby further enhancing microcrack resistance 106 (see FIG. 1A), minimizing structural degradation 108 (see FIG. 1A), and further improving humidity heat (TM) cycling performance (PERF.) 110 (see FIG. 1A) of the tailored modulus composite laminate system 10 and composite structures 94, such as aircraft composite structure 94a, including the tailored modulus composite laminate system 10.

[0062] When the tailored modulus composite laminate system 10 and a composite structure 94, such as an aircraft composite structure 94a, including the tailored modulus composite laminate system 10, are subjected to heat exposure 96 and moisture exposure 98 in a heat and humidity cycle 100, the various materials included in the tailored modulus composite laminate system 10 (e.g., the structural ply layers 22 of the composite laminate assembly 16, the surface layer 12 having the tailored modulus 25c, and the at least one secondary coating 70) exhibit different amounts of expansion 112 (see FIG. 1A) in thermal humidity (TM) expansion 112a (see FIG. 1A) and different amounts of contraction 114 (see FIG. 1A) in thermal humidity (TM) contraction 114a (see FIG. 1A). These differences in expansion and contraction are further accentuated by stresses 30 (see FIG. 1A), such as ply stress 30a, surface layer stress 30b, and / or secondary coating stress 30c. However, because the surface layer 12 has a tailored modulus 25c, i.e., is a tailored modulus surface layer 12d, there is only a minimal strain mismatch 32d (see FIG. 1A) between the tailored modulus surface layer 12d and the secondary coating 70, allowing the tailored modulus surface layer 12d to safely accommodate the magnitude change between the tailored modulus surface layer 12d and the overlying ply 24 of the structural ply layer 22. This results in a minimal occurrence of microcracks 116 (see FIGS. 1A, 2E, 3E, 4E) or cracks, resulting in a high microcrack resistance 106 (see FIG. 1A). By adjusting the modulus 25 of the surface layer 12 to obtain an adjusted modulus 25c and an adjusted modulus surface layer 12d, stress relief portions 31 (see FIG. 1B), such as laminar stress relief portions 31a (see FIG. 1B), are formed through the thickness of the adjusted modulus composite laminate system 10 and between the structural ply layers 22 of the composite laminate assembly 16, thereby relieving stress caused by thermally induced strain 32 (see FIG. 1A) due to an increase in temperature 38 (see FIG. 1A) during curing 34 or co-curing 34a.

[0063] The material design and tailored modulus 25 of the surface layer 12 improves humidity heat cycling performance 110 (see FIG. 1A) by providing stress relief 31, such as laminar stress relief 31a, through the thickness of the tailored modulus composite laminate system 10. This is evidenced by minimal or no microcracks 116 (see FIGS. 1A and 2E), including surface cracks (MC) 116a (see FIGS. 1A and 2E), in the secondary coating 70 and surface layer 12, and no or very few internal cracks (MC) 116b extending through the structural ply layer 22 of the composite laminate assembly 16. Additionally, this improvement is achieved while maintaining approximately the same level of adhesion of the secondary coating 70, such as paint film 72.

[0064] 1B, which illustrates exemplary one or more surface layers (SL) 12 that can be used in the tailored modulus composite laminate system 10 of FIG. 1A. Here, the one or more surface layers 12 can take the form of an uncured surface layer (SL) 12a, a partially cured surface layer (SL) 12b, and / or a cured surface layer (SL) 12c. The surface layer 12 includes one or more intermediate layers 12e (see FIG. 1B) that are laminated below the secondary coating 70 covering the surface layer 12, i.e., between the secondary coating and the surface layer, and above the plurality of structural ply layers 22 of the composite laminate assembly 16 located below the surface layer 12.

[0065] As further shown in FIG. 1B , in one embodiment, the surface layer 12 includes and functions as a surface layer (SL) shock absorber 13. The surface layer shock absorber 13 absorbs various stresses 30 (see FIG. 1B ) and various strains 32 (see FIG. 1B ) generated in the at least one secondary coating 70 (see FIG. 1A ) and the plurality of structural ply layers 22 (see FIG. 1A ) of the composite laminate assembly 16 (see FIG. 1A ) when the adjusted modulus composite laminate system 10 (see FIG. 1A ) is subjected to one or both of heat exposure 96 (see FIG. 1A ) and moisture exposure 98 (see FIG. 1A ). This improves the humidity heat cycling performance 110 (see FIG. 1A ) of the adjusted modulus composite laminate system 10 and a composite structure 94 (see FIG. 1A ) including the adjusted modulus composite laminate system 10, such as aircraft composite structure 94 a (see FIG. 1A ).

[0066] As shown in FIG. 1B, the surface layer 12 (or each of the surface layers 12, if there is one or more) has a surface layer (SL) modulus 25b as its modulus 25, which is adjusted to form an adjusted modulus 25c. The adjusted modulus 25c of the surface layer 12, which is the adjusted modulus surface layer 12d, is adjusted to be equal to or less than the ply modulus 25a of the ply 24. The adjusted modulus 25c of the surface layer 12 is adjusted to be equal to or greater than the secondary coating modulus 25d of the secondary coating 70 (see FIG. 1A), thereby obtaining an adjusted modulus surface layer (SL) 12d (see FIG. 1B). For example, if the ply modulus 25a is X, the adjusted modulus 25c of the surface layer 12 is Y, and the secondary coating modulus 25d is Z, then the relationship between these holds: X is greater than or equal to Y or Y is less than or equal to X, and Y is greater than or equal to Z or Z is less than or equal to Y. In one embodiment, the conditioning modulus 25c of the surface layer 12 is intermediate between the ply modulus 25a and the secondary coating modulus 25d.

[0067] 1B, the surface layer 12 (or each of the surface layers 12, if there is one or more) has a coefficient of thermal expansion (CTE) 26 (see FIG. 1A), e.g., a surface layer coefficient of thermal expansion (SL CTE) 26b, which is different from the ply CTE 26a.

[0068] As further shown in Figure 1B, the surface layer 12 (or each of the surface layers 12, if there is one or more) has a hygroscopic coefficient of expansion (CME) 28 (see Figure 1A), e.g., a surface layer hygroscopic coefficient of expansion (SL CME) 28b, which is different from the ply CME 28a.

[0069] 1B, the surface layer 12 (or each of the surface layers 12, if there are one or more) has or can have a stress 30, such as surface layer (SL) stress 30b, and a strain 32, such as surface layer (SL) strain 32b. As discussed above, the tailored modulus surface layer 12d forms stress relief regions 31 (see FIG. 1B), such as laminar stress relief region 31a (see FIG. 1B), through the thickness of the tailored modulus composite laminate system 10 when subjected to heat exposure 96 (see FIG. 1A) and / or moisture exposure 98 (see FIG. 1A) in a heat and humidity cycle 100 (see FIG. 1A).

[0070] 1B, in one embodiment, the facing layer 12 (or each of the facing layers 12, if there is one or more) comprises a facing film (SF) 64, such as a resin-based composite facing film (SF) 64a (see FIG. 1B). The facing film 64 can take the form of an uncured prepreg facing film (SF) 64b, a partially cured prepreg facing film (SF) 64c, and / or a cured prepreg facing film (SF) 64d. In another embodiment, the facing film 64 comprises a carbon fiber, fiberglass, or nylon fiber ply or scrim impregnated with a two-component epoxy resin that can be mixed at room temperature.

[0071] As shown in FIG. 1B, a surface film 64, such as resin-based composite surface film 64a, comprises a film layer 65 impregnated or infused with a prepreg thermoset resin 118. Film layer 65 may be, for example, a woven scrim or carrier medium having glass fiber 54 (see FIG. 1A) or fiberglass reinforcement, carbon fiber 50 (see FIG. 1A) or carbon fiber reinforcement, nylon fiber or nylon fiber reinforcement, polyester fiber or polyester reinforcement, or other suitable fiber reinforcement. Prepreg thermoset resin 118 may include one or more of adhesive 118a, epoxy 118b, phenolic 118c, polyimide 118d, bismaleimide (BMI) 118e, polyurethane 118f, fluoropolymer 118g, cyanate ester 118h, or other suitable prepreg thermoset resin 118. The prepreg thermoset 118 may further include a UV resistant prepreg thermoset, an aliphatic prepreg thermoset, a flame retardant prepreg thermoset, or other suitable prepreg thermoset 118.

[0072] 1B , in one embodiment, the surface layer 12 (or each of the surface layers 12, if present) comprises a lightning strike protection material assembly (LSPMA) 66 that is applied directly to the composite layup assembly 16 prior to curing 34 or co-curing 34a of the composite layup assembly 16 and co-cured with the composite layup assembly 16. The lightning strike protection material assembly 66 provides lightning strike protection (LSP) 120 (see FIG. 1B ) to the controlled modulus composite layup system 10 and to a composite structure 94, such as an aircraft composite structure 94a, that includes the controlled modulus composite layup system 10.

[0073] As shown in Figure 1B, lightning protection assembly (LSPMA) 66 can take the form of an uncured lightning protection assembly (LSPMA) 66a, a partially cured lightning protection assembly (LSPMA) 66b, and / or a cured lightning protection assembly (LSPMA) 66c. Lightning protection assembly 66 includes a lightning expanded metal foil layer 122 (see Figure 1B) that includes a lightning expanded metal foil 124 (see Figure 1B). As shown in Figure 1B, lightning expanded metal foil 124 has a lightning protection (LSP) coefficient of thermal expansion (CTE) 26d and a lightning protection (LSP) coefficient of hygroscopic expansion (CME) 28d.

[0074] As shown in FIG. 1B , the lightning strike expanded metal foil 124 includes a non-continuous metal foil (NON-CONT MF) 126, which may include one or more of a perforated metal foil (MF) 126a, an expanded metal foil (MF) 126b, a metal mesh 126c, a metalized fiber mesh 126d, a metal screen 126e, a metalized fiber fabric 126f, a metal fabric 126g, a wire mesh 126h, a metal foam 126i, an open-cell metal foam 126j, or other suitable non-continuous metal foil 126.

[0075] As shown in Figure 1B, the lightning strike expanded metal foil 124 further includes a metal material 128 or an alloy material 130. As further shown in Figure 1B, the metal material 128 includes one or more of copper 128a, aluminum 128b, titanium 128c, nickel 128d, gold 128e, silver 128f, or other suitable metal materials 128. As further shown in Figure 1B, the alloy material 130 includes one or more of copper alloy 130a, aluminum (Al) alloy 130b, titanium alloy 130c, nickel alloy 130d, gold alloy 130e, silver alloy 130f, bronze 130g, brass 130h, or other suitable alloy material 130.

[0076] 1B, in one exemplary embodiment, lightning strike protection assembly 66 further includes a resin-infused scrim layer 132 laminated to lightning expanded metal foil layer 122. As further shown in FIG. 1B, resin-infused scrim layer 132 includes a non-metallic scrim 134. As further shown in FIG. 1B, non-metallic scrim 134 includes one of a non-metallic scrim mat (SM) 134a, a fiberglass scrim mat (SM) 134b, a carbon fiber scrim mat (SM) 134c, a woven scrim mat (SM) 134d, a knit polyester scrim mat (SM) 134e, a non-woven scrim mat (SM) 134f, or other suitable non-metallic scrim 134. The non-metallic scrim 134 may further include a scrim weave or carrier medium having, for example, glass fiber 54 (see FIG. 1A) or fiberglass reinforcement, carbon fiber 50 (see FIG. 1A) or carbon fiber reinforcement, nylon fiber or nylon fiber reinforcement, polyester fiber or polyester reinforcement, or other suitable fiber reinforcement.

[0077] As further shown in FIG. 1B, the resin-infused scrim layer 132 includes a non-metallic scrim 134 infused with an infusion resin (IR) 136, also referred to as a lightning strike protection resin 138. As shown in FIG. 1B, the infusion resin 136 can take the form of an uncured infusion resin (IR) 136a, a partially cured infusion resin (IR) 136b, and / or a cured infusion resin (IR) 136c. For example, the uncured infusion resin 136a is in an uncured stage before cure 34 or co-cure 34a, the partially cured infusion resin 136b is in a partially cured stage during cure 34 or co-cure 34a, and the cured infusion resin 136c is in a cured or post-cured stage after cure 34 or co-cure 34a. The infusion resin 136 is initially heated during cure 34 or co-cure 34a, where the initial heating is sufficient to gel the infusion resin 136, such as the uncured infusion resin 136a, into a gel 140 (see FIG. 1A).

[0078] The infusion resin 136 may include a prepreg thermoset resin 118 (see FIG. 1B) similar to the prepreg thermoset resin 118 of the facing film 64, such as adhesive 118a, epoxy 118b, phenolic 118c, polyimide 118d, bismaleimide (BMI) 118e, polyurethane 118f, fluoropolymer 118g, cyanate ester 118h, or other suitable prepreg thermoset resin 118, as shown in FIG. 1B. The infusion resin 136 may further include a UV resistant infusion resin, an aliphatic epoxy infusion resin, a flame retardant infusion resin, or other suitable infusion resin 136.

[0079] In one embodiment, the infusion resin 136 has a tailored viscosity, a tailored cure profile, and a tailored rheology. As used herein, "tailored viscosity" refers to a viscosity of a resin, such as an infusion resin in a lightning strike protection assembly, that is controlled, tailored, and selected to be different from the viscosity of a structural resin in a structural assembly, such as a composite layup assembly, associated with the lightning expanded metal foil, to prevent or minimize intermixing of the infusion resin and the structural resin during co-curing or curing. As used herein, "tailored cure profile" refers to a curing profile of a resin, such as an infusion resin in a lightning strike protection assembly, that is optimized for cure time and cure temperature and that is controlled, tailored, and selected to be different from the curing profile of a structural resin in a structural assembly, such as a composite layup assembly, to allow crosslinking of the infusion resin to begin at a lower temperature and increase viscosity more quickly, thereby enabling the infusion resin to encapsulate the lightning expanded metal foil while minimizing intermixing and maintaining separation between the infusion resin and the structural resin. Additionally, as used herein, the term "tailored rheology" refers to the relationship between the flow behavior of a material's viscosity during curing (e.g., gelation or solidification) and the temperature of the material during curing, and means that the rheology of a resin, such as an infusion resin, contained in a lightning strike protection assembly is controlled, adjusted, and selected to be different from the rheology of a structural resin contained in a structural assembly, such as a composite laminate assembly, thereby preventing or minimizing mixing of the infusion resin and the structural resin during co-curing or curing, and increasing the microcrack resistance of a co-cured composite structure, such as a co-cured composite laminate comprising a structural component and a lightning strike protection component, when exposed to thermal humidity cycles.

[0080] The tailored viscosity, cure profile, and rheology of the infusion resin, in conjunction with the lightning strike expanded metal foil 124, provide chemical compatibility and bonding capabilities between the infusion resin 136 and the structural resin 60 (see FIG. 1A), such as the structural prepreg resin of a structural assembly 14 (see FIG. 1A), such as a composite laminate assembly 16 (see FIG. 1A) coupled or bonded to a lightning strike protection assembly 66.

[0081] The lightning protection assembly 66 is intended to be laid up 82 (see FIG. 1A) in a layup process 84 (see FIG. 1A) to be laminated to a structural assembly 14, such as a composite layup assembly 16, such as an uncured composite layup assembly 16a (see FIG. 1A), which in one embodiment comprises a plurality of structural layers 20 (see FIG. 1A), such as a plurality of structural ply layers 22 (see FIG. 1A), pre-impregnated with a structural resin 60. Alternatively, a structural assembly 14, such as a composite layup assembly 16, such as an uncured composite layup assembly 16a, may be laminated to the lightning protection assembly 66. In one aspect, the structural resin 60 has a structural resin viscosity, structural resin cure profile, and structural resin rheology that differs from the tailored viscosity, tailored cure profile, and tailored rheology of the infusion resin 136, thereby preventing intermixing of the structural resin 60 and the infusion resin 136 during co-cure 34a or cure 34, and allowing the infusion resin 136 to effectively encapsulate the lightning strike expanded metal foil 124 via encapsulation 142 (FIG. 1B) and form a defined resin boundary 144 (see FIG. 1B). Additionally, the tailored viscosity of the infusion resin 136 is higher than the structural resin viscosity of the structural resin 60, thereby preventing intermixing of the infusion resin 136 and the structural resin 60 during co-cure 34a or cure 34.

[0082] The tailored cure profile of the infusion resin 136 is tailored to initiate crosslinking at a lower temperature 38 (see FIG. 1A ), thereby rapidly increasing the tailored viscosity of the infusion resin 136, allowing the infusion resin 136 to effectively encapsulate the lightning-strike expanded metal foil 124 by encapsulation 142 while minimizing intermixing and maintaining separation between the infusion resin 136 and the structural resin 60. As used herein, "encapsulation" refers to solidifying the infusion resin, such as a thermosetting resin, around the lightning-strike expanded metal foil, thereby keeping the lightning-strike expanded metal foil separate from the structural resin and the composite laminate assembly or structure. For example, the infusion resin 136, infused into the non-metallic scrim 134, is forced into the multiple openings in the lightning-strike expanded metal foil 124 during co-cure 34a or cure 34b, such that the infusion resin 136 completely encapsulates or completely surrounds the lightning-strike expanded metal foil 124.

[0083] The tailored viscosity, tailored cure profile, and tailored rheology of the infusion resin 136 are tailored to be sufficiently different from the structural resin viscosity, structural resin cure profile, and structural resin rheology of the structural resin 60 so that the infusion resin 136 can fully and effectively encapsulate and protect the lightning strike expanded metal foil 124 for its lifetime. The infusion resin 136 has compatible properties with the structural resin 60 in relation to the heat and humidity cycle 100 so that resin intermixing between the infusion resin 136 and the structural resin 60 during co-cure 34a or cure 34 can be prevented or minimized.

[0084] In some embodiments, the lightning strike protection assembly 66 may further include one or more additional layers over the lightning strike expanded metal foil layer 122. For example, a thin layer of composite prepreg facing film 64 or other suitable layer may be added over or on the lightning strike expanded metal foil layer 122 to act as a protective or reinforcing layer over the lightning strike expanded metal foil layer 122, protecting the lightning strike expanded metal foil layer 122 during subsequent sanding, painting, polishing, or other secondary processing.

[0085] As further shown in FIG. 1B, in yet another embodiment, the facing layer 12 (or each of the facing layers 12, if present) includes one or more resin repair layers 68 applied directly to the composite layup assembly 16, such as the cured composite layup assembly 16c (see FIG. 1A), after curing 34 of the composite layup assembly 16. In one embodiment, the resin repair layer 68 includes a resin system 146 (see FIG. 1B) having reinforcing composite fibers 46a (see FIG. 1B). The resin system 146 preferably includes a prepreg thermoset resin 118 (see FIG. 1B) similar to the prepreg thermoset resin 118 included in the facing film 64 and the infusion resin 136 included in the lightning protection assembly 66. 1B, such prepreg thermoset resins 118 may include one or more of adhesive 118a, epoxy 118b, phenolic 118c, polyimide 118d, bismaleimide (BMI) 118e, polyurethane 118f, fluoropolymer 118g, cyanate ester 118h, or other suitable prepreg thermoset resins 118. Prepreg thermoset resins 118 may further include UV-resistant prepreg thermoset resins, aliphatic prepreg thermoset resins, flame-retardant prepreg thermoset resins, or other suitable prepreg thermoset resins 118. In another aspect, the resin repair layer 68 is comprised of a two-component epoxy resin that can be mixed at room temperature impregnated into a carbon fiber, glass fiber, or nylon fiber ply or scrim.

[0086] In another aspect of the present disclosure, an aircraft 200a (see FIG. 6) is provided having one or more aircraft composite structures 94a (see FIGS. 1A and 6) including the above-described adjusted modulus composite laminate system 10. The aircraft 200a includes one or more aircraft composite structures 94a. The one or more aircraft composite structures 94a include one or more of a fuselage 202 (see FIG. 6), one or more wings 204 (see FIG. 6), a tail section 208 (see FIG. 6) having one or more vertical stabilizers 210 (see FIG. 6) and horizontal stabilizers 212 (see FIG. 6), or other suitable aircraft composite structures 94a.

[0087] The controlled modulus composite laminate system 10 is incorporated into one or more aircraft composite structures 94a. The controlled modulus composite laminate system 10 includes a cured composite laminate assembly 16 that includes a plurality of structural ply layers 22 pre-impregnated with a structural resin 60. Each structural ply layer 22 has a ply modulus 25a (see FIG. 1A), a ply coefficient of thermal expansion 26a (see FIG. 1A), and a ply coefficient of moisture expansion 28a (see FIG. 1A), which are the same for each structural ply layer 22.

[0088] The controlled modulus composite layup system 10 further includes a surface layer 12 applied directly to the composite layup assembly 16 either before curing 34 or co-curing 34a of the composite layup assembly 16 or after curing 34 or co-curing 34a of the composite layup assembly 16. The surface layer 12 has a controlled modulus 25c that is less than or equal to the ply modulus 25a, a surface layer coefficient of thermal expansion 26b that is different from the ply coefficient of thermal expansion 26a, and a surface layer coefficient of hygroscopic expansion 28b that is different from the ply coefficient of hygroscopic expansion 28a.

[0089] The controlled modulus composite laminate system 10 further includes at least one secondary coating 70 applied directly to the surface layer 12 after curing 34 or co-curing 34a of the composite laminate assembly 16. The secondary coating 70 has a secondary coating modulus 25d (see FIG. 1A) that is different from the ply modulus 25a, a secondary coating thermal coefficient 26c (see FIG. 1A) that is different from the ply thermal coefficient 26a, and a secondary coating moisture coefficient 28c (see FIG. 1A) that is different from the ply moisture coefficient 28a. As shown in FIG. 1A, the secondary coating 70 includes one or more of a paint film 72, a basecoat film 72a, a topcoat film 72b, a primer film 74, an appliqué 76, a decal 78, or a repair paste film 80.

[0090] The tuned modulus 25c of the surface layer 12 is adjusted to be equal to or greater than the secondary coating modulus 25d of the at least one secondary coating 70 before the at least one secondary coating 70 is applied to the surface layer 12. The tuned modulus 25c is obtained by adjusting the surface layer modulus 25b (see FIG. 1A) of the surface layer 12, where the tuned modulus 25c is adjusted to be equal to or greater than the secondary coating modulus 25d, resulting in the surface layer 12 becoming a tuned modulus surface layer 12d (see FIG. 1B).

[0091] The surface layer 12 having the tailored modulus 25c, i.e., the tailored modulus surface layer 12d, forms a stress relief region 31 (see FIG. 1B), an integrated damping region 102 (see FIG. 1B), and an integrated buffer region 104 (see FIG. 1B) between the at least one secondary coating 70 and the plurality of structural ply layers 22 that make up the composite laminate assembly 16. This can enhance microcrack resistance 106 (see FIG. 1A), minimize structural degradation 108 (see FIG. 1A), and improve the humidity and heat cycling performance 110 (see FIG. 1A) of one or more aircraft composite structures 94a including the tailored modulus composite laminate system 10 when the one or more aircraft composite structures 94a are subjected to one or both of heat exposure 96 (see FIG. 1A) and moisture exposure 98 (see FIG. 1A) during a humidity and heat cycle 100 (see FIG. 1A) or a humidity and heat cycling event.

[0092] In one embodiment, the facing layer 12 of the controlled modulus composite laminate system 10 comprises a facing film 64 (see FIG. 1B) applied directly to the composite laminate assembly 16 prior to curing 34 or co-curing 34a of the composite laminate assembly 16, where the facing film 64 is co-cured or co-cured with the composite laminate assembly 16. In one embodiment, the facing film 64 comprises a resin-based composite facing film 64a (see FIG. 1A). The facing film 64 may also be any other suitable type of facing film 64. The facing film 64 comprises a film layer 65 (see FIG. 1B) impregnated or infused with a prepreg thermoset resin 118. 1B, the prepreg thermoset resin 118 may include adhesive 118a, epoxy 118b, phenolic 118c, polyimide 118d, bismaleimide (BMI) 118e, polyurethane 118f, fluoropolymer 118g, cyanate ester 118h, or other suitable prepreg thermoset resin 118. The infusion resin 136 may further include a UV resistant infusion resin, an aliphatic epoxy infusion resin, a flame retardant infusion resin, or other suitable infusion resin 136.

[0093] In another embodiment, the facing layer 12 of the controlled modulus composite laminate system 10 includes a lightning strike protection assembly 66 (see FIG. 1B ) applied directly to the composite laminate assembly 16 prior to curing 34 or co-curing 34a of the composite laminate assembly 16, where the lightning strike protection assembly 66 is co-cured or co-cured with the composite laminate assembly 16. In one embodiment, the lightning strike protection assembly 66 includes a lightning strike expanded metal foil layer 122 (see FIG. 1B ) including a lightning strike expanded metal foil 124 (see FIG. 1B ) and a resin-infused scrim layer 132 (see FIG. 1B ) laminated to the lightning strike expanded metal foil layer 122. The resin-infused scrim layer 132 includes a non-metallic scrim 134 (see FIG. 1B ) infused with an infusion resin 136 (see FIG. 1B ).

[0094] The lightning strike expanded metal foil 124 includes a discontinuous metal foil 126, which may include one or more of perforated metal foil 126a, expanded metal foil 126b, metal mesh 126c, metalized fiber mesh 126d, metal screen 126e, metalized fiber fabric 126f, metal fabric 126g, wire mesh 126h, metal foam 126i, open-cell metal foam 126j, or other suitable discontinuous metal foil 126, as shown in FIG. 1B.

[0095] The lightning strike expanded metal foil 124 further includes a metal material 128 or an alloy material 130. The metal material 128 may include one or more of copper 128a, aluminum 128b, titanium 128c, nickel 128d, gold 128e, silver 128f, or other suitable metal materials 128, as shown in FIG. 1B. The alloy material 130 may include one or more of copper alloy 130a, aluminum (Al) alloy 130b, titanium alloy 130c, nickel alloy 130d, gold alloy 130e, silver alloy 130f, bronze 130g, brass 130h, or other suitable alloy materials 130, as shown in FIG. 1B.

[0096] 1B, the non-metallic scrim 134 may include one of non-metallic scrim mat 134a, fiberglass scrim mat 134b, carbon fiber scrim mat 134c, woven scrim mat 134d, knitted polyester scrim mat 134e, non-woven scrim mat 134f, or other suitable non-metallic scrim 134. The non-metallic scrim 134 may further include a woven scrim or carrier medium having fiberglass 54 (see FIG. 1A) or fiberglass reinforcement, carbon fiber 50 (see FIG. 1A) or carbon fiber reinforcement, nylon fiber or nylon fiber reinforcement, polyester fiber or polyester reinforcement, or other suitable fiber reinforcement.

[0097] The infusion resin 136 may include a prepreg thermoset resin 118 (see FIG. 1B) similar to the prepreg thermoset resin 118 of the facing film 64, including adhesive 118a, epoxy 118b, phenolic 118c, polyimide 118d, bismaleimide (BMI) 118e, polyurethane 118f, fluoropolymer 118g, cyanate ester 118h, or other suitable prepreg thermoset resin 118, as shown in FIG. 1B. The infusion resin 136 may also include a UV resistant infusion resin, an aliphatic epoxy infusion resin, a flame retardant infusion resin, or other suitable infusion resin 136.

[0098] In yet another embodiment, the facing layer 12 of the controlled modulus composite laminate system 10 includes one or more resin repair layers 68 applied directly to the composite laminate assembly 16 after curing 34 of the composite laminate assembly 16. In one embodiment, the one or more resin repair layers 68 include a resin-based material 146 having reinforcing composite fibers 46a.

[0099] 2A-2E, which illustrate an exemplary controlled modulus composite laminate system 10 (see FIGS. 2C-2E) and an exemplary controlled modulus composite laminate assembly 11 (see FIGS. 2A-2B) of the present disclosure, in which a surface layer 12, such as controlled modulus surface layer 12d, is in the form of a surface film 64.

[0100] Referring first to Figure 2A, Figure 2A is a cross-sectional front view of an exemplary tailored modulus composite layup assembly 11 of the present disclosure having a facing layer 12, e.g., uncured facing layer 12a, in the form of a facing film 64, e.g., resin-based composite facing film 64a, applied and laminated to a structural assembly 14, e.g., a composite layup assembly 16 in the form of an uncured composite layup assembly 16a, to form a layup assembly 86, e.g., an uncured layup assembly 86a.

[0101] As shown in Figure 2A, the facing film 64, e.g., resin-based composite facing film 64a, includes a film layer 65 impregnated or infused with a prepreg thermoset resin 118. As further shown in Figure 2A, the facing film 64, e.g., resin-based composite facing film 64a, has a first surface 150, e.g., upper surface 150a, a second surface 152, e.g., lower surface 152a, and opposite ends 154. The facing layer 12, e.g., resin-based composite facing film 64a, has a tailored modulus 25c (see Figure 1B), and is a tailored modulus facing layer 12d.

[0102] As shown in Figure 2A, structural assembly 14 may be a composite layup assembly 16, for example, in the form of an uncured composite layup assembly 16a, including a plurality of structural layers 20, such as a plurality of structural ply layers 22 comprised of a plurality of plies 24. Each structural ply layer 22 is pre-impregnated with a structural resin 60 in the form of an uncured structural resin 60a, for example, a thermoset structural resin 62. Structural assembly 14 shown in Figure 2A may be a composite layup assembly 16, for example, in the form of a panel 18.

[0103] 2A, a structural assembly 14, such as a composite laminate assembly 16, includes four structural layers 20, such as four structural ply layers 22. Each structural ply layer 22 includes one ply 24, and multiple structural ply layers 22 include multiple plies 24. In other embodiments, a structural assembly 14, such as a composite laminate assembly 16, may include more or less than four structural layers 20, such as structural ply layers 22.

[0104] As further shown in FIG. 2A, the plurality of plies 24 includes a top ply 24a, a bottom ply 24b, and two intermediate plies 24c laminated between the top ply 24a and the bottom ply 24b. However, the plurality of plies 24 may include more than two intermediate plies 24c. As shown in FIG. 2A, the plurality of plies 24 is in the form of structural plies 24d. Each of the top ply 24a, the bottom ply 24b, and the intermediate ply 24c has a first surface 155 (see FIG. 2A), e.g., upper surface 155a (see FIG. 2A), a second surface 156 (see FIG. 2A), e.g., lower surface 156a (see FIG. 2A), and opposite ends 157 (see FIG. 2A).

[0105] As shown in FIG. 2A , a facing layer 12, such as a facing film 64, has its second surface 152, which is its lower surface 152a, laminated to, directly applied to, bonded to, and in continuous contact with a first surface 155, which is the upper surface 155a of a top ply 24a of a plurality of structural ply layers 22 of a structural assembly 14, such as a composite laminate assembly 16.

[0106] Reference is now made to FIG. 2B, which is a cross-sectional front view of the adjusted modulus composite layup assembly 11 shown in FIG. 2A during co-cure 34a in the interior 158 of a heating apparatus 88, such as an autoclave 90. The layup assembly 86 shown in FIG. 2B is, for example, a partially cured layup assembly 86b, with a surface layer 12 in the form of a partially cured surface layer 12b, such as a surface film 64, such as a resin-based composite surface film 64a, applied and laminated to a structural assembly 14, such as a composite layup assembly 16 in the form of a partially cured composite layup assembly 16b. The surface layer 12 is, for example, a surface film 64, such as a resin-based composite surface film 64a, having an adjusted modulus 25c (see FIG. 1B), and is an adjusted modulus surface layer 12d (see FIG. 2B). The structural assembly 14 is, for example, a composite layup assembly 16 in the form of a panel 18 (see FIG. 2B). FIG. 2B further illustrates structural resin 60 in the form of partially cured structural resin 60b.

[0107] A layup assembly 86, such as the tailored modulus composite laminate assembly 11, is co-cured or cured at an elevated temperature 38 (see FIG. 1A). As shown in FIG. 2B, the plurality of structural ply layers 22 includes different plies 24, such as a top ply 24a, a bottom ply 24b, and an intermediate ply 24c, which tend to expand and contract at different rates due to the anisotropy of the material. However, because the plies 24 are bonded together, this expansion 112 (see FIG. 2B) and contraction 114 (see FIG. 2B) manifests as stress 30 (see FIG. 2B). Specifically, FIG. 2B shows the expansion 112 and stress 30 in the intermediate ply 24c adjacent to the bottom ply 24b, as well as the contraction 114 and stress 30 in the top ply 24a.

[0108] Reference is now made to Figure 2C, which is a cross-sectional front view of an exemplary adjusted modulus composite layup system 10 of the present disclosure, in which a secondary coating 70, such as a paint film 72, has been applied to a layup assembly 86, which is a cured layup assembly 86c. This layup assembly includes the co-cured or cured adjusted modulus composite layup assembly 11 shown in Figure 2B. As shown in Figure 2C, the secondary coating 70, such as a paint film 72, has a first surface 160, e.g., an upper surface 160a, a second surface 162, e.g., a lower surface 162a, and opposite ends 164.

[0109] As shown in FIG. 2C, a secondary coating 70, such as a paint film 72, has its second surface 162, i.e., lower surface 162a, applied directly to, bonded to, and in continuous contact with a first surface 150, i.e., upper surface 150a, of a surface layer 12, such as cured surface layer 12c, in the form of a surface film 64, such as resin-based composite surface film 64a. Surface layer 12 includes an intermediate layer 12e positioned between the secondary coating 70 disposed on the surface layer 12 and a structural assembly 14, such as composite laminate assembly 16, i.e., cured composite laminate assembly 16c. Surface layer 12, i.e., resin-based composite surface film 64a, having a tailored modulus 25c (see FIG. 1B), is tailored modulus surface layer 12d (see FIG. 2C). Structural assembly 14 is a composite laminate assembly 16 in the form of, for example, panel 18 (see FIG. 2C). Figure 2C further illustrates a plurality of structural ply layers 22 made up of a plurality of plies 24. Figure 2C also illustrates structural resin 60 in the form of cured structural resin 60c.

[0110] Reference is now made to Figure 2D, which is a cross-sectional front view of the controlled modulus composite laminate system 10 shown in Figure 2C, illustrating the system undergoing thermal exposure 96, such as heat 36, and moisture exposure 98, such as moisture 165 from rain 165a. As shown in Figure 2D, the controlled modulus composite laminate system 10 includes a secondary coating 70, such as paint film 72, applied to a surface layer 12, which is a hardened surface layer 12c in the form of a surface film 64, such as resin-based composite surface film 64a. The surface layer 12 is, for example, a surface film 64, such as resin-based composite surface film 64a, having a controlled modulus 25c (see Figure 1B), which is controlled modulus surface layer 12d. As further shown in FIG. 2D, the controlled modulus composite laminate system 10 includes a structural assembly 14, such as a composite laminate assembly 16, e.g., a cured composite laminate assembly 16c, which comprises a plurality of structural ply layers 22 including different plies 24, e.g., a top ply 24a, a bottom ply 24b, and an intermediate ply 24c.

[0111] As further shown in FIGURE 2D , when a structural assembly 14, such as composite laminate assembly 16, e.g., panel 18, and a controlled modulus composite laminate system 10 including such panel 18, is subjected to thermal exposure 96, e.g., heat 36, or moisture exposure 98, e.g., moisture 165, some of the structural ply layers 22, e.g., plies 24, will experience expansion 112, e.g., hygroscopic expansion 112a, and some of the structural ply layers 22, e.g., plies 24, surface layers 12, e.g., controlled modulus surface layer 12d, and secondary coating 70, will experience different amounts of expansion 112, e.g., hygroscopic expansion 112a, and different amounts of shrinkage 114, e.g., hygroscopic shrinkage 114a. This becomes even more pronounced when stress 30 (see FIGURE 2D ) is present. Specifically, FIG. 2D illustrates shrinkage 114, such as thermal drying shrinkage 114a, and stresses 30 in the top ply 24a, and also illustrates expansion 112, such as thermal hygroscopic expansion 112a, and stresses 30 in the middle ply 24c adjacent to the bottom ply 24b.

[0112] Reference is now made to Figure 2E, a cross-sectional front view of the tailored modulus composite laminate system 10 shown in Figure 2D, illustrating microcracks 116, such as surface cracks 116a, after heat exposure 96, such as heat 36 shown in Figure 2D, and moisture exposure 98, such as moisture 165 from rain 165a. In Figure 2E, the microcracks 116 initiate as surface cracks 116a in the paint film 72, which is the secondary coating 70, and extend into the surface layer 12, which is the cured surface layer 12c in the form of a surface film 64, such as resin-based composite surface film 64a. The surface layer 12 is the surface film 64, such as resin-based composite surface film 64a, having a tailored modulus 25c (see Figure 1B), which is tailored modulus surface layer 12d.

[0113] As further shown in Figure 2E, the adjusted modulus composite laminate system 10 does not exhibit internal cracks 116b (see Figure 1A), and the microcracks 116 do not extend into a structural assembly 14, such as a composite laminate assembly 16 in the form of a cured composite laminate assembly 16c having a structural ply layer 22 comprised of multiple plies 24 underlying a surface layer 12 and a secondary coating 70. As shown in Figure 2E, the structural assembly 14 may be a composite laminate assembly 16 in the form of a panel 18, for example.

[0114] The tailored modulus surface layer 12d of Figure 2E minimizes the strain mismatch 32d (see Figure 1A) between the surface layer 12d and the secondary coating 70. The tailored modulus surface layer 12d of Figure 2E reliably accommodates the magnitude variations between the surface layer 12d and the underlying structural ply layer 22 of plies 24, resulting in enhanced microcracking resistance 106 (see Figure 1A), minimized structural degradation 108 (see Figure 1A), and improved humidity heat cycling performance 110 (see Figure 1A). In contrast, known structural panels and composite laminates that are not modulus tailored have a large strain mismatch between the structural panel or composite laminate and a secondary coating, such as paint, which can result in microcracking or other cracking of the structural panel or composite laminate when exposed to heat or moisture.

[0115] 3A-3E, which illustrate an exemplary controlled modulus composite laminate system 10 (see FIGS. 3C-3E) of the present disclosure and an exemplary controlled modulus composite laminate assembly 11 (see FIGS. 3A-3B) of the present disclosure, in which surface layer 12, e.g., controlled modulus surface layer 12d, is in the form of a lightning strike protection assembly 66.

[0116] Referring first to Figure 3A, Figure 3A is a cross-sectional front view of an exemplary tailored modulus composite layup assembly 11 of the present disclosure having a surface layer 12, e.g., uncured surface layer 12a, in the form of a lightning protection assembly 66, such as uncured lightning protection assembly 66a, applied and laid up to a structural assembly 14, such as a composite layup assembly 16 in the form of uncured composite layup assembly 16a, to form a layup assembly 86, i.e., uncured layup assembly 86a.

[0117] As shown in Figure 3A, lightning strike protection assembly 66 includes lightning expanded metal foil layer 122 including lightning expanded metal foil 124. As further shown in Figure 3A, lightning expanded metal foil layer 122 includes first surface 166, e.g., upper surface 166a, second surface 168, e.g., lower surface 168a, and opposite ends 170. Surface layer 12, e.g., lightning strike protection assembly 66, includes adjusted modulus 25c (see Figure 1B), and is adjusted modulus surface layer 12d.

[0118] 3A, the lightning strike protection assembly 66 further includes a resin-infused scrim layer 132 laminated to the lightning strike expanded metal foil layer 122. The resin-infused scrim layer 132 includes a non-metallic scrim 134 infused with an infusion resin 136 in the form of uncured infusion resin 136a, such as prepreg thermoset resin 118. As further shown in FIG. 3A, the resin-infused scrim layer 132 has a first surface 172, e.g., an upper surface 172a, a second surface 174, e.g., a lower surface 174a, and opposite ends 176.

[0119] 3A , the second side 168, or lower surface 168a, of the lightning-strike expanded metal foil layer 122 is laminated to, directly applied to, bonded to, and in continuous contact with the first side 172, or upper surface 172a, of the resin-infused scrim layer 132. As further shown in FIG. 3A , the second side 174, or lower surface 174a, of the resin-infused scrim layer 132 is laminated to, directly applied to, bonded to, and in continuous contact with the first side 155, or upper surface 155a, of the top ply 24a of the plurality of structural ply layers 22 of a structural assembly 14, such as composite laminate assembly 16.

[0120] As shown in Figure 3A, structural assembly 14 is a composite layup assembly 16, for example, in the form of an uncured composite layup assembly 16a, that includes a plurality of structural layers 20, such as a plurality of structural ply layers 22 comprised of a plurality of plies 24. Each structural ply layer 22 is pre-impregnated with a structural resin 60, for example, in the form of an uncured structural resin 60a, such as, for example, a thermosetting structural resin 62. Structural assembly 14 shown in Figure 3A is a composite layup assembly 16, for example, in the form of a panel 18.

[0121] 3A, a structural assembly 14, such as a composite laminate assembly 16, includes four structural layers 20, such as four structural ply layers 22. Each structural ply layer 22 includes one ply 24, and multiple structural ply layers 22 include multiple plies 24. In other embodiments, a structural assembly 14, such as a composite laminate assembly 16, may include more or less than four structural layers 20, such as structural ply layers 22.

[0122] As further shown in FIG. 3A, the plurality of plies 24 includes a top ply 24a, a bottom ply 24b, and two intermediate plies 24c laminated between the top ply 24a and the bottom ply 24b. However, the plurality of plies 24 may include more than two intermediate plies 24c. As shown in FIG. 3A, the plurality of plies 24 is in the form of structural plies 24d. Each of the top ply 24a, the bottom ply 24b, and the intermediate ply 24c has a first surface 155 (see FIG. 3A), e.g., upper surface 155a (see FIG. 3A), a second surface 156 (see FIG. 3A), e.g., lower surface 156a (see FIG. 3A), and opposite ends 157 (see FIG. 3A).

[0123] Reference is now made to Figure 3B, which is a cross-sectional front view of the adjusted modulus composite laminate assembly 11 shown in Figure 3A during co-curing 34a in the interior 158 of a heating apparatus 88, such as an autoclave 90. The layup assembly 86 shown in Figure 3B is, for example, a partially cured layup assembly 86b, and the surface layer 12 of this assembly is a partially cured surface layer 12b in the form of a lightning protection assembly 66, such as partially cured lightning protection assembly 66b. The surface layer 12, for example, is lightning protection assembly 66, and has an adjusted modulus 25c (see Figure 1B), and is an adjusted modulus surface layer 12d.

[0124] As shown in FIG. 3B, the lightning strike protection assembly 66 includes a lightning strike expanded metal foil layer 122 including a lightning strike expanded metal foil 124 and a resin-infused scrim layer 132, such as a non-metallic scrim 134 infused with an infusion resin 136, which may be a prepreg thermoset resin 118, for example, in the form of a partially cured infusion resin 136b. FIG. 3B illustrates the lightning strike protection assembly 66 in which the composite layup assembly 16 is laminated to a structural assembly 14, such as in the form of a partially cured composite layup assembly 16b. Specifically, the resin-infused scrim layer 132 is laminated to a first surface 155 (see FIG. 3B), e.g., top surface 155a (see FIG. 3B), of the top ply 24a of the composite layup assembly 16. The structural assembly 14 is a composite layup assembly 16 in the form of, for example, a panel 18 (see FIG. 3B). FIG. 3B further illustrates structural resin 60 in the form of partially cured structural resin 60b.

[0125] A layup assembly 86, such as the tailored modulus composite laminate assembly 11, is co-cured or cured at an elevated temperature 38 (see FIG. 1A). As shown in FIG. 3B, the plurality of structural ply layers 22 includes different plies 24, such as a top ply 24a, a bottom ply 24b, and an intermediate ply 24c, which tend to expand and contract at different rates due to the anisotropy of the material. However, because the plies 24 are bonded together, this expansion 112 (see FIG. 3B) and contraction 114 (see FIG. 3B) manifests as stress 30 (see FIG. 3B). Specifically, FIG. 3B shows the expansion 112 and stress 30 in the intermediate ply 24c adjacent to the bottom ply 24b, as well as the contraction 114 and stress 30 in the top ply 24a.

[0126] Reference is now made to Figure 3C, which is a cross-sectional front view of an exemplary adjusted modulus composite layup system 10 of the present disclosure, in which a secondary coating 70, such as a paint film 72, has been applied to a layup assembly 86, which is a cured layup assembly 86c. This layup assembly includes the co-cured or cured adjusted modulus composite layup assembly 11 shown in Figure 3B. As shown in Figure 3C, the secondary coating 70, such as the paint film 72, has a first surface 160, e.g., an upper surface 160a, a second surface 162, e.g., a lower surface 162a, and opposite ends 164.

[0127] 3C, a secondary coating 70, such as paint film 72, has its second surface 162, which is its lower surface 162a, applied directly to, bonded to, and in continuous contact with a first surface 166, which is the upper surface 166a of a lightning expanded metal foil layer 122 of a lightning protection assembly 66, such as a cured lightning protection assembly 66c. As shown in FIG. 3C, the lightning protection assembly 66 includes a lightning expanded metal foil 124, which includes the lightning expanded metal foil layer 122, and a resin-infused scrim layer 132, such as a non-metallic scrim 134 infused with an infusion resin 136, which may be, for example, a prepreg thermoset resin 118 in the form of a cured infusion resin 136c.

[0128] In this embodiment, a facing layer 12, e.g., cured facing layer 12c, is a lightning strike protection assembly 66 including an intermediate layer 12e positioned between a secondary coating 70 disposed on the facing layer 12 and a structural assembly 14, e.g., a composite layup assembly 16, e.g., cured composite layup assembly 16c. The facing layer 12, e.g., lightning strike protection assembly 66, has a tailored modulus 25c (see FIG. 1B) and is a tailored modulus facing layer 12d (see FIG. 3C). The structural assembly 14 is a composite layup assembly 16 in the form of, e.g., a panel 18 (see FIG. 3C). FIG. 3C further illustrates a plurality of structural ply layers 22 made up of a plurality of plies 24. FIG. 3C further illustrates a structural resin 60 in the form of a cured structural resin 60c.

[0129] 3C further illustrates the defined resin boundary 144 formed between the resin-infused scrim layer 132 of the lightning strike protection assembly 66 and the top ply 24a of the composite laminate assembly 16. The defined resin boundary 144 serves to prevent or minimize intermixing of the infusion resin 136 and the structural resin 60 during co-cure 34a or cure 34.

[0130] Reference is now made to Figure 3D, which is a cross-sectional front view of the adjusted modulus composite laminate system 10 shown in Figure 3C, illustrating the system undergoing heat exposure 96, such as heat 36, and moisture exposure 98, such as moisture 165 from rain 165a. As shown in Figure 3D, the adjusted modulus composite laminate system 10 includes a secondary coating 70, such as paint film 72, applied to a surface layer 12, such as a hardened surface layer 12c in the form of a lightning protection assembly 66, such as hardened lightning protection assembly 66c. The surface layer 12, such as lightning protection assembly 66, has an adjusted modulus 25c (see Figure 1B), and is adjusted modulus surface layer 12d (see Figure 3D).

[0131] 3D, the lightning strike protection assembly 66 includes a lightning expanded metal foil 124 that includes the lightning expanded metal foil layer 122 and a resin-infused scrim layer 132, such as a non-metallic scrim 134 infused with an infusion resin 136, which may be, for example, a prepreg thermoset resin 118 in the form of cured infusion resin 136c. FIG. 3D also shows a defined resin boundary 144 formed between the resin-infused scrim layer 132 of the lightning strike protection assembly 66 and the top ply 24a of the composite laminate assembly 16.

[0132] As further shown in FIG. 3D, the tailored modulus composite laminate system 10 includes a structural assembly 14, such as a composite laminate assembly 16, e.g., a cured composite laminate assembly 16c, which comprises a plurality of structural ply layers 22 including different plies 24, e.g., a top ply 24a, a bottom ply 24b, and an intermediate ply 24c.

[0133] As further shown in FIGURE 3D , when a structural assembly 14, such as composite laminate assembly 16, e.g., panel 18, and a controlled modulus composite laminate system 10 including such panel 18, is subjected to heat exposure 96, e.g., heat 36, and moisture exposure 98, e.g., moisture 165, some of the structural ply layers 22, comprised of plies 24, will experience expansion 112, e.g., hygroscopic thermal expansion 112a, and some of the structural ply layers 22, comprised of plies 24, will experience shrinkage 114, e.g., thermal drying shrinkage 114a. In addition, different materials, such as plies 24, surface layers 12, e.g., controlled modulus surface layer 12d, and secondary coating 70, will experience different amounts of expansion 112, e.g., hygroscopic thermal expansion 112a, and different amounts of shrinkage 114, e.g., thermal drying shrinkage 114a. This becomes even more pronounced when stress 30 (see FIGURE 3D ) is present. Specifically, FIG. 3D illustrates shrinkage 114, such as thermal drying shrinkage 114a, and stresses 30 in the top ply 24a, and also illustrates expansion 112, such as thermal hygroscopic expansion 112a, and stresses 30 in the middle ply 24c adjacent to the bottom ply 24b.

[0134] Reference is now made to Figure 3E, which is a cross-sectional front view of the tailored modulus composite laminate system 10 shown in Figure 3D, illustrating microcracks 116, such as surface cracks 116a, after exposure to heat 96, such as heat 36 shown in Figure 3D, and moisture exposure 98, such as moisture 165 from rain 165a. In Figure 3E, the microcracks 116 appear as surface cracks 116a in the paint film 72, which is the secondary coating 70, and extend into the surface layer 12, which is the hardened surface layer 12c of a lightning protection assembly 66, such as hardened lightning protection assembly 66c, which has a tailored modulus 25c (see Figure 1B), and is tailored modulus surface layer 12d.

[0135] 3E, the lightning strike protection assembly 66 includes a lightning expanded metal foil 124, which includes the lightning expanded metal foil layer 122, and a resin-infused scrim layer 132, such as a non-metallic scrim 134 infused with an infusion resin 136, which may be, for example, a prepreg thermoset resin 118 in the form of cured infusion resin 136c. FIG. 3E also illustrates a defined resin boundary 144 defined between the resin-infused scrim layer 132 of the lightning strike protection assembly 66 and the top ply 24a of the composite laminate assembly 16.

[0136] As further shown in Figure 3E, the adjusted modulus composite laminate system 10 does not exhibit internal cracks 116b (see Figure 1A), and the microcracks 116 do not extend into a structural assembly 14, such as a composite laminate assembly 16 in the form of a cured composite laminate assembly 16c having a structural ply layer 22 comprised of multiple plies 24 underlying a surface layer 12 and a secondary coating 70. As shown in Figure 3E, the structural assembly 14 is a composite laminate assembly 16 in the form of, for example, a panel 18.

[0137] The tailored modulus surface layer 12d of Figure 3E minimizes the strain mismatch 32d (see Figure 1A) between the surface layer 12d and the secondary coating 70. The tailored modulus surface layer 12d of Figure 3E reliably accommodates the magnitude variations between the surface layer 12d and the underlying structural ply layer 22 of plies 24, resulting in improved microcracking resistance 106 (see Figure 1A), minimized structural degradation 108 (see Figure 1A), and improved humidity heat cycling performance 110 (see Figure 1A).

[0138] 4A-4E, which illustrate an exemplary controlled modulus composite laminate system 10 (see FIGS. 4C-4E) of the present disclosure and an exemplary structural assembly 14, such as a composite laminate assembly 16, e.g., uncured composite laminate assembly 16a, with a facing layer 12, e.g., controlled modulus facing layer 12d, in the form of a resin repair layer 68.

[0139] Referring first to Figure 4A, Figure 4A is a cross-sectional front view of an exemplary structural assembly 14 included in an exemplary tailored modulus composite laminate system 10 (see Figure 4C) of the present disclosure, such as a composite laminate assembly 16 including an uncured composite laminate assembly 16a.

[0140] As shown in Figure 4A, structural assembly 14 is a composite layup assembly 16, for example, in the form of an uncured composite layup assembly 16a, that includes a plurality of structural layers 20, such as a plurality of structural ply layers 22 comprised of a plurality of plies 24. Each structural ply layer 22 is pre-impregnated with a structural resin 60, for example, in the form of an uncured structural resin 60a, such as, for example, a thermosetting structural resin 62. Structural assembly 14 shown in Figure 4A is a composite layup assembly 16, for example, in the form of a panel 18.

[0141] 4A, a structural assembly 14, such as a composite laminate assembly 16, includes four structural layers 20, such as four structural ply layers 22. Each structural ply layer 22 includes one ply 24, and multiple structural ply layers 22 include multiple plies 24. In other embodiments, a structural assembly 14, such as a composite laminate assembly 16, may include more or less than four structural layers 20, such as structural ply layers 22.

[0142] As further shown in FIG. 4A, the plurality of plies 24 includes a top ply 24a, a bottom ply 24b, and two intermediate plies 24c laminated between the top ply 24a and the bottom ply 24b. However, the plurality of plies 24 may include more than two intermediate plies 24c. As shown in FIG. 4A, the plurality of plies 24 is in the form of a structural ply 24d. As shown in FIG. 4A, each of the top ply 24a, the bottom ply 24b, and the intermediate ply 24c has a first surface 155, e.g., upper surface 155a, a second surface 156, e.g., lower surface 156a, and opposite ends 157.

[0143] Reference is now made to Figure 4B, which is a cross-sectional front view of the structural assembly 14 shown in Figure 4A, which is a composite layup assembly 16, such as, for example, a partially cured composite layup assembly 16b, included in the exemplary tailored modulus composite layup system 10 (see Figure 4C) of the present disclosure. This figure illustrates the composite layup assembly 16 being cured in the interior 158 of a heating device 88, such as an autoclave 90. As shown in Figure 4B, the composite layup assembly 16 includes, for example, a partially cured composite layup assembly 16b in the form of a panel 18. Figure 4B also illustrates a structural resin 60 in the form of a partially cured structural resin 60b.

[0144] During curing 34 (see FIG. 4B), the composite laminate assembly 16 is cured at an elevated temperature 38 (see FIG. 1A). As shown in FIG. 4B, the plurality of structural ply layers 22 includes different plies 24, such as a top ply 24a, a bottom ply 24b, and an intermediate ply 24c, which tend to expand and contract at different rates due to the anisotropy of their materials. However, because the plies 24 are bonded together, this expansion 112 (see FIG. 4B) and contraction 114 (see FIG. 4B) manifests as stresses 30 (see FIG. 4B). Specifically, FIG. 4B illustrates the expansion 112 and stress 30 in the intermediate ply 24c adjacent to the bottom ply 24b, as well as the contraction 114 and stress 30 in the top ply 24a.

[0145] Reference is now made to Figure 4C, which is a cross-sectional front view of an exemplary controlled modulus composite laminate system 10 of the present disclosure, including a surface layer 12, such as a controlled modulus surface layer 12d in the form of a resin repair layer 68, applied to a structural assembly 14, such as cured composite laminate assembly 16c, formed from the cured composite laminate assembly 16 shown in Figure 4B. The controlled modulus composite laminate system 10 shown in Figure 4C also includes a secondary coating 70, such as a paint film 72, applied to the surface layer 12.

[0146] As shown in Figure 4C, the secondary coating 70, such as the paint film 72, has a first surface 160, e.g., an upper surface 160a, a second surface 162, e.g., a lower surface 162a, and opposite ends 164. As further shown in Figure 4C, the resin repair layer 68 includes a resin-based material 146, such as a prepreg thermoset resin 118 having reinforcing composite fibers 46c.

[0147] As shown in FIG. 4C, a secondary coating 70, such as a paint film 72, has its second surface 162, i.e., lower surface 162a, applied directly to, bonded to, and in continuous contact with a first surface 150, i.e., upper surface 150a, of a surface layer 12, such as cured surface layer 12c in the form of resin repair layer 68. Surface layer 12 includes an intermediate layer 12e positioned between secondary coating 70 disposed on surface layer 12 and a structural assembly 14, such as composite layup assembly 16, i.e., cured composite layup assembly 16c. Structural assembly 14 is a composite layup assembly 16 in the form of panel 18 (see FIG. 2C). FIG. 4C further illustrates a plurality of structural ply layers 22 comprised of a plurality of plies 24. FIG. 4C also illustrates a structural resin 60 in the form of cured structural resin 60c.

[0148] Reference is now made to Figure 4D, which is a cross-sectional front view of the adjusted modulus composite laminate system 10 shown in Figure 4C, illustrating the system undergoing heat exposure 96, such as heat 36, and moisture exposure 98, such as moisture 165 from rain 165a. As shown in Figure 4D, the adjusted modulus composite laminate system 10 includes a secondary coating 70, such as paint film 72, applied to a hardened surface layer 12c in the form of a resin repair layer 68. The surface layer 12, for example, is a resin repair layer 68 having an adjusted modulus 25c (see Figure 1B), and is an adjusted modulus surface layer 12d (see Figure 4D). As further shown in FIG. 4D, the controlled modulus composite laminate system 10 includes a structural assembly 14, such as a composite laminate assembly 16, e.g., a cured composite laminate assembly 16c, which comprises a plurality of structural ply layers 22 including different plies 24, e.g., a top ply 24a, a bottom ply 24b, and an intermediate ply 24c.

[0149] As further shown in FIG. 4D , when a structural assembly 14, such as composite laminate assembly 16, e.g., panel 18, and a controlled modulus composite laminate system 10 including such panel 18, is subjected to heat exposure 96, such as heat 36, and moisture exposure 98, such as moisture 165, some of the structural ply layers 22, comprising plies 24, will experience expansion 112, such as hygroscopic thermal expansion 112a, and some of the structural ply layers 22, comprising plies 24, will experience shrinkage 114, such as thermal drying shrinkage 114a. In addition, different materials, such as plies 24, surface layers 12, such as controlled modulus surface layer 12d, and secondary coating 70, will experience different amounts of expansion 112 (see FIG. 4D ), such as hygroscopic thermal expansion 112a (see FIG. 4D ), and different amounts of shrinkage 114, such as thermal drying shrinkage 114a (see FIG. 4D ). This becomes even more pronounced when stress 30 (see FIG. 4D ) is present. Specifically, FIG. 4D illustrates shrinkage 114, such as thermal drying shrinkage 114a, and stress 30 in the top ply 24a, and also illustrates expansion 112, such as thermal hygroscopic expansion 112a, and stress 30 in the middle ply 24c adjacent to the bottom ply 24b.

[0150] Reference is now made to Figure 4E, a cross-sectional front view of the tailored modulus composite laminate system 10 shown in Figure 4D, illustrating microcracks 116, such as surface cracks 116a, after exposure to heat 96, such as heat 36 shown in Figure 4D, and moisture exposure 98, such as moisture 165 from rain 165a. In Figure 4E, the microcracks 116 initiate as surface cracks 116a in the paint film 72, which is the secondary coating 70, and extend into the surface layer 12, which is the hardened surface layer 12c in the form of, for example, resin repair layer 68. The surface layer 12 is, for example, the resin repair layer 68, which has tailored modulus 25c (see Figure 1B), and is tailored modulus surface layer 12d.

[0151] As further shown in Figure 4E, the adjusted modulus composite laminate system 10 does not exhibit internal cracks 116b (see Figure 1A), and the microcracks 116 do not extend into a structural assembly 14, such as a composite laminate assembly 16 in the form of a cured composite laminate assembly 16c having a structural ply layer 22 comprised of multiple plies 24 underlying a surface layer 12 and a secondary coating 70. As shown in Figure 4E, the structural assembly 14 may be a composite laminate assembly 16 in the form of a panel 18, for example.

[0152] The tailored modulus surface layer 12d of Figure 4E minimizes strain mismatch 32d (see Figure 1A) between the surface layer 12d and the secondary coating 70. The tailored modulus surface layer 12d of Figure 4E reliably accommodates the magnitude variations between the surface layer 12d and the underlying structural ply layer 22 of plies 24, resulting in enhanced microcrack resistance 106 (see Figure 1A), minimized structural degradation 108 (see Figure 1A), and improved humidity and heat cycling performance 110 (see Figure 1A). In contrast, known structural panels and composite laminates that are not modulus tailored have a large strain mismatch between the structural panel or composite laminate and a secondary coating, such as paint, which can result in microcracks or other cracks in the structural panel or composite laminate when exposed to heat or moisture.

[0153] Reference is now made to FIG. 5, which is a flowchart of an exemplary embodiment of a method 190 of the present disclosure. In another embodiment of the present disclosure, a method 190 is provided for imparting high microcrack resistance 106 (see FIG. 1A), minimal structural degradation 108 (see FIG. 1A), and improved humidity heat cycling performance 110 (see FIG. 1A) to a composite structure 94 (see FIG. 1A), such as an aircraft composite structure 94a (see FIG. 1A), using the controlled modulus composite laminate system 10 (see FIG. 1A) described above. The blocks in FIG. 5 represent processes and / or portions thereof or elements, and the lines connecting the various blocks do not imply any particular order or dependency relationship between these processes or portions thereof. The disclosure of the steps of method 190 in FIG. 5 and herein does not necessarily dictate the order in which these steps are performed. Rather, one exemplary order is shown, but the order of steps can be varied as appropriate. Thus, certain processes may be performed in a different order or simultaneously.

[0154] As shown in FIG. 5, method 190 includes step 192 of providing a controlled modulus composite layup system 10. As described above, controlled modulus composite layup system 10 includes a structural assembly 14 (see FIGS. 1A, 2A, 3A, and 4A), such as a composite layup assembly 16 (see FIGS. 1A, 2A, 3A, and 4A) in a cured or to-be-cured configuration. As shown in FIG. 1A, in one embodiment, composite layup assembly 16 is in the form of a panel 18 (see also FIG. 6), including, for example, a wing panel 18a (see also FIG. 6), a fuselage panel 18b, a horizontal stabilizer panel 18c, a vertical stabilizer panel 18d, or other suitable panel 18. Composite layup assembly 16 is preferably comprised of a plurality of structural layers 20 (see FIGS. 1A and 2A), such as a plurality of structural ply layers 22 (see FIGS. 1A and 2A), pre-impregnated with a structural resin 60 (see FIGS. 1A and 2A). Each structural ply layer 22 has a ply modulus 25a (see FIG. 1A), a ply coefficient of thermal expansion (CTE) 26a (see FIG. 1A), and a ply coefficient of moisture expansion (CME) 28a (see FIG. 1A). The ply modulus 25a, ply CTE 26a, and ply CME 28a are the same for all ply layers 22. The composite layup assembly 16 prior to curing 34 (see FIG. 1A) or co-curing 34a (see FIG. 1A) comprises an uncured composite layup assembly (CLA) 16a (see FIG. 1A). The composite layup assembly 16 during curing 34 or co-curing 34a comprises a partially cured composite layup assembly (CLA) 16b (see FIG. 1A). The composite layup assembly 16 after curing 34 or co-curing 34a comprises a cured composite layup assembly (CLA) 16c (see FIG. 1A).

[0155] Each structural ply layer 22 includes a composite material 42 (see FIG. 1A ) including one or more carbon fiber reinforced polymers (CFRPs) 48 or plastics, one or more glass fiber reinforced polymers 52 or plastics, one or more aramid polymers 56, or other suitable composite material 42. The composite material 42 includes composite fibers 46 (see FIG. 1A ), such as carbon fibers 50 (see FIG. 1A ), glass fibers 54 (see FIG. 1A ), aramid fibers 58 (see FIG. 1A ), or other suitable composite fibers 46. As shown in FIG. 1A , the structural resin 60 includes a thermosetting structural resin 62, such as one or more of an epoxy structural resin 62 a, a phenolic structural resin 62 b, a polyimide structural resin 62 c, a bismaleimide (BMI) structural resin 62 d, a polyurethane structural resin 62 e, a fluoropolymer structural resin 62 f, a cyanate ester structural resin 62 g, or other suitable thermosetting structural resin 62.

[0156] As described above, the tailored modulus composite laminate system 10 further includes a surface layer 12 (FIGS. 1A-1B) applied directly to a structural assembly 14, such as a composite laminate assembly 16, either before curing 34 or co-curing 34a of the composite laminate assembly 16 or after curing 34 or co-curing 34a of the composite laminate assembly 16. The tailored modulus 25c (FIG. 1B) of the surface layer 12 is adjusted to form a tailored modulus surface layer 12d (FIG. 1B). The tailored modulus 25c is adjusted to be less than or equal to the ply modulus 25a. The surface layer 12 further has a surface layer coefficient of thermal expansion (CTE) 26b (FIG. 1B) that is different from the ply CTE 26a. The surface layer 12 further has a surface layer coefficient of moisture expansion (CME) 28b (see FIG. 1B) that is different from the ply CME 28a.

[0157] As described above, the controlled modulus composite laminate system 10 further includes at least one secondary coating 70 (FIG. 1A) applied directly to the surface layer 12 after the composite laminate assembly 16 has been cured 34 or co-cured 34a, either alone or together with the surface layer 12. The at least one secondary coating 70 has a secondary coating modulus 25d (see FIG. 1A) that is different from the ply modulus 25a, a secondary coating coefficient of thermal expansion (CTE) 26c (see FIG. 1A) that is different from the ply coefficient of thermal expansion (CTE) 26a, and a secondary coating coefficient of moisture expansion (CME) 28c (see FIG. 1A) that is different from the ply coefficient of moisture expansion (CME) 28a. The controlled modulus 25c of the controlled modulus surface layer 12d, that is, the surface layer 12, is adjusted to be equal to or greater than the secondary coating modulus 25d of the at least one secondary coating 70 prior to application of the at least one secondary coating 70 to the surface layer 12. That is, the surface layer modulus 25b of the surface layer 12, which is the adjusted modulus surface layer 12d, is adjusted before the secondary coating 70 is applied to the surface layer 12 to obtain the adjusted modulus 25c of the adjusted modulus surface layer 12d.

[0158] Step 192 of providing a controlled modulus composite laminate system 10 further includes providing the controlled modulus composite laminate system 10 in which the facing layer 12 comprises a facing film 64 (see FIG. 1B) applied directly to the composite laminate assembly 16 prior to curing 34 or co-curing 34a of the composite laminate assembly 16, such that the facing film 64 is co-cured or cured with the composite laminate assembly 16 to form the controlled modulus composite laminate assembly 11 (see FIG. 1A). In one embodiment, the facing film 64 comprises a resin-based composite facing film 64a (see FIG. 1B).

[0159] 1B, a facing film 64, such as resin-based composite facing film 64a, includes a prepreg thermoset resin 118, which may include one or more of adhesive 118a, epoxy 118b, phenolic 118c, polyimide 118d, bismaleimide (BMI) 118e, polyurethane 118f, fluoropolymer 118g, cyanate ester 118h, or other suitable prepreg thermoset resin 118. The prepreg thermoset resin 118 may further include a UV-resistant prepreg thermoset resin, an aliphatic prepreg thermoset resin, a flame-retardant prepreg thermoset resin, or other suitable prepreg thermoset resin 118.

[0160] Providing 192 the adjusted modulus composite laminate system 10 further includes providing the adjusted modulus composite laminate system 10, wherein the surface layer 12 includes one or more resin repair layers 68 (see FIG. 1B ) applied directly to the composite laminate assembly 16 after cure 34 of the composite laminate assembly 16. In one embodiment, the one or more resin repair layers 68 include a resin-based material 146 (see FIG. 1B ) having reinforcing composite fibers 46 a (see FIG. 1B ). 1B , including one or more of adhesive 118a, epoxy 118b, phenolic 118c, polyimide 118d, bismaleimide (BMI) 118e, polyurethane 118f, fluoropolymer 118g, cyanate ester 118h, or other suitable prepreg thermoset resin 118. Prepreg thermoset resin 118 may further include a UV-resistant prepreg thermoset resin, an aliphatic prepreg thermoset resin, a flame-retardant prepreg thermoset resin, or other suitable prepreg thermoset resin 118.

[0161] Step 192 of providing a controlled modulus composite laminate system 10 further includes providing the controlled modulus composite laminate system 10 in a configuration in which the surface layer 12 includes a lightning strike protection assembly 66 (see FIG. 1B ) applied directly to the composite layup assembly 16 prior to curing 34 or co-curing 34a of the composite layup assembly 16, the lightning strike protection assembly 66 being co-cured with the composite layup assembly 16. In one embodiment, the lightning strike protection assembly 66 includes a lightning strike expanded metal foil layer 122 (see FIG. 1B ) including a lightning strike expanded metal foil 124 (see FIG. 1B ) and a resin-infused scrim layer 132 (see FIG. 1B ) laminated to the lightning strike expanded metal foil layer 122. The resin-infused scrim layer 132 includes a non-metallic scrim 134 (see FIG. 1B ) infused with an infusion resin 136 (see FIG. 1B ). 1B, the non-metallic scrim 134 may include one of non-metallic scrim mat (SM) 134a, fiberglass scrim mat (SM) 134b, carbon fiber scrim mat (SM) 134c, woven scrim mat (SM) 134d, knitted polyester scrim mat (SM) 134e, non-woven scrim mat (SM) 134f, or other suitable non-metallic scrim 134. The non-metallic scrim 134 may further include a scrim woven or carrier medium having fiberglass reinforcement 54 (see FIG. 1A) or fiberglass reinforcement, carbon fiber 50 (see FIG. 1A) or carbon fiber reinforcement, nylon fiber or nylon fiber reinforcement, polyester fiber or polyester reinforcement, or other suitable fiber reinforcement.

[0162] Step 192 of providing a controlled modulus composite laminate system 10 further includes providing the controlled modulus composite laminate system 10 as a controlled modulus composite laminate system having at least one secondary coating 70 comprising one or more of a paint film 72, a basecoat film 72a, a topcoat film 72b, a primer film 74, an appliqué 76, a decal 78, a repair paste film 80, or other suitable secondary coating 70, as shown in FIG. 1A.

[0163] Step 192 of providing a controlled modulus composite laminate system 10 further includes providing the controlled modulus composite laminate system 10 in a configuration in which the surface layer 12 is applied directly to the composite laminate assembly 16 prior to curing 34 or co-curing 34a of the composite laminate assembly 16, and the surface layer 12 and the composite laminate assembly 16 are co-cured in co-curing 34a (see FIG. 1A) with heat 36 (see FIG. 1A) in an autoclave 90 (FIGS. 1A and 2B) to provide the cured controlled modulus composite laminate assembly 11 (see FIG. 1A) prior to application of the secondary coating 70.

[0164] In this embodiment, the facing layer 12 and the composite layup assembly 16 are subjected to a snap-cure process 35 (see FIG. 1A) as disclosed in U.S. Patent No. 11,752,708, which is incorporated herein by reference in its entirety. The snap-cure process 35 further includes an initial heating step in which the uncured composite layup assembly 16a (see FIGS. 1A and 2A) is heated to an initial temperature 38a (see FIG. 1A), resulting in a partially cured composite layup assembly 16b (see FIG. 1A). This initial heating step is sufficient to gel the prepreg thermoset resin 118 (see FIG. 1B) of the facing film 64 or the infusion resin 136 of the lightning strike expanded metal foil layer 122, both of which are still in their uncured state, but is not sufficient to gel the uncured structural resin 60 (see FIG. 1A). The snap-cure process 35 further includes subsequently heating the partially cured composite layup assembly 16b to a final temperature 38c (see FIG. 1A) that is higher than the initial temperature 38a (see FIG. 1A), thereby curing the controlled modulus composite layup assembly 11 (see FIG. 1A). The combination of the initial and subsequent heating is sufficient to fully cure both the prepreg thermoset resin 118 of the facing film 64 or the infusion resin 136 of the lightning-strike expanded metal foil layer 122 and the structural resin 60. The snap-cure process 35 further includes at least one of applying pressure 92 (see FIG. 1A), for example, high pressure 92a (see FIG. 1A), to the uncured composite layup assembly 16a during the initial heating or applying such pressure to the partially cured composite layup assembly 16b during the subsequent heating.

[0165] As mentioned above, the lightning protection assembly 66 may further include a metal material 128 including one or more of copper 128a, aluminum 128b, titanium 128c, nickel 128d, gold 128e, silver 128f, or other suitable metal material 128, as shown in FIG. 1B, or may further include an alloy material 130 including one or more of copper alloy 130a, aluminum (AL) alloy 130b, titanium alloy 130c, nickel alloy 130d, gold alloy 130e, silver alloy 130f, bronze 130g, brass 130h, or other suitable alloy material, as shown in FIG. 1B.

[0166] As described above, the lightning strike protection assembly 66 further includes a resin-infused scrim layer 132 (see FIGS. 1B and 3A) laminated to the lightning expanded metal foil layer 122. The resin-infused scrim layer 132 includes a non-metallic scrim 134 (see FIG. 1B) infused with an infusion resin 136 (see FIG. 1B). The infusion resin 136 has a tailored viscosity, a tailored cure profile, and a tailored rheology. The infusion resin 136 includes a prepreg thermoset resin 118 similar to that included in the facing film 64, as shown in FIG. 1B, including one or more of adhesive 118a, epoxy 118b, phenolic 118c, polyimide 118d, bismaleimide (BMI) 118e, polyurethane 118f, fluoropolymer 118g, cyanate ester 118h, or other suitable prepreg thermoset resin 118.

[0167] 5, the method 190 further includes integrating 194 the tailored modulus composite laminate system 10 into a composite structure 94 (see FIG. 1A), such as an aircraft composite structure 94a (see FIG. 1A). The composite structure 94, such as the aircraft composite structure 94a, may include, for example, a fuselage 202 (see FIG. 6) of an aircraft 200a (see FIG. 6), a wing 204 (see FIG. 6) of the aircraft 200a, a horizontal stabilizer 212 (see FIG. 6) of the aircraft 200a, or any other suitable composite structure 94, such as any other suitable aircraft composite structure 94a.

[0168] As shown in FIG. 5, the method 190 further includes a step 196 of using the tailored modulus composite laminate system 10, including the tailored modulus surface layer 12, i.e., tailored modulus surface layer 12d (see FIG. 1B), to form a stress relief 31 (see FIG. 1B), such as laminar stress relief 31a (see FIG. 1B), an integrated damping portion 102 (see FIG. 1B), and an integrated buffer portion 104 (see FIG. 1B), between the at least one secondary coating 70 and the plurality of structural ply layers 22 of the composite laminate assembly 16, thereby This can enhance microcrack resistance 106 (see FIG. 1A), minimize structural degradation 108 (see FIG. 1A), and improve heat and humidity cycling performance 110 (see FIG. 1A) of a composite structure 94 having the adjusted modulus composite laminate system 10 when the composite structure 94 having the adjusted modulus composite laminate system 10 is subjected to one or more of heat exposure 96 (see FIG. 1A) and moisture exposure 98 (see FIG. 1A) during a heat and humidity cycling event 100 (see FIG. 1A).

[0169] In one embodiment, where the facing layer 12 and the composite layup assembly 16 are co-cured prior to application of the secondary coating 70 to form a controlled modulus composite layup assembly 11 (see FIG. 1A ), the method 190 includes providing an uncured structural assembly 14, such as a composite layup assembly 16, e.g., uncured composite layup assembly 16a (see, e.g., FIGS. 1A and 2A ), having a plurality of structural ply layers 22 pre-impregnated with a structural resin 60, wherein each structural ply layer 22 has the same ply modulus 25a, the same ply coefficient of thermal expansion (CTE) 26a, and the same ply coefficient of moisture expansion (CME) 28a. The method further includes providing a surface layer 12 having an adjusted modulus 25c that adjusts the surface layer modulus 25b, a surface layer coefficient of thermal expansion (CTE) 26b that is different from the ply CTE 26a, and a surface layer coefficient of moisture expansion (CME) 28b that is different from the ply CTE 28a; and providing at least one secondary coating 70 having a secondary coating modulus 25d that is different from the ply modulus 25a, a secondary coating coefficient of thermal expansion (CTE) 26c that is different from the ply CTE 26a, and a secondary coating coefficient of moisture expansion (CME) 28c that is different from the ply CTE 28a.

[0170] In this embodiment, when the surface layer 12 and the composite layup assembly 16 are co-cured to form the tailored modulus composite layup assembly 11 (see FIG. 1A) prior to application of the at least one secondary coating 70, the method 190 includes tailoring the surface layer modulus 25b of the surface layer 12 so that the tailored modulus 25c (see FIG. 1B) is equal to or greater than the secondary coating modulus 25d of the secondary coating 70. The method 190 of this embodiment further includes laying up 82 (see FIG. 1A) the surface layer 12 having the tailored modulus 25c, i.e., tailored modulus surface layer 12d (see FIG. 1B), on the uncured composite layup assembly 16a (see FIG. 1A). Alternatively, the uncured composite layup assembly 16a may be laid up on the surface layer 12 having the tailored modulus 25c, i.e., tailored modulus surface layer 12d.

[0171] In this embodiment, the method 190 may further include curing 34 or co-curing 34a (see FIG. 1A) the surface layer 12 having the adjusted modulus 25c laminated to the uncured composite laminate assembly 16a with heat 36 (see FIG. 1A) in a heating device 88 (see FIGS. 1A and 2B), such as an autoclave 90 (see FIGS. 1A and 2B).

[0172] In this embodiment, the method 190 may further include applying at least one secondary coating 70 to the surface layer 12 after curing 34 or co-curing 34 a to obtain a controlled modulus composite laminate system 10 .

[0173] In this aspect, the method 190 may further include integrating the tailored modulus composite laminate system 10 into a composite structure 94, such as an aircraft composite structure 94a. The composite structure 94, such as an aircraft composite structure 94a, may include, for example, a fuselage 202 (see FIG. 6) of an aircraft 200a (see FIG. 6), a wing 204 (see FIG. 6) of the aircraft 200a, a horizontal stabilizer 212 (see FIG. 6) of the aircraft 200a, or any other suitable composite structure 94, such as any other suitable aircraft composite structure 94a.

[0174] In one aspect, the method 190 further includes using the adjusted modulus composite laminate system 10 including the surface layer 12 having the adjusted modulus 25c to form a stress relief section 31 (see FIG. 1B ), an integrated damping section 102 (see FIG. 1B ), and an integrated buffer section 104 (see FIG. 1B ) between the at least one secondary coating 70 and the plurality of structural ply layers 22 of the composite laminate assembly 16, thereby enhancing microcrack resistance 106 (see FIG. 1A ), minimizing structural degradation 108 (see FIG. 1A ), and improving heat and humidity cycling performance 110 (see FIG. 1A ) of a composite structure 94 including the adjusted modulus composite laminate system 10 when the composite structure 94 including the adjusted modulus composite laminate system 10 is subjected to heat exposure 96 (see FIG. 1A ) and / or moisture exposure 98 (see FIG. 1A ) during a heat and humidity cycling 100 (see FIG. 1A ) or a heat and humidity cycling event.

[0175] Reference is now made to Figure 6. Figure 6 is a perspective view of a vehicle 200, such as an aircraft 200a, having one or more composite structures 94, i.e., one or more aircraft composite structures 94a, incorporating the exemplary controlled modulus composite laminate system 10 (see Figures 1A, 2C, 3C, and 4C). As shown in Figure 6, the vehicle 200, such as the aircraft 200a, has a fuselage 202, wings 204, an engine 206, and a tail section 208. As shown in Figure 6, the tail section 208 includes a vertical stabilizer 210 and a horizontal stabilizer 212.

[0176] The adjusted modulus composite layup system 10 (see FIG. 6) includes a cured composite layup assembly 16 (see FIG. 1A) in the form of a panel 18 (see FIG. 6), such as a wing panel 18a (see FIG. 6) of an aircraft composite structure 94a that is a wing 204 (see FIG. 6) of an aircraft 200a (see FIG. 6). The composite layup assembly 16 of the adjusted modulus composite layup system 10 is a cured composite layup assembly 16c (see FIG. 1A) in the form of a panel 18 (see FIG. 1A), such as a fuselage panel 18b (see FIG. 1A) of a fuselage 202, a horizontal stabilizer panel 18c (see FIG. 1A) of a horizontal stabilizer 212, or any other suitable panel 18, or in the form of a member that forms part of a composite structure 94, such as the aircraft composite structure 94a.

[0177] The composite structure 94, for example, an aircraft composite structure 94a, is preferably constructed of a composite material 42 (see FIG. 1A), such as a carbon fiber reinforced polymer (CFRP) 48 (see FIG. 1A) or carbon fiber reinforced plastic, a glass fiber reinforced polymer 52 (see FIG. 1A) or glass fiber reinforced plastic, an aramid polymer 56 (see FIG. 1A), or other suitable type of composite material 42. In the tailored modulus composite laminate system 10, the skin layer 12 has a tailored modulus 25c, and such a system can be incorporated into a composite structure 94, such as an aircraft composite structure 94a having an aerodynamic surface, for example, a wing 204, a fuselage 202, a horizontal stabilizer 212, or other aerodynamic surface in an aircraft 200a, to enhance microcrack resistance 106 (see FIG. 1A), minimize structural degradation 108 (see FIG. 1A), and improve humidity heat cycling performance 110 (see FIG. 1A).

[0178] While the aircraft 200a shown in Figure 6 generally represents a commercial passenger aircraft having one or more aircraft composite structures 94a, the teachings of the controlled modulus composite layup system 10 (see Figure 1A) and method 190 (see Figure 5) of the present disclosure may be applied to aircraft composite structures 94a in other passenger aircraft. The teachings of the controlled modulus composite layup system 10 (see Figure 1A) and method 190 (see Figure 5) of the present disclosure may also be applied to aircraft composite structures 94a in cargo aircraft, military aircraft, rotorcraft, and other types of aircraft or air vehicles, as well as aerospace vehicles such as spacecraft, satellites, space launch vehicles, rockets, and other aerospace vehicles. Furthermore, the teachings of the controlled modulus composite layup system 10 (see Figure 1A) and method 190 (see Figure 5) of the present disclosure may also be applied to composite structures 94 in watercraft, automobiles, trains, buildings, or other suitable vehicles or structures.

[0179] 7 and 8, Figure 7 is a flowchart of an exemplary aircraft manufacturing and service method 300, and Figure 8 is an exemplary block diagram of an aircraft 316. With reference to Figures 7 and 8, aspects of the disclosure may be described in relation to the aircraft manufacturing and service method 300 shown in Figure 7 and the aircraft 316 shown in Figure 8.

[0180] During pre-production, the exemplary aircraft manufacturing and service method 300 includes specification and design 302 of the aircraft 316 and material procurement 304. During production, component and subassembly manufacturing 306 and system integration 308 of the aircraft 316 occurs. The aircraft 316 then undergoes certification and delivery 310 and enters service 312. While in customer service 312, the aircraft 316 undergoes a schedule of routine maintenance and service 314, which may include modifications, reconfigurations, refurbishments, and other suitable maintenance.

[0181] Each step of aircraft manufacturing and service method 300 may be performed or carried out by a system integrator, a third party, and / or an operator (e.g., a customer). A system integrator may include, for example, without limitation, any number of aircraft manufacturers and major system subcontractors. A third party may include, for example, any number of vendors, subcontractors, and suppliers. An operator may include, for example, an airline, a leasing company, a military entity, a service organization, or any other suitable operator.

[0182] 8 , an aircraft 316 produced by exemplary aircraft manufacturing and service method 300 may include an airframe 318 having a number of systems 320 and an interior 322. Examples of the number of systems 320 include one or more of a propulsion system 324, an electrical system 326, a hydraulic system 328, and an environmental system 330, as well as any number of other systems. While described as being used in the aerospace industry, the principles of the present disclosure may be applied to other industries, such as the automotive industry.

[0183] Apparatus and systems embodied herein may be employed in one or more of the steps of aircraft manufacturing and service method 300. For example, parts or subassemblies corresponding to part and subassembly manufacturing 306 may be similarly manufactured or produced as parts or subassemblies produced during in-service 312 of aircraft 316. Additionally, one or more apparatus embodiments, method embodiments, or a combination thereof may be employed during part and subassembly manufacturing 306 and system integration 308, for example, to substantially increase the speed or reduce the cost of assembly of aircraft 316. Similarly, one or more apparatus embodiments, method embodiments, or a combination thereof may be employed during in-service 312 of aircraft 316, for example, but not limited to, during maintenance and service 314.

[0184] The disclosed embodiment of a controlled modulus composite laminate system 10 (see FIG. 1A) and method 190 (see FIG. 5) can provide a system and method for reducing microcracks 116 (FIGS. 1A, 2E, 3E, 4E) in the surface layer 12 (FIGS. 1A-1B) and secondary coating 70 (see FIG. 1A) during heat and humidity cycling 100 (see FIG. 1A), and can provide a controlled modulus composite laminate system 10 and a composite structure 94, such as an aircraft composite structure 94a, having the controlled modulus composite laminate system 10 that achieves high microcrack resistance 106 (see FIG. 1A), minimized structural degradation 108 (see FIG. 1A), and improved heat and humidity cycling performance 110 (see FIG. 1A). The disclosed embodiment of a controlled modulus composite laminate system 10 (see FIG. 1A) and method 190 (see FIG. 5) involves adjusting or modifying the modulus 25 (see FIG. 1A) of a surface layer 12 (see FIGS. 1A-1B) located on an uppermost surface of a structural assembly 14, such as a composite laminate assembly 16 (see FIG. 1A), to have an adjusted modulus 25c (see FIG. 1B) that is equal to or greater than the secondary coating modulus 25d (see FIG. 1A). By having the adjusted modulus 25c of the surface layer 12, different rates of expansion 112 (see FIG. 1A) and contraction 114 (see FIG. 1A) between the structural ply layer 22 and the secondary coating 70 of the composite laminate assembly 16 do not result in inelastic deformation. Such systems and methods include evaluating, selecting, designing, and treating the material of each individual material layer through its thickness to provide stress relief 31 (see FIG. 1B) for interlayer stresses due to thermal distortions during curing 34 (see FIG. 1A) or co-curing 34a (see FIG. 1A). Additionally, an approach that optimizes the tuned modulus 25c of various layers, including the surface layer 12, and a secondary coating 70 material that can be applied at lower temperatures can improve the efficiency of the overall system design to minimize surface cracks 116a (see FIG. 1A) and internal cracks 116b (see FIG. 1A) and improve lifetime thermal humidity cycling performance.Although strains 32 (see FIG. 1A) induced by heat exposure 96 (see FIG. 1A) and moisture exposure 98 (see FIG. 1A) differ between a composite laminate assembly 16 including multiple structural ply layers 22 and a secondary coating 70, by considering these characteristics during the design of the composite laminate assembly 16, tailoring the surface layer modulus 25b of the surface layer 12, and carefully selecting the secondary coating 70, the effects of the differential strains 32 can be reliably absorbed, thereby reducing or preventing the occurrence of surface cracks 116a or internal cracks 116b. In the tailored modulus composite laminate system 10, stress relief 31 is formed through the thickness of the tailored modulus composite laminate system 10 through the material design and modulus tailoring of the composite layers. This improves humidity heat cycling performance, as measured by surface cracks 116a and internal cracks 116b, in a structural assembly 14 such as the composite laminate assembly 16 (see FIG. 1A), and this improvement is achieved while maintaining approximately the same level of paint film adhesion.

[0185] Additionally, the disclosed embodiment of the controlled modulus composite laminate system 10 (see FIG. 1A) and method 190 (see FIG. 5) provides environmental protection to the secondary coating 70 and minimizes structural degradation 108 (see FIG. 1A) due to internal cracks 116b, thereby avoiding repair and expense during its lifetime. The disclosed embodiment of the controlled modulus composite laminate system 10 (see FIG. 1A) and method 190 (see FIG. 5) also minimizes premature aging of the product due to heat and humidity cycling 100 (see FIG. 1A) during use, thereby extending the interval between painting and other secondary coating maintenance during the product's lifetime. The controlled modulus surface layer 12d (see FIG. 1B) minimizes strain mismatch 32d (see FIG. 1A) between the surface layer 12d and the secondary coating 70. The tailored modulus surface layer 12d reliably accommodates the dimensional variations between the surface layer 12d and the underlying structural ply layer 22 of plies 24, resulting in enhanced microcracking resistance 106 (see FIG. 1A), minimized structural degradation 108 (see FIG. 1A), and improved humidity heat cycling performance 110 (see FIG. 1A). In contrast, known structural panels and composite laminates that are not modulus tailored have a large strain mismatch between the structural panel or composite laminate and secondary coatings such as paint, which can result in microcracking or other cracking of the structural panel or composite laminate when exposed to heat or moisture.

[0186] Additionally, the disclosed embodiment controlled modulus composite layup system 10 (see FIG. 1A) and method 190 (see FIG. 5) can mitigate, minimize, and / or prevent microcracks 116, improve microcrack resistance 106 (see FIG. 1A), and minimize structural degradation 108 (see FIG. 1A) to prevent further degradation of a composite layup assembly 16, such as cured composite layup assembly 16c (see FIG. 1A), thereby avoiding costly repairs and rework. Additionally, the disclosed embodiment of the tailored modulus composite laminate system 10 (see FIG. 1A ) and method 190 (FIG. 5 ) adjusts the surface layer modulus 25 b of the surface layer 12 of the composite laminate assembly 16 to be equal to or greater than the secondary coating modulus 25 d of the secondary coating 70 applied to the surface layer 12, thereby avoiding strain mismatch and preventing or minimizing cracks, such as surface cracks 116 a, in the secondary coating 70, thereby avoiding repairs and rework; preventing or minimizing the propagation of internal cracks 116 b to the underlying composite laminate assembly 16; minimizing structural degradation due to microcracks 116 and cracks; improving lifetime thermal humidity cycling performance; and allowing for differential expansion and contraction rates 112 and 114 between the structural ply layers 22 of the composite laminate assembly 16 and the secondary coating 70, providing advantages over known composite structural material systems and methods. Additionally, the disclosed embodiment of the controlled modulus composite laminate system 10 (see FIG. 1A) and method 190 (see FIG. 5) allows for a variety of face layer 12 configurations, including a face film 64 in one embodiment (see FIGS. 2A-2E), a lightning strike protection assembly 66 in another embodiment (see FIGS. 3A-3E), and one or more resin repair layers 68 in yet another embodiment (see FIGS. 4A-4E).

[0187] Many modifications and other embodiments of the present disclosure will come to mind to one skilled in the art to which this disclosure pertains having the benefit of the teachings of the foregoing descriptions and the associated drawings. The embodiments described herein are illustrative and not intended to be limiting or comprehensive. Although specific terms are employed herein, these terms are used in a generic and descriptive sense only and not for purposes of limitation. Functionally equivalent methods and apparatuses within the scope of the present disclosure, in addition to those enumerated herein, are possible from the foregoing description. All such modifications and variations are intended to be included within the scope of the appended claims. The present disclosure is limited only by the terms of the appended claims, along with the full scope of equivalents to which such claims are entitled.

Claims

1. a cured composite layup assembly including a plurality of structural ply layers pre-impregnated with a structural resin, each of the plurality of structural ply layers having the same ply modulus, ply coefficient of thermal expansion, and ply coefficient of moisture expansion; a surface layer applied directly to the composite layup assembly either before or after curing, the surface layer having an adjusted modulus less than or equal to the ply modulus, a surface layer coefficient of thermal expansion different from the ply coefficient of thermal expansion, and a surface layer coefficient of hygroscopic expansion different from the ply coefficient of hygroscopic expansion; 1. A controlled modulus composite laminate system comprising: at least one secondary coating applied directly to the surface layer after curing of the composite laminate assembly, the at least one secondary coating having a secondary coating modulus different from the ply modulus, a secondary coating thermal coefficient of expansion different from the ply coefficient of thermal expansion, and a secondary coating moisture coefficient of expansion different from the ply coefficient of moisture expansion; the adjusted modulus of the surface layer is adjusted to be equal to or greater than the secondary coating modulus of the at least one secondary coating prior to applying the at least one secondary coating to the surface layer; Furthermore, the surface layer having the adjusted modulus forms a stress relief and integrated damping between the at least one secondary coating and the plurality of structural ply layers of the composite laminate assembly, thereby enhancing the microcracking resistance of the adjusted modulus composite laminate system and minimizing structural degradation when the adjusted modulus composite laminate system is subjected to one or more of heat and moisture exposure in a heat and humidity cycle.

2. 10. The controlled modulus composite laminate system of claim 1, wherein each of the plurality of structural ply layers comprises a composite material including one of one of: one or more carbon fiber reinforced polymers, one or more glass fiber reinforced polymers, or one or more aramid polymers.

3. 10. The controlled modulus composite laminate system of claim 1, wherein the facing layer comprises a facing film applied directly to and co-cured with the composite laminate assembly prior to curing of the composite laminate assembly, the facing film comprising a resin-based composite facing film.

4. 4. The controlled modulus composite laminate system of claim 3, wherein the resin-based composite facing film comprises a prepreg thermoset resin including one or more of adhesive, epoxy, phenolic, polyimide, bismaleimide, polyurethane, fluoropolymer, and cyanate ester.

5. the surface layer comprises a lightning protection assembly applied directly to the composite layup assembly prior to curing of the composite layup assembly and co-cured with the composite layup assembly, the lightning protection assembly comprising: a lightning-strike expanded metal foil layer including a lightning-strike expanded metal foil; 10. The controlled modulus composite laminate system of claim 1, comprising a resin-infused non-metallic scrim layer laminated to the lightning-strike expanded metal foil layer.

6. 6. The controlled modulus composite laminate system of claim 5, wherein the lightning strike expanded metal foil comprises a discontinuous metal foil comprising one or more of perforated metal foil, expanded metal foil, metal mesh, metallized fiber mesh, metal screen, metallized fiber fabric, metal fabric, wire mesh, metal foam, and open-cell metal foam.

7. 6. The controlled modulus composite laminate system of claim 5, wherein the non-metallic scrim comprises one of a non-metallic scrim mat, a fiberglass scrim mat, a carbon fiber scrim mat, a woven scrim mat, a knitted polyester scrim mat, or a non-woven scrim mat.

8. 10. The controlled modulus composite laminate system of claim 1, wherein the surface layer comprises one or more resin repair layers applied directly to the composite laminate assembly after curing of the composite laminate assembly, each of the one or more resin repair layers comprising a resin-based material with reinforcing composite fibers.

9. 10. The controlled modulus composite laminate system of claim 1, wherein the at least one secondary coating comprises one or more of a paint film, a basecoat film, a topcoat film, a primer film, an applique, a decal, or a repair paste film.

10. 10. The adjusted modulus composite laminate system of claim 1, wherein the surface layer comprises a surface layer shock absorber that absorbs differential stresses and strains in the at least one secondary coating and each of the plurality of structural ply layers of the composite laminate assembly when the adjusted modulus composite laminate system is subjected to one or more of heat and moisture exposure, thereby enhancing the humidity heat cycling performance of the adjusted modulus composite laminate system.

11. 1. An aircraft having one or more aircraft composite structures including a controlled modulus composite laminate system, the one or more aircraft composite structures: The torso and one or more wings; a tail section including one or more vertical and horizontal stabilizers; the tailored modulus composite laminate system is incorporated into the one or more aircraft composite structures; a cured composite layup assembly including a plurality of structural ply layers pre-impregnated with a structural resin, each of the plurality of structural ply layers having the same ply modulus, ply coefficient of thermal expansion, and ply coefficient of moisture expansion; a surface layer applied directly to the composite layup assembly either before or after curing, the surface layer having an adjusted modulus less than or equal to the ply modulus, a surface layer coefficient of thermal expansion different from the ply coefficient of thermal expansion, and a surface layer coefficient of hygroscopic expansion different from the ply coefficient of hygroscopic expansion; at least one secondary coating applied directly to the surface layer after curing of the composite laminate assembly, the at least one secondary coating having a secondary coating modulus different from the ply modulus, a secondary coating thermal coefficient of expansion different from the ply coefficient of thermal expansion, and a secondary coating hygroscopic coefficient of expansion different from the ply coefficient of thermal expansion; the adjusted modulus of the surface layer is adjusted to be equal to or greater than the secondary coating modulus of the at least one secondary coating prior to applying the at least one secondary coating to the surface layer; and wherein the surface layer having the adjusted modulus forms a stress relief and integrated damping between the at least one secondary coating and the plurality of structural ply layers of the composite laminate assembly, thereby enhancing microcracking resistance and minimizing structural degradation of the one or more aircraft composite structures including the adjusted modulus composite laminate system when the one or more aircraft composite structures having the adjusted modulus composite laminate system are subjected to one or more of heat exposure and moisture exposure in a heat and humidity cycle.

12. 12. The aircraft of claim 11, wherein the skin layer comprises a skin film applied directly to and co-cured with the composite layup assembly prior to curing of the composite layup assembly, the skin film comprising a resin-based composite skin film.

13. the surface layer comprises a lightning protection assembly applied directly to the composite layup assembly prior to curing of the composite layup assembly and co-cured with the composite layup assembly, the lightning protection assembly comprising: a lightning-strike expanded metal foil layer including a lightning-strike expanded metal foil; 12. The aircraft of claim 11, including a non-metallic scrim including an infused resin, and a resin-infused scrim layer laminated to the lightning strike expanded metal foil layer.

14. 12. The aircraft of claim 11, wherein the skin layer comprises one or more resin repair layers applied directly to the composite layup assembly after the composite layup assembly has cured, each of the resin repair layers comprising a resin-based material with reinforced composite fibers.

15. 12. The aircraft of claim 11, wherein the at least one secondary coating comprises one or more of a paint film, a basecoat film, a topcoat film, a primer film, an applique, a decal, or a repair paste film.

16. 1. A method for increasing microcracking resistance and minimizing structural degradation in a composite structure using a tailored modulus composite laminate system, comprising: providing a controlled modulus composite laminate system, wherein the controlled modulus composite laminate system comprises: a cured composite layup assembly including a plurality of structural ply layers pre-impregnated with a structural resin, each of the plurality of structural ply layers having the same ply modulus, ply coefficient of thermal expansion, and ply coefficient of moisture expansion; a surface layer applied directly to the composite layup assembly either before or after curing, the surface layer having an adjusted modulus less than or equal to the ply modulus, a surface layer coefficient of thermal expansion different from the ply coefficient of thermal expansion, and a surface layer coefficient of hygroscopic expansion different from the ply coefficient of hygroscopic expansion; at least one secondary coating applied directly to the surface layer after curing of the composite laminate assembly, the at least one secondary coating having a secondary coating modulus different from the ply modulus, a secondary coating thermal coefficient of expansion different from the ply coefficient of thermal expansion, and a secondary coating moisture coefficient of expansion different from the ply coefficient of moisture expansion, wherein the adjusted modulus of the surface layer is adjusted to be equal to or greater than the secondary coating modulus of the at least one secondary coating prior to applying the at least one secondary coating to the surface layer; incorporating the tailored modulus composite laminate system into the composite structure; and using the composite structure including the surface layer having the tailored modulus to form a stress relief and integrated damping between the at least one secondary coating and the plurality of structural ply layers of the composite laminate assembly, thereby increasing the microcracking resistance and minimizing structural degradation of the composite structure having the tailored modulus composite laminate system when the composite structure is subjected to one or more of heat and moisture exposure in a heat and humidity cycle.

17. The step of providing the controlled modulus composite laminate system further comprises providing the controlled modulus composite laminate system, the surface layer comprising: a facing film applied directly to the composite layup assembly prior to curing of the composite layup assembly and co-cured with the composite layup assembly, the facing film comprising a resin-based composite facing film; one or more resin repair layers applied directly to the composite layup assembly after the composite layup assembly has cured, each resin repair layer comprising a resin-based material with reinforcing composite fibers; a lightning strike protection assembly applied directly to and co-cured with the composite layup assembly prior to curing of the composite layup assembly, a lightning-strike expanded metal foil layer including a lightning-strike expanded metal foil; the lightning strike protection assembly including a resin-infused scrim layer laminated to the lightning strike expanded metal foil layer, the resin-infused scrim layer including a non-metallic scrim including an infused resin; 17. The method of claim 16, comprising providing a controlled modulus composite laminate system having a configuration comprising one of:

18. 17. The method of claim 16, wherein the step of providing the adjusted modulus composite laminate system further comprises providing the adjusted modulus composite laminate system wherein the at least one secondary coating comprises one or more of a paint film, a basecoat film, a topcoat film, a primer film, an applique, a decal, or a repair paste film.

19. 17. The method of claim 16, wherein the step of providing the controlled modulus composite laminate system further comprises providing the controlled modulus composite laminate system wherein the surface layer is applied directly to the composite laminate assembly prior to curing of the composite laminate assembly, and the surface layer and the composite laminate assembly are thermally co-cured in an autoclave to obtain a cured controlled modulus composite laminate assembly before the at least one secondary coating is applied.

20. The step of incorporating the controlled modulus composite laminate system into the composite structure further comprises: aircraft fuselage, the wing of an aircraft, or 17. The method of claim 16, including incorporating into a composite structure including one of an aircraft horizontal stabilizer.

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